TU-16 AIRCRAFT AIRCRAFT SERVICE MANUAL

• Document Type: CREST
• Collection: General CIA Records
• Document Number (FOIA) /ESDN (CREST): CIA-RDP78-03066R000300070001-0
• Release Decision: RIPPUB
• Original Classification: S
• Document Page Count: 195
• Document Creation Date: December 19, 2016
• Document Release Date: October 30, 2003
• Sequence Number: 1
• Content Type: REPORT

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Attachment 3
DIA review(s) completed.
AIRCRAFT SERVICE-MANUAL
Book II
Navigation Equipment, Autopilot, Oxygen, Electrical,
Photo, and Radio Equipment
GROUP 1
Excluded from automatic
downgrading and
declassification
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A-RDP78-03066R0003000700)1-0
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TY-16 AIRCRAFT
SERVICE NIANUAL
Book Two
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1Y-16 AIRCRAFT
SERVICE MANUAL
Book Two
Navigation Equipment.
Autopilot. Oxygen, Electrical,
Photo, an Radio Equipment
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The book contains 261 pages.
Besides, there are seven inserts on seven sheets, and
figures on forty three sheets at the end of the book.
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CONTENTS
MATIgatiOn gniDvutand Engine instrument'
General
Aocess
Pre-Flight inspection
Probable Troubles of Navigational Inatruments amd Their
Remedies
Checking the Pitot-Static System for Tightness
Post-Flight Operation
Electrical Instruments of Navigation Equipment
General
Maintenance Instruotions
Pre-Plight inspection
Visual Inspection
Post-Plight Inspection
Checking, the instruments for Correspondence to Their Basin
Specifications
Engine Instruments and Gauges
General
Maintenance instructions
Pre-flight inspection
Checking the Instruments for Correspondence to Their Basic
Gpecifications
Troubles and Remedies
Elimination of Compass Deviation on Instruments URK-7
APE-5 No.1 ana No.2 and/S-12
Autopilot AA-5-2K
General
Checking Autopilot for installation =Aircraft and Operation
under Current
Paternal Inspection
Checking Operation of Energized Autopilot
Faults and Remedies
Oxygen Equipment
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General Specifications
Accessibility for Inspection
Preparation for Flight inspection
Charging RHI-30 Converters with Liquid Oxygen
Checking Operation. of Distant-Reading Liquid -
Oxygen level indicator, Type NUR
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35
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38
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46
49
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57
60-
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67
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Putting 0E40 Converters to Operating Condition
73
Checking Serviceability of 101I-30 Converters Before Flight
74
Checking Operation of 101-24M Economizer
74
Possible Faults of Oxygen System and Means of Their Elimina-
tion ?
76
Checking System for leakage
76
Effects of femperature Change during Check of System Gastight-
ness
77
Checking Shut-Off Valves
78
Faults of 1011-30 Converter
79
Washing the Vessel 101I-30 Converter
79
Care of 10II-30 Converter
Petits of I0-24 Economizer 82
Faults of KM-30M Mask
Faults of Tee-Pieces with Non-Return Valves.
Faults of 511-23 Parachute Oxygen Breathing Apparatus
Post-Plight Inspection
Pressure Release
Storage of Liquid Oxygen in 101E-30 Converters 85
Storage of Liquid Oxygen in Sealed Vessels of 511E-30
- Converters 85
Storage of Liquid Oxygen in 101W-3o Converters under Pressure 86
Precautionary Measures 87
Instructions for Fac.cing Parachutes with 511-23 Oxygen
Breathing Apparatus 87
83
84
84
4
85
Electrical Equipment
89
General 89
Aircraft Electric Mains 89
Operating Duties of Electric Mains 91
Protection of Electric Mains 96
Wiring 97
laying and Removing the Cables 99
Maintenance of Junction Boxes and Electric Control Boards
Specific Features of Aluminium Wire Maintenance
Regulation and Cheek-Out of Bonding Arrangement
Operation Peculiarities of D.C. Power Supply Sources
Generator Maintenance
Storage Battery Maintenance
Connecting D.C. Ground Supply Source
Control Over D.C. leiter Supply Source and Electric Mains
Operation Peculiarities of.A.C. Power Supply Sources
Connection of 1104500 Inverters and of Ground A.C. Power
Supply Source
Inverter Maintenance
Inverter Probable Troubles Constituting Reason for its
Replacement 117
Control Over A.C. Power Sources and A.G. Mains
101
101
104
105
105
107
110
111
115
115
116
118
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Adjusting and Checking the Operation of 1104500 Inverter lia
Electrically Heated Glass Panels 119
Maintenance of Heated Glass Panels 119
Dem of Glass Panel Heaters 121
Tail Empennage De-Icers 121
Brief 121
Checking Tail Empennage De-Icer System on the Ground 123
Instructions for Operation of Tail Empenrage De-Icer during
Flight 125
Warning System 125
Light Fitting C.IN-51 126
Care and Maintenance of Light and Sound Signal Units 134
Aircraft Interior Lighting . 134
Exterior lighting 142
Fire Fighting Equipment and Fire Warning Electric System 148
Checking Installation and Operation of Fire Fighting Equipment
Electric System 149
Fuel Shut-Off and Engine Fuel Cross-Feed Cocks 151
Possible Faults of Fire Fighting System and Their Elimination 151
Fuel Pumps Control and Fuel Gauge Electric System 152
Arrangement of Electric Units Included in the System 153
Checking Operation of Fuelmeter System on Aircraft 156
Checking Operation of Fuel Pumps Mammal and Automatic Control
System and Their Warning System 158
Feasible Faults of Fuel Ramps Control Electric System and Their
Elimination
Flap Control Electric System
acking Flaps Operation under Voltage
Tail Skid Control and Landing Gear Warning Electric Systole 169
Checking Operation of Tail Skid Control and yssAing Gear Warning
System under Voltage
Trim Tab Electric Control System
Voltage Check of Trim Tab Electric Control System 173
. Possible Faults of Electrical Parts of Trim Tab Control System
and Their Elimination 175
Brake System Pump Control Electric System 175
Checking Operation of Hydraulic System Electric Control 175
Cabin Heating Electric System 177
Feasible Faults of Cabin Heating Electric System and Their
Elimination 178
Pre-Plight Preparation 179
Fre-Flight Preparation Before Energizing Electrical :Equipment 179
182
182
183
162
165
168
171
172
Voltage Check of Electrical Equipment
Pest-Flight Inspection
Checking Instruments Serviceability
Photographic Equipment
General
Pre-Flight Preparation
Pre-Flight Preparation of Daytime Photography Cameras
Fre-Plight Preparation of AA Camera
185
185
189
189
193
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Pre-Flight Preparation of 012I-1 Camera
Atat-Plight Optratious
General
Radio Rouinment
Brief Infolmation
Pre-Flight Check 200
Poet-Flight Inspection 202
Visual Inspection of Radio Equipment, 202
. Checking of Live Radio Equipment 204
Interphone 83:atoms -CLIY-10 (feeder 27200-26) 204
Command Radio Set 1-PC5-70M (feeder 27200-24)
Radio Set FC2I-3M (feeder 27200-25)
Page
195
177
497
200
200
207
212
Radio Compass 12E-5 Hos 1 and 2 (feeder 27200-23) 214
lancer Receiver MPII-4811 (feeder 27200-23) 217
Radio Altimeter PB-17 (feeder 27200-22) 219
Low-altitude Radio Altimeter PS2 (feeder 27200-22) 222
Radio Range Finder CA-1 (feeder 27200-22) 225
localiser Receiver 112)114 and Glide-Slope Receiver rPII-2
of Instrument Lending System
Airborne Transponder CFO (feeder 27200,.20)
228
232
234
252
253
Radar Bombsight PEU-4
Poet-Flight inspection and Checking of Equipment
Troubles and Remedies
Measurement of Radio Boise Level 255
General Instructions 255
Operation of Electrical and Radio Facilities 2E1
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The Aircraft Service Manual comprises three books:
Book One includes: aircraft ground servicing; care and maintenance of
airframe, emergem.y and rescue equipment, aircraft control system, lAnAire gear,
hydraulic systems, power plants, high-altitude equipment; packing and Shipment
Of aircraft. . .
Book Two includes: case and maintenance of navigation equipment, autopilot,
oxycen, electrical, photo, and radio equipment.
Book Three includes: care and maintenance of bombing equipment.
qVir,RPT
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NAVIGATION EQUIPMENT AND ENGINE INSTRUMENTS
GENERAI,
1. The navit:ation equipment includes.:
(a) the Pitot-static system;
(b) the Pitot-static instruments, namely: type NYC-1200 airspeed (I. A.5.
and T./1..5.) indicators, type BA-20 altimeters, type BAP-30-3 rate-of-climb
indicators, type CCH-3 velocity bead warning units, type M0-1 machmeters
warning lights), tv7e BC-46 cabin pressure warning unite, type YBBA-15 cabin
altimeters;
Note: Apart from the above-listed instruments, the Pitot-statio system
actuates the T.A.S. trannmitter belonging to the H1-505 air posi-
tion indicator Sct and the altitude and speed transmitters of the
OB-11p optical bombsight set.
(c) the electrical Instruments, namely: type 1TUK-7 and type AAK-AB,-5
compasses, type H14-50B air position indicator, type ATT-2 gyro horizon, type
1UC-52 directional gyro, type TY3-48 tachometers. type 3Y1-53 turn indica-
tors;
(d) the autonomous Instruments, namely: type 101-12 magnetic compass,
type AK-53 hand-operated astrocompass, type VAC-51 aircraft sextant, type AI-1O
accelerometer, types AM10 and ABP-M clocks.
2. The engine instruments comprise the electric pressure gauges, thermo- .
meters, and tachometers.
The arrangement of instruments on the instrument panels is shown in Figs "1
to 7 inclusive.
MOT-STATIC NAVIGATIONAL INSTRUMENTS
GENINAL
The instruments, types NYC-1200 , BAP-30-3 , BN-20, MC-1, CO1-3,
YBBA-15 , and BC-46, are actuated by the Mot-static pressure system.
For installation on and removal of instruments from the instrument paris
refer to the Book "Repair of Aircraft".
Altimeters
Altimeters are mounted on the instrument paoels Of both pilots, navigator,
navigator-radar operator, and radio operator, i.e. five altimeters in total.
The BN-20 altimeters operate in the temperature range of +50 to-60?C, and
indicate a relative (barometric) flight altitude within the limits cf 0 to
20,000 m. One full revolution of the larger pointer corresponds to 1000 m.
full revolution of the smaller pointer corresponds to 20,000 16.
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NAVIGATION BQUIFUEN't AND ENGINE INSTRUMENTS
GENERAL
1. The navi,:ation equipment includes:
(a) the Pitot-static system;
(b) the Pitot-static instruments, namely: type NYC-1200 airspeed
and T.A.S.) indicators, type BA-20 altimeters, type BAP-50-5 rate-of-climb
imlicators, type 00/1-5 velocity head warning units, type MC-1 machmeters
warning lights), t^..-m BC-46 cabin pressure warning units, type YB1U715 cabin
altimeters;
Note: Apart from the above-listed instruments, the Pitot-static system
actuates the T.A.S. transmitter belonging to the AM-505 air posi-
tion indicator set and the altitude and speed transmitters of the
OBB-11p optical bombsight eet.
(c) the electrical instruments, namely: type 41113-7 and type AAA-2B-5
compasses, type BN-50B air position indicator, type 11T-2 gyro horizon, type
CUK-52 directional gyro, type 1Y3-48 tachometers, type BYA-53 turn indica-
tors;
(d) the autonomous instruments, namely: type VA-12 magnetic compass,
type AN-53 hand-operated astrocompass, type NAC-51 aircraft sextant, type AM,.10
accelerometer, types MO and ABF-M clocks.
. 2. The engine instruments Comprise the electric pressure gauges, therm- .
meters, and tachometers.
The arrangement of instruments on the instrument panels is shown in Flgsl
to 7 inclusive.
FITOT-STATIONAVIGATIONAL INSTRUMENTS
GENIT,AL
The instiuments, types ANC-1200 ? BLP-3u-3 , BA-20, m0-1, 00E-3,
SW-15 , and BC-46, are actuated by the Pitot-static pressure system.
For installation on and removal of instruments from the instrument pare1.3
refer to the Book "Repair of Aircraft".
Altimeters
Altimeters are mounted on the instrument paaels of both pilots, navigator,
navigator-ralar operator, and radio operator, i.e. five altimeters in total.
The BA-20 altimeters operate in the temperature range of +50 to -60?C and.
indicate a relative (barometric) flight altitude within the limits cf 0 to
20,000 m. One full revolution of the larger pointer corresponds to 1000 m.
full revolution of the smaller pointer corresponds to 20,000 m.
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?- 10 --
Airseeed Indicators
The PC-1200 airspeed indicators are installed on the instrument panels
of both pilots, navigator, navigator-radar operator, and radio operator. The
17CA.1200 airspeec indicators function in the temperature range between +50 and
-Wm and read I.A.S. from 100 to 1200 km/hr and T.A.S. from 400 to 1200 km/hr
at a flight altitude ranging from 0 to 15.000 m. The scale gradeation value is
10 km/hr, each 100 km/hr division being nuMbered.
Rate-of-Climb Indicators
The rate-of-climb indicators are mounted on the instrument panels of both
the left-seat pilot and the right-seat pilot. The rate-of-climb indicators operate
in the temperature range of +50 to -60?C and give the vertical component of the
rate of climb or descent within the range of 0 to 30 Il?/sec. both towards clitb
or descent.
gaehmeters
The machmeters are installed on the instrument panels of both the left-seat
pilot and the right-seat pilot. The machmeters function in the temperature range
of +50 to -6000 and read the Mach number within the limits of 0.5 to 1 at a
flight altitude ranging from 0 to 18,000 m. It Machnumher equal to 0.86 (the
instrument is adjusted for this value) the warning lights with red light filters,
mounted near the machmeters, go on. Under the warning lights there is a caption
"SPEED TOO HIGH". The warning light warns the pilot that the aircraft is approach-
he critical Mach number equal to 0.9 for this type of aircraft.
Veloci,v_Head Warning Units
TWO varying units are mountea on the aircraft behind the pilot s seats. The
warning units operate in the temperature range of -50 to +600C.By sending
electrical signals when the velocity head of qe2300 kg/m? or Mach 0.86 are reached.
the warning units warn the pilot that the aircraft is approaching the maximum
allowable flight speed.
Limitations for velocity head q and Mach number for various flying weights
versus flight altitude are given in the graph (See 1?ig.8).
Cabin Altimeters
Two cabin altimeters, type IBUZ-15, are mounted on the instrument panels
of the right-seat pilot and the radio opeiator.
The cabin altimeters operate in the temperature range of +50 to -60?C. They
'are intended to indicate the "altitude" in a pressurized cabin and the difference
between the cabin pressure and the outside air.
Pitot-Statio System
Schematic diagrams of Pitot and static pressure systems are shown in Figs 9
and 10.
Table 1 gives the necessary data on the connection of instruments to pressure
sources.
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Table1
Connection of Instruments to Pressure Sources
Nos
Name of pressure source
Connected instruments
1
2
3
4
5
7
OnAireraft (See Fig.9)
Pitot tube, front, left-side Type EIC-1200 airspeed indicator
and type MC-1 machmeter on left-seat
pilot's instrument panel; type KYC-
-12T! airspeed indicator on naviga-
tor's instrument panel; type CCH-3
velocity head warning unit, left-
side
Pitot tube, rear, left-side
(Flush-type) static vent, upper,
left-side
Static vent, medium, left-side
Static vent, lower, left-side
Pitot tubes right-side
Static vent, lower, right-side
T.A.S. transmitter of type 171M-50B
air position indicator set; *peed
transmitter of type 01113-11p optical
bombsight set; type KIC-1200 air-
speed indicator on operator's instru-
ment panel
Type KYC-1200 airspeed indicator
and type BA-17 altimeter on instru-
ment panels of navigator and operator
Type CCH-3 velocity head warning
unit, left-side; type ILC-1204
airspeed indicator, type B1-17 alti-
meter, type IMP-50-5 rate-of-climb
iedleator, type 1G-1 machmeter on
left-seat pilot's instrument panel
A.S. transmitter of type HB-50B
air position Indicator set; speed
and altitude transmitters of type
ORE-11p optical bombsight set
'Type UC-1200 airspeed indicator
and type MC-1 machaeter on right-
seat pilot's instrument panel; type
CCH-3 velocity head warning unit,
rieht-side; typo KYC-1200 airspeed
Indicator on radio operator's instru-
ment panel
Type KYD-1200 airspeed indicator,
type BA-17 altimeter, type BAP--50-3
rate-of-climb indicator, type YBRA-3
cabin altimeter on right-seat pilot's
instrument panel; type CCH-3 velocity
head warning unit, right-side; type
W17-17 altimeter, type nnA-3 cabin
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Nos
Name of pressure source
Connected instrument
8 Static vent1 upper, right-side
altimeter, and type NYC-1200 air-
speed indicator on radio operator's
instrument panel
For type An-54 automatic cabin-
pressure regulator (if type APA-50
automatic cabin-pressure regulator
is installed, the etatic pressure
outlet will be blanked off)
Inptellation_of Type 12-25_Speed_and Lltitude Recorder
1. Typo 12-75 recorder is mounted between frames No.9 and No.10 on tlm
right side (Fig.11).
The recorder is connected to the Pitot -static -sten in the following way:
insert tee-piece H7705-3/8 (available in the are parts set) between the
static line and the CCH-3 velocity head warning unit, connect the recorder hose
to the tee-piece.
The recorder supply hoses are connected to tee-pieces 1026A50-4 cut into
static line 117702-100-22 and Pitot line H7702-29-6 (the right-seat pilot's mains).
hanass
There are free acceeees to Pitot-atatic inetrunents mounted on instrument
panels. The inatrument panels of both pilots and navigator flap back thus
giving access to the rear side of the instruments.
Access to the Pitot-static system line is difficult in the following places:
(a) between frames Hos 5 - 9 on both sides;
(b) between frames Nos 9 - 12 on both sides. Moisture traps for collecting
moisture from the Pitot-static system axe located In this section on both sides
of the fUselage;
(c) in the P-5 fuselage section, starboard;
(d) in the F-4 fuselage section, starboard;
(e) in the region of frames Nos 49 - 571
? (f) in the F-6 fuselage section, starboard.
To
(a)
(b)
reach the line between frames Nos 5 -.9, starboard, proceed as follows:
open the access panels of the right-hand engine instrument board;
xenove the glass heating distribution box.
To reach the line beteeen frames Nos 5 - 9, port, do the.folloeing:
(a) flap back the left-seat pilot's instrument penal;
(b) remove the access panels of the left-hand engireat instrument board.
To reach the line between fru/m.6S No 9 - 12, starboard:
(a) remove the thyratron interrupters from the starboard rack in the
-operators cabin;
(b) remove the dynamotor of the PCB-70 aircraft radio set from the star-
board rack;
TO reach the line between frames Nos 9 - 12
rectifier of the radar bombsight.
In order to reach the line In the F-3 fuselage section, starboard, proceed
as follome:
port, remove the high-voltage
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(a) open the hatch door at the bottom section of frame NO.12;
(b) open the batches of the containers of fuel tanks No.1 and No.2;
(c) remove fuel tanks Uo.1 and No.2I
To get at the line in the F-4 fuselage section in the region of frames
gO5 27 - 34, utarboard, do the following:
(a) open the batch in the F-4?veelage section between frames Nos 27 - 29;
(b) remove starting feel tank H6154-120;
(c) remove air-cooler H5601-380:
(d) remove drain pipe H6152-38/1;
(e) slacken the yoke on pipe m152-38/3 and turn the branch pipe:
(f) remove bigh-altitude equipment pipes B7605-0/25.5.
Te read: the line in the F-4 fuselage section in the region of frames
Jo 49 - 75, proceed as follows:
. (a) open the batches of the containers of fuel tanks Nos 4 add 51
(b) remove fuel tanks lo.4 and Ho.%
(c) lift up the batches in the containers of fuel tanks.
To an access to the pipes in the 8-6 fuselage section, remove the PCBY
radio set equipment from the bottom section of frame Ho.69.
143-naririf IHSPBOTTOR
Prior to each
1. Remove protective oovers with red warning flags from the Pitot tubes.
2. Take the blanking plugs out of the static vents.
3. Make visual inspection of the instrument panels (check the instruments
for cover glans cracks, luminous paint for intactness, inetrunents for proper
attachment, etc.).
4. Drain moisture from moisture traps in rainy weather.
5. Check the position of the selector cock for switching the left-seat
pilot's instruments to emexgxnury supply and the presence of safety vire with a
seal on the cock. (The selector cock is installed on the left-hand engine instrm-
Rent board).
The cock must be set and sealed in the NORMAL position.
Before each flight, cheek the efficiency of the Pitot-stakic aystem in the
following manner:
1. Set the hands of tun-pointer altimeters to se= and the barometric scales
for tbe pressure check.
2. Wind up the clocks and see that they are in good repair.
3. Build up a pressure in the Pitot tubes equal to 60 - 75 mm Eg
corresponds to a speed of 400 to 550 ke/hx).
4. Connect a VUOuun source of 85 - 160 mmHg (which corresponds to an alti-
tude of 1000 - 2000 m.) to the flush-mounted static vents.
When the Mot-static system is serviceable the instruments gill teact to
supply as follows:
Alreeeed indicatore - with pressure increase in the Pitot line the hande -
will rotate clockwise.
Altimeters and cabin altimeters - eith vacuum increase the hands will rotate
clockeise.
Rate-of-clinb indicators - with vacuum increase the hands will deflect upwara,
ehile at constant vacuum of any mageitude the hands will return to zero.
.5. Check to see that the OCH-3 velocity head warning =it Benin a warning
signal. To this end:
(a) connect in turn a pressure source of the Kff..3 test set type to the
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TH-156 Mot tubes for both the left-hand and right-hind instrument panels of
the pilots;
(b) build up a'pressure in the Pitot-static system equal to 169 4 mm fig.
At this the c1}-51 ?warning lamp eill light upon a respective panel of the
pilot.
. 6. Check to see that the MC-1 macemeter sends a warning signal. To this end:
(a) disconnect the impact pressure lines from the CCH-3 velocity head warn-
ing units; blank off the pipe line ends;
(b) connect in turn a pressure source of the ET-3 test set type to the
TR-156. Pitot tubes for both the left-hand and right-hand instrument panels
of the Pilots;
Cc) build up a pressure in the Pitot-static system equal to 473 1. 19 mm Hg.
At this the warning lamp will light up on the pilot's instrument panel, whereas
the ma-1 machinator needle will be on the red line.
!Ate: If atmospheric pressure does not equal to 760 mm Hg, then during the
check create a pressure of 760 mm Hg In the static system.
PRORARLF TROUBLES OF NAVIGATIONAL INSTRNMENTS
IND TREIR REMEDIES
The Pitot-static system troubles include:
(a) unserviceable condition of Pitot-static instruments;
(b) leakage or clogging of the Pitot-static system proper.
To draw a conclusion on good or bad repairs of an instrument, if obvioun
defects axe not available, check the instrument as indicated below.
leakage of the Pitot-static system is eliminated by tightening the nipple
joints and replacing the rubberized hoses (in case the latter are worn out).
Clogging is eliminated be blowing the system.
Cheekier the Instruments
Altimeter
The altimeter check-up includes visual inspection of the instrument, check-
ing its readings for errors and its case for tightness. The altimeter case
tightness and the errors in altimeter readings can be checked in site by means
of the 1tUY-3 test set and master mercury barometer.
To check the altimeter, proceed as follows:
(a) set the pointers of both the master barometer and the altimeter under
test to zero;
(b) disconnect the altimeter to be checked from the aircraft static pressure
line and join it to a tee-piece connected with one end to the master barometer
(Fig.12) and with the other end to the KIIY-3 test set;
(0) using the KflY-3 test set, create a rarefaction in the altineter
corresponding to definite altitudes as read off the master barometer. Take into
account the altimeter instrumental corrections;
(d) record the readings of the altimeter ender test In the check list and
compute the errors. In doing so, take into consideration the corrections of the
master mercury barometer;
(e) compare the obtained corrections of the altimeter under test with the
corrections entered into the altimeter correction card.
If these corrections vary, compile a new correction card and use it in
flights.
The altimeter admissible errors (total instrumental errors) are given In the
altimeter Cettificate. If during the altimeter check it is found out that the
15 -?
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altimeter errors exceed the maximum permissible values, the instrument should be
replaced by a new one and the defective altimeter should be sent for adjustment .
to a special workshop.
After the errors and the instrument case tightness bave been checked, check
the lines for leakage.
LirsEeed Indicator
1. The operating efficiency of the airspeed indicator is determined by
visual inspection and check test.
2. The static gystem is checked for leakage at normal temperature by connect-
ing the instrument to a vacuum source. When rarefaction, corresponding to the
1200 kM/hr instrument reading, is created, the vacuum source is shut off with a
cock t By clamping the hose at the pipe cornection of impact pressure line, watch
the instrunent pointers, the readings of which should not change during one
minute.
3. Errors In the instrument readings are checked at nornal temperature in
the following manner (Fig.13):
(a) connect a pressure source to the instrument pipe connection with index
KA) and a vacuum source to the ;Ape connection with index 43 (0);
(b) cheek the error value for each numbered division of the dial by building
up a pressure (as read off a pressure gauge) corresponding to the dial readings;
loc) take the readings of the values to be checked both clockwise and counter-
clockwise at one and the same dial matk.
Maintain pressure at each dial mark being checked for not lesa than 1 minute.
*aim= pressure, corresponding to the 1200 km/hr dial mark, should be
maintained for not less than 15 minutes. Error value will be determined by cooper-
ing the readings of the airspeed indicator under 'beet with that of the master
pressure gauge;
(d) maintain vacuum (when checking the instrument at varioue altitudes),
corresponding to the altitude under check as read off the master baromet.r, taking
into account the calibration card given in the Service Manual of the airspeed
indicator;
(e) compare the data obtained during the check with that entered into the
correction card for speed and altitude. Correct the card should any difference
OCCur;
Cr) replace the airspeed indicator if the corrections obtained exneed the
permissible errors given in the instrument Certificate.
The correction cards are furnished with the speed and altitude indicators
mounted on the instrument panels of navigator, navigator-radar operator, and
both pilots.
The values to be determined by formula
Tindic =6 rinstr +6Taer +6 Vcompr
will be entered into column Vindi. ,
where Vindie is I.A.S. (tedicated airspeed);
Oinstr are the erroxs in instrument reading determined as stated above
Vaer
eVcompr
is an aerodenamic correction. It is a constant value for I.A.S.
and equals to 13 km/hr;
is a compressibility correction to be taken from tables.
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The values to be determined by formula
A Vtrue '6 Vinsbr Veer
will be entered into column A Tuve ,
where V
true
is T.1.8. (true airspeed), bVaer at standard atmosphere will be
taken from the graph (Fig. 14). The data obtained for ,Vaer, deter-
mined at an altitude of H = 8000 m., will be entered into the
third column, whereas the data for bVaer , determined at an
altitude of H = 12,000 m., will be entered into the fourth column.
Corrections for type W.20 altimeter mill be entered into column AIL
It is said on the reverse sides of the tables: "Aerodynamical and instru-
mental corrections and compressibility corrections are accounted for In AVinbie
Aerodynamical and instrumental corrections are taken into account in AVtraer.
For the table of aerodynamical corrections see the aircraft Service Log.
Bate-of-Climb Indicator
The instrument check-up includes visual inspection and airtightness check.
The instrument should be so tight that at a rarefaction of 380 mm Hg the rate of
pressure drop during ODP minute would not exceed 2 mm 4. Vacuum should be
created gradually without sharp jerks of the climb indicator's pointer.
Nachmeter
The instrument check-up comprises visual inspection, check of static pres-
sure line for tightness, and check for errors in readings.
The static system Should be so tight at a rarefaction of 380 mm Hg, supplied
to both pine connections, that the rate of pressure drop during one minute would
not exceed 8 an Hg.
The machmeter will be checked as shown in Fig.15.
The machmeter may be checked in situ. The machmeter is checked, by the
numbered divisions of the dial, namely 0 km. and at altitudes of 2, 6, 10, 14,
and 18 km. To check at these altitudes use the calibration card of the machmeter
Certificate. The Mach number readings will be taken both clockwise and counter.-
clockwise. The error value will be determined by comparing the readings of the
machmeter being tested with the reading of the master pressure gauge at an
altitude of 0 km. If atmospheric pressure does not correspond to 760 mm Hg, then
build up a pressure of 760 mm fig in the static system when checking the instru-
pent at an altitude of 0 km.
To check the machmeters at an altitude of 0 km., proceed as follows (Fig.15):
close cock 7, open cock 9, and using cook I supply the line with pressure which
should be read off pressure gauges 2 and 3 and which corresponds to the dial
divisions under check. In doing so maintain a pressure of 760 mm Hg as read
by the barometer. Simultaneously take the machmeter readings.
To check the machmeter at different altitudes" proceed as follows: close
cock 1, open cocks 7 and 9, and using cock 8 create controlled by the barometer
a rarefaction, corresponding to the altitude at which the machmeter should be
checked. Rarefaction should be read off barometer 4. This done, close cocks 7
and 8, and using cock I build up in the line pressure which should be read off
pressure gauges 2 and 3, and which corresponds to the main dial divisions
(according to the calibration table in the machmeter Certificate). Simultaneously
take the machneter readings.
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Upon completion of the machmeter check at a given altitude, close conk 1
at the final reading of the pressure gauge, gradually open cocks 7 and 9 and
than through cock 8 create a rarefaction corresponding to the subsequent
alti-
tinie.
Velocity Head Warning Dalt
The instrument check comprises visual inspection, checking the pitot-static
system for tightness, checking the operation of warning lamp for errors, as well
as checking the, electric circuit insulance.
1. The static pressure line is checked for tightness at normal temperature
by connecting the impact and static pressure pipe connections to a pressure
source of 300 mm lag (Fig.16). The pressure source will be blanked off with a
cock. Pressure differential rate per minute should not exceed 0.5 mm 06.
The impact pressure system (Fig.17) is checked by connecting the dynamic
pressure pipe connection to the pressure source.
Airtightness Should be preserved for 5 minutes at a pressure of .330 ma Hg.
No pressure differential is allowed during this time.
2. Operation of the warning lamp at normal temperature is checked for errors
in the following way. The warning unit is connected to the pressure source
(1!ig.17). By gradually increaeing pressure, watch the moment the circuit is
closed (the warning lamp goes on). When taking the reading counterclockwise,
gradually decrease pressure and watch the moment the circuit is open (the warn-
ing lamp goes off). The error value is determined by the -dings of the master
pressure gauge at the moment the warning lamp lights up.
3. Insulance of the current-carrying elements at relative humidity of 30
to 80% is checked by means of a megger, one wire of which is simultaneously
connected to three pins of the plug, while the other wire of the megger is
connected to the warning unit case. Insulance should, not be less than 20 megolos.
Cabin Altimeters
The instrument cheek includes Vislai inspection and testing the instrument
case for tightness. The instrument case is tested for tightness by connecting
the case pipe connections to a vacuum source. At a rarefaction corresponding.to
an altitude of 8 km. as read by the instrument, the vacuum source is blanked off
with a Cock. Then, by clamping the hoses at the pipe connections, watch rare-
faction decrease in the instrument case. The rate of pointer drop should not
exceed 400 m.per minute. ?
Airtightness of the instrument diaphragm, assembly is checked by connecting
the pipe connection with index C to the vacuum source. At a rarefaction
corresponding tc the instrument reading of 0.6 kg/c2, read off the excessive
pressure scale, the vacuum source is blanked off with a cock. Then, by clamping
the hose at the pipe connection, watch the pointer, the reading of which should
not change during one minute.
The altimeter readings are checked for errors using the method of checking
the instrument case for tightness by creaing rarefaction in the instrument
corresponding to the readings of the dial numbered divisions under check.
Rarefaction should be maintained at each dial mark being checked for not
less than 1 minute and at a maximum rarefaction - for not less than 15 minutes.
The differential pressure gauge operation Should be checked in the same
manner as the diaphragm assembly is checked for tightness.
To determine the instrumental errors, the altimeter readings are compared
with the readings of the master mercury barometer, while the readings of the
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differential pressure gauge are compared with the readings of the master mercury
pressure gauge which are both connected to the test set.
. Checking the Pitot-Static System
for Tightness
Testing the Static Lines for Tightness
'(a) Disconnect the rate-of-climb indicators from the static lines and blank
off the ends.
(b) instit'in turn the hose, connected to the vacuum source, into the holes
of all five static vents and create rarefaction (vacuum) corresponding to
700 km/hr as read off the airspeed indicator.
;kW It is allowed to check the static pressure lines for tightness with
the rate-of-climb indicators connected to the line. However, in this
cane create vacuum, corresponding to an airspeed of 700 km/br, and
equalize it with the atmospheric pressure grarbwil,y and for not less
than 2 minutes.
(c) Clamp the hose running from the vacuum source. Note the reading of the
airspeed indicator pointer and then determine the rate of airspeed drop per
minute.
Cd) Permissible leakage of the static pressure lines corresponds to a value
at which the rate of drop lathe readings of the airspeed indicators does not
exceed 5 km/hr per minute.
? Testing the Impact (Dynamic) Pressure
Iine for Tightness
(a) Fit a robber hose, connected with the pressure source, onto the Pitot
tubes (see to it that the drain hole is closed). Create a pressure in the line
corresponding to an airspeed of 700 km/hr read by the airspeed inA4rator.
(b) Clamp the hose running from the pressure source. Take the reading of the
airspeed indicator pointer and the- determine the rate of airspeed drop per
minute.
Permissible leakage of the impact pressure line corresponds to a value at
which the rate of drop in the readings of the airspeed indicators does not
exceed 2 km/kg per ;dente.
POST-TLIGHP OPERATIONS
If during the flight the Pitot-static instruments worked without failure,
then after the flight do the follbwing:
- (a) put the covers on the Pitot tubes;
(b) insert the blArlring plugs into the static vents;
(C) make visual inspection of the instruments on the inetrument panels
(instrument glasses, attachment of instruments, etc.);
(d) drain moistness from the moisture traps in rainy weather.
Should any malfunctions he detected in the operation of the instruments
during the flight, such as, for example, erroneous readings of the instruments,
fluctuation of pointers, different readings of identical instruments (for instance
type KYC-1200 airspeed indicators), installed on various instrument panels, etc.,
blow the Pitot-static system, check the system for airtightness and efficiency
as indicated above.
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HaECTRICAL 3NSTRUSENTS OF NAVIGATION
zcznammT
GENERAL
The electrical instruments of navigation equipment include: type ATU1t-7
remote-reading gYromagnetic compass, type AAR-1IB-5 remote-reading astrocompasn,
type 81-50 air position indicator, type kP5-2 gyro horizon, type ru-52
directicnal gyro, and type TY3-48 thermometer.
Type AP1(I-7 Remote-Reading Gyromagnetic
209PS?2
The MIN-7 compass is a basic magnetic compass on the aircraft. It is
intended to determine the magnetic and true courses of the aircraft.
The Aria-7 compass complete set (Fig.18) comprises:
type ll-3 transmitter 1 pc;
type P-2 gyro unit 1 pc;
type 7-10 amplifier 1 pc;
type YL M master indicatorbaavigator's indica-
tor) 1 pc;
repeater (additional indicator) 3 Pea;
type ER-55P5 erecting cutout 1 pc;
type CR-8 junction box 1 pc;
type 5k feet slave button 2 pea;
type BT-125 inverter 1 pc
1. Power supply
Basic apecifications
2. Power consumed from D.C. meins
with inverter
without inverter "
3. Power drawn from A.C. MPins
27 ? 2.7 V D.0.,
36 ? 3.6 V, three-
phame L.C.,
400 t 40 c.p.s.
not over 250 W
not over 25 W
not over 110 W
4. Navigator's indicator error by the scale of compass
not over 40
course
by the scale of true course (after measurement
method error, instrumental error, and compass
deviation have been eliminated) not over 10
5. Additional error :a; compass readings for each
net over 0.6?
not over 30
minute of tura
6. Error in repeaters" readings
2. Permissible angle of bank of aircraft, at which the
compass readings can be taken without using the
fast slave button
8. Temperature rang, (except for master indicator and
repeaters)
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9. iemperature range for master indicator and
repeaters from +50 to -35?O
If% Altitude limit . . .. up to 15.000 116
11. GOZIASEI is ready for operatim An 3 min. after power
supply is on
. Under unfavourable combination of flight corditions (bank with angular
speed less than 0.20 per second, altitude change, longitmiinal accelerationotc.)
the error in compass readings may reach 100.
All the assemblies, which go to make the ANK-7 compass complete set are '
interchangeable. In case the D7H-3 transnitter or master indicator are to be
replaced, correct an installation error and remove deviation on 24 compass points,
Air Position Indicator HI-50B
The HM-50B air position indicator is designed for continuous indication
of the aircraft position in rectangular axes, the drift being taken isto account.
The HZ-50B set (Fig.19) includes:
(a) T.A.S. transmitter - 1 piece;
(b) automatic course device - 1 piece;
(c) wind setter- 1 piece;
(d) P.R. computer - 1 piece;
(e) distribution box - I piece;
(f) supply-line filter CO-2 - 1 piece;
(g) supply-line filter CO-4 - 1 piece;
(h) inverter 1IA1'-10 (Fig.20) - 1 piece.
Basic apscifisations_
1. Power supply' D.C., 27
A.C., three-phase.
36 = 3.6 V,
400 = 40 c.p.s.
300 to 1200 km/hr'
2. Range of operating speeds
3. Range of wind speed 0 to 150 kM/hr-.?
4. Altitude up to 15,000 36
5. Coordinate system
rectangular with any
arrangement of the
axes
6. Maximum error at normal temperature (altitude up
to 8000 m., speed from 300 to 1100 km/hr) 5.5%a.
7.
Course indication error at 24 points (repeated
reams of. 1VE-7 compass main indicator) 1? max.
8. Power consumed:
direct current 25 1max.
alternating current In most loaded phase 35 W max.
The units of the H1-50B set are interchangeable. But in case of replace-
Lent .Of any unit except the inverter and filters, it is necessary to determine
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the total error of the set and the new correction to the change-over table of
tbe distribution box. Atter replacement of the distribution box or automatic
course device, do not fail to adjust the zero signal anew and to match the read-
ings of the A11m-7 compass main indicator with those of the H5-50B automatic
course device.
kro Morison AIS-2
The ATS-2 electric gyro horizon with slide indicators are designed to
determine the position of the aircraft In the space relative to the true horizon,
as well as to deter=ine aircraft sideslip.
The APB-2 gyro horizon makes it possible to check the following aircraft
acrobatics:
(1) aircraft circle turas with up to 800 banks;
(2) diving and climbing at angles up to 60?
Pb. peculiarity of the LFR-2 gyro horizon lies in the fact that the lateral
erecting mechanism is cutout at an angular velocity of aircraft turning exceed-
ing 0.2 deg/sec. In this connection, the krB-2 gyro horizon operates in conjunc-
tion with a BK-53-PB erecting cutout.
The gyro horizon and erecting cutout are supplied from the de1'-1. inverter.
1. Power supply
Basic Epecifications
alternating three-phase
current. 36 = 3.6T,
400 40 c.p.s.
1? max.
2. or In horizon determination
3. Time of initial erection at ambient temperature of:
3 min.eax.
3 min.max.
6 min.max.
4. Erection time from lateral and longitudinal tilts ... 6 to 12 min.
5. Time difference in gyro erection from opposite
tilts
6. Errors in circle turns and turns lasting not more
2?
3 min..ms.x.
than 6 min.
The ATB-.2 gyro horizon units are interchangeable.
Eregting Cutout_ER-53-PS
The purpose of the erecting cutout is to out out the erecting mechanism of
the gyro horizon when performing circle turns at an angular velocity exceeding
0.2 deg/sec.
The erecting cutout of the 4111K-7 compass cuts out the gyro erecting unit
at a turning velocity exceeding 0.3 deg/sec.
On some aircraft the cutout of the erecting unit of the gyro horizon and
IIIME-7 compass is performed with the help on one common. erecting cutout which
is adlusted to operate at a turning velocity exceeding 0.2 deg/sec.
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1. Power supply
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Rage Dacifications_
2. Maximum power consumed in A.C. circuit
3. Maximum power consumed in D.C. circuit
4. Sensitivity
5. Time of erecting cutout lag
alternating three-phase
clamant. 36 ? 3.6 V.
400 ? 40 c.p.s. direct
current, 271 2.7 V
0.45 A per phase
3W
0.2 or 0.3 deg/sec.
5 to 15 sec.
6. Mastanna erecting cutout lag time asymmetry 8 sec.
Notes. The erecting cutout should be so mounted on the aircraft that the
twin shock-absorbing springs are located on top and index "EP on
the erecting cutout casing is also on top.
ElectricResistance Thermometer 73-48_
The T78-48 resistance thermometer is designed to measure the outside air
temperature. The instrument includes the following units:
(a) indicator - 1 piece;
(b) transmitter - I piece.
_Bssic_Saecifications
1. Power supply
2. Range of measurement
3, Error of instrument does not exceed:.
at 20 t 500
at 50 t 5?C
at -60 t 5?C
The thermometer units are interchangeable.
27 t 2.7 V
-70 to 4.150?C
4?0
t700
MAINTENANCE INSTRUCTIONS
During service cheek the instruments of the navigation equipment before
and after the flight observing the instructions given below. Check also the
instruments in those cases which axe specially prescribed for each instrument
Pre-Flight Inapection
The pre-flight inspection comprises visual inspection of the aircraft and
a check of their readiness for operation.
Visual Inspection
In inspecting the instruments visually make sure that their outer surfaces
are not damaged, that the instruments are reliably secured to the instrument
beard or to the respective bracket and that the plug connectors or wires are
reliably connected to the respective terminal blocks. See also that the safety
fuses are in their places, that they are used in conformity with the diagrams
and reliably secured in their seats. Check the amplifier valves for proper
installation and the vireo for good condition, especially in places of attachment
to the plug connectors or respective terminal blocks. Make also sure that the
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respective knobs and spur racks rotate smoothly, that the dials move properly,
that the switches are reliably fixed in their positions, etc.
In performing visual inspection of the navigation equipment observe the
following sequence:
(a) examine, on the instrument board of the left pilot, the gyro horizon,
directional gyro indicator, indicator of the 1INK-7 compass, and fast slaving
button of the alLEC-7 compass;
(b) examine, on the instrument board of the right pilot, the gyro horizon,
directional gyro indicator, and the AAH-AB-5 compass course indicator;
(c) examine, on the navigator's instrument board, the AAK-45-5 compass
course indicator, track corrector of the Va-U-5 compass, main indicator of
the IIIMK-7 compass, fast slaving button of the rMX-7 compass, T.A.S.
transmitter, automatic course device, mind setter and D.R. computer of the
Iii-50B air position indicator;
(d) inspect the distribution box of the S8-50E air position indicator and
the amplifier of the AUK-7 compass;
(e) examine the nir-n , UT-70 and 0-125 inverters through which
the gyro horisens, the rag-52 directional gyro, air position indicator HE-50
and the 4:11E-7 compass are energized;
(f) inspect the computer of the AA1(-AB-5 compass;
(g) examine the transmitters of the AAI-B-5 compass, and the T1nc-7
compass;
(h) clean the transparent hood of the AA1-A5.-5 compass transmitter of
dust and dirt. To avoid scratches wipe the hood with a piece of soft fabric
soaked in alcohol;
(i) check the colour of the silica gel crystals in the dehydrator of the
11A1C-ITD-5 compass transmitter. If the silica gel crystals have turned pink or
brown, replace the dehydrator by a spare one.
Remote-Reading Guomagnetio Compass APUK-7
1. Switch on the ADIK-7 compass circuit breaker on the circuit breaker
panel of the navigator.
2. Cut in the switch of the )119L1(-7 set on the upper electric board of the
navigator.
3. In 2 - 3 min. after switching on power supply, press the fast slaving
button located on the instrument board of the navigator or left-seat pilot and
1" easethe button after 10 - 15 sec.
4. Check the readings of the main indicator compass course scale with those
Of the magnetic compasses. The difference in the readings must not exceed 100,
the magnetic compass corrections being taken into account.
5. Turn the main indicator magnetic variation scale to make sure that the
pointers of the auxiliary indicators repeat the readings of the main indicator
pointers, the error not exceeding 30.
6. Turn the compass transmitter card with the aid of a permanent magnet to
check the movement of the main and auxiliary indicator pointers, with the fast
sieving button pressed.
7. Release the slaving button and take the magnet away from the trananitter.
8. Check the follow-up rate of the navigator's indicator pointer with the
slaving button not; pressed. The follow-up rate should be within 1 - 4? per minute
9. Cut off the power supply from 2INK-7 compass.
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Air Position Indicator RN-50E_
1. On the circuit breaker panels of the navigator switch on the circuit
breakers of the AM-7 compaanand HM-505 air position indicator.
2. On the upper electric board of the navigator put the )170-7 compass
switch and two nwitches of the 111-505 indicator in ON.
3. In 5 minutes after the line has been energized,-turn the magnetic varia-
tion spur rack of the )111M-7 compass main indicator to make sure that the pointm
If the automatic course device follows the readings of the main indicator; then
cut out the VIIIC-7 compass power supply switch. When this is done, the pointer
of the automatic course device must not Shift. This will indicate to the fact,
that the zero signal has been adjusted correctly. Should the pointer of the
automatic course device shift, use a screw-driver to turn the adjustable resistor
screw located in the distribution box to the left (if viewed from the terminal
blocks, See Pig.21). The screw must be turned until such a position is found at
which movement of the pointer ceases.
4. Set the chart angle on the automatic course device just by 45? less than
the reading of the automatic course deviee pointer. Set the wind speed knob of
the wind setter to zero.
5. Use a K119-3 testing device or a special pressure producer which belongs
to the HM-5O testing installation to create gradually a pressure in the
dynamic system of the T.A.S. transmitter corresponding to a speed of 1150 km/hr.
As the speed changes from 300 to 1150 km/hr check the rotation of the D.R.
computer check indexes, the turning rate of the check indexes should change
smoothly without sharp jumps or binAleg.
6. At a speed of 1150 km/hr change gradually the value of the chart angle
of the T.A.S transmitter from 0 to 360?.
' The turning rate of the D.R. computer check indexes should change gradually.
7. Reduce the pressure in the dynamic system of the T.A.S. transmitter to
zero.
8. Shift the wind speed knob on the wind setter gradually from zero to
division 150 tan/hr. The turning rate of the D.R. computer check indexes should
change smoothly.
9. Change gradually the wind direction on the wind setter from 0 to 3600.
The turning rate of the check indexes should change smoothly.
10. Switch off A.C. and D.C. supply from the 1114-50B air position indicator.
2770 Horizon Ars,2
1. Switch on the power supply of the gyro horizon.
2. Turn the starting handle located on the front of the gyro horizon to
the left. This done, a red bulleye should appear in the zone of the port.
Not later than 3 min, after energizing the instrument, the horizon line
should assume the horizontal position, the permissible deviation being ?10.
Make sure that the skid indicator fluid contains no air bubbles.
Note: With the ambient temperature below zero, the gyro erecting time
Increase up to 6 min.
Electric Turn Indicator 8711-53
may
1. Switch on the power supply of the turn indicators.
2. Wait 2 - 3 Minutes, then press against the edge of the pilot's instru-
ment board to turn it about its vertical axis as far as the shock absorbers
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permit. When this is being done, the moving index of the tura indicator should
deflect from its central position.
5. Make sure that the skid indicator fluid contains no air bubbles.
4. Switch off the power supply from the instrument.
Postflight Inspection
The postflight inspection of the navigation equipment comprises the follow-
ing operations:
1. Exaeine visually the units of the navigation equipment in the same way
as during preflight inspection.
2. Cover the transparent hood of the 1.AK-4B-5 compass transmitter with a
protective casing.
3. Check the colour of the silica gel of the AAlin4B-5 compass transmitter.
If the silica gel crystals have turned pink or brown, replace the dehydrator with
a spare one, since pink or brown silica gel is not capable of absorbing moisture.
The silica gel can be reconditioned by drying, for which purpose it must be
poured on a metal sheet and dried on a moderate fire until ititUrnS blue again.
After the dehydrator is placed on the traeemitter, do not fail_ to open the bele
in the dehydrator bottom.
In addition to the visual inspection of the instruments, find 'out the
causes of the defects which have been revealed in the flight.'Sometimes the
defects and their causes may be found in the course of the check carried out in
the sequence adopted for the preflight inspection. Therefore this inspection
meat be performed immediately after the flight. Sometimes a more careful check
is required. The scope and sequence of this check is described below.
Besides, trouble-shooting is facilitated by the fault finding chart which
contains the most frequent defects of the navigation instruments, their causes
and remedies.
Checking the Instruments for Correspondence
to ''Their Basic Specifications
? ? ? ?
Such a check is to be carried out as soon as you begin to doubt whether
the readings of some instruments are correct, and not less than once every three
months.
Taking into consideration that special installations for checking some
instruments may not always be available under service conditions, the chmking
method has been so worked out as to reduce the number of the instruments to be
removed from the aircraft to the minimum and to carry out the entire check direct-
ly on the aircraft.
When special testing equipment is available it is used for checking the
instruments in accordance with the Instructions of the respective installation
(if available) or in compliance with the given Instructions.
In addition to the method of checking the instruments for correspondence
to their Specifications, this Section contains some special instructions on mount-
ing, care and maintenance of the navigation equipment instruments.
Remote-Reading Gyromagnetic Compass ma-7
1. Disconnect the plug connectors from the transmitter and check, using a
megger, the insolence between the terminals of the plug connectors and the
transmitter body. The insulance must not be below 20 megohms.
2. Atter having slightly tapped against the cover of the transmitter casing,
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note. the reading of the transmitter scale, then use a permanent magnet to
deflect the transmitter card by 100 to the right and take away the magnet. Take
again the readings off the scale of the transmitter. Tne difference between the
last and first readings mill be the card lag.
In the same way check the lag of the transmitter card when the latter is
deflected to the left. The absolute lag value of the transmitter card must not
exceed 7?.
3. Disconnect the double-terminal plug connector from the transmitter.
4, Disconnect the plug connector from the gyro unit and check the insulance
between the following pins of the gyro unit plug:
(a) the insulance between pins 3 and I must be from 100 to 130 ohms;
(b) the insulance between pins I and M must be from 400 to 600 ohms;
(c) the insolence between pins A and E ,E and B, A and B should be from
355 to 586 ohms;
(d) the insularee between pins 3 and I should be from 450 to 580 ohms.
5. Using a megger check the insulance between the following jacks of the
plug located at the end of the wire bundle: A and I , A and 5, A and r ? as
well as between jacks A and the aircraft framework.
The insolence must be not less than 1 megohm.
6.Check the insulance between jacks A and E E and B, A and B of the
gyro ,mit 'plug located at the end of the wire bundle. The insulance should be
equal (accurate within 120 ohms) and at least 100 ohms each.
7. Connect the plug connectors to-the gyro unit and to the transmitter.
8. Supply power to the Aruic-7 compass.
9. Wait 2 and 3 min; and press the fast slaving button. Release the button
after 15 - 20 sec. The readings of the compass course scale of the.main indicator
end the scale readings of the magnetic transmitter should agree within 30, whereas
the readings of the auxiliary indicators should agree with those of the main
indicator also accurate within 3?.
10. Using a permanent magnet turn the transmitter card and check, every
30 - 400, to see that the readings of the compass course scale and those of the
repeaters correspond to the readings of the transmitter scale and maim indicator
pointer respectively.
With the compass operating, oscillation of the main and auxiliary indic6.tor
pointers within 10.5? is permissible.
11. Check the follow-up rate of the navigator's main indicator pointer with
the slave button not pressed. The follow-up rate must be within 1 - 40 per
minute.
25X1
-27-
uswitter are airtight. Check also the sere signal and serviceability of the
1. The set is checked for airtightness as follows:
(a) use a IHY-3 teat set to create a pressure In the T.A.S trensmitter
umic system corresponding to a speed of 700 kn/hr. Pressure drop in the
et-m=1st not exceed 2 km/hr per one minute;
?dome a test -set to create a vacuum in the T.A.S. transmitter
%tic systom corresponding to a speed of 700 km/hr. With pressure supply cut
f, leakage must not exceed 5 km/hr per min.;
2. Ths sera signal and the serviceability of the set are checked in
=dance with the method adopted for preflight inspection.
3, The total error in the set readings is determined at four different
areas selected so that the error may be found by one of the selected courses
tthe intervals from 0 to 500, from 50 to 1800. from 180 to 2700 and from 270
' 360?-
4. Switch on the A.C. and D.C. power supply of the compass and HI-.:50E air
sition indicator.
5. Measure the voltage across terminals Ar, of the distribution box of the
111,50E air position indicator. The voltage in to be measured with a voltmeter
'ring reading corrections within the range of 24 to 30 V. Taking the corrections
to consideration, ensure exactness of voltage readings within =0.1 V.
6. Switch off the power supply from the 111-508 air position inAi,ator
ii the compass. Change over the internal wiring diagram of the indi,ator dis-
lbution box to a voltage of 27 V.
7. Switch on the power supply of the HZ-50E air position indicator and
=pass.
8. Create a pressure in the T.A.S. transmitter dynamic system corresponding
. a speed of 700 km/hr, which is to be checked by the RIC-1200 airspeed
Wieator having corrections for indication errors.
When applying the pressure, take into consideration the corrections for
mospheric pressure given in Table 2.
CAUTION: 1. It is strictly prohibited to use in the junction box a safety
fuse other than type 1E-0.15 A.
Table2
Speed of 700 km/hr with Corrections for Atmospheric
' Pressure
Atmospheric pressure,
mm of mercury .
2. Prior to cutting the compass into the electric mains after some
1
units have been replaced or defects in the aircraft diagram have been
eliminated, do not fail to check the insulance in conformity with
715
- 720
Item 6 of the given Section.
720
- 725
3. Prior to energizing the compass make sure that plug connectors
725
- 730
Nos 8 and 11 of the 00-11p sight computer are not confused to avoid
failure of the compass.
730
735
- 735
- 740
Air Position Indicator HZ-505
740
- 745
745
- 750
Checkinr total error of the set. Prior to checking the set for total
. 750
755
error, make sure that the static and impact pressure lines of the T.A.S.
755
- 760
Q CI-ErGuT1
Airspeed,
km/hr
2
718.7
718.45
714.25
715.0
709,9
707.6
705.4
703.2
701.05
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28 ???
2
760 - 765
765 - 770
770 - 775
775 - 780
780 - 785
785 - 790
695.95
696.8
694.7
692.65
690.6
688.55
9. Place a permanent magnet closely to the compass transmitter and turn the
main indicator magnetic variation spur rack to set the course automatic device
scale to a course equal to 450 or divisible by 10 deg, within an interval of 0
to 90?.
10. Set the mind setter to a wind speed of 60 km/hr and a direction divisible
by 100 or equal to 45?.
11. Set the chart angle on the course automatic device and on the wind setter
to zero.
12. Switch off the D.C. supply from the 51750B air position indicator and
set the D.R. computer pointers to zero. Send the pointers to zero position by
-moving them in a direction opposite to their usual movement (in this case - counter.
clockwise).
13. Switch on the D.C. supply of the 1111-505 air position indicator and
start simultaneously a stopwatch.
14. Wait 8 min. and 34 sec., then switch off the D.C. power supply of the
1111*505 air position indicator and take the readings of north and east pointers
of the D.R.computer.
Notes: 1. In case the voltage across terminals B of the distribution box
is other than 27 ? 0.1 V and if it is impossible to bring it to
this value, multiply the testing time (8 min.34 sec.) by
coefficient K:
K
vdxid
where V - is the voltage measured across terminals B of the
ind
' indicator dibtribution box.
2. During the test- maintain a pressure in the dynamic system of the
course automatic device which corresponds to a speed of 700 km/hr.
15. Using Table 3 determine the rated changes in the D.R. computer pointer
readings for a flight in calm weather.
16. Using Table 4 determine the rated changes in the D.R. computer pointer
readings depending on the direction of the %:ind.
17. Determine the rated changes in the readings of the north pointer LE
and east pointer LE for a flight with drift correction introduced.
LE 7 114 + LIE ;
Values lip 1E, 1' and l'E should be taken with the signs indicated in
Tables 3 and 4.
29 -
25X1
T?ble 3
Rated Changes in Readings of Pointers 1Nand IE Depending on Course in Calm Weather
Course I 0'
10'
_ .
50Yr
40'
45'
70"
80*
hi b?,
I, km.
,-100
0
.t98.5
+17 4
+94
+34.2
+86.6 '
+50
+76,6
+64.3
+70.7
+70.7
+64.3 +50
+76.6 +86.6
+34.2
+94
+17.4
+98.5
-
- - ------
Course
99
103?
110'
120
130'
130'
ifr I 15r
lur
170'
IN km.
LE km.
0
--l00
-17,4
+98.5
-34.2
+94
-50
+86,6
-64,3
+76.6
I -70.7
+70.7
-76.6
+64.3
-86.6
+50
-91
+34.2
98.5
+17.4
1 269'
Course
VW
190*
200*
210'
22! 1
___
225:1- 230"
F240-
250*
iN km.
/E l'"
-100
0
-98.5
- 17,4
-34.2
-86,6
- 50
-s-
-76.6
--64.3
-70.7
-70.7
-64.3
-76.6
-50
-86.6
-34.2
-94
I -17.4
1 --98.5
1 _
Comm.
270' i
280'
310*
315'
320' 330'
340"
I 350'
IN 1...
1E km.
0
-100
+17.4
-98.5
+34.2
-94
' +50
-66.6
+64.3
-76.6
+70.7
-70.7
+76.6
-64,3
-+86.6
-50
+94
-34.2
+98.5
-17.4
1
- change in readings of north pointer
1E - change in readings of east pointer
Table 4
Rated Changes in Readings of Pointers l' and 1' Depending on Wind Direction
Com.
a 10'
20' j
i 49'
45?
1 50"
60"
70'
80'
I'5 km.
1' E km.
+8,6 -4.8.4
0.0 +4.5
A.
-Ed
+2,9
+7.4. "
+4,3.
+6.6
+5,5
+6.1 .
+6.1
+5,5
+6.6
+4.3
+7,4
+2.9
+8
+1.5
+8.4
? Course
93? 100?
110'
lir
130'
lur
160"
Iga
170'
l's km.
0
-1,5
-2,9
- 4,3?. ..
-5.5
-6.1
-6,6
-7.4
-8
-8,4
l'E km.
48.6
+PO
+8
+7.4
+6.6
+6.1
+5.5
+4.3
+2.9
+1.5
- .
Comae
180'
190'
200"
210'
220'
225'
230*
240.
250'
260'
1'N4"?
-8.6
-8,4
-8
-7.4
-6.6
-6,1
-5.5
-4,3
-2.9
-1.5
l'E km.
0
-.-1,5
-24
-4,3
-5,5
-6.1
-6.6
-7.4
-8
-8.4
Co....?
270*
300'
310'
315'
320'
330'
340?
350"
0
'4-4.5
+2,9
+4.3
+5.5
+6,1
+6.6
+7.5
+8
+8.4
PEkm. i -8,6
-8,4
-8
-7.4
-6.6
-6..I
- 5.5 -4.3
-2.9
-1.5
PN - change in readings of north pointer
l'E - change in readings of east pointer
ORNHT
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18. By comparing the actually obtained charges in the D.R. computer read,
bags Ltiff and 10B with the rated values 111. and It determine the absolute errors
in the readings of the D.R. computer north pointer Lag and east pointer ear
4 IS = L'i 111
Alt = *
19. Using the graph presented in Fig.22 determine covered distance D by
the rated indications of the D.R. computer, i.e. IN and L.
For example a change in the readings of the north pointer Z11 is 82 km.
and that of the east pointer lu is 78 km. Day off 82 and 78 km. on the axes
and LE respectively. Pram these points erect perpendiculars to the axes until
they mutually intersect. Lay off the distance from the intersection point to ti
beginning of the coordinates an one of the coordinate axe e (in this example it
is axis log). This distance will determine in the adopted scale the covered dis-
tance L in km. (in the given example the covered distance is 114 km.).
20. Redding use of the graph given in Fig.23 determine the absolute error
by the covered distance.
Per instance, the absolute error of the north computer AIg is 6.15 kat.
and that of the east computer AIt is 4.7 km,
From these points draw lines until they mutually' intersect.
Lay off the distance from the intersection point to the beginning of the
amordinates on one of the axes and determine the absolute error of the set (in
this example the absolute error of the computer AL is 7.75 km.).
21. Determine the total error of the set A from the formula:
?
22. Following the same routine determine the complete error of the eet at
courses within the intervals from 90 to 180?, from 180 to 2700 and from 270
to 360?.
The total errors obtained during the tests should be with in the limits
given in Table 5.
Table
Permissible Amounts of Total Errors
AMbient
air
tempera
tyre,
oc
Total error of HR-50B Bet, %
altitude 0
altitude above
0 to 8000 m.
altitude from
8000 to 12,000 n.
speed
300 km/hr
speed
from 300
to 1000
km/hr
speed
up to
1100
kahr
speed
Aro to
100
kunthr
speed
no to
ii00
km/hr
speed
up to
1200
b-/hr
-
1
2
3
4
5
6
7
20 = 5
.50 1 5
40 = 5
7
9
9
5
8
a
5.5
-
a
7.5
....
9
6.5
-
8
7.5
-
9
25X1
-31-
Should the total error of the set nutted the permissible value, determine
the correction to the Inner diagram change-over table of the indicator distribu-
tion box.
Determination and Account of Correction to Inner Diagram
Change-Over Table of Indicator Distribution Box
"Correction r is the value which is to be algebraically added to the valve
of the air position indicator supply voltage (which is measured across terminalsB
of the distribution box) in order to determine the corrected voltage value and to
change over the distribution box inner diagram correppondingly.
The correction is accounted for by the formula:
Veor . 'supply t AV (volts)
- is the corrected voltage value, for which the distribution box
Vcor
inner diagraa is to be changed over:
supply - is the supply voltage to the air position coordinator, as
measured across the terminals B of the distribution box;
AV - is the correction, volts.
where:
If it is required, depending on the total errors obtained during the check,
to increase the change in the readings of the D.R. computer pointers, take the
correction with the in "-", and if it is required to decrease the value, take
the correction with the sign "+".
The correction ? lue must be divisible by 0.5 V. A correction value equal
to 0.5 V changes the readings of each D.R. computer pointer by 1.85%.
After determining the amount of correction for voltage, it is recommended to
make sure that r.he aelected correction to be introduced is correct, for which
purpose correct algebraically the previously obtained values IOE and L'E (See
Item 18), at which the total error proved to be in excess of the permissible
value, by the value 113.7 AV in %, where AV is the correction (in volts) to be
introduced.
Determine by the new corrected values L'E and L'E the absolute error values
of the north pointer ALE and east pointer ALE (See Item 18) and use them to
calculate the total error of the set (See Items 20. 21 and 22), which will be
obtained after this correction has been introduced.
If the calculated total error meets the requirements, change over the inner
diagram of the distribution box in accordance with the selected correction and
check the set again. ?
In case the set total error exceeds the permissible'value and cannot be
decreased to the value indicated in Table 5 no matter what value AV is taken,
it is required to check each unit of the set separately. The units must be
checked on a YHZ-50 installation only employing the method described in the
Operating Instructions of this installation. On detecting a faulty unit, re.place
it and check the set again for the total error and correction to the change-over
Table of the distribution box inner diagram. In changing over the diagram observe
the instructions which are placed on the inner side of the distribution box
Cover.
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Checking the PrilK-7 and HM-50 Instruments
for Synchresous 2peration
1. Switch on A.C. and D.C. power supply to the 3Fmic-7 compass and HH-505
air position indicator.
2. Walt at least 5 minutes, then adjust the zero signal employing the method
described in Section "Preflight Check".
3. Place a permanent magnet closely to the compass transmitter and rotate
the magnetic variation spur rack of the main indicator to check the readings of
the automatic course device for conformity to those of the compass main indicator
at headings 0, 15, 300, etc., every 150. The readings of the automatic course
device must not differ from those of the main pointer by more than 10. Otherwise,
match the readings by turning the respective deviation screws which are accessible
through the holes in the rear well of the automatic course device (Fig.24).
gyro Horizon A1'E-,2
. The APE-2 gyro horizon is checked on type HP-148 installation ensuring
a turn of the gyro horizon with respect to the three mutually perpendicular
axes: vertical, longitudinal and lateral.
The horizontal base of the turning table should be checked against a level.
In addition to the YUr-48 installation, checking of the AIS-2 gyro horizon
requires the employment of an electric panel whose diagram is presented in
Fig.25.
The check is performed in the following sequence:
1. Place the gyro horizon on the turning table and connect it to the
elen'ric panel.
2. Switch on the power supply to the gyro horizon and start a stopwatch at
the same time.
3. The line of the horizon should assume the horizontal position (accurate
within .110) not later than three minutes after the power supply has been
switched on.
4. In 5 or 6 min. after baying energized the gyro horizon measure the volt-
age between the inverter phases. The voltage should be 36 I 1 V.
If the voltage is other than specified, adjust it by changing the voltage
of the supply inverter, type nu-lo.
5. Watch the miniature airplane with the fixed luaices on the front flange
of the instrument.
G. Turn the casing of the gyro horizon about the longitudinal and lateral
axes to match the horizon line with the miniature airplane.
With the instrument in this position, tap it slightly to make sure that the
slip indicator ball is located between the two central marks made on the slip
indicator; see that there are no air bubbles in the slip indicator fluid. Remove
air bubbles, if any, by turning the instrument casing clockwise about the longi-
tudinal axis.
7. /n 5 minutes after complete erection of the gyro match the vertical mark
on the gyro horizon spherical shield with the zero division on the instrument
bank scale. The gyro horizon error is characterized by the misalignment between
the miniature airplane and the horizon line. This error must not exceed ?10.
8. Turn gradually the gyro horizon casing with respect to the longitudinal
axis until the gyro unit contacts the rest to create a lateral tilt exceeding 80?;
In this ease the tilt of the gyro unit will be in the longitudinal direction.
Then return the instrurent casing to initial position and turn it through 900
25X1
?
with respect to the vertical axis. Thus, the longituainal tilt of the gyro
wit is transferred into a lateral tilt (with respect tc the gyro horizon
casing.). The lateral tilt in this case should be at least 30?. If the tilt is
less than 30?, repeat tilting che gyro as indicated above.
9, Turn the gyro horizon casing through 30? wi.!-h respect to the longitudinal
axis (by the scale available on the turning table) in the same direction in which
the 7.0. A
gyro is
tshe
tilted.
::::nt the horizon line coincides with the miniature airplane
start the stopwatch.
The time from the moment the stopwatch is started to the moment the gyro
returns to its initial position is considered as the gyro erection time from .
the lateral tilt.
11. Check the time required for gyro erection from the opposite lateral tilt
in the same way.
The gyro erection time from lateral tilts must not exceed 4 - 8 min.
12. Check the gyro for erection time from longitudinal tilt. A longitudinal
tilt is established by creating a lateral tilt through 300 and turning sub-
sequently Oa instrument casing through 900 with respect to the vertical axis.
The time required by the gyre to erect from lonaitedinal 30? tilts should
be within 6 - 11 min. The difference in the erection time when the gyro erects
from opposite tilts must not exceed 3 min. ?ihen erecting from a longitudinal
tilt the gyro must not tilt in the opposite direction by more than 30, and when
erecting from a lateral tilt its pitch must not exceed 40.
Erecting_Cutout, Type RH-53PE_
The erecting cutout may be checked on type Yflr-48 installation uled to
check gyro instruments. The check is performed with the aid of an electric panel
whose diagram is presented in Fig.26. The erecting cutout is supplied from the
UP-11 inverter, which should be so adjusted as to produce a linear voltage
, of 36 ? I V, 400 t 10 c.p.s. in 3 - 5 min, after the inverter is energized with
D.C. 27 t 1 V current. The 3 - 5 minute time period is required for placing the
gyro under working load (well-racing gyro , warmed up instrument).
The voltage and frequency are regulated by means of a variable resistor
located in the inverter baseplate.
The erecting cutout is checked as follows:
1. Place the erecting cutout along with the shock absorbers on the turning
table of the Yilr-48 installation and connect it to the electric panel.
2. Set up a turning re for the installation table of 0.2 or 0.30 per sec.
depending on the adjustment of the erecting cutout to be checked. This adjust-
ment value is to be found in the cutout Certificate.
3. Turn switch 2 to supply power and start simultaneously a stopwatch.
Determine the time during which current in phase 1 will drop to 0.5 A. As this
is done, the circuit of button 9 must be open. This time period must not exceed
3 min.
4. In 5 min. after power has been switched on, check the 'current in phase 1.
The current must not exceed 0.45 A.
5. Use selector switch 11 to connect terminals A and E of the electric
panel to terminals 6 and 7 of the erecting cutout. Hake sure, with the aid of an
ohmmeter, connected to terminals A axaB . that the latter are disconnected
through the inner circuit of the erecting cutout.
In ndoonnecting terminals A and B respectively to terminals 8, 9, 10, 11
re
and 12, 13 of the erecting cutout plug connector, terminals A and E should be
short
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. 6. Make the installation table rotate at a rate of 0.2 or 0.3? per sec.
starting the stopwatch simultaneously.
. The time elapsed from the moment the table was started to the moment
terminals A and 3 close if they are connected to terminals .6 and 7 or open if
they are connected to terminals 8 and 9; 10, 11 and 12, 13 of the erecting cut-
out plug connector must be Within 5 to 15 sec.
7. Not earlier than 30 see. {after checking the erecting cutout for its
functioning time as it is rotated in one direction, check its functioning time
when the cutout is turned in the other direction.
The difference between the time found under Item 6 and that found under
Item 7 must not exceed 8 sec.
Resistance Thermometer. Type T73-48
The thermometer indicator is checked for accuracy of operation by cutting
it In a circuit eqpivalent to the resistance thermometer (Fig.27) in which the
transmitter is replaced by any resistance box, that will permit to cut into the
circuit resistance equivalent to the resistance of the transmitter accurate
within 0.1 ohm.
The transmitter resistance values for various temperatures are given in
Table 6.
T ? b 1 e 6
Transmitter Resistance versus Measured
Temperatures
Temperature,
ot
Resistance,
ohms
Temperature,
oc
Resistance,
ohms
-70
68.36
50
108.81"
-60
71.06
60
112.78
-50
73.86
70
116.96
- -40
/6.86
80
121.22
-30
79.90
90
125.56
-20
83.16
100
129.96
-10
86.56
110
134.41
-0
90.26
120
138.96
10
93.76
130
143.36
20
97.36
140
148.36
30
101.06
150
153.26
40
104.86
160
158.26
The error of the instrument is determined by the difference between the
indicator reading and the actual temperature corresponding to the resistance
cut into the circuit.
The indicator error at an ambient temperature of +20 ? 5?C must not
exceed 1.5?G.
25X1
-- 35 -
ENGINE INSTRUMENTS AND GAUGES
GENAL
The set of engine instruments and gauges included: T35-2 and T3-45 tacho-
meters, TBP-11 and TOT-29 thermometers, 3gY-3 pressure gauge and 3I31-3P
engine gauge unit. This section contains also information on type TOT-13 thermo-
meter of air' temperature in the wing de-icing system duct.
Remote-Reading Electric Tachometers
T35-2 and 13-45
The 1.35-2 and T3-45 tachometers are designed for continuous- -ursurement
of the aircraft engine and turbostarter shaft REM respectively. Each of the
instruments is a set consisting of a generator and single-pointer indicator.
Mncifications_
TachometeT;115-2
1. Menge of speed
2. Division value
3.
frau 0 to 5000 r.p.m.
50 r.p.m.
Instrument error should not exceed:
(a) at +20 t 5?C
500 - 3500 r.p.m. ?1% (t50 r.p.m.)
3500 - 4800 r.p.m. (iuclesive1Y) t0,5% (t25 r.p.m.)
4800 - 5000 r.p.m. Il% (1.50 r.p.m.)
(b) at +50 t 5?C
500 - 3500 r.p.m.
3500 - 4800 r.p.m. (nclusively)
4800 - 5000 r.p.m.
=1.6% (180 r.p.m.)
t0.8% (t40 r.p.m.)
t 1.6% (t80 r.p.m.)
(c) at -60 = 5?0
500 - 3500 r.p.m. 12.6% (1130 r.p.m.)
3500 - 4800 r.p.m. (inclusively) =1.3% (=05 r.p.m.)
4800 - 5000 r.p.m. 12.6% (1130 r.p.m.)
1. Range of speed
Tachome
ter T3-45
2. Division value e
3. Error at ambient temperature of +20 = 5?C at
divisions 600, 1000, 2000, 2600 and 3000 '135 r.p.m.
Exhaust Gee Thermometer TBP-11
The TBP-.11 exhaust gas thermometer is intended to measure the mean tempera-
ture of the gases leaving the air-jet engine nozile. It is a thermal electric set
comprising the following units:
indicator TEP-1 1 piece
from 400 to 3500 r.p.m.
50 r.p.m.
transmitter, composed of:
(a) thermocouples T-1
(b) connecting wires
4 pieces
1 set
SF.e14:kT
25X1
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25X1
QVC47:EGTMP
EAMIC Ep2cifications
1. Measurement range . 300 - 900?C
2. Menge of operating temperatures 450 - 750?C
3. The reading error of the set must not exceed:
- (a) at +20 ... 5ct
.450 - 640?0 I15?C
650 _ 75000 =120C
. an remaining portion of scale 12500
(b) at +50 50,0
450 - 750?C
on remaining portion of scale
(c) at -60 5?C
450 - 7500
on remaining portion of scale
4. Resistance of thermometer external circuit at
ambient temperature of -20 t 5?C 2.5 I 0.1 ohms
t18?C
136?0
122?C
t44?C
the indicators and transmitters are interchangeable within one graduation
group. The,conrecting wires are interchangeable as a single set.
Exhaust Gas Thermometer TCT-22
The TOT-29 exhaust gas thermometer -is designed for measuring the tempera-
ture'of the exhaust gases leaving the air-jet engine turbostarter.
The instrument is a thermal electric set comprising the following units:
indicator TOT-2 1 '4ece
thermocouple T-9 1 pi ce
connecting wires 1 eat
Basic ?pEcifications
1. Range of measurement o 900?C
2. Range of working temperatures coo - aoo?d
3. 20AC4,..g error of the set must not exceed:
(a) at +20 = 5?C
600 - 600?C ?20?C
on remaining portion of scale t35?C
(b) at +50 = 5?0
soo - 800?C
on remaining portion of scale
+300c
155oe
.. (0) at -60 = 5?C
600 - 800?C ?40?C
on remaining portion of scale '175?0 ?
Error within 0 - 200?0 not checked
4. Resistance of external circuit at +20 t 5?C 9 t 0.06 ohms
5. Indicator pointer oscillations with the engine
rurAing t1000 max
The indicators, thermocouples and wire set are interchangeable.
25X1
-37-
Thermoelectric Thermometer T1T-13
The TUT-13 thermometer is employed for nuasuring the air temperature in
the wing de-icing systen duct. The instrument est is composed of the following
units:
- thermocouple
indicator
compensating wires
1 piece
1 piece
1 set
Basic Deeifications_
1. Range of measurement from -50 to +350?C
2. Reading error of set must not exceed:
(a) at +20 t 50c
loo - 260?0 1800
on remaining portion of scale ?1600
(b) at +50 - 500
100 - 260?0 113?0
on remaining portion of scale 126?C
(c) at -60 t 5?C
100 - 260?C +18?C
on remaining portion of scale t3800
3. Resistance of thermometer external circuit 7.15 t 0.05 ohms
The indicator and set of compensating wires are interchangeable.
Three-Pointer Electric Engine
Gauge Gap ..3/113P
The 8M8-3P engine gauge unit is used for remote check of jet engine
operation. The purpose of the 90-31) gauge unit is to check:
(a) oil pressure in engine oil line;
(b) fuel pressure In idling rating manifold:
(c) oil temperature at engine inlet.
The SUB-32 set (Fig.28) consists of the following units:
- oil pressure pick-up unit 1 piece
- fuel pressure pick-up unit 1 piece
temperature pick-up unit 1 piece
- electric remote indicator 1 piece
The indicator comprises three metering gauges In one housing, each of which
constitutes along with its pick-up unit an independent metering circuit.
1. Power supply
2. Range of measurement:
oil pressure gauge from 0 to 10 kg/aq.cm.
fuel pressure gauge from 0 to 100 kg/sq.cm.
oil thermometer from 50 to +11500
3: Maximum reading errors at ambient air temperature
of +20 t 5?0:
(a) oil pressure gauge at divisions 0, 2, 4, 6
? and 8 10.4 kg/sq.crn
at division 10 t0.6 kgisq.cm.
Basic lip2cifications
27 V t 10%
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SECRET
(b) fuel pressure gauge at divisions.10. 20.
60. 80 and 90 13 Xg/sq.cm.
at divisions 0 and 100 15 kg/sq.cm.
40,
(c) oil thermometer at divisions
and 130
at divisions
-40, 50, 0, 100
tea
-50 and +150
..a. Permissible overload;
for oil pressure gauge pick-up unit
for fuel pressure gauge pick-up unit
15 kg/aq.ca.
120 kg/sq.co.
The units of the 30-32 set are interchangeable.
Remote-Reading Electric Pressure Gauge 97317-3
The 93317-3 pressure gauge is used to check the fuel pressure before the
bigDevressure fuel pumps. It is a set comprising a. pressure pick-up unit and
? remote-reading electric indicator.
Basic Decifications
1. Supply voltage
2. Range of measurement
' 3. Working portion
4. Maximum error on working portion at ambient tempera-
ture of +20 I 5?0-
5. Permissible overload
The units of the 8701-3 Set are interchangeable.
25X1
RemoteReadimg Electric Tachometers T35-2
and T3-45
Check the tachometers in service on a special tachometer installation at
east once every 6 months. The check consists in comparing the readings of the
odometer under test with those of a reference tachnmeter. The reading error
X the tachometer must not exceed the values given in the "Basic Specifications"
or the T35-2 and T345 tachometers.
Thermometer THr-il_
The connecting wires are checked for resistance as follows:
1. Disconnect the plug connector from the ineicator.
? 2. Using additional wires connect the terminals of the disconnected plug
onnector to terminals Rx of the YMB-49 electric bridge or some other bridge
asurin-g measurement of resistance within the range of 0 to 10-ohms with an
madness of =0.01 ohm.
3. Determine the resistance Rtatal of the wires connected to terminals
of the bridge.
4. Determine the resistance Redd of the wires by means of which the plug
27 = 2.7 V onnector terminals have been connected to the terminals Rx of the bridge.
0 3 waq.m. 5. Determine the resistance of the thermocouple connecting wires from the
0.6 -2.4kg/nor:Ida:
Moire Btotai Radd
4% from measure
?nt limit here Reim - resistance of thermocouple connecting wires.
up to 4.5 kg/4, The resistance of the IBP-11 thermometer connecting wires at ambient
emperatere of +20 = 500 should he 2.5 = 0.1 ohms. If this resistance is other
ban specified, it must be adjusted by changing the value of the series resistor
laced inside the socket of the indicator plug connector.
The indicator is checked for reading errors as follows:
Out the indicator into an electric circuit equivalent to the thermoelectric
0aennometer and employing a dry cell with a potentiometer as a source of electro-
iotive force (See Fig.29).
MILIRTEMANCE ISTRUCTIODS
In service the engine intruments and gauges dhould be checked before
flight in compliance with the methods described below, as well as in cases
specially specified.
',tonight Inspection The voltage fed to the indicator terminals, is measured with the aid of
The preflight inspectionof the engine instruments and gauges is confined , millivoltmeter of an accuracy class not less than 1.0. When performing this
.to visual examination of the units belonging to the sets of the instruments and beck, it is well to bear in mind that the resistance of the wires connecting
gauges. he reference millivoltmeter to the indicator under check depends on the ambient
During visual examination of the instruments make sure whether they axe ademperature as well as on indicator scale division to be checked. This resistance
damaged on the outside, that they are reliably secured to the instrument bear&wet correspond to the values indicated in Table 7.
)r respective attaahment brackets. See also that the plug connectors are well
:onnected and the wires are properly terminated. Check to see that the wiles
are not broken at places of termination, and that the compensating wires mats
a reliable connection at their joints, etc.
Oheckinenthe Instruments for Correspondence
to Their Basic Specifications
Table7
Wire Resistance as Function of Temperature
and Indicator Scale Division to Be
Checked
This check is to be carried out as soon as doubt arise:JAB to the correct -Scale division
mess of operation of some instruments, and at least once every three months. to be checked
Resistance of external circuits, ohms,
, at ambient temperatures of
In addition to the description of this check, the section contains a Table
dealing with the most frequent faults.their causes aud remedies.
.20%
.50oc
-60oc
2
3
2.6
2.6
2.4
QM'CITTGvP
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SECRET ?
1
2
_3
4
400
2.7
2.7
.,
2.5
500
2.7
2.7
2.5
600
2.6
2.8
2.6
700
2.8
2.8
2.6
800
2.8
2.8
2.6
900
2.9
2.9
2.7
The value of the electromotive force applied to the indicator terminals
depends on the scale division under check and on the calibration of the therm.
couple. This value is to be found in Table 8.
Table 8
Electromotive Force as Function of Indicator
Division under Check and of Thermocouple Calibration
No.
Calibre-
tion
index
Value of electromotive force in AV for temperatures of
300?C
406?C
500?C
600?C
700?C
aoo?c
960c
2
3
4
5
6
7
8
9
1
0
-
-
7.16
13.48
20.0
26.64
33.76
2
P.
0.96
4.16
10.24
17.88
25.08
33.04
41.08
3
T
-
5.46
11.31
17.76
24.26
.30.92
37.52
4
M
-
3.36
8.92
15.32
21.84
28.52
35.2
5
I
-
3.68
9.32
15.56
22.08
28.56
34.92
6
N
1.44
6.56
13.68
21.72
30.04
38.04
46.2
7
A
1.52
6.6
13.76
21.48
29.56
37.68
45.48
8
E
-
6.48
13.08
20.72
28.92
36.96
44.58
9
6
-
6.04
13.4
21.08
29.16
37.28
45.44
10
B
1.4
6.36
13.36
21.08
29.16
37.28
45.48
11
r
1.32
6.16
12.96
20.68
28.76
36.88
44.8
12
1
1.68
6.68
13.92
21./2
29.84
37.92
45.96
13
2
1.52
6.4
13.64
21.44
29.56
37.64
45.68'
14
3
1.36
6,12
13.36
21.15
29.28
37.36
45.4
15
K
1.84
6.72
14.12
22.16
30.64
38.92
47.12
16
N
6.12
13.36
21.32
29.56
37.8
45.96
25X1
-41-
Table 9
Permissible Reading Error of mr-ii Thermometer
Ambient
Indicator error, ?C
temperature,
oc
450 - 640?C
650 - 750?0
Non-working
range
+20 t 5
?10
17
?18
+50 = 5
113
=13
1'26
-60 t 5
115
115
=30
CAUTION :The TBF-11 are manufactured with transmitters (thermocouples) of
various calibration. Each calibration group has its own thermal electro-
motive force which differs from that of the other groups.
To distinguish one calibration group from another, each of them is
given its own index indicated on the scale, thermocouple cap as yell As
in the Certificates.
The indicators and transmitters are mutually interchangeable in one
and the same calibration group only, except groups S and B and groups A
and 2 which are mutually interchangeable.
In performing the check see to it that the indicator is kept for
at least 2 hours at the temperature at which the check is to be performed.
Thermometer TCP-29
The resistance of the connecting wires is checked as follows:
1. Disconnect the plug connector from the indicator.
2. Connect, with the aid of additional wires, the terminals of the dis-
connected plug connector to terminals Rx of the electric bridge, type YMB-45,
used for measuring the resistance or some other bridge ensuring measurements
accurate within 0.01 ohm.
3. Determine the resistance R
total of the wires connected to the terminals
Rx of the bridge.
4. Determine the resistance Rada of the 'fires by means of which the terminals
.of the plug connector are connected to the terminals Rx of the bridge.
5. Determine the resistance of the thermocouple connecting wires from the
'formula: ?
Rewire = Rtotal- Ram
The value of .he reading error is determined as the difference between the where Rwire - resistance of thermocouple connecting wires.
,indicator reading and the actual temperature value at the given electromotive ?
force. The indicator error must not exceed the values presented in Table 9. The resistance of the connecting wires at ambient temperature of +20 t 500
,should be 9 t 0.06 ohms.
The indicator error is checked as follows:
Cut the indicator into an electric circuit equivalent to the thermoelectric
thermometer and employing a dry cell with a potentiometer as a source of electro-
motive force (See Fig.30). The voltage applied to the indicator terminals is
measured with a millivoltmeter of an accuracy class net below 1.0. The resist-
lance of the wires connecting the reference millivoltmeter with the indicator under
Icheck must be equal to 9 = 0.6 ohms.
The value of the electromotive force applied to the indicator terminals
qM1(117GIT
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depends on the indicator scale division to be checked and on the ambient tempera-
ture at which the check is performed. This value is to be found in Table 10.
The error value is determined as the difference between the indicator read-
ing and the actual temperature at the given electromotive force. The inAirator
reading error mast not exceed t1200 within a reading range of 600 to 800?C and
t2700 on the rereieing portion of the indicator scale.
CAUTION: In performing the check see to it that the indicator is kept for
at leant 2 hours at the temperature at which the check is to be performed.
Thermoelectric Thermometer /4T-13
The resistance of the connecting wires is checked as follows:
1. Disconnect the plug connector from the icicator. ?
2. Connect, with the aid of additional wires, the terminals of the disconnect- el
ed plug connector to the terminals.R, of the electric bridge, type 70-49 , used :
for measuring the resistance or enrotber bridge ensuring a measurement accuracy E.
up to 0.01 ohm. ?
3. Determine the resistance Rtatal of the mires connected to the terminals
of the bridge.
4. Determine the resistance Radd of the mires, by means of which the plug
connector terminals have been connected to the terminals Rx of the bridge.
5. Determine the resistance of the thermocouple connecting wires from the
formula:
%ire e Noted - Radd
where Noire - resistance of thermocouple connecting wires.
The resistance of these wires at an ambient temperature of +20 t 5?0 should
' be 7.15 ? 0.05 ohms.
The indicator error is checked as follows: cut the indicator into an
electric circuit equivalent to the thermoelectric thermometer and employing a dry
cell with a potentiometer (Fig.50) as a source of electromotive force. The volt-
age of the potentiometer must be fed to a reference millivoltmeter whose accuracy
class is not below 1.0. The indicator to be checked is connected to the terminals
of the reference millivoltmeter through wires whose resistance is 7.15 t 0.05 ohms.
The value of the electromotive force which is determined by the mirivolt-
meter depends on the scale division to be checked, as well as on the ambient
temperature. This value can be found in Table 11.
The error value is determined as the difference between the indicator read-
ing and the actual temperature value at the given electromotive force. The
Indicator reading error must not exceed t5?C within the range of 100 to 260?C
'and =1000 an the remaining portion of the scale.
CADT/ON: In performing the check see to it that the indicator is kept for
? at least two hours at the temperature at which the check is to be
performed.
25X1
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4
45 -
25X1
25X1
Three-Pointer Electric Engine Gauge Fuel Pressure Gauge
Rait- Z!4.7.3.? _ ?
The fuel pressure gauge is checked in the same way as the oil pressure
The purpose of the check is to:
me reading r the fuel pessure gu
(1) find the error of the instru 5?Cl
ment at an ambient temperature of +20 t 'geuge. eror of rage at divisions 10. 20. 40, 60,
(2) find the inatlan00 of the electrical elements of the pick-up units eo and 90 kesq.cia. must not exceed =5 kg/s4.cm.
Oil Thermometer
(tranamitters) and indicator;
(3) determine the airtightness of the pressure gauge pick-up unit in; In order to perform the check, cut the temperature indicator into the sir-
(4) determine haw tilts of the indicator affect its readings: cuit shown in Pig.34. Set resistance on the resistance box which would correspond
(5) determine the resistance Of the boat-sensitive element ef the Pick-up to the indicator scale division to te checked. The resistance values are to be
unit at 0?0 end at +100?C.
1. Wire bundles for interconnecting the units belonging to the BUZ-3P set found in the respective table.
tare indicator and the temperature value for which the resistance has been select-
The following equipment is required for performing the check: The error is determined as the difference between the reading of the temners-
ed in the resistance box. The error must not exceed WC at divisions -40, 0, 50.
AD0 and 130?C, and .16?0 at divisions -50 and +150?0.
The resistance of the heat-sensitive element of the oil thermometer is
determined with the aid of an YUB wheatstone bridge or any other instrument
meamwermtiwhich will ensure a measurement accuracy within 1.0.2% . Submerge first the
temperature pick-up unit into a vessel with thawing ice measure the resistance
. of its sensitive element, then submerge it into boiling water and measure the
iresistance again. The resistance is to be measured not earlier than in 5 min.
later the pick-up unit was submerged into the respective medium. During the check
Ithe entire thin cylindrical portion of the pick-up it must be sUbserged in the
medium: under check.
The resistance of the pick-up it sensitive element mmet correspond to the
! resistance indicated in the Certificate of the given pick-up unit.
CAUTION: 1. Pressure must be 'supplied to the fuel pressure gauge pick-up
unit through a plate damper only. /nobservance of this condition leads
to premature-failure of the pick-up unit.
2. Prior to installing on the aircraft a damper that was already
in use, it must be checked as follows:
(a) =meat it to the compressed air system and apply a pressure
of 2 -3 atm.;
(b) if the air passes through the damper, the latter may be installed
on the aircraft.
in accordance with the diagrams exert in Figs 31, 32 and 33.
2. Reference pressure gauges up to 15 and up to 150 kg/sq.cm.
3. Beadstance box. type BU0 . or any other box anauring selection of
resistances socurato ithin 0.1 obis.
4s 'tonsure feud cocks.
5. Wheatstone bridge of 3113 type or any other bridge, ensuring
of the resistances accurate mithin 0.2%.
6. Wagger with a voltage of 500 V across the feelers.
7. Mercury pressure gauge rated for 1 kg/sq.cs.
8. Source of pressure up to 120 kg/sq.cm.
9. Source of direct current, 27 M.
10. Fittisgs (T-pleces pipes, etc.).
Oil IXessure_Gauge_
Time oil pressure gauge is checked as follows:
1. Assemble the BUK-3P set in accordance with the diagram given In
2 Malting use of the cocks create a pressure in the oil pressure gauge
AYstem of 0, 2, 4, 6, 8 and 10 kg/sq.ca. consecutively (the pressure is to be
checked by a reference pressure gauge).
3. Keep the system under a maxis= pressure of 10 kg/sq.cm. for 15 min.
4. Reduce the pressure in the system consecutively the reversed order.
5. The reading error is determined as the difference between the readings
of the reference gauge and those of the gauge under check.
Before taking the reading tap slightly against thu indicator casing and
-the casing of the respective pressure pick-up unit.
The influen-e of inclination of indicator 2 on the readings of the set ia
checked simultaneously with the check for reading errors. With the indicator
inclined through 900 to the right or left, the error must not emceed t0.4 kg/SgA
at divisions 0, 2, 4, 6 end 8 kg/sq.cm. and =0.6 kg/sq.cm. at division 10 kg/sqJ
In order to check the casing of the pressure gauge pick-up trait for air-
tightness, assemble the set in accordance with the diagram in Fig.35.
Open the inlet cock and create a pressure of 850 mm of mercury in the pick-
up unit casing and. in the sensitive element simultaneously. Close the inlet cock
aM watch the mercury level during one minute. Drop of the mercury level for ar
minute must not exceed 8 EH.
The insulance of the current-carrying elements of the pick-up units and
indicators with respect to their casings should be at least 20 megoines at an
alibient temperature of +20 t 5?0 and relative humidity of 30 to 80%.
SECRET
Remote-Reading Electrical Pressure
Gauge 3A1Y-3
The BINV-3 pressure gauge is checked In the same way as the oil pressure
gauge of the BUB-3P set. The wire bundle diagras for checking the 94W-3 oil
pressure gauge ia given In Fig.31. The error of the 34=-3 sat within the
meamamment range of 0.6 to 2.4 kg/sq.cm. must not exceed =4% iron the rated
value of the scale.
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TROUBLES AND RU-TDIES
25X1
25X1
-47-
2
-considerably from instru-
readings
1
2
3
TroUble Probable cause Remedy zient
Tachometers T3-45 and T35-2
With the melee running
the instrument pointer
will not start from zero
Indicator readings too
low
Indicator pointer
stands against zero divi-
sion when temperature in
inner cone differs
considerably from instru-
ment reading
Indicator pointer de-
flects to left from
zero
zedicator pointer
dances
Indicator reading too
low as compared with
actual temperature in
inner cons
Indicator reading
too high as compared
with actual tempera--
fuze in inner cone
Gauge pointer stands
against zero division
when temperature in
inner cone differs
. (a) Broken wires between
generator and indicator
(b) Wires from genera-
tor to indicator short-cir-
cuited
Faulty indicator
Thermometer TBP-11
(a). Thermocouple or con-
necting wires open-cir-
cuited
(h) Connecting wires
contact each other
(c) Thermocouple ends in
junction box are connected
in opposition
(d) Faulty indicator
Polarity in Junction box
or indicator plug connect-
or reversed
Poor contact in connect-
ing wires
(a) Wires of one or se-
veral thermocouples clos-
ed
(b) Polarity of one
thermocouple reversed
(c) Indicator circuit
faulty
(d) Poor contact in con-
necting wires
Faulty indicator
Thermometer TCT-29
(a) Connecting wires
open-circuited
(b) Indicator plug
connector open-cireuited
Instrument readings
Find faulty Place and
too low
remedy the wires
Find and remedy fault
Replace indicator
Find and remedy fault
Instrument readings
too high
Instrument readings
not stable
With thermocouple cut
in, instrument indicator
Find and remedy fault will not operate
Rearrange the thermo-
couple ends in accordane
with the diagram given
in the description
Replace indicator
Connect wires in accoM
Instrument readings
not stable
Instrument readings
ance with diagram given too low
in description
Find and remedy fault
Find and eliminate
fault
Connect wires in accord
ance with diagram
- Replace indicator
Find and eliminate
fault
Replace indicator
Find and eliminate
fault
Find and eliminate
fault
SECRET
Instrument readings
too high
When power supply is
switched on, pointer
remains in position OFF
At no pressure pointer
Shows 70
(c) Thermocouple open,
circuited
(d) Faulty indicator
(e) Connecting wires
sheateeirculted
(a) Faulty indicator
(b) Floor contact in con-
necting wires
Faulty indicator
Defective contact in
places of connecting
wire joints
Thermometer TAT-13
(a) Compensatory wires
open-circuited
(b) Faulty indicator
(c) Connecting mires
contact each other
(a) Defective contact
In places of connecting
wire joints
(b) Faulty indicator
(a) Poor contact in
places oi connecting
wire joints
(b) Faulty indicator
Faulty indicator
3
Replace thermocouple
Replace indicator
Find and eliminate
fault
Replace indicator
Find and eliminate
fault
Replace indicator
Find and eliminate
fault
Engine Gduge Unit 3M11-3P
Find and eliminate
fault
Replace indicator
Find and eliminate
fault
Find and eliminate
fault
Replace indicator
Find and eliminate
fault
Replace indicator
Replace iedicator
Fuel Pressure Gauge
(a)-No power in mains
(h) Faulty contact in
plug connector terminals
(e) Flexible wire in
pressure pick-up unit
broken
(d) Faulty brunt' contact
in pressure pick-up unit
Reversed polarity
Check power supply
line and remedy it if
it is broken
Remedy plug connectors
Replace pressure pick-
up unit
Replace pressure pick-
up unit
Reverse polarity
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-48-
1
2
3
At no pressure pointer
will not come to stand
? against were division
At no preesuse pointer
will not stand against
division 6 kg/sq.cm.
When pressure is in-
creased up to
100 kessq.cme pointer
stops against division
40 kg/sq.cm. and returns
to 6 kg/sq.cm.
At no pressure pointer
stops against division
190 kg/sq.en.
When pressure is in,
creased,- pointer goes to
division 60 kg/sq.cm. and
returns then to
100 kg/sq.cm.
At no pressure pointer
stands below zero divi-
sion
When pressure is in-
creased pointer moves
between divisions 0 and
100 outside the ficale
At no pressure pointer
stands at division
30 kg/sq.cm.
When pressure is in-
creased, pointer Shifts
to 70 kg/sq.cm. and then
returns to 30 kg/sq.cm.
When power supply is
switched on pointer
will not start from low
division, but this happens
when indicator is shaken
With power supply switch,
ed on pointer is pressed to
lower rest end will not
(a) Pickup emit paten,
tiometer turns shorted
(b) Membrane swollen
Wire running to plug
connector terminal 1
broken
Faulty contact in plug
connector terminals I
Wire naming to plug
connector terminal B
broken
Faulty contact in plug
connector terminals B
'lire meshing to plug
connector terminal r
broken
Faulty contact in plug
connector terminals r
Wire running to plug
connector terminal A -
broken
Faulty contact in plug
connector terminals A
21.1 Pressure_Gause_
(a) No power in supply
line
(b) Brush contact in
pressure pick-up unit
broken
(a) Reversed polarity
(b) Wire. running to
ping connector terminal
B broken .
25X1
25X1
1
2
Replace pressure pick nose away from it eh=
up unit indicator is shaken
Replace pressure piey
up unit 1 with power MAT switch-
Find and remedy fault ee oe, pointer is pressed
ito upper rest
In checking instrument
Find, and remedy ..auls for error corrections
iare beyond permissible
!limits
Find and eliminate
fault
Find and eliminate
fault
Find and eliminate
fault
Find and remedy fault
When power SSPPIY is
switched an pointer
sill not move away from
lower rest
Pointer leaves lower
rest when indicator is
shaken
With power supply switch-
ed on, pointer is pressed
to lower rest
With power supply switch-
ed on, pointer is pressed
to upper rest
Find and correct feel
Find and connect fast
Check power supply it
and remedy it if it is
broken
Replace pressure pich
up unit
Reverse polarity
Find and correet fauD
(c) Faulty contact in
plug connector terminals
Broken wire or faulty
contact-in plug connec-
tors
(a) Wrong adjustment of
pressure pick-up unit
(b) Amount of indicator
adjustment resistances
has changed
Oil Thermometer
No power in supply
line
Faulty contact in
indicator plug connect-
or
Broken wire or faulty
contact in plug connector
(a) Sensitive element
broken
(b) Broken vire
Find and correct fault
Find and correct fault
Replace pressure pick-
up unit
Replace indicator
Check Power supply line
and remedy it if it is
broken
Find and correct fault
Find and remedy fault
Replace temper..ture pick-
up unit
Find and remedy fault
ELDIIILITION OF COLWASS DEVIATION ON IESTRIZESTS
APIK-7 01-5 NO8 1 AND 2 AND 1E-12
Checking j1fFe7 Cogass for $1.7-fihronoue gperation with
111-55 Air Position Indicator
In case faulty compasses are replaced by new one, as well as when wrong read-
ings have been discovered in the flight or in cane of compass misalignment, check
the compass for operation on the ground and calibrate the compass.
Ground swinging is performed alit? When replacings
(a) engines;
(b) 2NA-7 compass transmitter;
(u) 711 navigator's indicators of 11rus-7 compass;
(d) frames of A.P8-3 compass Nos 1 and 2.
Compass calibration on the aircraft is performed in order to determine and
correct semi-circular deviations and to determine or compensate the residual
deviation.
The automatic course device of the 1111-501 air position indicator is checked
fss synchronous operation with the navigator's main indicator 92 of the A1'MA-7
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SP.CRET
compass in order to determine the mismatehing angle. If this angle exceeds 10,
Be the deviation adjusting acre= to synchronize the pointer readings. 'soul
leg
1. Before bringing the aircraft deviation correcting ground, cheek'
General
the operation of the instruments. which are to be used during elimination of
d4viations.
2. Operations related to elimination of deviations are to be performed on,
special ground located at least 300 m. away from steel structures, underground
power cables, metal tubes, buildings, H.T. lines, forests and other objects
causing a change in the earth magnetic field.
Lion should not use any tools made of ferromageetic materials (steel screw- Correction of compass deviation consists of
3. The attending personnel who takes pest in operations on correcting des;
drivers. flat pliers, etc.). (1) elimination of permanent and semi-circular deviation by means of the
4. Deviations are to be eliminated with the engines stopped. timbal/on instrument of the UAA-3 transmitter (first stage);
5, HH-12 and 1HW-7-compasses are to be calibrated sismaltaneounly (both--(2) elimination of quarter deviation and remote-reading errors of the
without fail). 3 transmitter with the aid of a mechanical compensator (profiling device)
compasses met be checked e located in the navigator's indicator (second stage).
Calibrat ADX-5 compasses Nom 1 and 2 separately.
6. ihnsply the electric mains of the aircraft from a ground source of posn
through a special plug connector. The mains voltage is 27.5 - 28 V D.C. and
115 0.5 V LC., 400 c.p.s. (through airborne inverter 1104500 ).
7. Before starting to correct deviation do the followings
(a) switch, on the CRY-10 interphone set;
(b) switch on the 3711-53 turn indicator;
(c) switch on two Ar6-2 gyro horizons (the
on the right-hand board and one of the APB-2 gyro
left-hand board;
(d) switch on the flux-52 directional gyro;
(e) unlock the controls, place the pedals and
position; ?
Cr) set the P1311-4 indicator of the navigator in stowed position;
(g) switch on the ill-5-211 autopilot;
(10 set up the clock (working);
(i) place the armament and sight post in stowed position;
(J) set the Cd1B-11p sight.
. 8. In addition to the above-said follow the inetructions given in the
dgscriptions er the 1/1,5 ?aux-7 and IU-50E instruments.
25X1
d describe the circumference marked off on the deviationgmound, the right
being inside the circurference.
4. The aircraft is turned by a towing truck with the aid of a drawbar.
Lizq: In case the truck and drawbar influence the readings of the compasses
they should be taken away from the aircraft each time the-aircraft
is turned to a new heading.
Blinination of_Deviation_of 1M0-7 Remote-
Reading Gyro-Magnetic CI:Inman
Elimination of 'Permanent and S?a
Circular De?iation
Permanent deviation is corrected as follows:
(a) switch on power supply to the compass;
(b) in two-three minutes after switching on power supply to the compass
Proceed to correction of permanent deviation for which purpose place the air-
?TE-2 gyro horizonmounted craft at magnetic courses of 0, 90, 180 and 270? in
on tha turn;
(e) at each of the magnetic courses determine tbe deviation as the difference
.in tits readings of the magnetic course of the aircraft and the compass course
control wheels in neutral male of the navigator's indicator.
Each time before taking the readings, align the navigator's indicator and
ithe transmitter by pressing the slaving button. Keep the slaving button pressed
not lase than 15 sec.
The algebraic deviation sum at all four courses divided by four will produce
ithe setting error.
If permanent deviation exceeds 2e, it should be eliminated by turning the
1112-3 transmitter, for which purpose ease off the screw and turn the transmit-
Iter casing in the ring with respect to the base through an angle equal to that
ssf the permanent deviation. The casing turning angle is counted by the scale
available on the ring.
Semi-circular deviation is corrected by permanent magnets of the deviation
instrument at all four magnetic courses (0, 90, 180 and 2700), The semi-circular
deviation at the Given four magnetic courses is determined in the same way as
when correcting permanent deviation. Here we usually encounter two cases:
ytrst Note: TO bead the aircraft it is advisable to use the method of "tail" case :::e.ailr:r:of:rsa:ss0er:?dar.ttst itIntietie:ItecnI:=Poi:ceex=nelt0)0; 1-nStts
direction finding at a distance of 150 - 200 m. In this case the M zero the deviation; at magnetic course 1800 turn the same extension piece to
direction finder must be placed at the above distance from the halve the deviation value; at mgnetie course 900 tura the extension piece
2. The sequence of heading an aircraft is izimetted=tat:lat magnetic course 2700 turn the same extension
rear cabin of the aircraft on a tripod.
given in the description of the
e Second case. At courses 0 and 900 the initial deviation is less than 100.
along with the Servic log.
deviation direction finder. This description is supplied with the instrument
3. When the aircraft is beaded the L.H. wheel of the main landing gear In this case determine and record the deviation at magnetic_courses.0 and 900.
At aestetie course 1800 bring the deviation value to 1/41-'0 el810) . At magnetic
COU 270? bring the deviation value to 2
(O90 * 270)
Installation of Direction Pinder on Aircraft and
Position of Aircraft When Turning to Assume
Required Headings
1. The aircraft is headed with the aid of a AO direction finder.
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Elimination of Total quarter Devia-
tion and Remote-Reading Error
The total error is corrected as follows:
(a) place the aircraft at one of the 24 courses (0, 15, 30, 450, etc.), at
which the compass error is the least;
(b) align the indicator with the transmitter by pressing the slaving butts;
(c) making use of the handle of the navigator's indicator, place the maga.
tic variation scale to zero, setting index S against zero;
(d) turn off the adjusting wrench and remove the pad to give access to the
24 adjusting screws:
(e) eliminate the compass error at the given course by turning that screw
at which the pointer of the navigator's indicator indicates. The screw is turned
with the aid of the adjusting wrench.
Example. The aircraft is placed exactly at 180?, the slaving button is
pressed, the indicator pointer shows 181?. Turn the screw at which the pointer
indicates to place the pointer exactly against division 180?, and the compass
error at the given point mill be eliminated.
The compass errors at the other points are eliminated after each turn oft;
aircraft through 150 in the same 'sequence as indicated above.
Atter elimination of deviation, the instrument pad must be put in place
and the adjusting screw turned into its seat.
Simultaneously with residual error elimination and making charts of the
navigator's indicator residual errors, make also charts of the repeaters resift
errors.
Notes: 1. When eliminating deviation, with the engines stopped, take the
indicator reading only after correcting the lag of the magnetic
transmitter which is achieved by tapping the instrument on the
casing.
2. The maximum error which can be eliminated with the help of any
Adjusting screw is 80.
Elimination of Deviation of 1E-12 Magnetic Compass
Prior to eliminating the deviation of the KR-12 magnetic compass, do net
fail to switch off the fans and the glass electric heaters. Elimination of
deviation of the EM-12 compass consists in determining and, correcting the
setting error, eliminating the semi-circular deviation and in determining. and
correcting the residual deviation.
Determination and -Elimination of
Setting . Error
(a) Place the aircraft at the four 'main magnetic courses (0, 90, 180 and
2700) and calculate the setting error as the algebraic sum of the four. devia-
tion readings divided by 4.
(h) Correct the setting error of the navigator and pilot d compasses by tam
log the brackets through the value of the setting error. Turn the bracket to Cr
left in case of a plus error and to the left in case of a minus error.
Elimination of Semi-Circular
Deviation
(a) Placa the aircraft at zero magnetic course. Use magnet N - 8 to make
the compass read zero.
SP.M.VT
25X1
25X1'
-53-
(b) Place the aircraft at magnetic course 1800. Use magn-A N - S to halve
the deviation
(Go c180)
2
(c) Place the aircraft at magnetic course 900. Use the N - W Magnet to set
the c Place the aircraft at magnetic course 270? and use the E - W magnet
t 9Q0
ma
to halve the deviation:
(090 + 0270)
2
Determination and Elimination
of Residual Deviation
Residual deviation is determined and corrected at eight points: 0, 45, 90,
135, 180, 225, 270 and 315?.
Correct the deviation charts. Residual deviation must not exceed ?50.
Elimination of Radio Deviation of Al-
Radio Compasses Nos 1 and 2
General
1. The azimuth rings of the APE-5 compass No.1 have black-painted numerals,
whereas those of APE-5 compass No.2 are painted red. On some aircraft the
numerals on the rings of both compasses are painted black.
2. The frames of LPE-5 compasses Nos 1 and 2 should be compensated In
accordance with Tables 12 and 13.
Note: The difference in the tables is a result of different installation
of the frame of AFK-5 compass No.2.
Table 12
Standard Correction Compensating Angles
at Eliminating Radio Deviation of APE-5
Compasses Nos 1 and 2
Radio
station
Averaged SP for
comptnsation, deg.
Radio
station
course
angle,
deg.
Averaged aP for
compensation,cleg.
course
angle,
deg.
AF-5 No.1
APE-5 No.2
APE-5 No.1
AP-5 No.2
0
0
0
180
0
0
15
+12
+6
195
+11
+6
30
+18
+11
210
+16
+12
45
+19
+14
225
+16
+15
60
+16
+14
240
+13
+16
75
+10
+10
255
+7
+10
90
+3
41
270
0
+1
? 105
-5
-6
285
-7
-6
L20
-9
-10
300
-13
-10
135
-la
-12
315
-16
-11
150
,13
-9
330
-16
-9
165
-10
-6
345
-12
-5
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Table 13
Standard Correction Compensating Angles
at allminatirmLiadio Deviation of APW+5
compasses Nos 1 and 2
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2bX1
*55*
id) residual radio deviation must not exceed '130. including the frame
setting error.
Note: At course angles 0 and 180? residual deviation must not exceed 11?.
Elimination of Radio Deviation of
--
Radio
station
course
angle,
deg.
Averaged AP for
compensation, deg,
Radio
station
Course
angle,
deg.
-
Averaged AP for
compensation. deg.
APK-5?Bo.1
APE-5 No.2
AM-5 No.1
APK-5 E04
the 1
0
15
? 30
45
60
75
90
105
120
135
150
165
0
+12 _
+18
+19
+16
+8
+3
-6
-9
-15
-13
-10
0
+12
+16
+16
+12 '
+7
+0
-9
-13
-18
+18
-13
180
195
210
225
240
255
270
285
300
315
330
345
0
' +11
+16
+16
+14
+8
0
-6
-16
-15
-12
an or
-+11ese
6
I.::
:1 6
+12
+6 ;Nos :
0 ,or a
'tunm
-14 ,0? m
-17 ,aust
-18
-12
3. In correcting radio deviation, see that the aircraft is placed from tn
RAP or OAP radio station at a distance of at least three wave lengths. It
the power of the radio station exceeds 10 kW, the aircraft should be placed at
a distance of at least 100 km. from the station.
4. Radio deviation must be performed not earlier than
A rise and not later than 1 hr 30 min. before sun net.
Elimination of Radio Deviation of APB-2 -
qopese
1. Determine and eckrect the setting error as follows: ?
(a) place the aircraft by the compass at a course which is approximately
equal to the magnetic bearing of the radio station which is used for deviation
correction;
. (b) using the deviation direction finder place the aircraft exactly at *
magnetic course equal to the radio station magnetic bearing;
(c) tune the AX8-5 compass in the radio station frequency, fine and recon
the inverse radio bearing. If the latter is not equal to zero, the framehase
petting error;
(d) to eliminate the setting error, turn off the six bolts and turn the
frame base through an angle equal to the setting error: in case of a plus erre:
turn the frame base clockwise, in case of a minus error turn it counter-eleekd
2. In order to determine and eliminate radio deviation, do as follows:
(a) position the aircraft at the magnetic course;
(b) adjust the radio compass receiver for the selected radio station, all
it to warm up during 5 minutes and correct the deviation at 24 inverse radio
bearings: 0, 15, 30, 45?, etc.;
(o) In case the deviation error exceeds =3? remove the frame of the tr1-5
compass No.1 and balance it. Then correct the deviation again at 24 inverse xi!
bearings;
0
1 hr 30 min. after
1
?
q PCP 'Ayr
Coepass No.2
1. Determine, in flight, the setting error with the radio station course
being equal to zero.
2. Correct the frame setting error on the ground, for 'which purpose remove
rams from the bracket, turn out the 6 bolts and turn the frame base through
le equal to the setting error: in case of a minus error, turn the frees
clockwise, and in case of a plus error turn it counter-clockwise, matching
same time the readings of the pointer of the AIN-5 compass No.2 on the
tor's selsyn indicator.
3. During the next flight (with the landing gear retracted) check compasses
and 2 for correct readings at eight points by the FUK-48 directional gyro
landmark (at any altitude). The difference in the readings of the compasses
to one and the same radio station at radio station course angles equal to
id 1800 must not exceed 2?, and at the remaining course angles this difference
not exceed =30, the radio deviation of APE-5 compass No.1 taken into account.
Notes: 1. The radio station course angle is to be taken by compass No.1 tuned
to a distant radio station (300 - 400 km.) or to an airfield homing
station of 1117-3B type (100 - 150 km.).
2. Take the readings off the navigator's seleyn indicator of compass
10.2 with antenna No.1 cut in and vice versa.
4. When approaching for landing (with landing gear extended check compasses
Nos 1 and 2 for differences at radio station course angles equal to 75, 120. 240
and 2850. See that the difference in the readings of the compasses tuned to one
and the same radio station does not exceed 13? with the deviation error of
compass No.2 being taken into account.
ppIe: The radio station course angle is to be taken by compass No.1 tuned
to a distant or homing station, and the results are to be entered
into the aircraft log Rook..
5. Check the entries in the aircraft Service log relating to the readings
of APE-5 compass No.2 at radio station COUrse angles equal to 75,120.240 an1285?
with the landing gear extended.
6. With the landing gear retracted, the residual radio deviat: for A1E-5
compass No.2 must not exceed =30; for course angles equal to 0? and ite the'
reaidnal deviation must not exceed =10.
Alignment of Automatic Course Device
Indicator of 1111-50 Air Position Indicator
with Navigator's Course Indicator of ATUI-7
2.02PE?2
After eliminating deviation of 11185-7 compass check the pointer of the
711VH-7 compass navigator's indicator for synchronous movement with the auto-
matic course device indicator of the 118-50 air position indicator. The courses
On both inAlCatOrS 'Should be aligned, the pointers should move in the same
direction. The alignment check is to be performed at courses from 00 to 3600
every 15?.
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Alignment is considered satisfactory, if the difference in the readings
of the automatic course device and the compaes indicator at courses divisible
br 150. does not exceed 10. otherwise align the courses with the aid of the
adjueting screws of the HN-50B air position indicator automatic course devb
(1i1.24). After adjustment is over, repeat the adjUstment cheek at all points
from 0? to 3600 every 15?.
25X1
ZOiN I
........
..... -----
AUTOPILOT 111-5-21
OREM
The in-5-211 autopilot serves:
(1) for automatic stabilization of the aircraft with respect to the three
axes in a straight flight;
(2) for performance of an automatic compensated turn and aircraft adationel
turns during lateral aiming;
(3) for ensuring stabilization of the sight in azimuth.
2ogp1ote Bet and Arrangement of A11-172K
Autopilot Obits an Aircraft
The directional stabilizer is arranged on a special bracket on freme BOA
tette front pressurized cabin (Fig.36).
The vertical flight gyro is installed on a special bracket in the front
presaurized cabin behind the seat of the left pilot (Fig.37).
The precession gyro unit and IM-10 inverter are positioned at the left
- loll between frames Nos 19 and 20 (Fig.38).
The servo units of the ailerons are located on frame No.33, those of the
rudder and elevator are positioned on a special wing on frame No.68. The elevator
servo unit is located to the left and the rudder servo unit is positioned to the
rigbt.
The control panel is located on tha upper electric board of the pilots
(ng.39).
The pilot director indicator (P.D.I.) is arranged on the instrument board
] of the left-seat pilot (See 1ig.110).
The turn remote control handle is located on the right-hand side of the
electric panel of the navigator-radar operator. .
The amplifier, 110-45 inverter, distributing box, relay box, resistance
box for changing the pitching moment are positioned on the left-hand ruck of the
1
! navigator-radar operator. -
Icartienci disengaging buttons are located on the spokes of ailerons control
i steering wheel of the left and the right pilots.
The 'formation 'stick and the control transfer are located on the smiielling
bracket on the middle panel of the pilots (See Fig.40).
The directional stabilizer attachment bracket is aranged on frame No.l.
The pitching moment limit switch, type B82-1416., is positioned on frame
10.33 in the mechanism of the bomb bay limit switches.
Epecifications_of AD-5-21I_Aut2pilot
1. The autopilot employs direct current, 27 ? 2.7 V.
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2. Plower consumed br the 411,5,2111 autopilot from the aircraft mains at
27 ? at rated load of the servo units (110 kg/cm. on the cable drum) does not
exceed 500 W.
5. Power consumed by heating hood e from the aircraft mains does not
exceed 250 W.
4. The operating temperature interval of the autopilot set is -45 to +50?0.
5. The autopilot operates normally within a temperature range from -20 to
-4590 only when its heating system is switched on.
6. Departure of the directional stabilizer from the course at any point
during 15 min. of operation Of the "yawing base" does not exceed 3.5?.
7. Total departure of directional stabilizer from the course at the to mat
points during 15 min. of operation at each point on the "yawing base" does not
sxnehd 8?.
8. Desistance between the contact bruedi and the aileron pick-up potentio-
meter centre tap on the directional panel of the directional stabilizer does not
exceed 5 ohms; the difference between the resistances of the rudder pick-up
potentiometer winding arms does not exceed 5 ohms.
9. The contacts of the vertical flight gyro erecting mechanism cutout close must
when the P.D.I. potentiometer pointer deflects 1 - 1.5? to the left and right
from the zero position.
10. The total defleotion of the vertical flight gyro rotor axis in both
directions from the vertical at normal temperature does not exceed 1.40.
11. breutheum deflection of the rotor axis from the vertical of the vertical
flight gyro is 1.20.
12. The erecting time of the vertical flight gyro cardan unit from 45? tilts
in each of tbe four qoadrants is 2 - 10 min., the difference between the maximum
and minimum erecting time from tilts at normal temperature must not exceed 4.5 el;
13. Power consumed by servo unit at normal temperature, 27 V and 110 kg/cm.
load moment on cable drum does not exceed 80 W.
14. The braking effort developed by the servo unit on the cable drum at
normal temperature is from 75 to 100 kg.
15. Servo unit potentiometer. Tension:
(a) potentiometer mita:ling brushes (total) - 25 to 45 gr;
(b) slip ring brushes (total) - 20 to 40 gr;
(c) limit switch plates - at least 150 gr.
Resistance of working portion of potentiometer winding is
25X1
2bAl
(c) ratio rheostat
600 t 60 ohs;
(d) trimming potentiometers 250 t 25 ohms;
(e) cospensating potentiometer 2000 t 200 ohms;
(f) control transfer potentiometer 2000 ? 200 ohms;
(g) turn control potentiometer ..... 10IND I 100 ohms.
19. Inverter 1101-45:
(a) output power .
(b) A.C. voltage
(c) A.3. frequency
(d) duty of operation
43.5 VA;
18.3 I 0.5 V;
*5
185 c.P.A.:
lone-tine
20. Invertar
(a) load current
(b) A.U. Voltage
(c) A.C. frequency
(d) number of phases
0.32 A;
36 t 4V;
.400 c.p.s.;
3
21. The alternating current in the gyro motor phase of the precession gyro
not exceed 0.35 A.
22. Precession gyre sensitivity must meet the following requirement: at an
immsular velocity not exceeding 0.1?/sec. voltage should appear and be registered
;by the voltmeter.
23. The precession gyro operates within an angular velocity range of
tesvc. +0.5?/sec.
24. Time required for formation stick to return to neutral position from any
extreme position is from 0.3 to 1.5 sec. both for the "aileron" and 'elevator".
25. The contacts of the formation central switch close before the signal
comes from the aileron potentiometer.
26. The button serving to switch off the aircraft control from the auto-
pilot has normally closed contacts. When the button is pressed the contacts open.
27. The brush surface contacting the commutator should constitute at least
85% of the brush section.
28. The time required for the navigator to additionally turn the aircraft,
with the autopilot coupling engaged at 4 to 60, must not exceed 18 sec.
29. With the autopilot in operation the control surfaces should deflect:
1-4 t0.5?:
16. Fewer supply of amplifier circuits:
(a) valve filament
(b) transformers
1100+2950 ohms.
-1
D.C., 27 t 2.7 V;
A.C., 125 t 15 c.p.s.,
17.5 t 2.5 V;
(c) voltage in secondary windings of bridge transformers at supply volt-
age of 17.5 V and frequency of 125 c.p.s. - 27 2 V;
(d) throttling voltage:
emmdblum "' ?2?V
minimum ... 25.5 t 7.5 V
17. Rarer consumed by amplifier:
(a) valve filament 30 V maximum
(0) transformers 40 VI; maximum
18. Control panel resistances:
(a) centring potentiometers 200 t 20 ohms;
(b) sennitivity potentiometers 0.33 0.066 megohms;
SECRET
(a) ailerons
(b) rudder L5 t0.5?:
(e) elevators
30. The elevator neutral posi:tion corresponds to a deflection of the elevator
by 2? downward.
31. Elevator deflection when the bosh bays are opened at a speed of 450 kw/hr
is equal to 20 t 5 angular minutes.
32. The AH-5-211 autopilot set employs the following valves:
1 piece;
3 pieces;
3 pieces.
15 to.5?;
(a) 6 x 5
(b) 611811
(c) WHY
33. The insensitivity zone of tte A11-5-211 autopilot with the sensitivity
handles shifted to minimum position is as follows:
(a) aileron
(b) rudder
(0) elevator
at least 1.5?:
at least 0.500;
at least 1.00.
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54. The insensitivity zone of the in-s-am autopilot with the
handles shifted to maximum position is as follows:
(a) aileron not in excess of 0.4?1
(b) rudder . not in excess of 0.25.3;
(o) elevator not in excess of 0.40.
35. The temperature inside the electric 'heating hoods i.e maintained by a
thermal relay within the range of from 10 to 45?C.
36. The minimum length of the brushes at which they should be replaced is
.given in Table le.
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sensittrill 5. Take out the course panel cover and make sure that there is no dirt or
dust inside the casing or on the potentiometers, :
6. With the sight mounted, check to see that the wire bundles do not inter.
ferewittefree movement of the sight with resptct to the drift angles.
, 7. Bake sure that the handles of the directional stabilizer bracket rotate
without binding and allow the stabilizer to be set by a level.
s. wee a dystammeeter cheek the tension of the autopilot, sight and drift
gear couplings in accordance with Table 15..
Table 14
Minimum Ieneth,at Which Brushes Should Be Replaced
Tablee 15
Opting Tension an Directional Stabilizer
Couplings
Spring tension by dynamometer, kg
Name of unit
Mark of
Length of tirush, mm Name of coupling
brush and
,-index
minimum
rated
minimum
memimum
1
2
3
4
6
Directional stabilizer:
Autopilot coupling ?
gyro
IIPC-7
22
28 Sight coupling
a
9
erecting motor
NIV-?
11 - 12
18.5 Drift gear coupling
4
5
Vertical flight gyro:
Control Easel
gyro
MTC-7
18
24 1. Check the handles for reliability of attachment on the shafts.
servo unit
Mr-4A-A6
15 - 16
24 2. Check the rotation of all control panel handles within their turning
Inverter 110.45:
limits. All handles should rotate smoothly and without binding except:
A.G. commutator
Kr-4A-A6
4.5 - 5
10.5 (a) turn control handle; when
the pointer approaches the shaded portion of
D.C. commutator
Mr-41-k6
6-6.5
13.5 the scale and the zero position, e
resistance to handle rotation should be left.
Inverter lar-itt:
(h) turn corpens_mtee_m=2). drive han411; when the
pointer approaches the
position "Pilot', rotation of the handle becomes more difficult.
14
3. Check the plate of the switches for correct functioning
D.C. commutator
- MC-6
10
and for proper
attachment to the switches.
ORECK/NG AUTOPILOT FOR INSTALLATION ON
AIRCRAFT AND OPERATION MOM CURRENT
EXTERNAL INSPECTION
Directional Stabilizer
1: Check the bracket for proper attachment and see that the directional gyre
is reliably secured to the bracket.
2. Engage and disengage the autopilot and sight couplings several times to
make sure that the engagement mechanism functions correctly. When disengaged, the
couplings should rotate freely on their drums without binding or dragging the
drums along with them.
3. v....mina the locking mechanism unit to make sure that the locking mochamia
plunger is locked with its nut. Shift the solenoid plunger several times downward
to make sure that the return spring returns the entire lever system and the plune
to the initial position. As the plunger goes downward, the lever of the autopilot
coupling must be pressed. Remove dirt and foreign particles.
4. Make sure that the locking mechanism does not interfere with the movement
of the autopilot coupling lever; see also whether there is a clearance between the
*jaws" of the locking mechanism and autopilot cuppling.
4. lake sure that all the pointers of the control panel handles are reliably
secured and more only when the Ebndles are turned.
Notes: 1. After the check handles ."RATIO", "TURN COMPENSATOR',?INCREILSE
AMP, 'TO DECREASE SKID' and "DP EISP: must be placed in position
determined in the air.
2. After the test is over set the drive control handle of the turn
compensator into position "Pilot".
Pilot Director Indicator (P.D.I.)
1. Check the instrument for reliability of attachment.
2. Check the condition of the glass.
3. Check the pointer and instrument scale for presence of luminous compound.
Alltapilat_Seitch-gt
Button_
l.necheonetkror wheels.ebutton for reliability of attachment in the splines of the
Luero
tu:itimes to make sure that they return to their
ti posjsj without binding.
RWIIPV.T
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Formation Stick after is
1. Maks Sur* that the stick after being deflected and released returns Mj.n d, the rods and levers of the braking solenoids return to
it. central position. ?
4. MOv sure that there is no dirt or oil on the potentiometer contact brush
2. Cheek to see that the buttons on the stick do not bind when being pre. and vimung. Check the tension of the potentiometer contact brush.
and return to the initial position. 5. Check the potentiometer brushes for reliability of attachment on braking
Formation Stick Control Transfer
1. Check the stick for reliability of attachment on its shaft.
2. Set the control transfer in all positions and make sure that control
transfer is performed without considerable efforts and that the stick is firsd
held in the selected position.
Vertical_Flight gyro
?
?
1. Cheek the attachment of the vertical flight gyro and see that there ex
no cracks on the shook absorbers.
? 2. InCalidne the plexiglass cover in the upper part of the vertical flight
gyro and sake sure that the cover is placed in a position at which the qertica
flight gyro is cmcaged. The plexiglass cover must net be cracked or have any
deep scratches.
timEllf.leE ?
Check the amplifier for reliability of attachment, good condition of she
absorbers and bonding.
.Inverters
Check the inverters for reliability of attachment.
Distribution Box
1. Check the distribution box in accordance with the requirements presnd
ed in Section "Aircraft Electric Mains" (See "Care of Split Boxes and Electric
Bearden). ?
2. Check the external condition and correctness of installation of serial
resistors on terminals 5-1 and 5-3 , as well as on terminals E-6 and E-8
which are equal to 400 ohms.
1,3AZ P.03
1. Check the relays for reliability of attachment in their seats.
2. Make sure that there are no metal chips, dust or any foreign objects
inside the boxes.
3. Examine tightening of the nuts and attachment shoes of the wires.
4. Check to see that all the wires and their insulation are in good condi.
tion.
Precession Gyro Unit
. Chick to see whether the precession gyro unit is reliably secured and that
its surface has no dents or scores.
Servo Units
1. Check tightening of the bolts which secure the servo units.
2. Deflect the rudder and elevators into both sides and make sure that ta
cable is wound around the drum and that it permits Shifting of the rudder and
elevators into both directions.
1. Press the tension springs of the braking solenoids and make sure that,
dss2.6. Placd the rudder and elevators in neutral position and make sure that
the Contact brushes of the follow-up system are centred, whereas the slide
moving along the potentiometer is in its down position.
Turn Remote-Control Button
1. Check to see that the handle is reliably attached on the potentiometer
shall Make sure that the handle rotates freely, without binsling, except in
position "Or and in the shaded portion of the scale where a resistanee is felt
in turning tile handle.
Box for Changing the Pitching Moment
1. Cheek the resistance box for reliability of attachment.
2. Check the wires for reliability of attachment to the resistance box.
Note. When inspecting the autopilot units, examine the condition of the
plug connectors of the units. The plug connectors should be tighten,
ed as far as they will go, have no considerable play and be looked
with safety wire.
Checking OReration of Energized Autopilot
CADTIOR: Prior to checking the autopilot for operation, do not fail to un-
lock the aircraft controls and remove the service ladders, covers and
other objects. Stop any operations on the aircraft controls.
1. Switch on the A3C-15 autopilot circuit breaker on the circuit breaker
board of the left-seat pilot and the A3C-2 "Servo" circuit breaker on the
circuit breaker panel of the navigator.
2. Actuate the plate on the panel to switch on the master switch and the
"Stab." switch; then make Sure that:
(a) the gyro motors of the directional stabilizer and vertical flight gyro,
the precession gyro unit, inverters and servo unit motors are already operating;
this is determined by the peculiar noise of the running motors;
(b) the erecting roller rotates properly.
3. After cutting in the master switch wait 5 to 8 minutes and then switch
=the "Servo P.D.I." switch on the control panel. This will cause the direc-
tional stabilizer erecting motor to operate.
4. Disengage the autopilot coupling and shift it to the left and right to
check whether shifting of the P.D.I. potentiometer brush causes the P.D.I.
Pointer to deflect. When the potentiometer brush is moved to the left,the P.11.1.
pointer should gradually deflect to the right and vice versa.
5. Place the F.D.I. potentiometer brush at zero. The P.D.I. pointer should
also be at zero. Then engage the autopilot coupling..
6. Shift the aircraft control surfaces (rudder, ailerons, elevators)
manually from one extreme position to the other. Repeat this movement several
times. Make sure that the controls move freely.
7. In shifting the control surfaces manually check operation of the control
Psnel pilot lamps. When the control surfaces are in the neutral position, the
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lamps must not light, when they are in any other positions one of the lamps 17.Turn the turn control knob on the control panel to the right so that
burn without blinking.. Blinking of the imps when the control surfaces are not knob indicator should be at the beginning of the shaded portion of the
in the neutral position means that the servo unit potentiometers are dirty. /a de and sake sure that the steering wheel has turned to the right and the right
this case clean the potentiometer of the respective servo unit. Lai hen displaced forward. Turn the control knob to the same position, but to
left and wake sure that the steering wheel bas turned to the left and the
Innl: On some aircraft the slide of the servo wait potentiometer may come ,
off the potentiometer :Binding when the control surfaces are in the Pedal has displaced forward.
extreme positions. In this ease both pilot lamps will not burn. Set the knob in position "0", first to the right and then to the left, to
ce sure that the solenoid of the directional stabilizer locking mechanism is
8. Turn the centring knobs on the control panel to the right or to the
'left; the position of the control surfaces at wbich both lamps go out will be ;aged and locks the autopilot coupling lever, whereas the top erecting roller
17.aksconbasiendp:::::io:g
changed. :Centre", and then to the right
9. Set the knobs "Centering" so that the pointers face upward.
I left-hand positions "OP. This done, place the knob in position "Centre" again
10. Without centering, switch on the following switches on the control ef:iu:1112%:7n::o
he steering wheel and pedals should not move) and 'make Sure that the solenoid
then:ersteitc:lh
panel: 'AILERON", "RUDDER", and "EIEVATOR". When some of the mentioned switches the lockingmechanism has disengaged and set the autopilot coupling lever
are cut in, the respective stabilization pilot lamp must light up and then go ee, whereas the top erecting roller of the vertical flight gyro begins to
out in some time since the control surfaces will be set in the neutral positiontate,
by the operating servo unit. 19. Set the turn control transfer on the control panel into position
11. As the centering knob is slowly turned to the right or to the left avinatorP to make Sure that the transfer position indicating lamp is on. Use
as far as it will go, the respective stabilization channel control surface e turn remote-control knob to carry out the checks described in Items 17 and 18.
should deflect. 20. Place the aircraft controls in neutral position as indicated by the
The control surfaces should deflect with interruptions, but evenly, each lot lamps on the control panel.
time the control surface displaces, the respective pilot lamp on the control 21. Set the formation stick control switch in position "ON".
panel should blink. 22. Shift the formation stick to the right as far as it will go and make sure
Terming the centering knob "AILERON" clockwise should cause the right- at the steering wheel has turned clockwise and, that the right pedal has displaced
hand aileron to go up (the steering wheel rotates clockwise)and mice versa: turzrward. With the formation stick in this position the locking mechanism
the knob "AILERON" counter-clockwise will cause the left-hana aileron to go up (;lenoid of the directional stabilizer should be engaged and lock the autopilot
steering wheel rotates counter-clockwise). ?upling lever.
Turning the centering knob "RUDDER" clockwise will cause the rudder tn 25. Release the formation stick and make certain that it returns to the
shift to the right-hand turn position (the right pedal goes forward). Terming mtral position. With the formation stick in the neutral position, the solenoid
of the same knob in the opposite direction should Cause the rudder to shift to the directional stabilizer locking mechanism must automatically disengage
the left-band turn position (the left pedal goes forward). ualock the autopilot coupling lever in 5 to 9 sec. after the stick has
?
Turning the centering knob "ELEVATOR" clockwise should cause the elevate:turned to the neutral position.
to move upward (the control column moves backward). When this knob in turned 24. Shift the stick to the extreme left-hand position and make sure that
counter -clockwise, the elevator will go down (the control column moves forwardhe steering wheel has turned counter-clockwise and that the left pedal has
With the knob "ELEVATOR" in the extreme positions, the pilot lamp may not burn. ved forward. Release the formation stick. Make sure that the stick, steering
12. Set the centering knobs to "Pointer Up" position. teel and pedals return to the neutral positions. With the formation stick in
le neutral position, the solenoid of the directional stabilizer locking mecba-
Note: An the centering knob is turned, not more than two simultaneous
mot must automatically disengage and unlock the autopilot coupling lever in 3
blintings (pulses) of both pilot lamps are allowed on separate
sections. 9 sec. after the formation stick has returned to the neutral position.
25. Shift the formation stick backward as far as it will go and make certain
13. Disengage the autopilot coupling on the directional stabilizer and twamt the control column has moved rearward. Release the formation stick and make
the coupling lever into the extreme left-hand position. The steering wheel shod
u-e that the stick and column return to the neutral position.
turn to the right, the right pedal should go forward and the P.D.I. pointer 26. Push the formation stick forward as far as it will go to make Sure
should deflect to the right. at the control column moves forward. Release the formation stick and see
14. Set the lever of the autopilot coupling in the extreme right-hand1"-' si
aether the formation stick and control column return to the neutral position.
tion. The steering wheel should turn to the left, the left pedal should move 27. Set the switch in position "Only Elevator 011". Deflect the formation
forward, whereas the P.D.I. pointer should displace to the left. Act forward and rearward to make sure that the control column follows the
15. Return the autopilot coupling lever into the central position; after ormation stick. When the formation stick is moved to the right and to the left
the contact 'hush of the P.D.I. potentiometer assumes its zero position, engage he pedals and steering wheel must not move.
the autopilot coupling. 28. Press the autopilot disengaging button of the left-seat pilot for
16. Set the control transfer to position "Pilot" and make sure that the - 2 sec. and displace the steering wheel, control column and pedals to make
pilot lamp does not burn, are that the autopilot servo units are disconnected and that the aircrnft
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controls move freely. Using the plate cut out the master switch on the contmg
panel and cut it in again. Cut in the switches "AII.F.a0N"? "RUDDER" and "ELraa
Displace the bopster control handle to make sure that the handle actuates the
aircraft controls.
29. Press the autopilot disengaging button of the right-seat pilot for
1 - 2 sec., with the booster control switch in position "OFF", and shift the
steering wheel, control column and pedals to make sure that the autopilot sea
controlled by auto-
units are disengaged and that the aircraft controls move freely. Using the
that the controls are actuated by the knob. control surface movement
21$ilot master switch on the control panel and cut it immediately given course
in mi.,
- Note: If the autopilot disengaging button is released after it was ;rem
arm
Cut in the switches "AILERON", "RUDDER" and "ELEVATOR" and check to see that;
the servo units should rot engage the aircraft controls before the Perio(lical jerks in
master switch on the control panel is switched off and on again.
. with the aid of the plate.
When
operation of each stabilization channel Separately, for which purpose: aircraft is control-
30. Check the sensitivity adjusting knobs on the control panel for o ai rm
r' "ELEVATOR" on
with
"ON", the lock-
led by foation stick,
switch placed in
leg mechenism fails to
(b) set the centering knob of the cut-in stabilization so that its point
control panel;
(a) cut in one of the switches "AILeRCIre, "RUDDER" or
knob counterclockwise; operate or operates twice
will should face upward (both pilot lamps ll be out); turn the sensitivity adjustWhen the centering knob
(c) turn the sensitivity adjusting knob on the control panel clockwise.
The steering wheel, control column and pedals will begin to oscillate, whn.?: is actuated, the control
the pilot lamps on the control panel will begin to blink in turn, surface lags
31. Check operation of the pitching moment counteracting mechanism for WE
(a) adjust the neutral position of the steering wheel by the elevator; solenoid does purpose:
time (3 to 9 sec.) after
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Fault
1
- 67 -
Faults and Remedies
Probable cause
2
Remedy
3
(b) open the bomb bays. When this is done, the steering wheel should mo
forward. When the both bays are closed, the steering wheel should again as=
its neutral position. With the bomb bay open the elevator should be deflected
9 7 10 min with relation to the inner face of the trailing edge, the blade bee
not taken into account. The may of measuring the deflections of the aircraft
control surfaces is described in Section "Controls" (See Book I).
32. Check operation of the autopilot heaters, for which purpose cut in 0
I30-10 circuit breaker of the autopilot heater. The circuit-breaker is locate
on the circuit-breaker panel of the left-seat pilot. Then switch on the heatig
aystea switch on the upper electric panel of the pilots and the A3C-10 sins
breaker of the rudder and elevator servo unit heaters. This done, make sure 0
the lower covers of the servo units and vertical flight gyro warm up.
33. After the check is over cut out the switches on the autopilot control
panel making use of their common plate, cut out the A3C-10 circuit breakers:
the rudder and elevator servo unit heaters, as well as the heating switch on
the upper electric panel of the pilots.
the formation stick has
been turned
Unequal bank when
control is exercised
through directional
stabilizer
Aircraft turns sponta-
neously
gFeRRT
Aircraft comes out of
turn spontaneously
Directional stabilizer
ger?
unbalanced
Improper braking mo-
ment of servo unit
braking solenoids
(a) Unstable opere-
tioa of time relay
(b) 'hong adjustment
of lockina mechanism
jaws on directional
stabilizer
No contact on servo
unit potentiometer
Wrong adjustment of
tine relay
heplane directional
stabilizer or balance
its gyro on a stand
Adjust the braking
roment of the servo unit
braking solenoids by even
distribution of the
efforts to the left. and
right sides
Replace time relay in
relay box
Adjust the lockina
meeneniam of tte direc-
tional stabilizer
arong centering of po-
tentiometers on direc-
tion panel of directio-
nal stabilizer
(a) Wrong tension
adjustment of directio-
nal stabilizer, coupling
(b) Erecting mechanism
of directional stabilizer
and vertical flight gyro
out of order
Locking mechanism on
directional stabilizer
is loose
Clean surface of potentio
meter with a brush out of
the AR-5-2U autopilot
set. The brush shoeld be
soaked in clean gasoline
Replace time relay in
relay box
Centre potentiometer
slide on direction panel
with the aid of an ohm-
meter
Adjust tension of
directional stabilizer
coupling
Eliminate fault of
erecting mechanism-
Adjust tightening of
the locking mechanism on
the directional stabili-
zer.
Note: When the autopilot
coupling lever is tight-
ly locked, there should
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3
When autopilot is engaged
the aircraft tends to turn
eherply
Serve unit braking sole-
noids will not engage
Gyro of directional
stabilizer 'or vertical
flight TO is out of
order
Autopilot is discon-
nected from buttons on
aileron control steering
wheel
be a cleasexce4
2 - 3 MIK beteal
the bottom and
armature of the
locking nechasiii
solenoid
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..... -----
Remove directional
gyvd OXYGEN EQUIPMEHT
flight e an check
stabilizer and yenta
their operation 'on a
stand The oxygen equipment ensures norms/ oxygen supply for the aircrew during
ash-altitude flights and bailing cut by ejecttng tie geat.
(a) Reins the mane The oxygen system (Fig.41) includee the follovicg elements:
gens
emitch on the antopih 1. Two liquid oxygen converters, type 1011-30 , designed fcr stnni nd
ga a
control panel en
zasification of livid oxygen and delivering gaseous oxygen to the line which
"aPietautoeplidland autopilot dinenneupplies the aircrew (the arrangement diagrem of the 01-30 liquid oxygen
the
converter is given in Fig.42).
cot(t)::::::=:!:: 2. Six oxygen stations. Each of them includes:
autopilot for aircraft (a) stationa type
ry oxygen regulator, KR-24M;
sowing to engage the
(t) pressure gauge, type 0-1311
control .(c) oxygen-flow indicator;
(d) excessive pressuxe gauge, type M-1000;
(e) oxygen valve, type KB-5;
(f) oxygen hose, type 11-24;
(g) oxygen meek, type KM-3011, with the mask-to-face tightness compensator
and a lock;
(h) yellow warning lamp (only in four stations).
3. The aircraft charging system consisting of aircraft charging connection
.
and pipe lines of the AHril-T12x14 material connecting the changing connection
,to the 11131-30 liquid oxygen converter.
, 4.S1x tee-pieces with nen-return valves. The tee-pieces ensure oxygen supply
ito each working station from both KIII-30 converters simultaneously and prevent
qcompn from being released from both converters simultaneously in case me of
ithe sections of the oxygen system is damaged.
5. Sim parachute oxygen apparatus, type KU-23 , with hosee.
All these elements are connected to one another by means of pipe lines of
ithe Alr1-16x8 material and aircraft fittings.
The vessel of
amount of liquid oxygen filled into both converters is 64 kg. The amount of
CQfl8me1 oxygen is ::
ch:onverter,type 11X-30 ? has a capacity of 28 litres. The
Bur. Ti:elnitzmx:a.ting pressure is 8 atm. gauge, the maximum operating Pres-
.
11.8
The
The pressure reqyired for the operation of the safety valve is from 11.0 to
oxygen consumption from one converter is 6 kg per hour.
The evaporativity of each converter must net exceed 250 gr per hour.
GENERAL SPFCIFICATIONS
AGCESSIBILTTY FOR INSFECTIOR
Oxygen panels with instruments both in the front and rear pressurized cabins
are easily accessible for inspection.
RP.C11:2Thrr
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To inspect the pipe lines in certain places it is necessary to remove tbe
neighbouring units. Such places are:
(1) Pipe line in the area of frames Nos 4 - 9 on the starboard and port
sides.
To obtain access to the pipes, remove from this section: from the starbon
side - the shut-off valves control panel, from the port side - the interphone
set panel of intercommunication system COY-10 , and the IFF transponder pan
To obtain access to the pipes on frame No.4, remove the fuse panel and
flap back the navigator's instrument board.
(2) Pipe line in the area of frames Nos 9 - 12 on the starboard and port
sides. To obtain access to the pipes of this section remove the following free
the starboard side:
(a) converter, type PCS-?O on the operator's panel support;
(b) thyratron interrupter on the operator's panel support (starboard).
from the port side:
(a) block, type P-6;
(b) P-6 block panel.
(3) Pipe line on the bottom of frame No.12.
To obtain access to the pipes, it is necessary to detach the operator's
electric panel (central panel).
(4) Pipe line in the 0-3 cabin on the starboard and port sides.
To obtain access to the pipes, do as follows:
(a) open the container hatches of fuel tanks Nos 1 and 2;
(b) remove tank No.1;
(c) remove tank No.2;
(d) retove hatches In the containers of tanks Nos 1 and 2.
(5) Pipe line in the 0-4 cabin from frame No.26 up to frame No.34 on the
starboard and post sides.
To obtain access to the pipes, proceed as follows:
(a) open the hatch in the ('-4 cabin between frames Nos 27 - 2e;
(b) remove the starting fuel tank between frames Nos 27 - 29;
(c) remove the cooler between frames Nos 30 - 31.
Besides this, to obtain access to the pipes of the starboard side, do as
follows:
(a) remove the drain pipe;
(b) eemove the pipe of the high-altitude equipment;
(e) loosen the yoke on the drain pipe and turn the branch pipe.
To obtain access to the port side pipes, remove the drain pipe.
To obtain access to the pipes in the area from frame No.34 up to frame
No.49, it is unnecessary to remove the neighbouring units.
(6) Pipe line in the 0-6 cabin.
To obtain access to the pipes, it is necessary to remove th. PONY-3M
radio set from the bottom of frame No.69.
(7) Aircraft charging pipe line in the area of frames Nos 19 - 22.
To obtain access to the charging pipe union and charging pipe line:
(a) open the aircraft oxygen charging hatch;
(b) open the batch in the nose wheel well on the starboard side between
frames Nos 19 - 22.
(a) The pipes laid In the area of frames Nos 49 - 57 are not accessible.
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(e) To Obtain access to the liquid oxygen converter, type 101E-30 , at
frame N0.13 and the pipe line at the converter use the hatch on the bottom of
frame No.12.
PREPARATION FOR FLIGHT INSPECTION
The pre-flight inspection consists in thorough examination of all accessible
units and pipe lines of the oxygen system. For this, do as follows:
1. Open the entrance hatches in the pressurized cabins (if they are closed).
2. Open the doors of the beef, bay hatches and of the hatch between frames
Nos 27 - 29,
3. Remove cases from the oxygen regulator, type m-24 .
4. Examine the oxygen panels with instruments and make sure that they are
not de:gaged.
9. Check to see that the instruments are securely fastened to the panels.
6. Examine the pipe line in all accessible places; make sure that the pipe
lines are securely connected to the oxygen flow indicators, BK-1311 pressure
gauges, 11-1000 excessive pressure gauges, 111-241( economizer, KB-5 oxygen valves,
Belem (EM3 ) and TOP (BEPI ) transmitter pipe unions, pipe unions of the KB-5
valves on the Klii-30 converters and check the pipe lines for secure attach-
ment.
7. Examine the 1III-30 converters (Fig.43); check whether the safety valve
case is securely attached and whether the KB-5 valves and the pressure release
valves of the 1111-30 converters open easily.
8. Examine tee places of connection of the E11-24 hoses to the K11-24U
econoeizer.
9. Check whether the oxygen adapter in the operator's seat turns easily
(Eig?44)?
10. Check the presence of liquid oxygen in the 111I-30 converters. Add
liquid oxygen if necessary.
11. Put the 111I-30 converters to the operating condition and check them
for serviceability.
12. Check the operation of the 10I-244 set.
Margin:, 1(111-30 Converters with /Squid
en
To save liquid oxygen, prior to each flight the E11-30 converters should
be charged with the amount of oxygen necessary for the flight only.
The amount of oxygen required for the flight is to be determined by the
formula:
Gna Ge + tq ,
Where Gn is the required amount of liquid oxygen in kg;
Ge is the stock of oxygen not taken into account In. kg (6 kg per each
converter, type 111I-30 );
is the rated oxygen consumption for all the aircrew in kg per hour
(5 kg per hour for all the aircrew):
t is the time of flight in hours.
Calculate the amount of oxygen necessary for the flight and then measure
the amount of oxygen on the aircraft turning on the switch, type 20-250, with
the 110-4500 inverter operating. Add the readings of both indicators and
ceseere the amount obtained with the amount of oxygen required for the flight.
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If the measured amount is less than the rated value, charge the cOnverters vita (f) eater the adapter pipe
Of the vacuum flask has been connected to the
additional amount of oxYgen. nion, slowly open the valve on the cylieder with gaseous
.
mums Flight with ovvel supply* below the rated v create 'value is prohibited. milt . ng pipe u.
m liquid manner: gen, ate excessiVe pressure of not less than 4 atm. gauge and fill the
re::::.
I. Open the pressure release valves (if they are closed) on the iffile30
co:rrerters and equalize the pressure in the converter vessels with the atmosphezt .
-:
The conveeters should be charged with oxygen in the following fo ?barge the 0-30 converter to full capacity, it is necessary to
pressure (116.450. E 11. bele eix full vacuum flasks.
We: To uniformly charge the IIII-30 converters with liquid oxygen. open ' n sure that the vessels are charged completely, close the valve on
or on
pressure release valve on the converter at frame No.13 by 0.5 of the cylinder with gaseous oxygen, if charging is done frontlet
the uum flasks.
turn only and that on the converter at frame No.22 conpletely. 4 12. in two or three minutes after the liqeid has been filled disconnect
2. Oloos the valves, type IB-5, before the automatic pressure increase unitigtteteheo2anrair7aftand :3,7.au7t:: Idt:0::ie:::PipPiePeliluielle connecting the vacuum flask
and after the evaporators. 13. Screw a plug on the aircraft charging pipe union and close the batch.
3. Open the cover of the aircraft charging pipe union hatch end unscrew the 14. Switch off the 21111-250 ezygen level indicator switch, the inverter
plug on the pipe union (1118.46). .
4. Wipe the charging pipe union with a piece of gauze soaoke with aicohO7.1 15. The oterging completed, disconnect the filler pipe from the adapter pipe.
5. To the charging pipe union connect the pipe line from the .,e,V4 and aove the filler pipe from the vacuum flask, plug its upper end and fit a rubber
the vacuum flask CCA-15 ). end.
d itch and the A3C-2 circuit breaker.
lele: When unscrewing the plug end connecting the pipe lize, see that no p on the lower
moisture and dust get into the filler pipe. Disconnect the low-pressure hose from the filler pipe and plug its free end;
6. Switch on the 13C-2 circuit breaker on the operator's electric panel ose the vacuum flask with a plug.
in the inverter circuit, set tto inverter switch on the generator panel to the Checking (titration of Distan Lioeid
t -Readipz
013X071 141
OfiffttfING ( PABOVII ) position and turn on the oxygen level i he ndicator switch, 1LIndieator. O
TID 1173
e MA-250 . an the pilot oxygen panel. Check the operation of t omygen level indicator, type AFIK ,
typ when
the BI-30 converter with liqei oxygen.
7. Pressure in the "TaUC' reaching 3 or 4 atm. gauge, open the valve on the largitg Id
'Tank" and fill the vessels of the RE1-50 converters with liquid oxygen.
8. /111 the converters until liquid appears in the drain holes which indica? to 52 kg'
With the Tassels filled completely, the indicator pointers must be within
With the pressure increased, the indicator pointers may fluctuate
tea that the vessels are charged completely (32 k. in each vessel). Besid?s. ihieksext,..!2?
cheek filling the vessels with liquid, by the pilots," oxygen level irleutors. i:::1?:o7e7::::7f:f:::
Irfe::::: rpir::s:::trdils:f:Lfri::i::::: ITien:e::
With. the vessels charged completely, the indicator pointers mmst be within the eve towards increase in readings, this is indicative of leaks in the Oxygen
s f 28 to 32 kg. evel /ndicator Toe"
. connections; if they move tovards decrease in readings, this
limit o
.
9. If the converters aril. not charged with.oxygen to their full capacity, s indicative of leaks in the "Oxygen. Level Indicator Bottom" connections. If on
check the amount of ogygen.filled by the oxygen level indicators only. reseure increase the pointers do not move at all, this shows that the AUK
en level indicator circuit i de-energized. This being the came, ring out
Naas When charging the converters with liqpid oxygen, see that excessive he circuit and eliminate the damage.
xygs
preesure is oval to 3 or 4 atm. gauge as decreased excessive pressen
mAy cause the non-return valve to get frezmn in the KRA-30 converter Putti MI-30 Converters to Operating
ng
Condition
filler neck. In this case stop charging and kmock on the valve or
mith a wooden etiok.
iionee follows:
, Put the charged IDI-30 convertors (See Fig.45) to the operating condi -
ID. If the converter is filled with liqeid oxygen from vacuum flasks,? prior
1. Close the essure elease valves (if the nverter are put to the
to connecting the pipe from the Taeuum flaEsk to the aircraft charging pipe unioat)perating condition immediately after charging . Id the pressure release valves
pr r cos
do as follows.:
(a) bring tbe cylinder with gaseous oxygen to tbe air 2. Open the
craft, connect the ? ahead of th
lv .
automatic pressure increase nits.
cam ape reducer to the cylinder and check the presence of oxygen in the eninderl 4re open).
1035 vaes e u
(b) open. the valve on the cylinder and scavenge the reducer .with the hose 3. Open the after the evaporators.
X3-5 valves
with gaseous oxygen; ? at 4. Pressure in the converters must filet increase rapidly and then stop
. 8.3 or 8.5 atm. gauge. If the converter contains not less than 26 or 27 kS.
(c) remove the plug from tto vacuum flask, it a rubber packing gasket over the time of pressure increase to 8.3 or 8.5 atm. gauge most not exceed 10 minute
the vessel neck, insert the filler pipe and fasten the extension deapere to ths (usually this time is eeual to 3 to 5 minutes).
Dandles of the vacuum flask;
when premiere increase stops, begin to wine up.
(d)eannect the low-preesure hose to the filler pipe; The pipes before the aptomatic unit get frozen and then in 10 or 15 minutes
(e) connect the adapter pipe to the vacuum fleekl. Increase pressure not less than 30 minutes before flight.
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Checking Serviceability of KUK-30 Converters Before
Flight (See ig.45)
1. When opening the EB-5 valves ahead of the pressure increase automatic
units, pressure in the converters must rise up to 8.3 or 8.5 atm. gauge durdig
or 10 minutes, after this pressure increase must stop (the. pipes ahead of the
automatic units get warmed). If the pipes ahead of the automatic pressure incm
valves fail to get warmed during 10 or 15 minutes, this testifies to a loose
valve-to-seat fit. This being the case, reduce pressure and then increase it s.
If this does not cause the valve to close, it is necessary to close the valve
ahead of the defective automatic unit and then open it by 1/8th of a turn afte:
it has warmed up. .
Note: When no gaseous oxygen is consumed. pressure in the system increase
due to evaporation of oxygen in the converter vessel and in some th
reaches 11.5 cr 11.8 atm. gauge, that is the safety valve relief gn
sure. If the automatic units are in good repair, this must take WI
less than 45 minutes.
2. Cheek the gastightness of all the converters connectione accessible fm
inspection. Pay,attention to the gastightness of the cbargiog pipe union.
3. Open the oxygen valves, type KB-5, (See schematic ctiegram 6 Fig.45)
ahead of the economizers, type 1CJ1-2441 , at all stations.
4. Check oxygen delivery to oxygen stations by the pressure gauges, type
NK-13N, which must indicate operating pressure in the supply line from 8.3 to
10 .Am. gauge.
5. Check oxygen delivery to the supply line from each vessel of the K1X-3C
converter separately. For, this close in turn the KB-5 valves after the evaport
of the KIII-30 converters. Check oxygen delivery by the pressure gauges, type
1K-1311, installed at oxygen stations. Prior to checking oxygen delivery from es
vessel release pressure in the supply lime through the emergency cocks of the
KM-24M economizers.
Checking Operation of KII-2411 Economizer
1. Check the gastightness of the high-pressure system of each oxygen stet;
For this open the KB-5 valve on the oxygen station and then close it. If the
pressure gauge pointer does not show pressure drop during not less than 2 mint
the system is gastight.
2. Check the gastightness of the low-pressure system of the E1I-24M era?.
mixer (from the economizer valve up to the plug on the KM-24 hose). For this
purpose do as follows;
. (a) release the remaining oxygen from the system by means of the manual
regulator on the KIK-24M apparatus;
(b) set the handle of the air dilution sritch to the CLOSED ( 31PJjTO )
tion;
(c) remove the plug on the KM-24 base and make an inhalation. If it is
Impossible to make an inhalation, the system is gastight;
(d) after checking the gastightness, set the handle of the air dilution
switch to the OPEN ( OTKPUTO ) position.
3. Check oxygen delivery of the mak by the K11-2461 economizer without
excessive pressure and with excessive pressure setting the economizer cock to d
NIXED ( Web ) and PURE ( TETA ) positions successively.
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?75?
Checking Operation ofKII-24N
ECOn0MiLer without Excessive
Pressure
1. Put on the mask, type HM-30U. and caamect it to the hose, type KE-24.
2. Open the 1B-5 valve of the oxygen station.
3. Close the manual regulator on the KM-24U apparatus as far as it will go.
4 Waite several inhalations and exbalations setting the manual switch of the
sir dilution automatic device first to the OPEN ( ?MUT? ) and then to the C1OSED
(3mauTo ) positions. If this causes the flow indicator flaps to get together.
the EI-Z4M apparatus functions properly.
Checking Operation of KM-2441
tconomizer with Excessive Pressure
1. By means of the manual regulator build up excessive pressure equal to
250 msl.G. on the KII-2461 economizer (watch the pressure by the N-1000 excessive
pressure gauge).
2. Put on and remove the plug from the 1cM-24 hose several times. If this
mums the flow indicator blinkers to get together and depart, the KM-24M
am:miter operates zormally.
let: When the apparatus operation is checked under excessive pressure on the
ground1 the flow indicator flaps might fail to operate.
In this case watch the pointer of the excessive pressure gauge. If
during the opening and closing of the K1-24 hose plug the pointer of
the N-1000 pressure gauge oscillates, the apparatus functions properly.
3. Check emergency oxygen delivery by setting the emergency cock of the KH-24N
ncomonizer to the OPEN ( OT(PUTO ) position.
Note:,Nben checking emergency delivery oxygen, pressure in the system must
not drop below 8.3 atm.
Check emergency oxygen delivery by listening, bolding the end of the KM-24
hose close to the face and opening the plug in the hose.
After the entire oxygen system has been checked and if there is a sufflcient
supply of liquid oxygen available, the technician reports.to the commander on tne
nmdies of the aircraft for flight.
If the flight is cancelled for some reason, do as follows:
(a) close the KB-5 valves ahead of the pressure increase automatic units and
after the evaporator on the 1COK-30 converters;
(b) release pressure from the pipe lines through the emergency cocks c,f the
CI-2414 economizers. This done, close the valves, type KB-5, on.the aircrew
owen panels.
CAUTION. To avoid fire and accidents, it is necessary to observe the follow-
ing rules:
1. When releasing oxygen from pipe lines all the entrance hat,hes
and ports must be open.
2. The clothes of the personnel handling oxygen equiamont must be
free from oil and grease.
3. When releasing oxygen, it is prohibited to perform any other
work in the aircraft.
4. It is strictly prohibited to smoke.
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POSS/DIE FAULTS OF OXYGEN SYSTEM ARD MEANS
OF THEIR EllMINATION
The comp= system may have the following defects:
(a) leakage, clogging of the system;
(b) faults of the instruments included In the system. /eakage in the :atima &ring 30 or 60 seconds. A strong jet of gas will clean the system from
system Should be eliminated by additional tightening of the union nuts in plaeltrogen. ?
of oxygen leakage while clogging of the eunbem eboolm be eliminated by =even pPlying pressure. the pressure release valve of the MA1,30 converter
ing. est be closed while the valve after the evaporator must be open.
Checking System for leakage : 11. When scavenging the entire system with medical oxygen, cheek the
Perstion of the KRI-30 converter safety valves. For this:
To check the system for leakage, do as follows:
1. Open the hatch of the aircraft chargin pipe union. (a) close the KB-5 valves after the evaporators on the 01-30 converters;
g
2. Unscrew the plug fro the end of the aircraft c p.p. on and (b) smoothly build up a pressure of 11.0 to 11.8 atm. gauge in the KUI-30
m
onvexters from the cylinder with medical oxygen. Under this pressure the
screw the union nut of the charging hose on the charging pipe union.
afety v k
3, Connect the pipe line from the stand (a cylinder with nitrogen) to ti- alves must open operate);e-- -
other end of the -hose. (c) reduce pressure in the Kfla-30 converter vessels to one atm. gauge;
ipening the pressure release valve on one of the KRI,30 converters make sure
4. Charge the vessels of the KRI-30 converters and the pipe line with
g
nitrogen !chose purity and humidity correspond to those of medical listening that the vessels are gastight; at thin pay attention to the gas-
creasing pressure to 10 atm. gauge. Charge the system slowly from forty litre ightness of the AM oxygen level indicator pipe lines.
12. Release pressure from the taix-38 vessels and disconnect the hose from
high-pressure (150 kg/sq.cm.) cylinders. During charging the pressure release
valves of the DIX-30 converters must be closed, while the valves after the ,he aircraft charging pipe union.
evaporators ust be
13. Screw the plug on the pipe union and close the hatch.
m open.
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10. After gastightness has been cnecked up, scavenge the system. For this
. 6oee through the aircraft charging pipe union apply a pressure of 10 ate.
mmoe to the system from the ground cylinder with pure oxygen. Open the
eimiwer emergency valve at each oxygen station in turn and scavenge each
cmnacs. 1. When scavenging the gyetem with oxygen, use of fire (smoking,
EgteL With the pressure increased, take particular care to see that the
Lighting up matches, etc.) and presence of oil on pipe unions, valves
pipe lines joining the AFIK oxygen level indicator and the KIII-3t
and oxygen system units are absolutely prohibited.
converter are tightly connected.
2. To avoid accidents, open the oxygen valves slowly.
5. The charging over, disconnect the pressure source and note the indica- 3. After the test and the elimination of faults in the system,
tions of one of the pressure gauges. thoroughly wipe each connection with a Piece of clean gauze moistened
6. In no less than 12 hours measure pressure In the system again by means with rectified alcohol.
of the same pressure gauge. 4. Scavenging the system with pure oxygen should be done out-of-
The gestightness of the system is considered to be satisfactory if with doors.
the KB-5 valves open, the pressure in the system decreases not more than by
6.5 kg/sq.cm, during 12 hours or with the KB-5 valves closed, not more than
by 5.2 kg/sq.cm. of System Gastightness
Effect of Temperature Change during Check
Note: 1. When checking gastightness, take into consideration the effect d When determining the system gastightness, take into consideration pressure
- temperature. change in the system caused by a change of gas temperature in connection with
!
2. To avoid errors due to possible hysteresis of the mech ambient air temperature change.snism,
slightly-knock on. the instrument case with the finger prior to Gas pressure is directly proportional to the absolute temperature at
taking readings ,constant volume, that is on condition of complete gastightness of the system,
iressure increases at temperature rise and detreases at temperature drop. This
7. If the system is found leaky, detect the leaky place first of all. For relation is expressed by the formula:
this smear all the places of connection of the pipe line to be replaced with
.soap-suds. In the event of considerable leakage the leaky place can be detectd P1. T1 r
by the hissing of emerging gas. Examine all the connections and mark the detect Pg -TT.
ed leaky places to eliminate the leakage.
therefure
8. Slight leakage can be eliminated by careful tightening of the threaded
tightening the threaded connections of pipes made e aluminium alloy oft:enema
3amming and stripping of thread. there T and T P2 ' N-
r112
connections without releasing nitrogen from the system. Bear in mind that over
9. ln the event of considerable leakage of nitrogen completely release - absolute temperature equal to:
1 2
frjamming of threaded connections and to ensure their gastightness, use special T1 = 273 + t100; T2 . 273 +
pressure
from the system. This done, start eliminating the leakage. To elistint
oxygen-proof lubricant.
21 and p2 - gas pressure at temperatures T1 and T.2.
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ore mat be released from both.vessels in turn.
To take into ? account the influence of temperature upon pressure in the Fres:t
system, using the above given formula determine pressure to the moment of 2. To check the converters for shutting off the Vessels, do as follows:
secondary pressure (in 12 hours after the measuring of initial pressure) tui the, 10-30 converters vessels with nitrogen ander a pressure of 10 atm.
compare it with the original indications of the pressure gauge. If the diff! to
ere owzo. nen sharply release pressure from one converter through the release
of these pressures with the valves closed exceeds the amount of le -.:
e pe 0 atm. gauge. Pressure in the second Converter must not drop more than
for this time (5.2 kg/sq.cm. per 12 hours), the gastightness is insufficient.
Example of Calculation Rote: Checking the operation of the Shut-off valve should be performed with
''' atm. gauge during 5 minutes.
y 4.
the valves after the evaporators closed.
ralve
Example 1. Tnitial pressure in the system: pl . 10 kg/sq.cn.
Pressure by the reference pressure gauge: pk 6 kg per sq.cm.
Initial temperature: 11 = 20?C.
Temperature in 12 hours: t2 = 10?C.
The system is to be tested with the valves closed.
Pressure In the system with no leakage:
P T
1 2 10(273+10)
9.65 kg/sq.cm.
Pressure difference:
op . 9.65 - 6 3.65 kg/sq.cm.
, The difference thus determined is less than the permissible va
of. leakage (5.2 kg/sq.cm. per 12 hours). The system gastightness is
satisfactory.
Dmample 2. Initial pressure: pi . 10 kg/sa.sv6
Initial temperature: ti 10?C..
Temperature in 12 hours: t2 500.
Pressure as indicated by the reference pressure gauge:
pk = 3.5 kase?ca?
The system should be tested for leakage with the valves
Pressure In the system with no leakage:
T2
P2 = P1
Pressure difference:
(273+5) 10 27-?. _ 10.57 kg/sq-cm.
(273-10) 263
AP = P2 - Pk = 10.57 - 3.5 s 7.07 kg/sq.ca.
open.
Faults of E11-30 Converters
l. If the connection is found leaky, tighten it up. In case tightening is
of no affect, replace the gasket.
? 2. If the safety valve is leaky at 10 atm. or tails to operate at 11.0 -
11.8 att. gauge, remove and replace it by a new one.
3. If the automatic pressure increase valve fails (fails to close at 8.3
to 8.5 ata. gauge), replace it by a new one.
Prior to fitting the emergency valve, yeah the line up to the automatic
*tit with liquid oxygen. For this do as follows:
(a) remove the oxygen valve;
(b) plug the line after the valve running to the receiver;
lue
(c) by means of the second automatic unit increase pressure in the converter
and force out liquid oxygen through the 15-5 valve ahead of the removed automatic
considerable evaporativity of oxygen from the 10II-30 converter is
'detected, check the converters for evaporativity using the Description of the
4111-30 converters.
5. If oxygen coming out of the converter has an unpleasant smell, the
converter must be removed and washed.
6. Prior to the installation of a new converter on the aircraft in the
event of replacement of the 1111I-30 converter, check the new converter according
to the Description of the 1111-30 converter.
Washing the Vessel of 1111I-30 Converter
On detecting unpleasant smell of oxygen coming out of the 111I-30 converter
wash.end degrease the converter. For this do as follows:
1. Disconnect the pipe of the line from the KB-5 valve after the evaporator,
the pipe of the shut-off valves from the cross-piece, pipes from the pressure
release valve, the safety valve arid the oxygen level indicator transmitter.
2. Remove the converter from the-aircraft having unscrewed the attachment
thec
28: Fill the vessel with 6 litres of tetrachlorated carbon or pure gasoline;
Disassemble the converter and remove the vessel and the evaporator from
tilt the vessel and turning it round its axis during 10 minutes wash the vessel
wm14.
5. Force out the liquid with nitrogen through each pipe in turn.
6. Fill the evaporator completely with tetrachlorated carbon and then blow
out the carbon. Repeat the procedure three times.
7. After the vessel has been washed with tetrachlorated carbon wash it with
alcohol as described above. Washing with alcohol should be done not less than two
tires until the alcohol coming out of the vessel is quite transparent.
The pressure difference obtained Ap . 7.07 kg/sq.cm. exceeds the permie-:
Bible leakage (6.5 kg/sq.c.m. per 12 hours), therefore, the system is insufficied
ly gastight.
Checking Shut-Off Valves.
After the leakage test, prior to scavenging the system, check the operation
of the shut-off valves.
Check the shut-off valves with nitrogen for shutting off the 11111-.30
converters and for equalizing pressure in the converters.
1. To check the converters for pressure equalizing, do au follows: fill WA
converters with nil,rogen under a press,re of 10 atm. gauge. Then slmp
IT release pressure from one converter through the pressure release valve by an
or two atm. gauge; in this case pressure in the second converter must become
equal to that in the first one in one or two minutes.
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oxygen or nitrogen (free from oil) till DC smell is felt more. Sca "1 Uncut press ure is the amount of oxygen in grams lost per hour during the
8. After washing the vessel with alcohol thoroughly scavenge it with di7
vessel through each pipe in turn plugging the rest. due to Warm air coming from the
10. During scavenging check whether the washing is done properly. For th:s
i
vene:0 Neorage of oxygen under atmospheric pressure
. 96 The fittings and pipes should be washed with alcohol and easy de. To c k eVaporativity, fill the converter with 30 or 32 kg of liquid
hold a piece of white linen cloth against the stream of emerging gas and cheelasond weighing
eneaevadporaintfaviourti is increased due to the thermal
weigh it for the first time (during the first four
thermal capacity of the vessel). The
it for dark deposit. of the vessel should be done in 16 or 20 hours after the first
:erndoifg::::::::weight in crams divided by the time in hours between the
12. Apply paste, grade K10-22 , to all threaded connections. Dilute the Zes
ed connections. Paste, grade A710,-22,i is the evaporativity of the converter. The evaporativity must not
11. The washing completed, assemble the converter.
paste just before the assembly of threaded connections.
Service log.
To prepare the paste, fill a mortar with glycerine, add.dextreen and When checking evasorativity, take particular care to plug the non-
thoroughly grind the mixture. Then add litharge and grind it to obtain unifo recommended to fit conical plugs r: -- urn valve and the OXYGEN LEVEL INDICATOR BOTTOS (YPOSHFAEP HS)
Iota:
contains 15 gr of glycsrine, 4 gr of dextreen and 32 gr of litharge. ret
compound. pipeco union because leaks greatly increase evaporativity. It is
Cover the pipe union thread (but not the nut thread) with a thin rnif ugs of aluminium foil packing over the
layer of paste. This done, assemble the threaded connection. the operations
surface of the pipe union or flat plugs of the AUrK material.
When screwing the threaded connection again, remove the old paste from the 17. 111 rations and test results must be recorded in
13. The assembled converter should meet the following requirements: Care of EBK-30 Converter
face of the thread.
(a) the evaporator must be arranged concentrisally inside the case. The ;o calr:Lie:atmoin:Itittf:71 m:::::i2eohltewiatgei.? necessary
to subject the converters
(c) the pipes must not come. in contact with the nearest parts and each ,f 14;1:02::: the vessels the KBI-30 converters contain not less than 2 kg
rubber stops must uniformly expand the evaporator relative to the vessel;
(b) the vessel must not move crosswise or lengthwise inside the case;
other; 3. It is not recommended to leave * small amount of liquid oxygen. in the
(d) after assembly check the converter for leakage by means of dry Medice:1:2
a:III:use during evaporation in the remaining oxygen there are concentrated
oxygen or nitrogen under a pressure of 10 atm. gauge. Check the gastightssess k by in particular substances of unpleasant smell which will be absorbed
_
cloth moistened with rectified alcohol. newly filled oxygen.
4. Take care to protect the converters from oil and grease.
means of soap - suds. This done, wipe the connections with a clean piece of
' 5. Idsuid oxygen always contains lubricating oil which gets into oxygen
14. Fill the completely assembled converter with liquid oxygen and check during the production of the latter. During the service of the converters this
15. When checking the serviceability of the converter, do as follows: o:g::::leeds) Zni.01:1:essel walls therefore the vessels Should be washed
the operation of the automatic pressure increase valves.
(a) plug the non-return valve and the pipe unions QXYGESI LEVEL INDICATOR
connect a pressure gauge and the pressure release valve to the pressure releas Faults_ot Distant-BeadinG Livid-Oxygen
Indicat(z:Ze WU
TOP (PPOBHEREP BEPX ) and OXYGEN LEVEL INDICATOR BOTTOM ( YPOSHSSIFP }3j3 );
pipe communicated with atmosphere;
:
: 1. With the power supply switched on and pressure drop changed in the
(b) increase pressure in the converter. For this close the C
? n ,
_ressure rates transmitter, the indicator pointer does not move. This may take place if there is
valve and open the valves ahead of the automatic pressure increase valves;
(C) watch pressure increase; in the coverter filled by not less than
90 per cent, a pressure of up to 8.3 to 8.5 atm. gauge is reached during 3 or .
[no proper contact in plug connectors. To eliminate the fault, check the supply
Lime and repair it if broken. .
In 10 or 15 minutes after presser's increase the pipes ahead of the auto- . 2. If the instrument reading errors exceed the permissible values, tighten
5 minutes (but not in excess of 10 minutes); then pressure stops increasing. ,
. defect.
valves are closed;
up the union nuts where the pipe lines are connected to the vessel and check
the system for le or check the bunched coaductors lines and eliminate the
matte unit must get warm as at such a pressure the automatic pressure increase
(d) note the time up ? to the moment the safety valve starts bleeding; with.
Checking _ KIII-50Sonverters for
the sound automatic pressure increase valves this time must be not less than
45 minutes;
(e) in an hour after the beginning of the safety valve bleeding set the .. Eveporetivity
amount of consumption to 0.5 kg per hour. At this pressure must drop to 10 ala After the KR1-30 converters have been installed in the aircraft, as vell
gauge and lesVage through the valve will stop;
(f) set oxygen consumption to 6 kg/se.cm.; pressure In the converter mst .L. zy three months and on expiration of the guaranteed period of service life,
the tax-30 converters should be tested for evaporativity.
device as follows:
not drop below 8 atm. gauge.
16. Check the evaparativity of the converter without pressure. EVaporatiO the
. ? evaporators,
SECRET for evaporativity by
1. Close the IB-5 valves ahead of the automatic pressure units and after
means of the KY-4 testing
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2. Open the pressure release valve of both vessels of the 1111,30 conver
3. Mill the KON-30 converters with the amount of liquid oxygen required!
the next flight as prescribed in the Section "Filling the K11,30 Converters
with Liquid Oxygen',
/6 Atter filling, disconnect free the cross-piece the Ifl1,30 converteri.
pipes connecting the eint-off valves to the M1,30 converters and plug the ca
piece pipe unions.
5. From! the tee-piece in the pressure release line disconnect the pipe
rutuaing to the pressure gauge and connect the pipe union of the tee-piece to h.
of the KY-4 testing device rheometer.
T 6. In four burs after fitting the vessels with liquid oxygen close the
pressure release valve and during two hours every 15 minutes measure by the nh
meter tie amount of gas coming out of the converter (in litres per minutes).
Avemege the results of all measurings.
7. Convert the average capacity of losses tkms obtained (in litres per
minute) to units of weight (in gramma per hours) using the graph of Fig.55 tat
into account the ambient air temperature. The permissible amount of losses cam
by evaporativity is not In excess of 250 grams per hour at a temperature Of
15 ? 50C. At a temperature of 30 to 50?C the amount of losses increases by 50
to 90 grams per hour. while at temperature of -20 to -30% it decreases by
or 60 grams per hour.
8. After checking the converter for evaporativity, open the pressure role
valve1 connect the pipes joining the shut-off valves with the cross-piece on
111I-30 converter and connect the pipe running to the pressure gauge with the
tee-piece in the pressure release line.
Note: On completing the test of the vessels for evaporativity make record
in a special Log; indicate the number of the aircraft, the number d
the Ink-30 converter vessels and the amount of evaporativity.
Faults of 1(11-241 Economizer
1. If the bieh and low-pressure cavities are out of repair, replace the
apparatus by a new one.
2. In cane leakage is detected in the valve of the economizer, connect th
1116308 mask to the apparatus and make several deep inhalations. If leakage
persists, rapists? the apparatus by a serviceable one.
MAte: In the event of replacement of the 111-241 economizer prior to
installation on the aircraft cheek the economizer by the Descripth
of the 111-245 apparatus.
?
Faults of 1M-2101 Set
Repair of the 1Cn-241 economizer set involving disassembly and adjustmmit
is not permitted in field conditions. In this case the items to be repaired
should be replaced by new ones; the rems'.fed items must be sent to repair shops
Leakage in High-Fressure System
If during the bib-pressure system leakage test the pressure gauge indica!
tions decrease, the system is leaky.
Detect leaky places by means of soap-suds.
As a rule, leakage is detected by tightening up the union nuts,. However,
leakage in the ? system is sometimes caused by a leaky economizer valve. This
being the case, replace the faulty apparatus by a new one.
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LOA I
!!. 83 ?
Leakage in Low-Pressure System
In case the low-pressure system is leaky, the best way of detecting the
lea kis to divide the low-pressure system into several sections. Suppose that
the
le akr low-pressure :system is divided into three sections:
let section- from the mask (mith the mask-to-face tightness compensator
oahnectedto jt to the excessive pressure limiter.
2nd section - from the pipe union of the hose running to the Efl-23
apparatus and to the elbowed pipe union ;with a union eut of the 15,24 hose.
3misection - 11-241 economizer.
Stan checking the lst section, close the hole in the excessive pressure
li.iter lock with a hand and make a long but not deep inhalation. If it is
hmessible to inhale,the 1st section is gastight.
The check over, connect the excessive pressure limiter lock to the pipe
'mime the hose running from the K11-23 apparatus.
To cheek tee 2nd section, disconnect the K5-24 hose from the Kfl.-241
apparatus, close with a hand the hole in the hose elbowed pipe union and make.a
long but not deep inhalation. If it is impossible to inhale, the 2nd section
le gastight.
Mortise check-up, connect the 15-24 hose to the 111-241 apparatus.
Name doing so, check to see that the apparatus valve is closed, the air dilu-
tion switch handle is set to the CLOSED ( 311PUTO ) position and the 82-1311
pressure gauge pointer is at zero.
When Checking the 3rd section, make a long but not deep inhalation. If you
=mot do so, the 5r,d section is gastight. If it is possible to make an inhale-
tion, the lowpressure cavity of the Kfl-241 apparatua is leaky. This being
the came, replace the defective apparatus by a new one.
Bear in mind that the gastightness of the low-pressure system depends to a
greet extent on the condition of the rubber gaskets fitted in each joint.
Themefore, pay special attention to the joints and replace unserviceable gaskets
by new ones in due time.
Faults of 1011308 Mask
1. A faulty exhalation valve. In most cases the leakage of the exhalation
valve is caused by dust, sand and other foreign objects getting under the valve.
On detecting leakage, mash the valve with a pad moistened with clean water or
blow it with oxygen (mithout dismantling the valve and the mask). This done,
retest the valve for leakage. If the valve is still leaky, replace the mask by
a new one.
2. Leaky connections of the mask with the mask-to-face tightness compensator
aM the hose running from the 1fl-23 apparatus. In such cases replace the gaskets
and than retest the connections for leakage.
3. Leakage in the meek body. corrugated hose and mask-to-face tightness
emenermator. In sudh canes replace the mask and the mask-to-face tightness
coepsmator by serviceable ones.
4. leakage in the excessive pressure regulator valve. This being the case,
replace the meek by a serviceable one.
Faults of KM-24 Hese
Unainge in the hose (Fig.49) closed with a plug. This being the case, re-
place the gasket of the plug. If the leakage is not eliminated replace the hose
b7 a serviceable one.
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Vaults of En Plow Indicator of EK-101 Pressure
Gauge and 11-1000 EXcessive Pressure Gasge
The Blass is ticket., the body is cracked, the luminous compound baa
come off.
2. The indicator blinkers fell to react to inhalations and exhalations.
In case one of these faults is detected, replace the flow indicator or the
pressure gauges by serviceable ones.
Faults of 12-1 Valvas
1. leakage in the valves cavities.
2. Leakage in:the valves flap.
If at least one of the faults is detected, the valves dhould.he replaceda
new ones.
Vaults of Tee-Pieces with Non-Return/Nivea)
Leaky num.-return valves. To eliminate leakage, disassemble the unit with '
non-return valves, wipe the valves and seats with a piece of gauze moisteied eft
pure gasoline (without oil); at this take care to see that all, foreign particle
(white or brown deposit) are removed from the valves and seats. Next, wish all
the parts of the disassembled unit in pure gasoline (Without oil), blow them
with oxygen and assemble. Check the newly assembled unit for leakage.
If the unit with non-return valves is still leakY, do as follows:
(a) replace defective valves and seats in the wait by new ones or
(b) replace the entire unit by a serviceable one.
r 101,23
ute
. ,
Breath.
Apparatus
1. The apparatus is leaky. 4
2. Tho dieconnector operation is improper (the box of the KU-248 economi-
ser is disconnected with difficulty).
3. The nonewetraztvelve of the change-over switch is leaky (oxygen leaks
out after the disconnector has operated).
If. leant one of these faults is detected, replace the apparatus by a
serviceable one.
POS2-FLIGH2 INEIWTION
1. Open hatches (if they are closed) to obtain access to the oxygen equip'
meat.
2. Check comma pressure in the line by the pressure gauges, type ISE-1511.
3. In accessible places examine pipe lines awl their attachment, oxygen
panels and instruments.
4. check the amount of liquid oxygen remaining in the IUX-30 converter
by means of the AYH livid =men level indicators. In case of necessity
add oxygen.
5. Record pressure in the vessels of the KIII-30 converter by the pressure
gauges mounted near the KUI-30 converter.
6. Check the gastightness of all connections on the KRI-30 converters
(without releasing pressure).
7. Make sure that the safety wave is serviceable. If pressure in the appa-
ratus amounts to 11 or 11.8 atm. gauge the safety valve lust be open. If ',visor
in the apparatus is below 10 atm, gauge, the valve must be tightly closed.
SECR.Prr,
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85 -
The permissible leakage through the safety valve at 10 atm. gauge is not
iaexcess of 200 cu.cm.per minute.
s. Cheek the KR-24M economizer set. For this:
(a) carry out outside examination of the 101-248 economizer mask, mask-to-
r:Iv:tightness compensator and the P11.-24 regulator; check to see that the .
items are free from damage and moisture;
(b) check. the operation of the K11-248 economizer.
9. Close the KB-5 valves ahead of the automatic pressure increase units and
after the evaporators on the KRI-30 converters.
10. Release oxygen from the pipe lines and converters by opening the emergen-
cy cook of the KU-24M economizer at each oxygen station.
11. Close the KS-5 valves at oxygen stations.
12. Wipe the masks with a piece of gauze soaked with alcohol and place it
weber with the mask-to-face tightness compensator and the corrugated hose
into special bags located at the working stations of each member of the air-
crew.
13. Examine the parachute apparatus and send them out for storage at depots
or to, special workshops.
14. To save liquid oxygen, do not release pressure from the I01E-30 converters.
15. If it is necessary to add liquid oxygen to the IHI-30 converters, release
pressure from the apparatus.
16. If any faults are detected during the flight or inspection, eliminate
them in compliance with the Section "Possible Faults".
17. Fast eases on the 10I-248 economizers.
PBESSURE PRTRARR
Open the pressure release valves by 1/4th of the knob turn and then slowly
(during 3 or 5 minutes) open the valves completely.
Determine complete pressure release by the pressure gauge.
Note: During pressure release intensive evaporation of livid oxygen in
the vessel takes place. The amount of oxygen which has evaporated is
directly proportional to the amount of warmth absorbed by the liquid
oxygen. The maximum amount of liquid oxygen which may evaporate
during pressure release is equal to 10 kg. This corresponds to pressure
release when the apparatus vessel contains 25 or 27 kg of liquid
oxygen completely heated to the boiling point at a pressure of the
safety valve releasing. If the apparatus vessel contains less liquid
oxygen, the amount of evaporating oxygen during pressure release will
be proportionally less.
STORAGE OF 'awn OXYGEN IN 1011-30 COMIRITITS
It is permitted to store liquid oxygen in the Ent-So coavorters in sealed
Imasels under pressure and without pressure.
Storage of Liquid Oxygen in Sealed Vessels of 1U15-30
Converters
If the 1011-30 converters are filled with liquid oxygen 12 or 16 hours
before flight, it is recommended to close the KRI-30 converters in an hour
elms filling. For this close the pressure release valves and the valves after
the evaporators (if they are open).
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Leave the converters in this condition as at evaporativity of 250 grams Per
hour no emygen willbe lost during this tire because the entire arount of the
esrutb coming to the vessel from the outside will be spent for warming up liquid
oxygen.
Pressure in the vessel will increase gradually. The storage period of livid
oxygen in the lam-30 converters closed vessels without losses is from 27 to 46 hour:,
Ejle:
The less liquid oxygen is contained in the apparatus, the less is
the time of its storage in closed vessels without losses, because
less heat is required for warming up oxygen to the boiling temperature
at a pressure of safety valve releasing.
The approximate time required for increasing pressure in a closed vessel
to a pressure of safety valve releasing depending on the amount of liquid oxygen
in the converters is given in Table 16.
Table 16
Time Reqeired for Safety Valve Releasing Verona
Amount of liquid Oxygen and
Nemoorativity
Weight,
kg
Amount of
earmtn
cal.
Time required for increasing pressure to
10 atm. gauge at evaporativity, grans
per hour
150
200
250
25
360
46 hours
35 hours
27 geurs
zo
280
35 hours
25 hours
21 hours
15
210
26 hours
20 hours
16 hours
10
140
16 hours
13 hours
10 hours
Here q is the amount of warmth in calories required for beating liquid
oxygen to 10 atm. gauge.
Storage. of L__Ag.dftsgalQui in - exinserters
wider Pressure
If the 01-30 converters are in the operating condition, livid oxygen
in the 111-30 converters vessels can be stored under pressure. For this close
the KB-.5 valves ahead of the automatic pressure increase unit and after the
evaporators. Do not open the pressure release valves as during pressure release
losses may amount to 10 kg.
When storing liquid oxygen on the converters under pressure,
looses do not exceed 6 kg a day.
Storeee of Liquid Oxygen in MI-30 Converters
without Pressure
If the vessels of the 101I-30 converters are filled with liquid oxygen
tme days before flight it is recommended to leave the pressure release valves
open for 16 or 20 hours and then close the vessels, that is close the pressure
release valves.
ere,
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PRECAUTIONARY MEASURES
1. Protect all pipe line joints and apparatus elements from oil and grease.
2. When filling liquid oxygen, fence the place where oxyeen is drained
(from the pressure release valve).
3. The overalls of the personnel engaged in filling the E1Y-30 coneeeters
en and in testing the system should be clean and free free,
wggi:haseUcstainsquivhednoxy:glheling
ground near the aircraft nest be clesned from oil and
::::e:::hibited to smoke, to light matches, etc.
the apparatus with liquid oxygen and testing the system,
5. Be careful when filling the apparatus with liquid oxygen and see that no
liquid oxygen gets onto the skin to avoid frost biting (barns).
6. Prevent moisture from getting into the vessels of the 01,30 converters
mai pipe lines as on filling the vessels with liquid oxygen water is turned into
, ice, which might cause failure of the apparatus.
7. Prior to filling liquid oxygen remove the ease of the fuselage compart-
ment (in the 0-3 cabin) in the area of the pressure release drain holea.
8. Take care not to spill liquid oxygen as all organic substances moistened
with livid =Men are explosive and inflammable until oxygen is completely
evaporated.
Quality of Oxygen
Pill the vessels of the 111I-30 converters only with medical liquid oxygen.
Omeen must have a Certificate indicating whether it meets the requirements
specified by Item 2 of State Standards ( FOCT ) 6332-52
INSTRUCTIONS FOR PACKING PARACHeTES WITH
0,23 OXYGEN BREATHING APPARATUS
rig.50 shoes the position of the K1-23 oxygen breathing apparatus in rela-
tion to the seats of the aircrew members, which ensures safe and reliable dis-
, eanmection of the 0-23 apparatus disconnecters during ejection.
To prevent the breathing apparatus hoses from being broken place them into
the seats very carefully. In doing so observe the following order:
1. On the navigator's seat lay the short oxygen hose of the 0-23 breath-
fag apparatus through the weight lightening bole in the seat right-hand arm
i
That as shown in Fig.51.
COTTON. It is strictly prohibited to pass the oxygen hose through clamp 3
(Fig.51) as during ejection the snap hook of the 11-23 apparatus locking
pins may stick in the clamp. As a result the oxygen disconnecter will
fail to get disconnected and the supply will fail to change over from the
aircraft mains to the 0-23 apparatus.
2. On the pilots' seats when connecting the apparatus hoses to the aircraft
?
hoses pass the short oxygen hoses Into the seat arm rests through the
Slits in the arm rests to prevent the apparatus hoses from being broken.
3. Prior to placing the parachute on the navigator-operator's seat pass the
short hose of the 1(11-23 breathing apparatus through the hole in the rear part of
the seat pan right-hand side. If the hose is passed Into the hole of the side
atter the parachute is placed on the pan, the hose mrat be sharply beat which
emmes its rapid wear.
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4. On the gunner-radio operator's seat see that the parachute does et
towards the seat back otherwise the oeygen hose will be crumpled by the
hand arm rest.
The parachute with the KU-23 breathing apparatus is freely arraal,
the gunner's seat, &nano special instructions Oa Paski-D? are required.
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TIECTRICAL jfltilT
GEIERAL
The electrical equipment of the aircraft, model TY-16 , consists of D.C.
nd 1G. power supply sources, aircraft electric mains and electric power
Gua5ZsMaior D.C. power supply sources of the aircraft are four generators,
two PCP-18000 ? of 18-kW power each; the generators operate in parallel and
are connected to the aircraft mains to produce a total power of 72 kP, 28 - 28.5 V.
Apart from the generators, the aircraft is provided with a starter-type
storage battery, type 12CAM-55; the battery operates in parallel with the
generators and serves as a stand-by power supply source.
For A.C. power supply the aircraft is equipped with two HO-4500 inverters
Which invert direct current into alternating current of 115 V, 400 c.p.s.
The aircraft electric mains consists of wires gauging from 0.33 to 95 eq.mun
and incorporates switching equipment, as well as control and protective devices.
The mains uses mainly non-shielded and shielded wires,mark BOA, the air-
fMus being used as the minus wire. In order to lighten the weight of the
electrical equipment the D.C. electric power distribution lines are made of
ahrednium wire, mark ERR.
Direct and alternating currents are consumed by various instruments and
ulnas provided with remove control facilities, as well as by complex automatic
'systems (the autopilot, cannon system fuel quantity and flow gauging equipment,
etc.), signalization means, heating, de-icing, illumination equipment and radio
equipment.
The aircraft electric mains is connected to ground power supply sources
through two ground-supply plug connectors; one of the plug connectors is used
for connecting D.C. ground supply sources, whereas the other - for connecting
A.C. ground power supply sources.
laacivaor illthatIC MAINS
The entire electric mains system of the aircraft consists or two major sections:
1. The D.C. circuit of 28-28.5 V supplied from the PCP-18000 generators
and the storage battery, which is connected for buffer operation with :1-,e genera-
t
ore
.
2. The single-phase k.C. circuit of 115 V, 400 c.p.s. which is supplied
from the operating or stand-by inverter, type HO-4500.
To ensure effective all-condition operation of the aircia_ft, the D.C. cir-
cuit is divided into three subcircuits:
(a) the normal supply circuit;
(b) the emernency supply circuit;
(e) the dual supply circuit.
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As a rule, connected to the normal supply circuit are all the four genera.
tors and the storage battery (Fig.52). The generators and the storage battery
are connected separately and therefore they may be connected to the normal
supply circuit in any combination, for example: one generator and the storage
battery or two generators and the storage battery, and so on.
Connected to the emergency supply circuit can be only one generator (cite
generator 2 dnstalled on the left engine, or generator 3 installed on the rigm
engine) and the storage battery.
With the aid of svdtchiee contactors, type RH (KU-200A R1I-400A), the
dual supply circuit is automatically connected either to the normal supply cir.
cult (in case it is energized) or to the emergency supply circuit if the normW
supply circuit is de-energized.
The schematic distribution diagram of D.C. supply system (of the aircraft
mains system) is presented in Fig.53.
The normal, emergency and dual supply mains provide power supply to three
groups of distribution busbars;
1. The normal supply busbars which are connected only to thenormal STIPA
circuit.
2. The dual supply busbars connected to the dual supply circuit.
3. The triple supply busbar which is usually connected through a special
change-over switch to the dual supply circuit and, consequently, is energized
free the normal or emergency supply circuit. In case of failure of the normal
and emergency supply circuits this busbar is manually reset for direct supply
from the storage battery.
The distribution busbars have no direct connection to the emergency supply
circuit.
The normal supply busbars feed such power consumers which are necessary
for normal operation of the aircraft but which can be done without in emergency
conditions. Such power consumers are: the autopilot, de-jeers, heaters, venti-
lators, camera equipment, part of the illumination system, etc.
The dual supply busbars feed such-consumers which make it possible to .
fulfil the mission and to return to the home airfield even in case of the faulty
normal supply circuit. Such power consumers are: the bombing system, flight
control and navigating instruments, fuel system pumps, landing flap actuator,
L.G. warning system, part of illumination system, etc.
The triple supply busbar (the busbar rhich provides battery supply of the
instruments with the mains de-energized) supplies voltage only to such power
consumers which are absoletely necessary for accomplishment of a forced lending
of the aircraft in case of failure of the normal and emergency power supply
circuits. These consumers are: the main gyro horizon, bank-and-turn indicator
of the pilot, remote indicating astrocompass, type AAK-AE-5 , heater of the
upper left pitot tube, type 111-156 , circuit No.1 of the interphone syster
and the emergency illeeleetion system (the ultra-violet illumination lamps of
the pilot's and navigator's instruments penels, the receptacle of the pilot's
extension lamp and the illumination system of the KM-12 compasses), automatic
brake control unit, drag chute system, engine blow-off band control system, CO2
bottle control gystem, fuel shut-off and stopcock control system and radio
station,type P057-311.
Three consumers: the feeder of the in-flight engine starting, the feeder of
the top emergency bomb dropping system and the radar transponder destructor feein
are connected directly to the storage battery and may be used at any moment vita
additional switching and change-ever operations on the power supply sources.
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additional switchieg and change-over operations on the power supply sources.
Operating Duties of Electric Mains
Iselew of the necessity of voltage supply to some power consumers even in
conditions when separate sections of the electric supply mains are damaged the
D.C. electric power distribution system is designed to allow three operating
dutieS
- emergency;
- de-energized mains duty, when only consumers of vital importance are
coreit eb:::::I.
In operating duty the electric mains, as a rule,
se:mecca all the four generators and the storage battery. /n this case energized
are the busbars of the normal supply circuit, the busbars of the dual supply
.eiedtotitie
circuit and the busbar which supplies the instruments from the battery with
the mains de-energized (the triple supply busbar).
To select the normal operating duty, it is necessary that the switefies and
selectors located on the generator control panel (Nig. 55) at the radar operator's
station should be placed to the following positions:
1. The switches of all the four generators and the battery-to-normal supply
. circuit blocking switch should be ON.
2. The storage battery change-over switch should be thrown to NORMAL
( HOPILINIO ).
3. The voltmeter change-over switch should be turned to NORMAL SUPPLY CIRCUIT
(HOPALLEHAA CETI, ).
4. The emergency supply circuit switch should be in the OFF position.
5. The change-over switch connecting the generators to the emergency supply
system (bearing the inscription FROM GENERATOR COT PEREPATOPA ) should be
placed to LEFT No.2 ( )
6. The change-over switch bearing the inscription BATT= sums OF
ElEIRGENCY riSTRUCTIMD3(BKAMEHME ABAPOINI OFREOPOB HA !MAW OT AKKAMTOPA)
should be thrown to OFF.
7. The switch with the label GROUND sum (ASPOZPOMUCE WAVE ) should
be OFF.
Note: The storage battery blocking switch is rigidly fixed to the generat,,
emergency switch connecting bar; this means that when at least one
of the generator switches is ON, the storage battery blocking switch
is also engaged.
In case of failure of part of the generators, connected to the normal supply
circuit may be three, two or even one generator in combination with the storage
battery. When connected to the normal supply circuit are three generators plus
the storage battery, the number of connected consumers is unlimited, that is,
the flight weo
::2s:nc:ntinued in the same conditions, as if all the four genera-
torsIn case the normal supply circuit connects only two genera-
tors plus the storage battery connected simultaneously may be either the cannon
'Totem with continuously Operating consumers or the tail unit de-icers with
continuously operating power consumers. /t is forbidden to connect the cannon
gstem and the tail unit de-icer system simultaneously. When it is only the
codbination of one generator and the storage battery which is connected to the
noreal supply circuit, the total number of power consumers connected should
ensure that the total load does not exceed 600 A.
Rvcip
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Emergency duty. In case a shorting' appears in the normal supply mains (Um
trouble will be indicated by beyond,scale movement of the ammeter needles and k
decreased,voltage indications of the voltmeter) or in case of another trouble
which requires disconnection from the normal supply circuit, the radar operator
should quickly select the emergency supply circuit which is de-energized in
the normal operating duty serving as a stand-by circuit.
When flying with the supply males in emergency duty, the circuit in opera.
tion connects one of the two generators (generator 2 on the left engine or
generator 3 on the right engine) and the storage battery. In this case energized
are: the emergency supply circuit, the dual supply basbar and the triple etriply
busbar. The normal supply circuit and its busbars are disconnected and de-ener.
gized.
To change from the normal to the emergency operating duty the following mrk
log should be done on the generator control panel at the radar operator's static:
(Fig.54):
1. Operate the generator emergency disconnection lever to disengage all Us
four generators and the storage battery from the morsel supply circuit.
2. Turn the emergency supply circuit switch ON.
3. Place the voltmeter change-over switch to the EMERGENCY SUPPLY CIRCUIT
position.
4. As a result (See the Diagram in Fig.54):
(a) the storage battery will get disconnected from the normal supply cir-
cuit;
(b) all the four main differential undercurrent relays, type 201P-600 ,
Will disconnect the generators from the normal supply circuit;
(c) generator No.2 will become connected to the emergency supply circuit
through its additional relay, type 1MP-600.
When sure (by the ammeter and voltmeter readings) that the emergency suplkt
circuit and generator No.2 operate normally, the storage battery change-over
'switch should be placed to the EMERGENCY POSITION (ABAPHRAO ); this action will
connect the. storage battery to the emergency supply circuit for buffer operatim
with the generator.'
Notes: 1. In case le 't generator No.2 or its circuit is faulty, the
generator change-over switch should be turned to the RIOT No.3
?MEE le3 ) position. In this position connected to the
emergency supply circuit instead of generator No.2 (installed as
the left engine) will be generator No.3 located on the right
engine.
2. At the moment of the emergency supply circuit selection it is
necessary to disconnect the inverter, type 110,4500 , so as not
to overload the generator with large starting comments during
its connection to the circuit. Upon engagement of the generator
it is necessary to re-engage the inverter.
In the course of emergency-duty flying it is allowed to use only those
power consumers which are connected to the dual supply busbers (See Table 17)
and to the triple supply busbor (See Table 18).Under these conditions the MIX
time has no specific limitations.
In case the emergency suprazr system is faulty it is necessary to select the
de-energized mains operating duty.
De-energized maim: operating duty. Under the headlined duty conditions the
normal and emergency supply circuits will be de-energized, and the storage
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battery sill supply only those consumers which are vitally important for flight
Estimation (See Table 18). The following operations should be carried out on
us generator control panel at the radar operator's station to select the duty
ingl:T
ticanur on the change-over switch labelled BATTERY SUPPLY OF MERGENCE
uszwaras (BEEMEHME ABAPEZEUX HPEEOPOB EA MITAHRE OT AKEYEYAHTOPA ).
2. Turn the emergency supply circuit switch off.
3. Turn the storage battery switch off.
4. Turn off the switches of the four generators and the blocking :Twitch of
the storage battery.
5. Turn the voltmeter change-over switch to STORAGE BATTER! (AKEYWIHTOP).
CAUTION: The storage battery, type 12-CA.W.55, is capable of supplying the
instruments listed in Table. 18 for not longer than two hours.
Table 17
Consumers Connected to Dual Supply Busbar
lo.
Description
Protector of
consumer and
type of fuse
Narking of
feeder
2
3
4
1
Noel flow controller, left
130-5.
AT
2
Fuel flow controller, right
130-5
AA
3
Boni) emergency dropping control
130-5
BA
A
Electric bomb release supply (release
of boebs armed)
130-15
BB
5
ARMED-SAFE system
130-10
BB1
6
Armed emergency dropping system
130-10
BB2
7
Fuze circuits, left front
011-5
BBs
8
Fuze circuits, right front
011-5
BEd
9
Fuze circuits, left rear
011-5
BBr
160
Fuze circuits, right rear
011-5
BBr
3.1
Boob emergency dropping control relay
130-2
BE
12
Bomb emergency dropping control relay
130-2
BE
13
Armed bomb release blocking relay
130-2
BA
14
Armed bomb release blocking relay
130-2
BM
15
Emergency bomb dropping system
supply
E-50
BE
16
Sight supply
130-15
BA
17
Supply of bomb release variant
selector box, type 11305-48
130-5
BP
18
Rear adapter disconnecting relay
130-2
130
19
Starting system supply
k30-25
+3
20
Air cock of left engine
130-5
13A
21
Left engine starting system
A30-15
13B
22
left engine starting system control
130-5
13H
23
left engine ignition system
ASC-20
1311
24
Air cock of right engine
A3C-5
.
2aA
25
Right engine starting system
A30-15
23B
RF.CTR.RT
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1
2
3
4 '
`61'1Ntt
33
34
35
36
37
38
39
40
41
42
43
44
45
46
.47
48
48
50
51
52
53
54
55
56
57
58
59
60
63.
62
Might engine starting system control
Right engine ignition system
Inverter. type 110-4500, stand-by
Fuel pump of left teak No.19
Flie/ pump of left tank No.16
Pixel pump of left tank No.10
Fuel pump of left tank No.2
Fuel pump of right tank No.3
Fuel pump of left tank No.4
Fuel pump of right tank No.5 .
Fuel puRp of left tank No.6
Fuel pump of right tank No.6
Fuel pump of right tank No.10
Fuel pump of right tank No.16
Nei pump of right tank No.19
Fuel, stopcock of left engine
Fuel stopcock of right engine
Fuel abut-off cock
Air position indicator (dead reckon-
ing computer system , type HI-50B
Flap actuator, electric motor No.1
Flap actuator, electric motor No.2
Ultra-violet illumination of pilot's
instrument penal and overhead
electric control board
Directional gyro of pilot
Gyro horizon set of pilot
Gyro horizon set and directional gyro
of co-pilot
Three-pointer indicator, type 3A8-3P,
of right engine
Fuel quantity gauge of left engine
tanks
Fuel quantity gauge of right engine
tanks.
Fuel flow gauge of left engine tanks
Fuel flow gauge of right engine
tanks
Wel pressure gauge
Three-pointer indicator, type mg-3p,
of left engine
Rank-and-turn Indicator of co-pilot
Flap position and free air tempera-
ture irdicator
Range-tinder, type CA-1
Radio compass, type APH-5, No.].
Radio compass, type AP-5, No.2
13C-15
I3C-20
tH-200
161-45
08-50
IH-75
II-75
II-75
IH-75
HP-75
101,50
II-50
HI-75
HI-50
NE-15
120-5
ABC-5
I3C-5
ABC-5
011-150
.111-150
130-2
130-5
130-5
13C-5
13C-2
13C-2
A3C-2
A3C-2
130-2
13C-2
13C-2
13C-2
2311
MBa
MEd
06e
MHz
O0500100
UI
us
ur
113
ill
Us
110
111
un
PH
PH
PR
Crorvcrh-wri
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2
63
1.18 equipment
A3C-10
PM
64
Aircraft transponder
130 -5
PO
65
Radar bomb sight, type PEE-4,
(control)
A3C-20
PM
66
Command radio station, type PCMY-3M
13C-5
PB
Antenna duplexer of radar altimeters,
types FB-2 and rs-17
130-2
PH
68
Left tank group fuel pump warning
system
13C-2
CS
69
Bombing equipment warning system
ABC-5
CB
70
Right tank group fuel pump warning
system
ABC-2
CA
71
Hydraulic system warning unit
130-2
cr
72
Cabin sound warning system
130-2
CB
73
Mack limit warning system
A3C-2
CM
74
Differential pressure warning unit of
front cabin
13C-2
CO .
75
Fire warning unit of left tank group
13C-15
CU
76
Fire warning unit of right tank group
13C-15
CF
77
Follow-the-leader bombing procedure
ABC-15
co
Lampe
78
Colour flare bomb normal release
system
ABC-20'Cl
79
Colour flare bomb bay doors warning
system and release control interlock
IBC-2
CII
80
Colour flare bomb emergency dropping
system
0E730
Cq
81
Colour flare bomb station status
indicator
130-2
Cr131
62
L.G. warning system'
130-2
CII
83
Colour flare bomb emergency dropping
control
ABC-2
en
84
Heaters of Pitot tube of co-pilot,
radar operator, radio operator,
110-50E air position indicator
and ORB-lip sight
ABC-10
Tn
85
Control of stand-by pumps of tanks
1BC-2
YEA
No.16
86
Control of stand-by pumps of tank
13C-2
EEC
No.6
87
Remote-indicating compass
ABC-2
YR
88
002 bottle control
ILBC-10
YE
89
Emergency fuel jettison valve system
ABC-2
YI
go
Control of stand-by inverter,
type 1104500
130-2
? YV21
91
Bomb bay doors control (normal)
13C-5
71
32
Bomb bay doors control (emergency)
ABC-5
sn
93
Flap controltelectric motor No.2
13C-5
EY
94
Fuel flow control
130-2
Yq
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3
4
95
Flap control, electric motor No.1
ASC-5
Control of first fuel pump group of
left engine
130-5
Yn1
Control of first fuel pump group of
right engine
A3C-5
782
98
Control of second group fuel pumps
ASC-5
783
99
Control of third group fuel pumps
A3C-5
734
100
Control of fourth group fuel pumps
13C-5
735
Table 18
Consumers Connected to Triple Supply Busbar for Storage Battery
Supply of Instruments in Case of De-Energized Mains
No.
Description
Protector of
consumer and
type of fuze
? Marking
of
feeder
1
Emergency ultra-violet illumination of
front cabin and illumination of
A3C-5
OA
K11-12 compasses
2
Gyro horizon set, master
A3C-5
SE
3
Bank-and-turn indicator of pilot
A3C-2
nn
4
Interphone system channel No.1
13C-5
PA1
5
Interphone sets 0117-10
13C-2
PA-20
6
Heaters of TE-156 Pitot tube of pilot,
navigator and velocity head warning
unit CON-3
A3C-5
TM
Automatic brake control unit
A3C-10
a
Engine blow-off band control system
9
CO2 bottle control system
A3C-10
YE
10
Drag chute control system
A3C-5
IC
11
Fuel Shut-off and stopcocks control
system
130 -5
M6
12
Radio station, type PCMY-3M
A3C-5
P7
13
Radio transponder destructor
No protec-
tion
3131
14
De-energized mains bomb release
No protec-
tion
30
15
In-flight engine starting system
No protec-
tion
311
Protection of Electric Maine,
The electric mains of the aircraft is built up of separate feeders.
Termed "feeder" is a single consumer or a group of power consumers supplied
through a separate protective device (a circuit breaker or fusible cutout).
The following protective devices are used for protection of the aircraft
mains and power consumers:
(1) Automatic circuit breakers of 130 family,
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(25) Glass fuses of CII family.
Deleyedct1On fuaee of 1111 family.
(4)1ggh-heat fuses of TS family.
Automatic circuit breakers of 120 type (Fig. 56) are employed for automatic
dUcomuction of electric power consumers, as well as for protection of electric
cagaiSet dangerous over-loads and short circuits in electric circuits. The
'circuit breakers can be used for manual on-off switching operations on electric
!circuite, in which case they function as ordinary single-pole switches. However,
-the limpet part of the circuit breakers installed in the aircraft act as fuses,
cad
therefore they Should be always turned on before each flight and held in
de minion throughout the entire flight. The automatic circuit breaker is
manually by its operating handle, In overload and short-circuit condi-
time the circuit breaker is cut out automatically; minder normal loading condi-
thaw the circuit breaker is disengaged manually.
The circuit breakers are mounted In D.C. circuits with nominal voltage of
28T,elia rule, in locations where they are easily accessible in flight. The
/Anodize range of automatic circuit breakers is used on the aircraft: 13072,
130-5, 130-10 A30.15 130-20 . 130-25 130-30 . 13C-41) and e30-50
(the hyphenated figure indicates the nominal voltage the circuit breaker is
rated for).
ems, types an, in and TS (Fig.57), are designed for protecting
electric units from short-circuit currents and continuous, although small over-
loads. Delayed-action fuses ensure normal protection and at the same time with-
Steed instantaneous current surges(300% and even 600% of rated currents) which
are characteristic for the operation of some electric units.
Immo, type CH are installed in A.C. circuits, in permanent-load D.C.
circuits, and at places difficult for 1w-flight access.
Fuses, types Ian and TM , are installed in electric actuator supply cir-
Mae and are also used for group protection of the electric power distribution
Fetes and for the generators protection (See Figs 53 and 5e).
Fiume of all the usable types are mounted on the aircraft in various-type
boxes. The following ranges of fuses are used on the aircraft: CH-la , CH-2a,
1111-5, 11E-10 HH-15.1411-30 14E-35-2, 1111-75, 1111-100
; E-200, 1411-250, T11600 and T11-900 (the hyphenated figure demotes
the nominal voltage the fuse is rated for).
les.: Fuses, type EU, which have polarity marking should be installed in
cespliance with the polarity identification, i.e. attaching the fuse
to the supply busbar with its book lug which corresponds to the plus
sign marked on the fuse cap. This is a must, as the operating
characteristic of these fuses depend on the polarity of the current
applied to them.
; For the arrangement and layout of the protective devices on the panels and
1:091,13 see Figs 58, 59, 60, 61, 62 and 63. The general layout diagram of the
airc149it protective devices is presented in Fig.64.
1116
The electric mains of the aircraft consists of wires, marks MX and
Inm, coated with coloured insulation, and, of aluminium wire, mark EHBAA,
mithiffiite insulation.
All the wires belonging to the armament system are of red colour, thosc
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of the radio equipment system - of light blue colour, the A.C. mains wire
coloured yellow, and all the other wires are of white colour.
To ensure radio interference suppression, part of copper wires usede
Shielded (wire, mark EID3X3 ). For the same reason, part of copper wires e
encased in common anti-interference braidings.
The wires are fitted in the terminal lugs, individual connectors me,
connections by means of upsetting, while their connection to plug conneee
terminals, to warning light fittings, miniature relays and other bistro:le
effected by soldering, use being made of 1100-40 or HOC-50 solders and m
For the types of wire fittings and terminations used ou the aircraft seek
in Fig.85.
The wires of the aircraft electric mains are coded in letters and fie.
Bach Wire Should be coded over its entire length every 400 - 500 mm, and
bear at least six code markings every 50 mm by the wire end. Wires, markl
are coded only at their ends: three code markings every 50 mm. Apart from;
put on the end of each wire prior to its fitting are vinyl pipes carrying;
wire identification marking.
Sires and vinyl pipes are marked in 181-52 paint with the aid of meh.
stamps, the markings procedure being as follows:
1. Prior to marking an electric wire or vinyl pipe, clean the wire oe
surface from moisture and duet using a clean cloth for this purpose.
2. Stir up the KU-52 paint and pour it on to a felt pad (State Steak
rOG1 288-53 ) contained in a metal case.
3. Inspect the stamp and in case it is fouled wash it in rectified ab
- 4. Coat the stamp with the paint covering the pad and mark the wirer
vinyl pipe.
5. The wire or the vinyl pipe marked, dry it during 20 to 30 minutest
a temperature of 15 to 20?C.
The plotted markings should be well discernible. The marking may bek
with use of special devices, or with the aid of an automatic wire marker, it
available.
Note: It is allowed not to mark the following wires:
(a) in bonding jumpers;
(b) in internal wiring jumpers of control boards, boxes, inse
panels and other unite if the wire does not run out of
respective unit and if the wire length does not exceed le
(c) all wires whose length does not exceed 200 mm;
(d) wires connecting electric units to the airframe if it lap
to trace than over their entire lengths from the unit to 0
structural member they lead to.
In conditions noted in Points (b) and (c) it will be the vinyl pipes%
the end of each wire which are to be marked.
Separate wires of the aircraft electric system are ganged in bunches0
"lunched conductors") with the aid of thread bandages. The bunched condue:
numerical or compound numerical and letter mareenes which are placed on me
rings fitted around the bunched conductors.
Metal tags are provided at points where the bunched conductors are be
out of the electric units and over the entire length of the bunched condee
at points most accessible for inspection. No tags are attached to bunched
conductors of smaller-than-10-me diameter.
Used as connecting links between separate wires and bunched conductor!
_
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cables. The term "cable" is used for e single wire or a group of wires which
irtercomeee any two electric units. Cables have letter and numerical markings;
identification letters stand for: .
I - cables of the front pressurized cabin;
U _ cables of the centre plane and the non-pressurized section of the
fuselage;
II - cables of the left outer wing panel;
III -cables of the right outer wing panel;
C -cables of the roar pressurized cabin;
HI - cables of the left engine;
MO - cables; of the right engine;
- cables of the tail unit.
The figure which follows the identifying letter denotes the ordinal number
of the cable for the given electric unit of the aircraft.
The abovementioned cable designations are indicated in all the feeder and
Wm:natio miring diagrams available, but as a matter of fact these designations
are present on the aircraft only in case of a single-cable conductor: the
maductor tag in this case reads the cable designation. In all other cases, when
bunched conductors consist of several cables, the identification tags carry only
mmerical data to indicatcethe_line number of the given bunched conductor on the
aircraft.
lazingeandeRemoving the Cables
Then laying or removing cables, keep it in mind that the electric system
is built up as a single-wire circuit, the airframe being used as the minus wire.
The single-wire circuit sets fourth the following requirements:
1. The plus wire should be insulated with utmost thoroughness. Any' contact
of an energized current-carrying element (wire lugs, plug connector terminals
and the like) with the airframe results in short-circuiting.
2. The minus vire of the electric equipment should be reliably connected
to the airframe. The connection should ensure minimum contact resistance (not
inercess of 100 microohms) whi,da is accomplished by cleaning the contact points
fromedelectric coatings and by secure attachment of the minus wire lug to the
airframe.
3. The insulator maximum resistance of the aircraft Mars relative to the
eeereme is the requirement to be fulfilled. For each feeder the insulator
resistance of the plum wire eat the relative air humidity of e0e) should not
be mealier than:
(a) 10 megohms if the feeder supplies up to three consumers;
(b) 8 megohms if the feeder supplies more than three consumers;
(c) the insulator resistance of the electric power distribution system wires
should not be smaller than 1 megohm.
CAUTION. BEVER lay or remove wires when the electriC system is energized.
Wires with damaged insulation are subject to replacement. To replace a
vire:
1. Disconnect the damaged wire from the equipment.
2. Slacken the bunched conductor attachment yokes end loosen all the
thread bandages on the section of the wire to be replaced.
3. Withdraw the damaged wire and lay the new one. The gauge, colour and
its Earl:bag of the newly laid wire should be identical to those of the replaced
vire.
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4. Re-bandage the conductor with Mackay threads and fasten up all the
Blackened yokes of the bunched conductor.
In case of rupture or partial replacement of wires gauging from 0.35 to
8.8 eq.mm it is allowed to joint the wire ends by means of fixed connections
shown in Fig.66. It is not recommended to butt-joint wires gauging over 8.8e
As an exceptional measure, it is allowed to couple the wires bv way of fitte
the wire ends in terminal lugs with successive jointing of the lugs with thst
of a bolt and a nut; the jointing over, the connection should be thoroughly
insulated with a vinyl pipe and-vinyl tape.
In case all the wires of a bunched conductors are damaged, and the dame
portion of a separate mire constitutes not less than 100 mm, the defective
bunched Conductor 'should be removed and replaced. The new beached conductor
should be made according to the respective Drawing or to the model of the
bunched conductor to be rep/aced. To make a new bunched conductor:
(a) prepare and mark the required quantity of wires of corresponding gav
and colours;
(b) collect together and bind the wires in a bunch according to the moi
of the damaged bunched conductor;
(o) put vinyl pipes with respective marking on the wire ends:
(d) carry out termination of the wire ends.
The wire or the bunched conductor replaced, identify it with the aid Off
testing lamp or a voltmeter; then, referring to the feeder diagrams, check th
Insulator resistance of each feeder comprising the repaired bunched conductor.
For examples on circuits, for testing separate sections- of the electric sYsta
see Figs 67 and 68.
Note: nom checking by the diagram presented in Fig.68 the method of
connecting the megolommeter is the same as when testing with employ-
ment of the circuit presented in Fig.67.
The capacitors the puncture voltage rating of which is smaller than the
voltage developed by the megohmmeter should be disconnected and tested separat
ly.
When testing the continuity of the electric circuit of any electric uaR,
it is necessary to insulate the circuit from all the other electric circuits.
Before connecting the minus wires to the airframe, the contact place on
the structural member should be thoroughly cleaned from its protective coati*
this done, the lugs of the mimes wires should be tightly bolted to the airfras
and painted red.
Electric Wire Maintenance
Atter every two or three flights all the electric wires must be inspecte
and all the faults detected should be corrected.
The electric wire maintenance procedure consists of the following operinMe
1. Wipe dry the wires covered with oil or hydraulic mixture. Fasten up CO
loose attachment fittings of shielded bunched conductors to prevent radio Late
ference which is likely to appear due to insufficient tightness of attachment.
2. Check plug connectors for secure coupling and lock their union nuts.
3. Check the through bolts in power leads, pressurized cabin bottoms and
contact blocks for secure-fastening.
. 4. Check all the minus wire-to-airframe contact points. If the red lock-
ing paint is deteriorated, it is required to tighten up the attachment screw,
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et,check the Contact resistance value which should not exceed 100 microohms,
e4ee1y red paint to the contact point.
Rote: When wires gauging 5.15 sq.mm and heavier are attached to the air*
frame, the wire lug se/face contacting the airframe structural member
should be coated with a layer of anti-corrosion paste used in alumi-
nium wire fittings; this done, it is necessary to reliably secure the
lug, to wipe the place dry all around, to check the contact resistance
value and to apply red locking paint.
5, When replacing a separate aircraft unit, make sure that the contact
mistime? of the newly installed unit does not exceed the value specified in
Maintenance of Junction Boxes and Electric Control
Boards
Electric power is distributed within the aircraft electric system through
different distribution arrangements (electric control boards, panels and junction
ems) whie.h are provided with various-kind switching, control and protection
smipment. The layout of electric control boards and panels, as well as of
jtmetion boxes, is presented in Fig.70, (a) and (b).
knee a prolonged period of operation or parking of the aircraft it is
meessar7 to check all the junction boxes, as well as electric control boards
=droll panels, the check-out procedure running as follows:
1. Check the cover locks for intactnets and reliability.
2. Check the condition of wires' insulation at points where they are inserted
into their boxes and electric control boards; inspect for adequate wire tereina-
tion.
3. Check the contacts for reliable coupling. Use a nut wrench to tighten up
the nuts on contact bolts of plus and minus connections.
4. Check and, if such a necessity arises, tighten up the contact connections
on the on-off and change-over switches, circuit breakers, etc.
5. Check the switching arrangements (on-off switches, change-over switches,
deostats, relays, buttons, contactors and the like) for secure attachment and
seed operation.
6. Remove dust, dirt or moisture from the junction box or the electric
control board and wipe it with a dry cloth.
7. Use a dry cloth to clean those portions of the supply busbars which bear
traces of oxidation or dust.
8. Check all the fuses indicated in the attached diagram for availability,
their integrity and for meeting the current intensity rating requirements, as
Bell as for secure fitting of the Cal -type fuses in their holders. If it is
revealed that some fuses are missing or faulty, mount or replace the fuses.
9. Inspection over, close the cover of the box, panel or electric control
hoard and lock the cover, if it was not locked before the inspection.
eAUTION1 Never repair or check units mounted in junction boxes, electric
control boards and panels when the aircraft electric system is energized.
Specific Features of Aluminium Wire 'Maintenance'
With the view to lightening up the aircraft weight, the electric power
distribution system is wired principally with aluminium wire, mark MIA ,
from 35.0 to 95.0 sq.mm. The current-carrying core of -.Lose wires is
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rade of the material, grade AT, and consists of separate (twisted together)
mimes the gauge of which (1.08 to 1.4 sq.mm) is much heavier than that of co
, In clean and dry air it is characteristic of aluminium wires to get coo(
with a thin non-conductive oxide film which prevents the metal from further
oxidation. However, moisture and gas contaminated air may become a favourable lira gauge'
medium for intensive electro-chemical corrosion of aluminium. Apart from thu, sq.nal A
when in contact with some metals or alloys (copper, for example), aluminium
makes up a couple prone to intensive corrosion. 5.6
Oxide film on the aluminium surface adversely affects the contact betwes 35 12
5.2 50 :
12 6.8 7.2
the wire conductors and between the wire and the lug which may result in vol M 70 16 7.2 - 7.6
drop and exceasiwe overheating at the wire termination point.16
95 8.2 - 8.6
In order to preclude the probability of oxide film formation and corrosie
the ends of aluminium wires are sealed by upsetting wire ends in special cop A 7. Measure the contact resistance between the upset lug and the wire using
lugs which are hot-soldered (to obtain a heavier coating) and are filled with ameter and millivoltmeter according to the circuit diagram presented in
special anti-corrosion paste (a mixture of petrolatum with zinc powder). gl'l To
:(b) using an excitation rheostat, determine the intensity of current
71 wire under check to a D.C. power supply source with rated
CADTIONi NEVER use electrolytically treated copper lugs with holes for
adtage of 26 to 28.5 V and power not exceeding 2.5 kw;
aluminium wire termination.
Upon upsetting the terminal lug and checking the contact resistance, ta
:through reference to the ammeter) flowing in the wire. The intensity should not
tare portion of the wire is to be wrapped with sealing tape, mark 720A.
!need 140, 180, 200 and 225 A for wires gauging 35, 50, 70 aui 95 sq.mm,
Next to their terminal lugs the aluminium wires are provided with identt
tion marking: a red ring on a vinyl pipe or a red vinyl pipe fitted on theWh'esitTeplja:
e one of the probes of the milliveltmeter in the middle of the
If it occurs that in the course of operation the terminal lug
of an alu'lipset portion of the lug and connect the other probe to the yoke fitted around
wire brc ks or the lug gets out of contact with the wire, the repeated wire M
he bum section of the wire:
should be carried out as follows: ; (d) calculate the contact resistance according to the formula: R = V//,
1. Remove the Wire from the aircraft. 'here I is the current passing in the wire at the moment the measurement is
Note: It is allowed to terminate (fit) alnminium wires directly on the eirgtaken (as read by the ammeter), and V is the voltage drop in the wire
aircraft only in top urgency eases, for example, when removal of terehmtion point (as indicated by the millivoltmeter). The contact resistance
wire from the aircraft calls for large-scope domairiting operatimutould mM exceed the limits specified in Table 20. If the contact resistance
other equipment. till amasses the indicated limits, the lug should be cut off, the wire should
e terminated anew, and contact resistance should be checked once more.
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Table 19
Key to Values Indicated in Fig.71
Dimension, ma
13 - 15
15 - 12
16 - 18
17 - 20
2. Cut off the broken conductors of the wire or the defective lug, and
remove the insulation from the wire end, having previously shifted the vinyl p:
with the label along the wire. The insulation should be removed from the win
only with the aid of an electrothermal tool since no cuts and other mechanical
Tolerated Contact Resistance Values for Aluminium Mire Lug
Table 20
Terminations and Tolerated Bend Radii of These
damage
tolerated on the wire conductors.
are
3. Having stripped the wireend, coat it from outside with a thin layer_
of anti-corrosion paste and then Clean it with a special metal brush to mule
nre gauge,n
oxide film from the wire conductors.
4. Half-fill the lug sleeve with anti-corrosion paste (to expel' air fra
it) and fit the lug onto the wire.
5. Using a special device for fitting aluminium wires of the given gauge,
upset the lug on the wire.
6. Check the degree of lug upsetting; the dimensions of the pressed real
(Fig.71) should be within the limits specified in Table 19 below.
om
35
50
70
95
Wires
of wire, mm
Contact resistance Tolerated bend radius
t(empinmieractrurooshrisa) at
20 to 22 C 1 2
up to 20
up to 15
up to 12
up to 10
50
60
100
150
30
40
60
100
8. Use a clean piece of cloth or gauze to remove superfluous anti-corrosion
este from the portion to be tapesealed. Tightly wrap tape, mark 20 A, around
t,ca Portion of the wire until completely covered, and then use 'a 10-mm
alle tale to wrap around the lug and the insulation so that the tape would
!erlsp them 2 to 3 mm. Cover to tape surface with talc and fit the vinyl pipe
"tit the tag over the lug.
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Das Due to the fact that aluminium is an eenily corroded material, au
the operations covering the lug-to-wire fitting and sealing of the /t should be always remembered that poor bonding in any aircraft system
termination should be completed during not longer than one hour. aeates even heavier radio interference (due to appearance of additional wart-
9. Record the work done in the aircraft Service LeGs with fadicat oftible contacte) than a completely mnbonded system. Therefore, onlike other
content resistance of the fitted lug. 'pes of aircraft mountings, the bonding system requires constant care. This
10. Mount the wire on the aircraft. In view of the fact that al um wilocasayrefers to the engine bonding systema which should be paid the great-
ate much leas flexible than copper wires,. no sharp bends should be invol ust.ettention since the engines carry large masses of metal and mount a great
their mounting. of tate which are sourcia.ef radio interference.
the course of operation some bonding jumpers *y get broken, or the
Small-radius bends of aluminium wires result in displacement of conduct= iumper-to-airframe contact may become loose. The other problem is absent-
point.
the wire at the fitting point and in increased content resistance at the vg?nding
l t of Providing bonding arrangemehts for all the newly installed
point. Therefore the bend radii. of aluminium wires ahould be not smaller than ----"-g ec
those specified in column 1 of Table 20. ? df-ehelf items. There are other problems, too.
levier of all this; the aircraft bonding system should be systematically
If it proves impossible to maintain the specified radii during mounting
twanghly checked and maintained throughout the entire service life of the
operations (at the inlets into boxes, control panels and the like), resort mai
drcraft. The maintenance procedure consists in the following:
be made to the radius values specified in column 2 of Table 20, In the latter 1- Checking all the electric cables of the engine group for secure attach-
Case the lug should be put onto the wire bent to the radius indicated in
Table 20, and the wire should not be bent after the fitting operation.
Regulation and Check-Out of Bonding Arrangements
mot ant reliable contact with the engine body.
a. Cheering the integrity of all the bonding jumpers, installed on the air-
Mat; special attention should be paid to the bonding jumpers installed on the
Wendt engines.
Due to the fact that the airframe is used in the function of the adinusvi, ). Tightening-up loose jumpers and check-out of static dischargers for
all the units and items of aircraft equipment are reliably bonded to ensure yiemaeg.
normal operation of electric power consumers, to reduce to the minimum radio 4. Replacement of all unusable or broken bonding jumpere with due anti-
interference, as well as to eliminate the probability of local overheating
The following bonding methods are used: and:orrosimprovisions.
electric corrosion of separate unite and joints. lamnInstalling a bonding jumpers
(a) use an end cutter or emery paper No.00 to clean bright the contact sur-
1. Connection of all the aircraft structural members and equipment into MOO of the bonding jumper lugs and of the bodies to be bonded;
an integral system by means of rivets and bolts. (b) mount the bonding jumper seeing to it that its resistance value and length
? 2. Provision of special bonding jumpers which interconnect separate streeere mum as those of the replaced bonding jumper; make sure that the bolts attach..
tural members of the airframe and connect the aircraft equipment to the airfruing the bonding jumper to the airframe element are tight;
The maximum allowable values of content resistance between separate air- (c) measure the contact resistance;
craft structural members are indicated in Fig.65. The marl +m allowable contact (d) apply red paint, mark A-67, to the cleaned portion of the structural
resistance for all the other structural members and equipment units of the attache; to the jumper lug and the bolt head in the same manner as it was done
cyan is divided into the following major groups: with the replaced bonding jumper.
(a) 50 microohma - at installation points of ballast resistors, type 111-
5. For effective inspection, it is necessary to take regular selective
measurements of contact resistance with the aid of low-resistance meters, type
(b) 10.0 microohms - for points of direct coupling of all the ignition 8pagm_3 If it is revealed in the course of inspection that the actual contact
screens and for points at which the manifold pipes are connected to the engim resistance values considerably differ from the rated ones, actions should be
taken to normalize the bonding system.
Away;
(c) 200 microohms - at installation points of decoupling capacitors and
filters;
(d) 600 mdervohns - at points of direct coupling of parts and units;
Loll: Tolerated in some cases for directly coupling, parts is contact
resistance as high am 2000 microohms (for coVers, access panels,
doors, etc.).
(e) 2000 microoluns - for bonding jumper connections of parts and units.
However, in COMO cases it proves possible to obtain smaller contact rest*
ante values which considerably improves the aircraft bonding characteristics.
Contact resistance is chocked with the aid of low-resistance meters, tSla
IMC-3, or with special microolometers of high accuracy Class, with division
value not mere than 100 microohme.
OPERATION PECULIARITIES OF D.C. POWER SOW= E01MCES
Generator Maintenance
Aircraft generators, type PCP-18000 operate in heavy vibration conditions
and therefore need systematic and thorough care and inspection.
generator ra1tennnee procedure consiats in the following:
I. Checking the bolts of the generator lead-out vises and all the threaded
connections for tight fastening.
2. Checb-,pg the pipeline for secure attachment to the generator
bunml pipe.
air delivery
3. Checking the cap for proper attachment to the commutator end shield of
the generator; tightening up the nut attaching the cap with the branch pipe to
ths comastator end shield in case of necessity.
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regulator. The faulty voltage regulator should be removed from the
4. Checking the commutator end shield for play-free attachment. b. _lndjuBted
is likely to play in case of loose attachment of the air pipe to the geee4rcraft *24 "'it over for adjustment to the repair workshop.
air delivery branch pipe or due to excessive length of the free portionge: UTI /: BEVER tre to adjust the P7r-82 voltage regulator on board the
lead-out wires.by changing the air gap of the electromagnet or by taxying the
-
5. Checking the commutator and brushes for condition.
......-
....-, , garbee Pile pressure.
To inspect the commutator and those brushes which are accessible, l
xe ey
necessary to remove the cover band. When remounting the cover band, m ipart
et, from voltage regulator adjustment checks, attention should be paid
that the body-mounted pin by all mems coincides with the reference hole eqthe course
of the aircraft service life to the integrity (condition) of the
band. roe running from the second socket of the regulator plug connector to the
If it has been revealed in the course of inspection that the commutat;te:u.ty
resistor, type 130-20, and from the resistor to terminal I of the
terminal xcr of the stability transformer
severely burned and the brushes have been worn out down to a length of 1810ouilhe txannfkereaeript"e roperllundingg. Any breakage. in the above-mentioned
the generator should be removed from the aircraft and there:agar impacted, .
commutator should be cleaned with sandpaper, the faulty brushes should beA7c
uit results in a 'Sheep generator voltage increase, in disturbances in the
placed, and the generator - tested on a laboratory stand. el operating system, and In burn-out of the generator field winding.
voltage regulator is ensured by the stability
surface of the brush. 6::::,ery::11:1.8:f the
Note: The length of brushes should be measured from the side of the hmee
CAUTION. It is forbidden to overate the voltage regulator, type Pyr-2,
Maintenance Of Voltage Regulator and of Equipment
without the stability transformer, type TO-8.
eperstieg in_Setth Voltage Generator The differential undercurrent reley, type IUP-600, the stability transformer,
In the course of operation of voltage regulators, type PYP-82 , theee me- .
, TC-8 the-external resistor, type 80-20, and the Capacitor, tips KEU-31,
level adjustment and distribution of loads among the regulators operating- '-
,net require special maintenance. In operation, they will be checked only for
in parallel are effected by means of external resistors, type BC-20. ItUV--
atact tightness of their connected wires and for secure Attachment.
bay forbidden to employ any other method for adjusting the regulator.
relay
Due to wear of the carbon pile and possible sagging of the springs tb CitIT/ON: NEVER clean the contacts of the IMP-600 relay or adjust the
1,
in operating conditions.
pressure applied to the regulator carbon pile is likely to becomeweakened.'
I
incorrect operating conditions the wear of the carbon pile of the Pir-82 1 Storage Batty Maintenance
tion when the regulator begins to pop. It is forbidden to operate the regOiemmam7
Then installing the capacity-charged storage battery on the, aircraft, it is
age regulator may be so severe that the regulator will be maladjusted to
compound, terminals, group
popping conditions since this leads to burning out and desinteexation of there end to inspect it for condition of the sealing,
bon pile. plugs. There should be no cracks in the sealing compound and peep bees.
en as the surfaces of the terminals contacting the bretbaxs ehould be cleaned
he terminal bolts should have intact thread, and the output barber lugs as
To prevent popping operating conditions of the voltage regulator, tile
later should be subjected to regular inspection (approximately after everjmmmoildmm.
11
50 hours of operation) on board the aircraft. The regulator adjustment tad Paternal inspection over, the plugs (free from fouling) are screwed into the
carried out at high generator speed, and therefore it is advisable to conbAtterY,
generator check with the engine maximum r p m testing. and the functioning of the valves is checked. Never install plugs which
Ilea the generator is operated with the storage battery disconnected,
' o not open when the battery returns to the normal operating position after being
f.lted through 160 and 150g.
cut-in and cut-out tate place, the load variation being not less than 50%:. .
the nominal generator rating. If the regulator, due to maladjustment, ?pod Battery Discharge Level Test
The 1
popping conditions, thi 1 to which the storagebattery has been discharged can be ronghly
s operational instability will be detected by ?millets d ' '
'
of the voltmeter needle. ey she voltage produced by the battery under load or by the density
The voltmeter check of the regulator allows to determine popping opent
f the electrolyte, the second eethed being the most correct one.
The battery voltage is mead with the battery connecting (generators
Conditions of the regulator. However, checking by this method fails to reveloing ,,,) one of the aim
raft tions allowing, it is best advisable to carry out the check by listening Udeoharge level me Table 21 below.
For the battery volt electrolyte density as functions of the battery
power cons re rated for a current close to ,22 A.
Close-to-popping conditions since in this case popping is present only.0.1me,-e --
transient operating conditioconditionsand disappears quite rapidly. Therefore, cod
electrolyte density is gauged with a densimeter.
operation of the carbon pile through high-resistance earpieces. For operaiM
the earpieces should be connected to the wires running from the first andi
pins of the P71"-.82 regulator plug connector (in a place easiest for acomol
t:.-, effect several load on-off cycles. /f the regulator functions normally,1
load cut-out is accompanied by a single click and changed tone in the ear10
If the regulator load cut-out is characterised by wheeze, this testifies to
_
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Table
Battery Voltage and Electrolyte Density as Functions of
Battery Discharge level
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ware/ON: IT IS ABSOLOTELY FORBIDDEN TO STORE ELECTROUTL-FREE 12-CAU-55
21
sT0RAGE BATTERIES WHICH HAVE BELT IS OPMATION OR RAVE 2151:0 LIEC/hICAL
Battery discharge
level relative to
nominal capacity
Battery volt-
age (in volts)
at 20 A load
Electrolyte
density in
cells, regu-
eed to 25 0
-.
Note
Charged battery
25 to 24
1.260 I 0.005
Battery ensures 6 enal
startings
Battery discharged
14 29%
25 to 24
1.200 - 1.210
Battery ensures 3 to
engine startings
Battery discharged
by 5011
24 to 23
1.170 - 1.160
Battery may fail to lem
engine
Battery discharged
DI 75%
23 to 22
1.120 - 1.110
Engine starting failure
to be expected
Battery fully die-
charged
22 to 21
1.080 - 1.010
No engine starting
After each flight it is necessary to check the battery discharge level. If
the battery has been discharged completely or partially (by over 25%), it is
necessary to send it for charging to the charging station in not longer thane
eight-hour period. After each flying day (night) it is necessary to check the
battery discharge level by the electrolyte density. All the charging cycles re
the number of engine startings effected by the battery should be recorded bid
Service Log of the storage battery.
Inoperative storage batteries should be additionally charged with a cured
of 3.5 A at least once in a month.
Once in every three months all the storage batteries (both operating and
inoperative batteries) should be subjected to a procedure charge-discharge egh
as a measure against sulphating. The results of the operation should be entered
in the Service log of the storage battery.
In the course of operation it is necessary to regularly check the leveled
density of the electrolyte and add distilled water to the cells. It is forbiel
to add electrolyte or acid in the cells unless it is known for sure that the
level decrease in due to electrolyte spilling. In the latter case it SEI =tee'!"
to add battery sulphuric acid solution of the same density as the density of de
electrolyte contained in the cells.
Never expose storage batteries to direct sun rays or place them one onto
another.
If cracks are detected in the sealing compeund, eliminate them by the
melting method. Hot-treat the sealing compound only with the battery discharged
and. plugs removed, taking use of a soldering torch, hydrogen flame or other
means.
Storing the Battery
Storage batteries which are in aotive service and which have been in CPO
tion for not longer than half the guaranteed service life period, as well as
storage batteries which have passed the Manufacturer's electrical tests (rsII'l
with a red strip on the group bar) should be stored with electrolyte in the
charged state.
rssrs.
stoxage batteries should be placed for storage as follows:
1. Charge the storage battery to capacity.
e.eheckand carry out necessary operations to obtain the normal density
eelmml of electrolyte.
3. Install the vent plugs in all the battery cells and wipe the battery
relate with rags soaked in a solution of soda or ammonia hydroxide.
4,1ash the battery surface mith water and wipe the whole battery dry with
ale rags.
.3.01ean the clamps and intercelltonnections of the battery and coat them
tma thin Layer of petrolatum or grease. This done, the battery may be consider-
tready for storage.
6. Every month it is necessary to give the battery an additional charge with
'anent of 3.5 A till there are indications that the battery is charged to
opacity. it least once in every three months the battery should be subjected to
.p000seure operating cycle.
Prior to beginning the operation of a storage battery just removed from
step, it is necessary to give it an additional charge with a current of 3.5 1
tabedn constant electrolyte density and voltage.
The storage battery can be storGd with electrolyte charged for not longer
tan six months.
Then there is no possibility of storing the battery with charged electrolyte,
deetwmee batteries, type 12-01M-55, which have been in operation for some time
edam not intended to be used during long period of time may be stored dis -
Coed, without electrolyte. Before the storage battery is placed for storage,
Hiseubjected to one procedure operating cycle, and then it is discharged with
gement of 11 A till the voltage in one of the battery cells drops to 1.7 V.
feedMehalsed batteries are turned with their plug holes down and are left in
Obimmition during three hours. For complete removal of electrolyte from the
tater), it is necessary to slightly tilt the battery and give it light shake-ups.
Bleforbidden to wash the battery out with water before placing it for storage.
Eateries are placed for long-time storage with their blank plugs tightly
=elk in and with their surfaces thoroughly wiped dry with clean rags. To
pent bulging of the sealing compound during its storage, the cells should be
dosed with blank plugs at a temperature of 30 to 45?C inside the battery, for
tkhpurpose the battery should be either placed in the corresponding embient
teeerature condAtions, or warmed up with hot water from the outside.
corioN1 One-time used storage batteries, type 12-CAM-55, can be stored
without electrolyte for not longer than three months.
lietia_StoEage_Betteey_Troebleq
in the troubles which are probable to develop in the storage battery can
belvided into three categories:
I. ?reales of electrochemical character which can be eliminated by electro-
*Meal methods (by using specially selected charging-discharging conditions).
2. Mechanical troubles which can be eliminated on the spot, by available
^te.
ult!Zolitlzer0=1:hdoep:e.ctive plates and group bars; these faults are
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Troubles in the storage tottery can to detected either by external inaN.
tiom or by pertinent measurement:I during electroehemical tests.
Detected by visual (external) inspection are: cracks in the vessels aada
bars, leakage of electrolyte, cracks or softened spots in the sealing coepetai
fouling of the external surfaces, breakage of the output pins and interceR;
connections, poor contact between the output pins and intercell connectiorm,a
seal of the covers, as well as breakage or fipling of the plugs. The.najorila;
these troubles is eliminated right in the using unit.
The troubles mentioned under Item; 1 and 3 above can be detected by the
battery voltage and voltages of separate cells in the course of the charge-4,
charge cycle, by the density and temperature of electrolyte and by gas evolve;
during the charging half-cycle.Thesetroubles can be eldeduated only at spedd
repair workshops or at a charging station.
Pollowing should be the characteristics of a sound battery by the end d
its charging:
(1) voltage at each cell - 2.45 to 2.6 1' (when alive);
(2) specific weight of electrolyte - 1.260 = 0.005;
(3) eleetrolyte temperature - not over 4500;
(4) almoet simultaneous 'boiling' and gas formation in all betteryeelle
(5) neutral-colour, transparent electrolyte, free from any sediment.
When test-discharged, a sound storage battery should.manifest a capacia-
which is not smaller than 75% of normal capacity.
LamalkaAttlaa_elatornzer_liettta Omeation in 8.1fplano
Temaeratures
fa these cases when the storage half-battery cella are left in their
'containers with the aircraft parked at temperatures down to minus 40?C, prior
to flight it is necessary to engage the electrical heater system of the
containers.
The electrical heater system of the containers can be 'energized only foe
4 ground supply. source (See the diagram. in Fig.54) which is connected to the
ground supply plug connector of the aircraft.
When already in flight, i.e. when the storage battery is connected for
buffer operation with the. PCP-18000 generators, there is no need in sleets
heating of the containers even when the ambient air temperature is below mom
and down to minus 60?C. This is explained by the fact that while in flight *1
temperature of the electrolyte in the storage battery cells remains above mo'
dme to operation of the storage tottery.. The effect of the ambient air tope,"
ture is considerably reduced due to the use of-heat-insulator which liner We'
interior of each container.
To engage the container heater system, proceed as follows:
1. Connect a D.C. ground power supply source to the aircraft mains.
2. Tightly close the covers of both storage battery containers.
3. Tura on the heater, switch located above the left storage battery odd,-
The heater system is disconnected automatically by means of thermal Vitt
'Me 7778, which are connected to the minus circuit of the heaters of each
container. The switches operate as soon as the temperature at the surface of0
heating plates reaches 80 = 10?C.
Connecting D.C. Ground Supply Source
To energise the aircraft electric Balms at parking and for engine statil
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pmeme, the aircraft is equipped with a ground supply plug connector the plug
dristich is secured in the nosewheel leg well, port side, at frame 10.16.
The plug and the mating detachable receptacle of the ground supply plug
gector have three pima and three sockets. Two thicker pins are power pine,
?Atha are longer than the third (thinner) pin which is used as a guide element.
sale construction ensures that the power contacts are energised only after
;befall contact is obtained, which precludes burning of the power contacts when
?meeting the receptacle.
To connect the ground supply source to the aircraft electric mains, act as
follows:
1. Couple the ground supply receptacle (with the ground supply source
=meted to it) with the ground supply plug.
2. Place the voltmeter change-over switch on the generator control panel
at the radar operator's station to GROUND SUPPLY RECEPTACLE (PAD:),
3. When sure (through reference to the voltmeter) that the voltage across
t10,tomdeals of the ground supply plug connector is aortal, select the NORM
SOPPLI CIRCUIT ( HOMAN:RAH 08Th; position of the voltmeter ohangz-ower switch.
4. Turn on the ground supply switch.
5. As soon as the voltmeter begins to indicate that the aircraft mains is
aergfzed, it is allowed to begin connecting power consumers,, checking their
operation by the ammeter and voltmeter.
Ground supply sources axe connected to the aircraft electric mains through
a uotactor, type I-400A (See Ref.Ro.32 in the diagram of 716.54) which
aerates only with the ground supply switch cut in (See Ref.No.,1 in the same
Tiger.).
To disconnect the ground supply source:
1. De-energise all the power consumers.
2. Turn off the ground supply switch on the generator control panel at the
pular operator's station.
3. Disconnect the ground supply receptacle.
OiDTION. When the aircraft mains is energised from a ground supply source,
it is not advisable to impose a simultaneous load which would exceed 500 A.
In case the aircraft mains requires a current larger than 5004 it in
necessary to withdraw the fuses from the storage battery ammeter and ground
supply circuits which are installed in the storage battery junction box.
In overload conditions use should be made of a special ground ammeter with
a scale range exceeding'500 A.
Control over D.C.POWer Supply SOV=08 and
Ileetrie Maine
Control over the operation of the power supply sources and over the continui-
tfof the electric circuits is effected by means of five ammeters. Four armeters,
taxi i-a, with scales reading to 100 - 0 - 1.000 A are installed in the generator
circuits, while the fifth ammeter, type A-2, with its scale reading to 50 - 0 - 500 A
18 Provided in the aircraft storage battery and ground supply circuit.
The ammeters are provided with extension shunts which are located on the
distribution panels of the engine compartments and in the storage battery junc-
tion bon (Fig.72).
The operation of the power supply sources and functioning of the electric
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circuit are checked by a voltneter, type 8-1, rated for 30 V. By me= of
selector switch, type , the voltmeter can be connected to each of theft
generators to the pommel supply circuit, to the emergency supply circuit, ea%
ground supply plug connector and to the storage battery.
In flight, under normal electric power supply conditiona the voltmeter
should be connected to the normal supply circuit, and in amargenoy power co
conditions it should be connected to the emergency supply dirouit; in de-
enexgized mains conditiona the voltmeter should be connected directly to Um
storage battery.
All the above mentioned instruments, as well as the selector switch, te,
1146, of the voltmeter are mounted on the generator control panel insta1lt!
at the radar operator's station (See Fig.55). In addition, the radio operatme
instrument panel mounts a voltmeter, type 8-1, which measures the Bernal see
cimeelt voltage in the rear pressurised cabin.
?
Basic Technical Characteristics of Ammeters. T715118 A-l. A-2. Ae3
and of Voltmeter. Type B-1
Descrlp-
tion
Type of
instru-
Measuring range
Graduation
value
Scale gn4
ation nineRent jag
Ammeter
A-1
40 - 0 - 400 A, with shunt
rated for 300 V
20 A
0, 1, 2,)
and 4
Ammeter
A-2
50 - 0 - 500 A, with 11-2
25 A
0, 1, 2,1
Shunt rated for 500 A
4 and 5
Ammeter
L-j
100 - 0'- 1000 A, with 111-3
shunt rated for 1000 A
50 A
0, 2, 4, i.
8 and 10
Voltmeter
8-1
0 - 30 V
1 V
0, 1, 2st
1. The main error of the ammeter without shunt under normal condieioued
at nominal resietance of the connecting wires does not exceed 4% of the 52
-total of the nominal scale values.
2. The shunt is accurate within 10.5% of the shunt nominal current retie
3. The main error of the IS-1 voltmeter under normal operating conditions
Should not exceed 12% of the nominal scale value.
4. The additional error for every 10?C ambient air temperature varietim
within plus 50 to minus 60?C should not exceed 10.5% at the sum total of tte
nominal scale values for the ammeter, and of the nominal seals value for tie
voltmeter.
Naintenance of Ammeters and Voltmeters
When the power supply sources are disconnected, the needles of the item
smuts should indicate sere.
If the instrument needle does not respond to the connection of a power
sway source, it is necessary to check the mires for condition and to check
ebethar the contacts at the wire-to-instrument (or to shunts in case of acnte
connections are reliable.
In short-circuit conditions the ammeter needles swing to the extreme ri4,
positier, (beyond the scale range) and the voltmeter needle indicates reduced
voltage. If, at the moment of connecting a power supply source, the instrse
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?sae torus in the reverse direction, it is necessary to change the places of
tbe else ends leading to the indicating instrument.
laces, troubles develop inside the indicating instrument, it should be
"pieced. There is no need in removing the ammeter shunt (if it is intact) since
? the ammeter shunts are interchangeable.
If in the course of operation there appears a necessity to replace the
=meeting wires in a certain section between the indicator and the ammeter shunt,
ew length and the gauge of the newly selected wires should be identical to those
d the replaced wires. Changes in the length and gauge of the wires result in
changed resistance of the connecting wires, and other-than-nominal resistance
lege to additional instrument errors.
Adjustment of D.C. Power Supply Sources
ma generator system adjusting procedure should be started from individual
voltage adjustments on each generator with the view to obtaining a voltage of
03.5 V with the aid of the external resistors, typo 80-20, and the 8-1 voltmeter
smmted on the generator control panel at the radar operator's station (See
11L.55).
emeTION. It is allowed to connect the generator to the &Jewett mains only
after it has been adjusted for the voltage of 28.5 V.
Generator Voltage Adlustment ___ 2round Conditions
In ground conditions the generator voltage will be adjusted with the engines
naming; in the course of the adjusting procedure, the power consumers of the
geeing accessories group should be energized from a ground power supply source.
To adjust the generator voltage:
1. Place the voltmeter change-over switch to the position corresponding to
the generator subject to adjustments.
2. Obtain the engine speed of 3750 r.p.m.
3. Obtain the voltage of 28.5 V by rotating the knob of the 80-20 external
mister of the generator to be adjusted.
4. For a short period of time advance the engine speed to 4100 r.p.a. la a
nmult, the generator voltage should not vary by more than 0.5 V.
The voltage adjusting procedure for all the other generators is absolutely
Identical to that described above.
Connecting Generators to Aircraft Mains
To connect the generators to the aircraft mains act as follows:
1. Disconnect the power consumers leaving the minimum number of connected
emmumers which ensure normal operation of the engines.
2. Disconnect the ground supply source and quickly connect all the four
generators, one after another.
3. Cut in all necessary power coneumers.
CAUTION. Before connecting the generators, see to it that the 12-CAM-55
storame battery is installed in its container.
Adjueting earallel_Opeeatien_oe Generators
The parallel operation of the generators will be adjusted In flight, in 30
to 40 minutes after the take-off, i.e. as soon as the voltage regulators and the
generators are warmed up sufficiently.
The adjusting procedure runs as follows:
rr ? ? ? ? ? ? ? / 1 ?
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1. Cermet all the de-icer and beater devices. The current load in this
case will total: .
peretenently connected consumers 430 A (approx.)
electric heaters of cabins 340 1
glass panel beaters 190 A
tail unit de-icers 470 A (approx.)
amplidynes and dynamotor of cannon armament
system
250 A (approx.)
Total 1680 A (approx.)
Hence, the average load per one generator amounts to approximately 420 A.
To avoid dangerous overloading of any one generator, all the large loads
(de-jeers and heaters) should be applied in turn.
Upon connection of a power consumer. it is necessary to check, through
reference to the generator ammeters, whether the current is equally dietributed
among all the generators. In case the generator current is unbalanced by more
than 120 A, it is nacesnary to level off the generator loading with the aid ot
the B3-20 extereel registers; the voltage of the generatore bearing the sealle:
load should be increased, and the voltage of the heavier-loaded generators
should be reduced.
2. Place the 11-46 selector switch of the voltmeter to the NORMAL SUPple
CIRCUIT (HOPMAILHAR GEM ) position and check US aircraft mains voltage; th
voltmeter should read within 28 to 28.5 V. In case of other readings, the voltee
level of all the generators should be tither rained or lowered by the required,
magnitude. This is effected by rotating the 33-20 resistor control knobs throe
one and the same angle.
3. Disconnect the power consumers which are not required for normal flight
procedure and chock the generator loading by the ammeters. Unbalanced loadings
of the generators in small-load conditions is no problem to bother about; how=
all the gonoratore should supply current to the aircraft mains. In condition:
vet e small loading some generators can be disconnected by their respective robe,
tni AMP-600 This ;resents no trouble, since,as the load increases, the
1P-600 relay will reconnect the generator to the mains.
Bete: It is necessary to adjust the parallel operation of the generators
In each flight. The adjustment should be repeated only if the
generator current is unbalanced by more than 150 A. at a load amount-
ing to 25 -.50% of the nominal loading, and 120 A at loads exceediX
hal/ the nominal rating of each separate generator.
?
In flight, all the generators should be connected. A generator may be dis-
connected in flight only in case a trouble has developed in it. In this case
the radar operator should report his actions to the aircraft commander.
If fire breaks out on the engine or in the engine nacelle, the fire -fight-
ing system of the aircraft is engaged into operation automatically. In synchnr
niam with the fire-fighting eysten actuation the engine cowl vent pipe to auto-
matically shut off which stops the generator blowing. Therefore all the genera-
tors installed An the engine located in the fire area should be quickly dis-
connected from the aircraft electric mains. Raving disconnected two of the
generators, make sure that the total load applied to the operating generators!'
not in excess of their performances. In case of excessive loading, part of the
power consumers should be dieconnected.
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gaga! In conditions when two generators are dieconrected from the air-
craft mains, it is forbidden to effect simultaneous connection of the
. cannon syste tend the tall unit de-icer system.
Disconnecting Generators from Aircraft Mains
To disconnect the generators from the aircraft mains prior to stopping the
eeMes, act an follows:
1. Disconnect three generators from the aircraft mains.
e. Disconnect all the power consumers from the aircraft electric mains but
fa chennel No.1 of the intercom.set, the stand-by pumps and the engine control
instruments.
5. Disconnect the storage battery from the aircraft mains.
4. Stop the engines.
5. Disconnect all the consumers which were left connected.
6. Disconnect the fourth generator from the electric mains.
OPERATION P33U1.IAR1TIES OF A.C. POSER SUPPLY t30UR1ES
Connection of 110-4500 Inverters and of
Ground A.G. Power Supply Source
Connection of 110-4500 inverters is effected from the generator control
Icel at the radar operator's station (See Fig. 55) by means of a change-over
such, type 3HDR-45 ? which precludes simultaneous connection of both /neer-
tue.
The operating inverter is supplied with direct current through the storage
Weary junction box from the normal supply bnehar, and the stand-by inverter is
mergized through the dual supply circuit Junction box (mounted at frame No.17)
teethe dual supply busbar.
For the key circuit diagram of A.C. power supply sources refer to Fig.73.
When connecting the inverter for operation from a ground D.C. power supply
once, see to it that at the inverter starting moment the voltage across its
tninals is not lower than 20 volts.
CAUTION. NEVER start the 1101-4500 inverter for operation from a ground D.C.
power supply source which reduces the voltage across the inverter
terminals to below 20 V at the inverter starting moment.
The inverter connecting circuit (See Fig.73) makes it impossible for the
beater to be engaged with its voltage regulator, type P-258, disconnected.
If (in ground operating conditions) the inverter fails to get disconnected
*lithe 3111111-45 change-over switch is turned off and the respective ASC-2
drcuit breaker is opened, and goes on operating, it is required to de-energize
Re D.C. circuit, i.e. to disconnect the ground supply source.
CAUTION. It is FORBIDDEN to uncouple the plug connectors until the 04500
inverter is de-energized. ?
With the inverter disengaged, it is necessary to check the external supply
cheuits: if they are faulty, remove the inverter from the aircraft end send it
mer to the repair workshop.
For A.C. supply of the aircraft electric mains on the airfield, provided
It the nosewheel well, starboard, at frame No.16, is a ground A.C. supply rip-
Mit junction box with a two-pin plug connector of 11,28112HM7 type. The greund
aWay source is connected by mane of a switch, type B-45. leeated on the
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generator control panel at the radar operator's station. The design of the gs.paied
supply circuit (Seal Fig.73) makes it impossible to connect the ground suPP17
after one of the aircraft inverters. type 0-4500, has been engaged.
;evertor Maintenance
To ensure reliable operation of the inverter, type 110-4500, it is recoil*,
to carry out its inspections after every 100 operating hours; the inspection 1
procedure will include condition checks of the commutator, slip rings, brushes,
holders and scavenging with compressed air to remove duet from the brushes.
In case traces of burning are detected on the slip rings Or commutator, the
elements Should be cleaned with sandpaper ]o.00. During this step of the main..
tenonne operations it is necessary to differentiate uniform dark-colour deposit
from real burning; the deposit in question has no adverse effect an the inverter
operation and therefore it is not subject to removal.
Should the length of the Commutator brushes be worn out down to 16 my and
the slip ring brushes - down to 14 mm, the brushes are to be replaced. The new
brushes should be lapped to the commutator and slip rings with the aid of sand-
paper Io.00 or ground in at idle running of the inverter during 5 to 6 hours.
The inspection of the inverter over, it is necessary to push the centrifugal
switch return button as far as it will go, and to make sure that the inverter is
ready for starting.
In case of failure of the voltage stabilizer (which is indicated by higher.
than-nominal output voltage and absence of glow in the voltage stabilizer) the
faulty stabilizer should be replaced.
The voltage stabilizer replacement procedure is as follows:
1. Open the access hole in the top part of the box having previously un-
sealed. the access panel fastening screw and turned it by 900.
2. carefully lift the voltage stabilizer. To replace the voltage stabilizer:
with your left hand pull back the cap which holds down the voltage stabilizer.
Exerting pressere with the index finger of the right hand, move the voltage
stabilizer aside, holding it up while doing this. Then, operating with the left
baud, install the new stabilizer and fit the cap on.
When replacing the voltage stabilizer, see to it that the coil springs wind
secure the voltage stabilizer are not expanded excessively and that they axe
positioned correctly. The axis of the springs should run normal to the horizontal
plane; the position of the Springs is adjusted by turning the cap on the voltaI
stabilizer to one or another direction.
3. Close the access hole panel of the box and seal.
4. Enter the reason for the replacement and the number of pre-inutallatict
operating hours of the new inverterr, in the Service Log of the inverter.
If after the replacement it proves impossible to obtain the nominal
. output voltage value (115 V) with the aid of the voltage level adjusting rheostid
the inverter should be replaced and subjected to thorough inspection at the
repair workshop.
C1DTIOI. &EVER adjust 110-4500 inverters on the aircraft.
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Inverter Probable Troubles Constituting Reason for Its
Replacement

Trouble
Indication
In freouencv control circuit
Electric connection of magnetization
141.0tg of Z0-25-170 choke with
Saftwlizing winding
breakage of one A.C. winding of choke
Breakage of magnetization sinning of
0-26470choks
!Soakage of neutralizing winding of
0-26-170 choke
Warted turns of AZ-11 mesh choke
Ondased ends of control winding
Is A.C. voltage in the circuit
Sudden r.p.m. drop, Somewhat reduced
voltage
? Somewhat increased frequency at idle
running (approximately 435 c.p.a)
Frequency increases to 500 c.p.s.
Frequency drops to 300 c.p.s.
Increased frequency
High frequency, large current of
electric motor, heavy starting
High frequency, large current inten-
sity of electric motor
In voltage control circuit
Electric connection of sovietisation
winding of A0-12-25 choke with
neutralizing winding
Shorted (punctured) capacitor rated
for 2i0.5 sT
Breakage in magnetization winding
circuit of Z0-12-25 choke
Loose contact in voltage stabilizer
panel
Breakage in neutralizing winding cir-
cuit of Z0-12-25 choke
Wrong connection of A.C. coils of
110-12,25 choke
Confused interconnections of magneti-
zation and neutralizing windings
Confused connections of TC-11 stabili-
tamnsformer
Mixed polarity in connections- of
Isametization and neutralizing 'windings
Of A0-12-25 choke
Confused polarity in connections of
untralizing winding of AO-12-25
Chaim
Caufemed polarity In connections of
bilmstization winding of A0-12725
WWks
Voltage drops to 70 V
Voltage stabilizer fails to fire,
low voltage level
Voltage increased to 150 V at idle
running and to 130 V under load
Voltage increases to 135 V under
load and to 165 V at idle running
. Voltage drops to 57 V both at idle
running and under load
Inverter voltage is 70 V
Low voltage (55 V), high frequency.
Voltage elevated to 160 V. negative
drop up to 10 V
High voltage - up to 130 V
Low voltage (55 to 60 V), high
speed
Low voltage (60 to 65 7), high
speed
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2pntrol over A.C. P'pp end
MUM
For control ever A.O. voltage of 115 V the radar operator's generator *egg
Panel (Mee ViS.55) is provided with a ferrodynanie voltmter. type 80-150.
apecifications of Voltmeter. Trpe st-po
I. Measuring range
0 to 150
2. Main error ? 12.5%
3. Additional error for every 1000 temperature variation
from normal (+2090) 24.0
4, Additional error for a t50 c.p.s. frequency variation
from mean frog:ler:my of 400 c.p.s. =1.2%
5. Power consumption 2.5 W
6. Operating temperature roved; from -60?C to +50%
7. Weight of instranont 400 gr '
12111
Voltmeter errors are given leper cent of full scale.
Maintenance 2f-Ya1Inlaanz -
To ensure correct operation of the instrument, the pre-flight preparation
procedure should include a check-up of the voltmeter needle for correct zero
position. The check should be carried out before energizing the instrument. 16
voltmeter needle should be zeroed by moans of the corrector screw located onto
face panel of the instrument; in the course of seroimg the instrument must be k
energized.
CASTIOV. The 80-150 voltmeter should be checked for its needle position
before each flight.
If the voltmeter needle fails to respond to the connection of the 110-45C?
inverter, it is necessary to check the condition of the connecting wires, as tM
as the integrity and reliability of contact connections. Should other trouble; Ca
detected in the course of the voltmeter operation, the faulty instrument ahmal
be removed and replaced.
Adjusting and Checking the Operation of 110-4500
Inverter
It ia allowed to connect the inverters for operation with the aircraft A.C.
mains only when the aircraft D.C. mains is energized from a ground power suppil
source or from aircraft generators, type 1'CP-18000.
The
1. wit:Ter:C-2
adjusting chocking procedure is as follows:
t breakers of the operatieg and staul-by invert*
open, place the generator selector switch on the generator control panel to
OPERATING (PAE0V(X11 ). This action should engage the operating (starboard)
inverter.
2. At least 5 minutes after, check the inverter voltage by the aircraft La
voltmeter. If the gauged voltage differs from the nominal voltage of 115 V DJ?
than 0.5 V, operate the respective inverter voltage rheostat to obtain the TOM
of 115 ? 0.5 V.
3. Connect the permanently engaged A.C. power consumers (the radar bomb Ng
gun sights).
4. Check the voltage of the operating inverter, and correct it if the vait'
age is other than 115 t 0.5 V.
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5. Disconnect the A.C. power consumers.
6. Throw the generator selector switch to the STAID.HY 0163EPBHUI)
Ionian; check and adjust the voltage of the stand-by inverter repeating steps
(2) 3 expenditure of the service lives of the inverters it is
advisable to connect them to the aircraft mains during ground adjust-
ments and to engage them in flight in turn.
' (1o)te
grmcTRICALLY HEATED CUSS PANELS
Electrical heating of glass panels prevents their external icing and inter-
ma fogging thus providing adequate visibility conditions under any flight condi-
Clans encountered.
Electrically heated are the two forward glass panels, type 11-13 , (right
and left) of the pilots and the lower glass panel, type 11-13 . of the naviga-
Um.Fachglass panel is an assembly of two hardened silicate glasses butwar-
mertmibetween which is a heater element consisting of thin constantan wires.
The power requirement of the heaters depends on the heated area and
constitutes 0.5 to 0.64 VI per one square em. of the heated glass surface. This
specific power (0.5 to 0.64 W/sq.cm.) is so large that should there be no
sufficient heat dissipation, the operating heater would raise the glass
tenerature to such a degree which might result in deterioration of the glass.
tome: these conditions, provisions are made for temperature regulation. This
work is done by thermistors press-fitted in the glass panels and by an automatic
toperature controller, type 100 -8111. The AOC -81N controller is installed at the
starboard side of the front pressurized cabin in the area :1 frame N0.5. The
electric heaters of the glass panels of both the pilot and co-pilot are engaged
by weans of two B-45 switches located on the uierhead electric control board of
the pilots (See Fig.90), while the heater of the navigator's glass panel is
at in by the B-45 switch mounted on the navigator's overhead electric control
board (See Fig.63). The current energizing the pilots' glass panel heaters is
Supplied through two K-50)1 contactors (See the diagram in Fig.74)? while the
midistor's glass panel heater is energized through a IT-100A contactor. The
pmmr Supply lines are protected by delayed action fuses of NH type (two
hmesratsd for 75 A each and one fuse - for 100 A).
The fuses and contactors are housed in the glass penal heater junction
bcd(Fig.75) which is installed at frame No.6, starboard.
Maintenance of Heated Glass Panel's
When replacing equipment items of the electrically heated glass panel sets,
11144 sure that the equipment is mounted and wired correctly. Special attention
Should be paid to correct connection of glass panels which have additional lead,
iota (the navigator's glass panel, type 11-13 ). The continuity check will be
carried out in compliance With the existing rules, using the methods of identi-
fication and measuring the insulation resistance value.
The connections of the thermistor circuits are of no lesser importance.
Mere/stop No.1 of each glass panel is its operating thermistor. Thermistor No.2
la a stand-by instrument and it is engaged only in case of failure of thermis-
tooled.
Ettea Thermistor No.1 corresponds to lead 1 on the terminal block, and the
operating lead for thermister No.2 is lead 3. Lead 2 is common for
both.
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If the terminal lead-outs are not nuabered, the left lead Shouldh
considered as the lead of thermistor No.1? and the right lead (ea
viewed from the down-oriented block side) should be taken as corr.*.
ponding to thermistor No.2.
Adjusting the Glass Panel Heating Degree
When through with the circuit continuity check, it is necessary to adjust
(regulate) the degree of glass panel heating; this procedure consists in teeth;
the channels for correct connection and in adjusting the automatic temperatua
controller, type 10C-8111. The AW-Bill controller has three independent channai
each of mbion controls its connected glass panel.
The adjustment procedure runs as follows:
1. Disconnect the wires from the terminal block of the navigator'e glaw
panel.
. 2. Connect a teat lamp between the plus wire and the airframe.
3. Engage the NAVIGATOR'S GLASS MEL HEATER ( 0501TEB 0TEKKA EMMA )
circuit breaker, type A30-2 , on the navigator's circuit breaker control pmm1
, 4. Engage the NAVIGATOR'S miss PANEL HEATER ( OBOITEB CTEXZA MTYPUIRA)
switch, type 71-45, on the overhead electric ? control board of the navigator. la
a result, the test lamp connected to the plus wire of the glass panel heater
should flash.
5. Close (through a resistor of 1000 to 2000 ohms) the wires disconnected
from the thermistor of the navigator's glass panel heater. This action should
result in going out of the, leap connected to the plus wire.
Notew Used in the function of the resistor may be a calibrating resist=
rheostat (Fig.76).
6. With the navigator's glass panel heater switch disengaged, connect the
thermistor mires and the plus wire to the terminal block without disconneettm
the test lamp.
7. Mace the slide of the navigator's glass panel heater channel rheostata
the ADC-81:8 controller to the extreme left position and turn on the switch arta
respective glass panel beater.
8. Moving the rheostat slide- to either side, check to see if the test Ime
flashes up when the rheostat slide is turned to the right and goes out as some
the slide is moved to the left from the centre.
9. By turning the rheostat slide to the left, and then slowly returning ;
it to the right and farther on, determine the position in which the lamp flash;
It this moment the K-100) contactor will be engaged and the heater will start
its operatioh. After a certain lapse of time the A0C-8111 controller will dis-
connect the navigator's glass panel heater. As soon as the glass cools di:finite
the pre-established degree, the A0C-81M controller will re-engage the heater.
All the time during the check it is necessary to watch the temperature of
the outer surface of the glass panel referring to the thermometer, type NM
2045-43 . The thermometer ball should be applied to the hottest spot on the
glass (See Fig.74), holding it tight against the glass by means of a piece of
cotton wool or felt.
10. After two or three operating cycles of the contactor, the glass sues%
temperature can be considered to be stable. If the temperature is other than
32 t 2?C, it it necessary to carry out additional adjustments which is done W
turning the slide of the respective rheostat on the AOC-8111 controller.
11. Adjustments of the A00-8111 controller over, determine the value of to
resistance the A0C-8111 controller is adjusted for, and make the correspondif3
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,try iathe controller Certifidate with the indication of the system resistance
odd the glass surface temperature. To do this:
(a) disconnect the wires from the terminal block, and connect a test lamp
ta the glass panel electric supply cable and the calibrating resistance rheostat
Lathe other two wires;
(b) set a resistance value of 9000 ohms on the rheostat, and then gradually
decrease the resistance until the test lamp goes out. The resistance at which
the p.m goes out will be the resistance for which the A0C-81/1 controller is
elated; the value of this resistance should be within 1500 to 8000 ohms. In
owe the obtained resistance value is other than 1500 to 8000 ohne, it is
,amary to regulate the glass surface temperature by means of the stand-by
thensistor, since the conditions indicate failure of thermistor No.l.
12. The beater channels of the ACC-SIM controller for the glass panels
of the pilot and co-pilot will be adjusted with the same methods an those
described above.
13. Once all the three controller channels are completely adjusted, seal
it, covers closing the rheostats of the A00-8/11 controller.
CAUTION. 1. During the check, NEVER short-circuit the wires running to the
thermistor and NEVER set a resistance smaller than 1000 ohms on the
rheostat, since this will result in failure of the automatic temperature
controller, type ACC-8111.
2. IT IS ABSOLUTELY FORBIDDEN to engage the glass panel heaters if
the thermistor is disconnected or the AOC-81M controller is maladjusted;
same is true when thermistors have internal breakdowns.
3. The automatic temperature controller, type A0C-8111, should be
adjusted with employment of a thermometer at ambient air temperatures
of minus 10 to plus 25?C as at lower temperatures it may occur that glass
surface temperature measurements will be erroneous, while at higher-than-
specified temperatures the glass cools down very slowly after its heater
has been automatically disconnected.
Use of Glass Panel Heaters
In flight the glass panel heaters will be engaged if icing or fogging
seditious are about to be encountered (before aloud-breaking, in haze and mist).
when flying In adverse weather conditions, it is advisable to keep the heaters
engaged throughout the entire flight.
At parking sites and when taxiing to the take-off position, the glass panel
heaters should be engaged only in icing or fogging conditions.
Before going down for landing, as well as prior to taxiing to the take-off
position it is recommended to switch off the glass panel heaters for when the
Waters are engaged, the additional strains (deformations) resulting from
bumpiness (vibration) of the aircraft may render the electrically heated glass
panels unserviceable.
CAUTION. Never engage the glass panel heaters with thermistor disconnected
or A00-8111 automatic temperature controller maladjusted.
TAIL EMONNAGE DE-ICERS
Brief
The fin and stabilizer leading edge sections are provided with electrically-
oPerated de-icere. Each de-icer consists of sections, ports and heaters.
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The Btu:blazer de-icer is divided into two sections:
(a) inner section heating the basic (butt) parts of the left and right.
stabilizer leading edge sections;
(b) outer section heating the end parts of the stabilizer leading edge
sections.
The ftn de-icer has one section consisting of one part. The inner and outs,
sections of the stabilizer consist each of two parts located on the left- and
right-hand surfaces of the stabilizer. In each section the parts of the stabili.
zer left- and right-haea surfaces are connected to each other in parallel (tee
diagram in Fig.93). Each part consists of several heaters connected to each
other in series.
? Each part of the de-icer sections is provided with bimetal thermoswitches,
type 777 B, which cut out the de-icer section when at least one of the parts is
heated to a temperature of 80 = 1006 in the place where this thermoswitch is
mounted.
The de?icer sections. are energized during 40 seconds and de-energized durir
80 seconds. The cycle is ensured by the distributing electrical mechanism, type
Mha-3L, switching on the de-icer sections in turn through the K-60011 contactma
The beginning of switching on the de-icer sections varies in time and
depends on the position of toe 18E1-3A contact device at the moment the MEA-31
electrical machanivon stops; the order of switching on the sections is always the
same. The switching on of the inner section of the stabilizer de-icer is always
followed by that of the stabilizer de-icer Atter section, than by that of the .
fin de-icer, then again by that of the stabilizer de-icer inner section,etc.
The 111(1-3/1 mechanism is mounted on the port aide of the fuselage tail un-
sealed section at frame No.63. The power contactors, type 1-60071, and fuses,
type TH-600 , of the de-icer sections power circuits are located in the juretim
box of the tall empennage de-icers (Fig.78) which is also mounted on the port
side of the fuselage tail unsealed part between frames Nos 63 and 63a.
The de-icers are switched on by means of the 8-45 switch on the pilots'
upper electric board. The de-icer operation is checked by a white lamp, type
CU-51 which every 80 sec. flashes up for 40 sec. thus warning of the
operation of the stabilizer outer section de-icer. The lamp is mounted on the
right-seat pilot's instrument panel.
IPE=2 EF2cificativEs_ MICH.4_ -
Electric Mechanism
1. Nominal voltage
2. Operating voltage range
3. Nominal current at the moment the commutator
contacts open
4. Nominal current consumed by the mechanism motor
5. Duty of mechanism operation
6. The electric mechanism must operate normally under
the following conditions:
(a) at relative humidity of ambient air
(b) at change in ambient air temperature
(c) at vibration and shaking with acceleration
(8) at sea level altitade
P. Mechanise service life
27V
27 = 2.7 V
5 A (of inductive
load)
0.8 A
continuous
up to 98 per cent
from +50 to -at
4g
up to 15,000 v.
300"hours of
continuous Opera
-
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?123-
8. weight og the electric mechanism
9.Comatator ensures switching on the contactor
whaiings under voltage according to the follow-
ing cycles:
(a) two 60 i 9-sec.resitchings with an interval
between them not in excess of 4 see.
Zane
tied in the course of
two years from the
moment the mechanism is
Installed on the air-
craft
not in excess of 2.4 kg
(b) three 40 = 6-sec. switohings with an inter-
val between them
(0) six 20 -3-sec. switchings with intervals
between them
same
Notes: 1. The connection diagram of the KKA-3A electric mechaniam ensures
the latter operation only according to the cycle indicated in
Item "b".
2. The interval.' between the switchings is included in the time
during which he commutator contacts are closed. Two contacte of
one cycle cannot be closed simultaneously.
Care of MEA-3A Electrical Mschanica
During service the KKA-31, electrical mechanism does not require aijustment,
wdutenance or special care. The attending personnel suet only periodically
check the quality and reliability of the electrical mechanism attachment, as well
lade locking of the plug connector and attaehment bolts. On detecting a:fault
lathe 18E8-3A electrical mechanize r;place it by a new one.
Checking Tail Empennage De-Icer System on the Ground
The ground epeck of the tail empennage de-jeers should be carried out only
vid the aircraft mains supplied from the ground D.C. power source connected
through the ground supply plug connector.
CAUTION. To avoid overheating of the Akin and' denage to the protective
coating, it is not permitted to switch on the eleetric de-icer with the
aircraft mains supplied from the mr-18000 generators and with the
, engines running on the ground.
The de-icer ground test makes it possible to checks
1. The condition of the circuit and the serviceability of the de-jeers.
2. The sequence of switching on the de-icer sections.
3. The duration of the de-icer sections switching cycles.
4. Current consumed by the de-icer separate sections at a voltage of 26
across the heater terminals, that is, provided the de-icers are supplied from
tba aircraft generators.
Check to see that the surface of the de-icer boots is heated and the de-
icer sections are switched on in correct sequence by hand feel. Besides this,
the surface of the de-icer boots can be checked for warning up by moans of a
'Pedal instrument which is a thermocouple mounted on a tubular rod adjustable
lie length. Inside the rod a wire running to the tenperature indicator is laid.
Approved
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When cheoking temperature by soaps of the special instrument (thermocouple) +,
??? Forth, fin de-icer the value I26 must be within 480 A t 10 per cent, for
temperature on the surface at any point of any de-icer boot section must art"
me ide section of the stabiliser de-icer - within 450 A t 10 per cent and
ambient air temperature approsimatel7 by 30 to 5000 during one cycle of t116
th. outside section - within 494 A I 10 per cent.
for
de-icer section operation. Iffor any section of the tail empennage de-icers the value 126 exceeds
Check the duration of the switching cycles by means of a stopwatch, et%
...,,,pdttedlimits, make sure that the taken measurements and calculations
the TU-600 fuses in the power circuits of the de-icer sections removed. ' correct and only after that replace the defective leading edge sections by
Consumed current is checked by means of the_ 2 aircraft ammeter connec ters
seems.
to the storage-battery and ground power supply source.
It is not permitted to eliminate the defects caused by short circuit and
Complete chock up of the tail empennage de-icer system on the.ground 1t1sopen circuit of the heaters during service.
be performed in the following manner:
1. Remove the TH-600 fusee from the power circuits of all the three Instructions for Operation of Tail Empennage
sections in the junction box of the tail empennage de-iceri. Dc-mere during Flight
2. With the de-icer 13C-5 circuit breaker (on the right-seat pilot's
The tail empennage de-jeers should be switched on during flight prior to
circuit breaker penal) switched on, turn on the tall empennage de-icer switch
open the zone of probable icing. The de-jeers must be switched off if the
on the pilot's upper electric board. In t -is cane the MIA-3A electrical
warning Igo:1w of the tail empennage leading edge sections is quite free from ice.
mechanism most be brought into operation which is indicated by the
Tee: When checking the serviceability of the tall empennage de-mere
an the right-neat pilot's instrument panel flashing up periodically.
during flight in case no ice formation takes place. it is permitted
3. Check the operation of the K-600/1 contactors by means of pilot lar',
connected to the winding terminals of all the three contactors. The lamps cast to switch then on for not more than 5 minutes.
periodically flash up.
I Ede being the case, the de-icer operation is checked by the warning lamp
4. Check the duration of She switching cycles by the pilot lamps. rel =mused current (as measured by the generator ammeters).
5. Turn off the de-icer switch at the moment the warning lamp-on the MO
seat pilot's instrument panel goes out. WARNING sin=
6. Install the T1I-600 fess in the power circuits of all the three do-tee
sections.
The aircraft is provided with light and sound warning system. The light
,
rualegasten consists of warning lamps, type 0)111-51, of various colour mount-
?. Connect a voltmeter between the aircraft body and the plus terminal
?n teaks, boards and instrument panels.
(power bolt on the terminal block) of the fin leading edge heater.
Er light warning devices are designed for signalling:
8. Turn on the de-icer switch and measure the current consumed by the
e;
section of the. fin leading edge heater and voltage at this section. 1. preparation of th engines for starting;
9. As soon as the fin leading edge heater is out, turn off the de-icer 2. oil pressure in the turbostarter;
switch.
3. operation of fuel pumps and determination of the order of fuel consunp-
10. In a similar way measure current and voltage on the inside and outei4edm4
,
sections of the stabilizer de-icer. 4. fire and opening the fuel shut-off cocks;
11. When checking according to Items 8, 9 and 10, simultaneously check 2m release of the brake parachute;
hand feel the serviceability and sequence of switching on the de-icer sectims
,
5. pe
tic
0. pressure drop in the hydraulic system and oration of the brake auto-
Wake sure that the inside or outside sections of the stabilizer leading edge a milt;
heaters start operating simultaneously. Asymmetric operation of the dc-Leer' 7. landing gear and tail Support position:
is not permitted. 8. trim tab neutral position;
%SPEED TOO HIGH (CKOPOCTB.BE)IWKA );
CAUTION. It is prohibited to switch on the tail empennage de-Leers on lb 10. switching on the APE-2 gyro horizon;
ground for a longer period than one operation cycle of the MICA-3A
11. armament position prior to aircraft landing;
electric mechanism.
12. camera ti un
lting mot and camera hatch position;
12. Determine the current consumed by each section of the de -icers.at a
14. charging
13. operation of the tail empennage de-icers:
voltage of 26 V across the heater terminals. The current is to be determined 1*.
mit;
the formula:
I=
26 I measured '26
V measured
-be OUZ:n
13. pressure drop in the pressurized cabins;
M. outside signalling by signal flare launcher.
cocks open position and FUEL DELIVERED (TOMB? flOAAHO )
where 126 is the current consumed by the de-icer section at 26 V across t
Tr:aindi::::idnin=7.: their arrangement on the
heater terminals; aircraft and nature of
'measured is the current measured during the test of the given sectieb
=enured is the voltage measured during the tent of the given section.
V
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126.-
Wept Pitting Sligg51
The light device, type CSU.-.51 (Pig.79) is designed for signalling inft
this'aircraft or for use at a light indicator of the aircraft separate unite a
mechanisms operation. The light device ensures the possibility of operation
under day- and night-tine conditions. The reduction of signal brightness by
turning the davits cover to the right as far as it will go during the night th
makes it possible to preserve the adaptation of the aircrew member's eyes to
the low brightness background. The position at which all the four holes Lath
and face walls of the device head and cover are open (the cover with the ligh
filter is turned to the left as far as it. will go) corresponds to the day tut
conditions. Besides the two extreme positions, the cover with the light fines
can be set to any intermediate position which corresponds to partial openinge
Vas three triangular holes in the end face walls of the device head and coven
2ho aircraft is provided with devices of five colours: red, green, yens
blue and white.
The C11r5Lfitting is intended for use with a special aircraft lamp (ram
for 28 V and 0.17 I), type 016-30, with a single-contact base 10-9-1.
2be sound signalling is performed by continuous and intermittent buzzing
of aircraft horn, type 0-1 (718.80). The aircraft has two 0-1 horns; the arm
sent, use and nature of operation of the horns are given in Table 23.
The transmitters of intermittent sound signals are cabin pressure wards;
units, type B0-46.
The trannnitters of continuous sound signals axes
(a) in the event of aircraft take off with the flaps retracted or arterial
by an angle below the rated value - mechanism, type MKB-2, mounted on the.f/g,
traneniesion shaft (at trace 80.33) and limit switches, type B12-142 (front),
mounted an the right-seat pilot engine control panel;
(b) in the event of throttle control retraction (aircraft laamag) withn
landing gear =extended - blocking contacts, type 812-142 (rear), mounted man
right-seat pilot engine control panel and the landing gear extended positiod
limit switches.
becifications_ of 0-1 Berne
1. Voltage range
2. Nominal voltage
5. Nomdmal duty of operation
4. Maximum current consumed at 26 V
5. Sound intensity at 26 V (as measured at a
distance of 1 m. from the horn)
6. frequency of contact opening at 26 V
7. Horn weight
20 to ya v
26 V
intermittent (one-dO
operation, one-minuts
interval)
0.85 I
not less than 80 db
200 to 310 c.p.s.
not in excese of 1.11
The cabin pressure warning unit is designed for closing the electric eine
Of the sound and light signal mates warning the aircrew of the necessity of
using oxygen apparatus.
- The warning unit, type 26-46, is a unit of four aneroid boxes connected"
the electric circuit moving contact. If pressure drops below the rated value,
the boxes unit closes the circuit contacts and sends electric signals to tie
P8-12 buzzer relay. Phe warning unit is adjusted for signal transmitting at
altitudes from 1000 to 5000 m.
25X1
- 127 -
Epecificatione of 26-46 Cabin Pressure Warning Unit
fb,, cabin pressure IMMIng unit must continuously
Bog out light and sound signals from the moment
plfseure decreases in the cabin to a value correspond-
to the altitude set at the dial.
2.6sn6e of adjusting the unit for the beginning of the
signal transmission according to pressure in the
pressurized cabin corresponding to altitude in compliance
with the international standard atmosphere
3.Instrwmint temperature range
4.instrument error during signal sending out at normal
teuperature at scale marks: -; 2; 5; 3; 3.5; 4: 4.5:
5 km.
from 1000 to 5000 m.
free +50 to -60?C
5. Wit must operate during vibration within the
frequency range
6. Reliability of the electric contacts must ensure
7. Instrument weight .
not in excess. of
1150 m.
from 20 to 80 c.p.s.
and overload of
2.5 db.
up to 1000 switch-
ings
not in excess of 450 gr
(with the plug connec-
tor)
Cabin pressure warning units are mounted in the front cabin at frame 10.5
(starboard side) and in the rear cabin at frame No.75 (starboard side).
To obtain intermittent light and sound signals, connected to the circuit
of each cabin pressure warning unit is a buzzer relay, type P1-12 , with two
aqmoitors,typeK3-11-50 .22. - V. The relay and capacitors are installed in
the boxes of the sound signal relays (Figs 81 and 82).-The.relay boxes are mounted
intim front pressurized cabin on the navigator-radar operator left-hand rack and
in the rear pressurized cabin on the starboard side at frame No. 73.
The relay is switched on by the operation of the cabin pressure warning r,nit
calumets and ensures intermittent duty of operation.
1.
2.
Decifications of P1-12 Belly
Nominal voltage
Frequency of relay operation at nominal voltage
3. The relay normally under the
e
(a) ambient air temperature
(b) atbient air relatifs htmlidity
(c) sea-level altitude
(d) aircraft vibration
4. OPerating voltage
5. Lmmture attraction voltage
operation)operating sillier load
6.
'
Inatura drop-out voltage
7. Relay weight
following
5
26.5 V
3 to 5 switchings
per second
from +50 to -60?C
up to 98 per Cent
up to 15000 m.
not in excess of 14 V
not in excess of ]2 V
2 to 5 V
not in excess of
195 gr
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Table 22
Light Signalling
Nos
Purpose
Number
of
limps
Signal devices
Note
Type. of
device
Conditione under which.
'blab device operates
Nature of
operation
Location of device
1
2
3
4
5
6
7
8
1
2
3
Signalling of
the engine readi-
nese for starting
Signalling of
oil pressure in
the turbostarter
Signalling of
fuel consumption
sequence
2
1
A
0211-51,
green
014-51,
green
CA11,51,
blue
With the turbostarter .
exhaust gas shutters open
At oil pressure in the
turbostarter exceeding
3.5 atm.
let group lamp flashes
up when the fuel consump-
tion control switch is
set to the AUTOMATIC
(ABTOMATM. ) position
and the amplifiers
switches turned on. The
rest of the lamps flash
up in sequence after
200 lit, of-fuel is left
in the previous tank
6....3)
Constant glow
of the lamps
Constant burn,
ing of the lamps
Constant glow
of the lamps
On the engine start-
jog panel on the left-
seat pilot's engine
control console
?
On the turbostarter
control panel
On the fuel supply
panel
1
2
3
4 ,
. :5
.6
7 ?
a
.
'
4
Signalling of
fuel available
4
024-51,
red
Two lamps flash up with
fuel available for a
Constant glow
of the lamps
On the fuel supPLY
board
for 30, or 15-30
minute flight
-minute flight, the
other two lamps flash up
. on the left-
seat pilot in-
,
,with fuel available for
a 15-minute flight '
.
. .strument panel
.
5
Signalling of
fuel pumps opera-
tion
12
04-51,
green
With the pump operat,.
Log and pressure in the
system reaching 0.3 or
Constant glow
of the lamps
On the fuel supply
board
,0.33 kg per sq.cm.
6
Signalling of
engine shut-off
cocks open posi-
tion
2
02111-51,
green
With the engine shut-
off cocks open from the
beginning of engine
starting to its stopping
Constant glow
of the lamps
On the fuel supply
board
7
Fire signalling
6
Rod
With the fire-fighting
system switched on, when
temperature in the area
of the fire-sensitive
units reaches 140 to
Constant glow
of the lamps
On the fuel supply
board
'
17000 or on pressing the
button
f;:tnalling of
the brake para-
chute release
2
021LI-51,
green
With the brake pare-
chute released
Constant glow
Of the lamps
On the instrument
panels of the right-
and left-peat pilots
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1
1 2
3
4
5
6
7
9
3,
liailoron
Signalling of
preseur? drop in
the normal and
emergency hyd-
vault? systems
Signalling of
the brake auto-
natio unit opera-
tion
Signalling of
the armament pa-
sition during
landing
Signalling of
the tell suppOrt
and leading gear
position
Signalling of
the rudder and
trim tabs
position
2
1
2
8
3
07111-51,
red
c2i4-51,
blue
mul-51,
blue
CU-51,
five
green
and three
red larP6
034-51,
white
At pressure in the
normal hydraulic system
less than 100 kg per ao.cm.
and in the e0ergenc7
hydrulio system lees than
130 kg per sq.cm.
With abrupt braking of
the wheels, with the
brake automatic unit
switch turned on
With the armament bar.
role of the lower and
rear sighting stations
in the lowered posi-
tion
The three green lamps
indicate the landing.
gear lege extended
position, the three red
lamps - the landing gear
lege retracted position
With the rudder and
aileron trim tabs in
the neutral position
Constant glow
01 the lamps
Blinking of
the lamp
Constant glow
of the lamps
Constant glow
of the lamps
Constant glow
of the lamps
On the pilot's'
cientmal electric
board
On the pilot's
central electric
board
On the left-seat
pilot instrument
panel
On the pilots' ,
central electric
board, on the gunner -
radio Operator's and
gunner's eleotrio
board
On the left-seat
pilot's instruments
panel, in the aileron
onmalaseminatri4n ye.441 ,
13
1 'f 2
,
4
5
a
7
a
14
SPEED TWO HIGH
(CHOROCTI, BERRA)
signalling
2
C2111-51,
red
With presnura bead
amounting to 2300 kg per
sq.m. at low altitudes;
at 160.86 at high alti-
tudes
Constant glow
of the lamp
On the pilot'.
instrument penele
15
Signalling of
the stand-by gyro
horisoA switch-
tug
1
0)I4-51,
red
With the stand-by
gyro horison switched
on by the left-seat
Pilot or navigator*
radar operator
Constant glow
of the lamps
On the left-seat
Pilot's instrument
wail
16
Signalling of
the camera hatch
in the open poet-
tion, camera
tilting mount
position during
air survey and
checking
3
cm-51,
green,
white,
7?11"
(a) The green lamp is
on with the hatch open
(b) The white lamp
flashes, when the oamera
tilting mount passes the
zero position, in SURVEY
MODS OF OPERATION
(P118BEAXA)
Constant glow
of the lamp
811?kirg of
the lamp
On the navigator's
right-hand console
(o) The yellow lamp is
on with the camera tilt-
tug mount operating in
Constant glow
of the lamp
CHECK mode (KOHTPOIL )
at tilt angles of 0, 10,
15, 20 and 250?
17
Signalling of
1
031a-51.
On switching on the
Blinking of
On the right-seat
the tail empennage
de-leers operation
white
dc-loop outer sections
the lamp (40-
sec. glow fellow-
ed by 80-sec.
interval)
pilot instrument
panel
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1
2
3
4
5
6
7
8
18
19
Signalling of
the fuel supply
cocks open posi..
tion and FUEL
DELP/MEE (IM)M-
BO ROAM)
signal
Signalling of
pressure drop in
the cabins
5
5
O14.61,
four green
lampsone
yellow
lamp,
CAH-51,
yellow
.
(a) The green lamps
are on when the fuel
supply cooks are switon-
ed on for opening
(b) The yellow lamp
is on when the FUEL
DEIIVEHED (TOMO
1104A80 ) warning unit
is switched on
When pressure drop
corresponds to the
altitude set at the
dial within 1000 to
3000 m.
Conetant glow
of the four lamps
Constant glow of
the lamp
Blinking of the
lamps
On the fuelling
control board
On the navigator'
oxygen panel, on the
left-seat pilot in,
strument panel, the
navigator-radar
operator's oxygen
panel, the gunner-
radio operator's
instrument panel,
the gunner's elect-
rio board
T a b 1 ? 23
eeoad inguallinm
Nos
Purpose
Signal devices
Note
Number
of
devices
Type of device
.
Conditions under
which the device
operates
Nature of
operation
Location of
devices
l'
2
3
4
5
6
7
8
1
2
3
Signalling of
flap position
during take off
Signalling of
lending gear re-
traoted position
during landing
Signalling of
pressure drop
below the value
rated for pros-
surized cabins
1
1
2
Horn,type 0-1,
of the front
pressurized cabin
Horn,type 0-1,
of the front
pressurized cabin
Horn, type 0-1,
of the front
pressurized cabin
Horn, ono 0,1,
of the rear pros-
eurised cabin
With the flaps not
extended to the take
off angle, that is
to a value from
19 + 10 to 23 + 10
and with both throt-
tle controls set to
the position corres-
ponding to the air-
craft take off
When at least one
landing gear leg is
not extended and at
least one throttle
control is set to
low throttle (during
landing)
When air pressure
in the pressurized
cabin is below the
value set at the dial
of the cabin pressure
warning unit,type
B4-46
Constant buzzing
of the horn
Constant buzzing
of the horn
Intermittent
buzzing of the
hem
On the port side .
at frame No.9 of
the front pres-
surized cabin
On the port side
at frame No.9 of
the front pressu-
rized cabin
On the port side
at frame No.9 of
the front pressurized
cabin.
On the starboard
side at frame No.71
of the rear pressurized
cabin
.
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? 134 ?
Care and Maintenance of Light and Sound Signal
Chits
The light, type ca.51 , horn, typo 0-1, cabin pressure warning mats.
type 110-46, and the buzzer relay, type 11-12, de not require special care ex
maintenance. During service take care to see that these units are securely i
attached, and the contacts of the connected wires are in good condition and
properly tightened.
Maintenance of the light signal fittings in the vain consists in replack,
burnt out lamps. To replace e lamp in the 0.111.61,fitting, it is necessary to
unscrew the held with the light filter, fit a new lamp into the holder and acrl
up the head again. When soldering the wires to the fitting, observe polarity
(the plus wire meet be soldered to the fitting central contact). /f polarityh
incorrect, short circuit may take place during the replacement of the lamp
lender voltage.
In cane the C-1 born is being replaced or the supply conductors are bet%
connected to the horn during installation of the horn cap, when the attachmmt
screw is being tightened, ensure not only proper attachment of the cap but tin
normal sounding of the horn. The horn is adjusted by the Manufacturer. Durtz
servica the horn does not require any additional adjustment.
As the signal fittings under voltage are checked only in conjunction with
the operation check of the mechanisms included in the trysts= provided with ti
signal fittings, the data concerning the methods of adjusting the limit evitct.;
are given in the sections of the Instruction dealing with the operation of
corresponding mechanisms, units and devices.
As a rule, various types of light signal devices are cut in when the
corresponding units and mechanisms are switched on by means of cutout and Wane'
over switches located on the instrument panels and electric boards.
Sound signal devicee are switched on by means of switches, type B-45, ant
ed on the rheostat panel (Fig.83) of the right-seat pilot engine control Paul
and on the gunner.radio-operator's electric board (Fig.84).
? AIRCRAFT INTERIOR LIGHTING
For aircraft interior lighting the following fittings are used:
(a) dome lights, type 0-45 and UCM-51
(b) light fitting, type HICPK-45;
(c) ultra-violet scale illumination lights, type 111001].45
In addition to this fitting extension lamps, typo fl-1O-36, are used to
illuminate dark places on the aircraft.
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? 135 ?
halation to the dome lights, t5pe W-45 , for general illumination of
? gearcompartments the aircraft is provided with small dome
oga ,.hich ere used by the aircrew or attending personnel during repair or
'00mmee operations in the landing gear compartments. The FIC1-5I damn lights
ditfar from the 0-45 dome lights in that they are not provided with protection
fp. the blinding effect of direct lamp rays as these dome lights are within the
sight of the aircrew for a short time.
he main parts of the smaller um-51 dome light (Fig.86) are: body, Iprztec-
? eavarent glass' and reflector whose neck mounts a single-contact socket for
601-24 lamp of 28 V and 20 W. The second pole for the ncm-51 dome light an
wale for the nc-45 dome light is the dome light body and aircraft frame.
TO replace the lamp in the WH-51 dome light or the entire lnminalre,first
,dia ream* the protective glass. When mounting the protective glass, take
protease that the spring retaining the glass and the reflector is mounted
gamely. The spring must pass between two lugs on the glass bowl. To retain
&Elms in position , bend the spring when it is weakened in the central part
*wore reliable attachment of the cap. For maximum glow keep the protective
eumalenity wiping it regularly with a clean piece of cloth.
Why sang the 110-45 and ncm-51 dome lights, check to see that the union
ant of the nipple on the dome light inlet pipe union is tightened up at all
dm, as loosening of the nut may result in poor contact and flickering or dying
at of-the lamp.
The location of the dome lights, types 110-45 and WM-51, as well as
*location of the switches designed for switching these dome lights on and off
Mindicated in Table 24.
Dome Lights, Tlpes 110-54 and . UCH-51_
The dome light, type 110-45 . (Fig.85) without a special lens but withs
reflector and a aingle-contact holder for the Cl1e.25 lamp of 28 V and 20 W is
lenteendedwfvot illumination the press:::
compartments. The bulb of the 084-25 lamp has a bowl of plate glass to protect
red
aluminium
the aircrew from the bliedieg effect of the lamp ray::::::::: ietn:unpresein'jzehisr::EE: i
recommended to use luminaires with other lamps (with open bulbs).
ing
The inner surface of the dome light body which serves as a reflector is
prope operation of the luminaire, it is recommended to wipe the inner surfecee
the reflector with a clean moist piece of cloth or cotton wool.
Light Fitting, Type OPCK -45_
For illumination of panels, boards, dark places and instruments the aircraft
Is provided with fittings, type KICPK-45, with a rheostat and a button, which
fate CW-30 lamps of 28 V and nominal current of 0.17 A.
Tie aircraft has ten cabin /amps altogether (two of them are mounted at
the navigator's seat, three lamps - at the pilot's seats, two - at the navigator-
mac operator's seat, two - at the gunner-radio-operator's seat and one - at
Cm gunner's seat).
Cabin lamps, type 1012011-45 , are mounted on special hinged brackets (Fig.87)
diet sake it possible to use these lamps for directed illumination of several
*COS. On some of the hinged brackets the cabin lamps are mounted together with
alte.dolet illumination fittings. If necessary the K1lPC11-45 fitting can be
Mewed from the hinged bracket or from its base and used as an extension light-
? device for temporary illumination of some section in the cabin within the
ImgMeof the cord.
The rheostat for adjusting the lamp light and the button by means of which
lbs rheostat can be short circuited temporarily and the lamp caused to flush
81ell glow are located on the fitting case. By changing the distance between
tlalens and the lamp which is ensured by moving the cylindrical nozzle on the
fitting it is possible to obtain more dissipated and more directed lighting.
The cabin lamp, type KIWK-45 , is switched on ard off and its filament is
tklvated by means of the rheostat handle made of colour plastic and mounted on the
Siting b,A3,. Replace lamps in the HAFCH -45 fitting in the following manner;
(1) turn out the stop screw fastening the cylindrical nozzle;
(2) remove the cylindrical nozzle;
(3) replace the lamp;
(4) taunt the cylindrical nozzle and turn in the stop screw. When fitting
lovtop screw, see that a metal washer is placed under the screw head.
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T a b 1 ? 24
location of HC-45 'and ncm,51 Dome Ideate and Their
Seitohee on Aircraft
Nos
TY" ?f
dome light
Place of dome light installation
_
Pleas of switch installation
Note
1
2
3
4
5
1
RC-45
On the ceiling of the front pres-
surised cabin between frames Noe 4
ana 5
On the dome light mounting
penal
For the navigator
-
2
IEC-45
On the ceiling of the front pram-
surieed cabin at frame No.9
On the dome light mounting
panel
For the pilots
3
110-45
On the oeiling of the fuselage un-
On the navigator-radar opera-
For lighting the Wiraulic
pressurized portion at frame No.14
tor electric panel
panel
4
110.-45
On the ceiling of the fuselage un-
pressurised portion at frame No.20
On the starboard side on the
bracket of the No.4500 inverter
support
5?
110-45
On the ceiling of the fuselage un-
pressurized portion at frame No.34
On the navigator-radar opera-
tom's electric board
For lighting the bomb ba7
6
110-45
On the ceiling of the fuselage un..
pressurised portion at frame No.38
On the navigator-radar opera.
tom's electric board
For lighting the bomb bar
7
110-45
On the ceiling of the fuselage un-
pressurized portion at frame No.42
On the navigator-radar opera-
tom's electric board
For lighting the bomb bay
8
110-45
On the ceiling of the fuselage un-
pressurised portion at frame No.46
On the navigator-radar opera..
tor's electric board
For lighting the bomb bey
9
110-45
On the ceiling of the fuselage un-
preesurised portion at frame No.49
On the navigator-radar opera-
tler*e electric board
For lighting the bomb bay
110-45
n0-45
On the starboard side of the fuse.;
lege unpressurlsed portion at frame
No.66
On the ceiling of the rear pressuriz-
ed cabin at frame No.71
On the ceiling of the rear pressuris-
ed cabin at frame No.71
In the landing-gear port leg well
In the landing gear starboard leg
well
In the nose leg well on the star-
board aide at frame No.20
In the landing gear nose leg well
on the port side at frame No.20
On the starboard side at frame
No.62
On gunner-radio-operator's
electric board
On the gunner's electrio board
On the navigator-radar opera-
tor's electric board
5
For the gunner-radio-operator
Por the gunner
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ft:tension 'saps. Type 31:10t-36__
132 additional lighting of dark places on the aircraft the latter is mei& 1
*death extension lamps, type 1U1,10-36, (asaba). The 32-10.36 extennion lani;
has a 10-2, cord and a filament lamp, type OM-15, of 26 V and 10 11, The extender
lamps' type DI-10-36, are switched on and off by means of the switch mounted ce
the handle of the lamp carbolite body.
The aircraft has throe extension lamps altogether kept in special bags
located in the following places: on the rear well of the pilots' central console;
an the wall of frame No.9 (starboard elide) and in the rear pressurized cabin an
the port side at frame No.73.
kr:tension lamps. type 111-10-36, are connected to the aircraft mains by
means of two-pin plugs. Receptacles, type 47K, for these lamps are mounted in
various places of the aircraft. The aircraft has 13 reoeptecles, type 47g,
altogether; four receptacles are mounted in the front pressurised cabin, four
receptacle.- in the fuselage unpressurized portion, two receptacles - in the ago
nacelles compartments, two receptacles - in the landing gear main legs wells eng,o;
receptncle- in the rear pressurized cabin. The location of receptacles, typo 47:0
described in Table 25. .
Table 23
Location of Plum Connector Receptacles 47k for
111-10-36 Portable Lamps
Nos
Place of receptacle installation
Note
1
2
3
5
6
.7
8
On the navigator's right-hsiut console
On the pdlot's central comma.
On the port side at frame 130.5
On the well of the upper past of frame
NO.9
On the bracket of the II0-4500 interter
rack (starboard side)
On the fuel pumps starboard junction
box at frame No.33
On the junction box of the extension lamp
mounted on the distribution board of the
port engine nacelle
On the junction box of the extension lamp
mounted on the dietribation board of the
starboard engine nacelle
Nor illuminator of the
might; it is switched on
through a special rheostat
from the set of this sight.
The rheostat is mounted on
the left-seat pilot's in-
etrument panel
For illumination of the
aircraft sextant
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1
2
9
10
12
13
On the junction box of the extension lamp
mounted on the port side at frame No.5 of
the landing gear starboard leg well
On the junction box of the extension lamp
mounted on the starboard aide at frame
No.3 of the landing gear port leg well
On the fuel pumps junction box at frame
go.49 (starboard side)
On the starboard side at frame No.62
On the port Side of the rear pressurized
cabin at frame No.72
11-12 CO2Ralls Illumination
For illumination of two compasses, type 131-12 , mounted at the navigator's
met ia the upper part of frame No.1 and at the pilots' seats on the cabin
canopy frame, special lamps are provided. The lamps are mounted right into the
boolg of these compasses. Each compass illumination lamp is switched on by means of
the switch. type 3-45. The navigator's compass illumination switch is mounted on
de navigator's upper electric board, while the pilots, compass illumination
,witch is on the pilots' central electric board. Both lamps are supplied by the
triple-duty supply basher through the automatic circuit breaker. type A3C-5,
smarted on the right-seat pilot circuit breaker panel, that is the ;01-12 compass
illumination is ensured when the aircraft mains; operates in all duties.
The triple-duty supply busbar supplies through the same 23C-5 circuit
belabor one of the receptacles, type 472, mounted on the pilots' central panel.
The rest of the circuits of the light fitting are supplied from the normal
awlybusbare and are protected by the automatio circuit breakers mounted an
pmels and boards of the front and rear pressurized cabins The 412M%4210U lamps
neeptaelee mounted outside the pressurised cabins and done lamp. are protected
by glees fusee.
Luminaire, Type 0700E45
The cabin luminaire, type INOOM-45 , is designed for ultra-violet illumine-
ticm of the instruments and the control units (electric boards and instrument
Innele) to cause luminescence of the luminous compounds as well as for lighting
porposes. Ultra-violet illumination is performed by means of special aircraft
tadneecent mercury lamps of low pressure with rated power of 4 N. type V00-42.
The luminaire, type 21160M-45 ,' is used in conjunction with the PY00-45
Seostat designed to switch on the ultra-violet illumination lamp and to control
ha light intensity. The 11100M-45 fitting is provided with special twin-conduc-
tor in a common copper braiding which serves as a third conductor.One of the
0MM of the twin-conductor has white insulation, the other end has white insulation
With a black thread. The conductoz having white insulation with a black thread is
mmluded from the lamp connection circuit and is insulated. The braiding of the
fitting cord in connected to the aircraft framework either directly or through
the aircraft conductor, type 11B1, connected to this braiding.
The lamp plastic body bas a cylindrical nozzle with two light filters of
bleak uviol and a hinged base for the luminaire. The upper light filter may turn
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together with the nozzle ring within 90?. By rotating the nozzle ring it is
possible to:
(a) match the slots of the both light filters; in thiscase white lighted
the lamp passes through these slots;
(b) overlap the slots; in this case only ultra-violet rays pass through
the uviol light filters. The ultra-violet rays cause luminescence of luminous
compounds.
Most of the luminaires, type APY0091-45 , are mounted on special hinged
brackets (See F1g.87) which make it possible to use this lamp for directed
illumination of several places. On some of the hinged brackets the luxdraires
are mounted together with the NIPCK-45 fitting. If necessary, the fitting mark
removed from the hinged bracket or from its base used as an extension luminaire
for temporary lighting of some area in the cabin as far as the cord permits.
The aircraft is provided with APYTOM-45 fittings with rheostats, type
PY00-45 . Three APY8OM-45 fittings are mounted in the rear pressurised cabin,
the rest of them are installed In the front pressurized cabin.
The arrangement of the ultra-violet illumination devices and PY00-45
rheostats is shown in Table 26.
Table 26
Arrangement of IFY8OM-45 Luminaires
and 1700-45 Bleostats
Nos
Place of luminaire
installation
Place of rheostat
installation
Note
1
2
3
4
1
On the starboard side
On the navigator's
of the front pressurized
cabin on frame No.2
upper electric board
For lighting the
sight, the instrunn
2
On the ceiling of the
front cabin on frame
No.3
Same
panel and the Davits.
ton's right-band
console
3
On the ceiling of the
front cabin at frame
liC.4
Same
On the right-seat
pilot's steering wheel
On the-right-seat
pilot's engine
control panel
For lighting the
Samepilots'
Bane
instrument
D4461
On.the left-seat
pilot's steering wheel
On the left-seat
pilot's engine contra
panel
i
7
Same
'Same
a
On the port side at
frame No.8
Same
Together with the
UCH-45 fitth4 9
serves to Melt 12
the pilot's inatzt-
rent panel
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141 -
1
9
2
4
10
11
12
13
14
15
16
On the ceiling of the
front cabin on frame No.8
On the starboard aide
at frame No.8
On the fuel supply
panel
On the left-hand side
of the upper blister
On the right-hand side
of the navigator-radar
operator's central panel
On the port side of
the rear pressurized
cabin at frame No. 74
On the starboard side
of the rear cabin at
frame No.70
On the rear cabin cir-
cuit breaker panel on
the port side at frame
No.71
On the left-seat
pilot's engine control
panel
On the right-seat
pilot's engine control
panel
On the right-seat
pilot's engine control
panel
On the navigator-
radar operator's
instrument panel
Same
On the gunner's
electric board
On the gunner-
radio-operator's
electric board
Same
Together with the
5I075-45 fitting it
serves to light the
pilots' upper
electric board and
the fuel supply panel
Together with the
gla5-45 fitting
it serves to light
the right-seat
pilot's panel
To light the
pilot's central
panel
Basic apecifications of the 140 Set
1. Total resistance of the 7700-45
2. Rheostat resistance in the cutoff position
3. Normal current of lamp operating duty
4. lamp resistance with the rheostat cut off
5. Igadnous compound brightness adjustment range of the
1'10045 rheostat
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Together with the
EFUN-45 fitting it
serves to light the
instrument panel
and the navigator-
radar operator
central panel
Same
For lighting the
instrument panel,
board and the gun,-
roes panel
Same
Same
35 ohms
22 ohms
0.35 A
0.5 to 0.6 A
from 150 to 30%
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6. Y40,-4A lamp flashes up
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100% is azAt-
to be tbe
nese of imitl
compound jj
noted by All
operating er4,
nominal comli.!
tionm (at ?35
in no more thl.
12 seconilatk
switching
7. Lamps which were on for not less than 10 min, at
nominal current of 0.35 A flash up again on switching in no less Um
2 minutes hm
the =Mead
switching
Switching On_ Y00 Lapp
The 700 lamp is switched on by means of the rheostat, type 7Y01045. In
the idle position the rheostat handle must be set to the OFF ( BUMPIER? )ya,,
tion at all times. The 5111.4A lamp is switched on automatically in no more 1
than 12 sec. after the rheostat is set to the ON (RIONEHO ) position; it is:
necessary to set the PY006.45 rheostat handle to the right as far as it will;,
On switching on the lamp, adjust as required the brightness of the scales
and stencils covered by luminous compound or the degree of illuminance duriemt
operation with the light filter open. The rheostat handle being turned to the
left, the brightness decreases, the handle being turned to the right, the
brightness increases.
Replacement of Lamp in APYOOM-45 Fitting
1. Remove the cylindrical nozzle with the light filters. To detach the
cylinder from the body, press with the finger the lower part of a special pa
riveted ts the cylinder. This done, turn the fitting cylinder counter-elm/oda
and remove it from the body.
2. Replace the lamp. When fitting a new lamp, type Y00-411 bear in milt
that the lamp has a bayonet base with pins (type 2c-15A-1); the base pins are
located at various height due to which the lamp cam be inserted into the socket;
only in one definite position ensuring correct polarity of switching.
3. Fit the cylindrical nozzle with the light filters. When fitting it oz
the body turn the cylinder clockwise until the elastic pin clicks in the hole.
EXTERIOR LIGHTING
The aircraft exterior lighting consists of the following
(a) taxiing lights, type OP-100;
-(b) landing lights, type 100B-45;
(e) formation lights, type 11000-45;
(d) navigation lights, type EAH045 , and I0-39.
light fitting,:
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TesiLng "Apts.. Type OP-100
aircraft tailing lights, type OP-100, are designed for illumination of
earrouddluing taxiing in the night time. For the landing gear nose leg the
1110.01.p in mounted right on the landing gear strut while for the landing
gwvialegs the leading lights are mounted on the struts.
MS taiding light, type OP-100, (F143.09) consists of a case with a reflector
sda :angle-contact socket for the C8,21 lamp of 27 V and 2.7 A. The landing
uippeteetive glass which is a colourless transparent disperser ensures angle
domes:ion in the horizontal plane equal to 30?. The narlerun lusinona intenxi-
of the taxiing light is 3000 candles.
With the aid of the adjustable bracket each taxiing light can be set to the
ideidzaiddch ensures illumination of part of the leseisg strip at a distance
,c1.5 or 20 L. from the pilots' cabin in the direction of flight. The tartiug
Rosie fixed in the required position by means of a nut and a lock nut.
jji the three tailing lights are switched onsimultaneously with one switch.-
wep-45, mounted on the pilots' upper electric board (Fig.90).
landing bights.. Dv 100B45
Te illuminate the place of aircraft landing in the night tine, the aircraft
Is provided with two retractable lending lights, type IO0B-45 , installed in the
hsacpart of the Mose unpressurized section of the fuselage at frame No.13. The
*mutable part of the landing light consists of a casing and a xpecial reflector
hasp, type 0114-2111 , of 28 V and 600g. The CM0-21 lamp consists of a filament
Imp msper, a reflector and a protective glass.
The landing light control electric mechaniam, type MEO,..2 consists of a
sausible electric motor of series excitation, a reducer and a cutting off
=tact device.
The landing light is supplied from a single-line mains; the second pole
fathe light and for the electric drive is the landing light body and the air-
mart true. The light is switched on automatieally when the landing light is
saimatid. Switching off is performed automatically too wham the lasaleg light is
ntcacted.
The landing lights are controlled by means of a switch, type 21111-45, from
hap:lots' upper electric board (Fig.91).
gpecifications of landing Lig*. Tue 100-45
1. Tad= luminous intermit)" not less than
350,000 candles
2. lending light angle of dispersing:
Lithe horizontal plena not less than 120
in the vertical plans . net lens than 8?
3. leading light extension angle 06?30, ? 30,
t.tbse required for extension or retraction of the
landing light not in excess of
12 sec.
1 Ammiesible continuous glow of the landing light not in excess of
5 min.
6. CO-21.9.Lalerpoturr
: by life guaranteed the
7.0plIziania%:oltage range for the NUO-2 electric
!
5 hours of burning
24 to 30 V
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2.8 A
8. Current consumed by the electric mechanism
9.-Maximum moment 220 kg-ca
10. Duty of the electric mechanism operation intermittent
11. Weight of the landing light with the electric mechanism 3.5 kg
The landing lamps, type 10013-45, installed on the aircraft in the extended
position ensure illumination of part of the landing strip at a distance of
60 m. from the pilots' cabin in the direction of flight.
Maintenance of landing /ightsi Type 1GCB-45
To check the operation of the landing lamps on the ground, it is permitted
to switch them on for not more than 5 minutes. The landing light may be switch4
on again only after they have been cooled during not less than 5 minutes.
Avoid shaking and knocking to prevent crack formation and failure of the
landing lights.
Chen checking the landing light In a workshop the supply voltage must not
exceed the rnm1n,a1 value of 28 V. otherwise the lamps may burn out.
To prevent decrease in the landing light luminous' intensity, clean the pail
of the lamp surface which serves as a protective glass.
pglacament af_qq0:21 .....
The CUT-211 lamp is a changeable element of the landing light and is rep1a0,
by the unit technician In the event of failure. To replace the lamp in the J0345
landing light, do as follows:
1. By means of the 21111-45 switch extend the landing light with the burnt
out lamp; at this the circuit breakers, types A80-5 and A3C-30, mounted on the
left-scot pilot circuit breaker, panel for the serviceable landing light must be
switched off.
2. Switch off the supply circuit breaker, type i3C-50, of the unsr.:-.viceath
landing light on the left-seat pilot circuit breaker panel.
3. Unscrew four screws 3 (Fig.91).
4. Remove retaining ring 4 and draw the lamp out of the streamlined case.
5. Disconnect the supply conductors from the lamp terminals and remove the
lamp.
6. Installation of a new lamp is performed in the reverse order.
When replacing the lamp, see that the rubber shock absorbers supporting CM
lamp bowl in the case are intact.
Re-adjustment of Turning Units of the mni-e Electric
Mechanism Sector
The M110-2 electric mechanism is adjusted for a turning angle of the sector
(landing light) equalling 780 ? 30' For the aircraft, model TY-16 , the landia
light extension angle of 86?30' ? 50. is necessary. Therefore, when replacing
the FOCB-45 landing light during service, bear in mind that it is impossible to
meant a new light on the aircraft without preliminary re-adjustment of the sector
turning angle of the 11118.-2 electric mechanism.
Re-adjustment of the electrio mechanism is performed in the following manna'
1. BeMOVe the seal and the check wire, turn out the screws fastening the
cover to body 2 and turn the cover aside according to Fig.92.
2. Loosen screw 3 and by moving limit switch 4 together with plate 5 along
guide grooves iee that the contacts of limit switch 4 are opened by stop 6 at a
angle of the sector (landing light) turning equal to 86?30' t 30'.
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- 145 -
5. Tighten up screw 3 again.
4, glaok the operation of the electric mechanism.
5. Place cover 1 in position, turn in the screws fastening the cover to the
body owl seal the electric mechanism again.
6. Wake corresponding entries In the Certificate of the M18-2 electric
lanc4nni".
abe re-adjuptment of the electric mechanism should be made by trained
personnel, DO damage to the inner connecting conductors, limit switches and other
&agate of the electric mechanism is permitted.
Installation of 100-45 Landing
Light on. Aircraft
Install the landing light on the aircraft so that the body of the landing
lightclectric mechanism is pointed in the direction of flight while the dimmer
located inside the lamp and designed for screening the direct rays looks with
its prominent portion towards the aircraft centre line.
The landing light is fastened in position with 15 screws 4 mm in diameter
parsing through the holes in the landing light flange.
Formation Lights, Tyle _ BCCO.45.
Formation lights, type ncco-4.5 , are used during group flights in the night
hiss or under conditions of poor visibility to allow the aircraft flying in the
rear to form up and to keep their proper places in formation.
When forming up above the leading aircraft, the upper formation lights are
used; when forming up below the leading airdratt, the lower formation lights are
used. The upper and the lower formation lights are installed over the axis along
the fuselage and over the wing span on the landing gear fairings so that during
the flight the illuminated aircraft resembles the letter T. The formation light
attachment is made flush with the skim by means of bolts and self-locking anchor
ate. The upper and lower formation lights are switched on by meant of correspond-
ing matches, type 8-45, mounted on the pilots' upper electric board (See Fig.91).
The DCG045 formation-light (Fig.92) consists of the following main parts:
alesinium body, the inner electrically polished surface of which serves as a
reflector of the socket holder with a single-contact socket mounted in it for the
CI-30 lamp of 28 V and 0.17 A, and a prismatic light refractor of blue polystyrene
serving as a light filter at the came time.
Epecifications of UCCEE4?_Fommation Light!
1. Maximum luminous intensity
2. lights visibility range in the direction of
maximum luminous intensity in the night time
/melees. loather
SECRET?
not less than 5.5
colour candles, with
the formation light
in the horizontal
position; it is
directed backward at
an angle of 45 to
500 up from the
direction opposite
to the direction of
flight
about 3 km.
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3. uzular width of lrminens beaa about 20?
This arrangement of formation lights enables the aircraft flying Dahill ,
smeethat higher to keep its pines in the formation taking- bearing .on the hrigt. I
est of the leading aircraft.
Iiintenanee of Aircraft Formation Lights,
1721012C045
To avoid decrease in the formation lights luminous intensity, clean tho
reflector and light refractor from dirt with a clean piece of soft cloth oe
cotton wool. It is prohibited to wipe the aluminium reflector or plastic
refractor with COATCO material.
CAUTION: The 1100045 formation lights are designed for operation on
during flights. To prevent the plastic light refractor from overheat*
and damage, switching on the lights for a long time during parking
not permitted.
Polyetyrene of which the light refmactor is manufactured swells and dime.
wee almost in all solvents - acetone, ethyl-acetate, ether, chloroform, bend,
benzine (containing benzol), toluene.
The vapours of these solvents also produce a harmful effect on plastics:
the colvents are absorbed by polystyrene and then slowly evaporate causing the
appearance of irregular aleeirg in the light filter and the formation of call
cracks near the surface.
If such a light filter is held up against the light, some silvery brillire
may be noticed in the plastics.
These phenomena considerably reduce the coefficient of the light filter
total passing capacity, its transparency and, therefore, change the light dile
tribution of the R00045 formation lights. Therefore, to avoid harmful inflow
of 'solvent vapours upon the plactics do not paint and dry the bomb bay doors it
the light refractors of the UCCO-45 formation lights mounted on them. Ifth
formation lights are already mounted, prior to painting remove the light
=tractors and tightly close the reflectors with some plug. To protect polystew
retractors, take care to see that they are not splashed with solvents.
Replacement of...1=0645 Fitting and Replacement
of Lane in_the_Fitting
To remove the ECO-45 fitting, turn out the fitting attachment screamed
disconnect the plus conductor. ?
When mounting a new formation light, see that the installation is pratism
the plans passing through the leap axis perpendicularly to the refractor priri
Sorb be parallel to the aircraft longituaSeel axis and the socket holder mut
face dammed (upward for the lower lights) and forward with flight. The foratien lights are fastened in position with five screws 3 mm in diameter (1iadr3
simultaneously through the holes and anchor nuts in the aircraft frame) in the
body of dome lamp 1 (See Fig.95), rubber gasket 3, light refractor 4 and rattle
Leg ring 5 holding the refractor in place. Thus, the formation light is
assembled simultaneously with its installation on the aircraft. The aspeushr14
location of attacheent holes in.the formation light excludes incorrect postitics
of the blue light refractor in relation to the reflector, nevertheless, mice
that the refractor prisms look inside the formation light. If the locating
diameter of the UCC045 formation light done lamp does not correspond to the I
cut in the aircraft frame recess, it is permitted to fit Vadhers measuring 1
25X1
- 147 7
upelmadMr the dame lamp attachment bolts and the recces bottom. In this cane
the ire lamp my project in relation to the akin outer surface by up to 1 we.
The dome lamp installed, fill the clearance between ring 5 and the aircraft
shewith Packing sealing thlokol putty.
es replace the lamp in the UCCO-45 fitting, do as follows: .
a. Tura out the attachment screws of the formation light fitting, remove the
"seeing ring and the light refractor.
2. Replace the lamp.
3. Icamt the light refractor and the retaining ring in place, screw in and
tighten up the attachment screws.
Iavigation Lights, Trpe EAH0-45
The navigation lights, type BAH0-45, are designed to be :Shown by aircraft
tithe air during flight and on the ground during taxiing.
The fairing of each wing mount front and rear navigation lights. Two red
lights, type ELHO*45 , are located an the port wing-tip fairing, two green lights
as located on the starboard wing-tip faixing. The navigation lights are installed
In recesses closed with plexiglass and are fastenmd in position with three bolts,
32112 diameter. The fitting is provided with lamps, type CII-22, of 28 V, 24 I
aUalueiroun intensity of 21 candles.
The navigation light, type BAH0.45, has asymmetrical light distribution.
Tm mudeum luminous intensity in the direction of flight in not less than. 33
etourcenelea which ensures the visibility range in the night time of about 5 km
mder normal conditions. In the horizontal plans light is emitted within 110?
outside from the direction of flight: in the vertical plane - within =900 up and
don from the horizon.
Replacement of lamp In EKHO-45 Fitting
1. Turn out the plexiglass fairing attachment screws and remove the fairing.
2. Turn out screw 7 fastening the light filter (Fig.94) and remove the glans.
3. Unscrew the lamp and replace it.
When replacing the lamp, take into consideration that the pi= of the lamp
hum are located at various.heightel this makes it possible to insert the lamp
Into the socket only in the definite position: the amalgamated surface of the
hth must face screw 7.
4. Mount the glass light filter and fasten it in position with screw 7.
When mounting the light filter, it is recommended that the glass end face
should be slightly covered with sealing tbiokol putty to peevent moisture from
setting inside the device. During the assembly take care to see that packing
gasket 2 is fitted under the glass and lead washer 6 is fitted under the head
Of attachment screw 7, otherwise the glass might break when the screw is being
tightened. After the light filter has been installed, it is recommended to cover
Um head of screw 7 with putty or paint.
5. Mount the plexiglass fairing and fix it with screws.
Removal of BAH0-45 Fitting
1. Remove the plexiglass fairing. ,
2. Turn out the light filter attachment screw and remove the glass.
3. Turn out the three fitting attchment screws.
4. Unscrew socket union nut 4 and disconnect the supply conductor.
5. Remove the ELH01-45 fitting,
Installation of the navigation light is performed in the reverse order.
SE&MT
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148 -
Vri_x2 -29
In the fuselage tail section the rear fairing lower part mounts a ta'
navigation light, type 10-39, with the CM-15 lamp of 26 I, 10 M. The tail lign
is switched on by means of the same switch, type 8-45, which is designed for
switching on wing-tip navigation lights, type EAH0-45 . The switch is mounInd
on the pilots' upper electric board (See Fig.90).
Re2lacement of_L5m2 in the_IC-29_Fittim
1. Unnerve the attchnent screws of 'the wire Lattice and remove the Lotto:,
2. loosen the glass shade attachment screws 7 (Fig.95) and remove the Eton
3. Replace the lamp.
4. Monet the glass shade in position and tighten up the screws.
Fit the wire lattice.
5.
Removal of ______________
1. Remove the wire lattice.
Z. Reece* tail light fitting attacnuent screws 7 and remove the fitting nft
the recess in the fairing bracket.
3. Disconnect the supply conductors.
Installation of the tail light fitting is performed in the reverse order.
FIRE FIGHTING EQUIEUENT AND FIRE WARNING
IIECTRIC SYSTEM
With the aid of the electric System:
(1) CO2 is delivered to the area where fire ?emus in the aircraft;
(2) fuel delivery to the engines is cut off;
(3) the fuel system in being filled with neutral gas.
The aircraft electric gysten includes the following units:
fire-sensitive unit TO - 28 pieces;
- electromagnetic shut-off cocks unit (unit 635900) - 2 pieces;
- push-button type lamp with a red light filter - 6 pieces;
- electromagnetic air valve 2512800 - 2 pieces;
- relay, ten* P11-2 - I piece;
- warning Lapp with a greeen light filter - 2 pieces;
- fuel cross-feed cock with the 1(35-2 electric mechanism - 1 piece;
? - firing mechanism - 10 pieces (four of them are intended for CO2 cyltan
and six - for neutral gas cylindnrs);.
CO2 cylinder switch button 5K - 1 piece;-
- feel shut-off cock with the 1135-2 electric mechanism - 2 pieces.
The units of the system are located in the following places:
1. lire-sensitive units on special brackets:
(a) in the area of the fuselage nose section fuel tanks: two-fire-sanaitki
units are located on frame No.17, one unit is located on frame No.22, one mat
is located on frame No.25 and four unite are located on frame Ro.33;
(b) in the area of the fuselage tall section fuel tanks: two units are
located on frame No.50 and two - on frame No.56;
(0) on the engines under the collapsible cowls: four fire-sensitive woin
are located an each engine;
(d) in the area of fuel tanks between ribs Nos 3 - 4, 6 - 9, 8 - 9 and
12 - 13 along the rear wall of the wing second longeron - four fire-sensittm
=its are located in between each pair of the ribs.
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2. gbe electromagnetic fuel shut-off cocks units are located on the ceiling
dtheNb44 ayballindlwetwrrariftresechingl"nande3warng
7anani38guall-button type lamps (Fig.112) are
Bonged oaths pilotb, upper electric board.
e. The electromagnetic air valve 2512800 is mounted an the engine.
5. Tbe relay, type Pfl-2 , and the warning lamps with green light altars
a 1scated on the pilots' upper electric rebl;:!:6. The fuel cross-feed cock with the electric mechanise is located
"fra7:2T: The mechanism designed for opening the CO2 cylinders is installed
avow 10.22, that designed for opening the neutral gas cylinders is mounted
on irpee8:11hoe.511Loard
enar electric board.
*button for switching an the 002 cylinders is located on the pilots'
and port engines fuel shut-off fire cocks are installer.=
the engines behind the fire wall.
Specifications of Electric Units Included
in_Snstem
1. DUm-eensitive unit TR (Fig.113):
operation range 140 to 17000
insulerce mot less than
2 megohms
2. Fuel shut-off cocks unit:
operating pressure up to no kg/cm2
nominal voltage 27 V "
moment on switching not in means
of 7 A
current with cocks open 0.5 A
time unit is energised not in excess of
20 min.
daimon pulling effort at the beginning of travel
with 6.5 = clearance, at nominal voltage and time
unit is energized not exceeding 15 sec.
9 kg
The Specifications of the M3K-2 mechanism are given In the Section "Fuel
Dnms Control and Fuel Quantity Measuring Electric System".
Checking Installation and Operation of Fire-
Fighting Xeuinnent Electric System
Carry out the outside inspection with the aircraft mains de-energized.
Daring the inspection make sure that:
1. The fire-sensitive units attachment is in proper condition, the diaphragm
ere free from dents and are not deformed, there are no foreign matter and no
aoisture between the body and the fire-sensitive unit diaphragm.
2. The glass of the push-button type lamp is not broken and is securely
Mod in position.
3. The 002 cylinder button, type 5K, operates without jamming and is in
good repair.
4. The firing mechanisms on the discharge bonnets of the CO2 and neutral
tino cylinders are screwed up and locked with the 0-3 explosive charges
immrted (P18.98)*
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fiVCIPMTT,
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CAVT/07: During service it is prohibited to touch the fire-sensitise mit
contact screw with the nut, to re-adjust the contact screw and preasth.
alaPhzegm?
5. The plug connectors of the fuel shut-off cocks units must be connected
in compliance with the marking (the plug connector bodies and the fuel shult.qe
cocks units are marked) and locked with vire.
Checking 0peration_of Energised _System
CAUT/ON: The neutral gas switch, type B-45, is turned on only when neutnd
gas is delivered to the system.
Neutral Gas HP_
The system is checked by connecting an extension pilot lamp to the firls;
mechanisms. For this:
1. Unscrew all the six firing menhenisme from the neutral gas cylinder din.
charge bonnets and remove the explosive charged.
2. Switch on the neutral gas circuit breaker on the right-seat pilot et=
breaker panel and turn on the B-45 switch on the pilots' upper electric patud.1111
this case the extension pilot lamp whose one conductor is connected to the het [
and the other conductor is connected to the mid contact of the firing sechanhe
nest be n.
3. The check up over, turn off the neutral gas switch, type R-45, on the
pilots' elver electric board.
4. Charge the firing mechanisms with explosive charges, screw them to tie
neutral gas cylinder discharge bonnets and look the firing mechanism nuts ama '
mire.
1
Fire-Fighting System
1. On the right-seat pilot circuit breaker panel switch on two starboard
and port engines tanks fire warning circuit breakers, type 13C-15.
CAUTION: The CO2 cylinder opening circuit breaker, type A3C-10 I must be.
in the OFF. ( RUEBREHO ) position.
2. On the pilots' upper electric board turn on the fire-fighting system
switch; the lamp must be dead.
1
electric board. Each push-button type lamp must flash up and the cook of the
3. Pries in turn all the push-button type lamps mounted on the pilots' 1014,
fuel shut-off cocks unit corresponding to this lamp most operate. SimultenemslI 1
the 002 cylinder opening relay, type P11-2, must operate: Prior to each pre88l4
of the push-button type lamp, turn off the fire-fighting system switch and JAI
or 2 seconds turn it on again; the push-button type lamps must go out. On P000214
the push-button type lamp of the starboard and port engine besides flaehingtf
the push-button type lamp and operation of the P11-2 relgy, the electronagtetin
air valve of the shutter in the system designed for scavenging the space unda
the engine cowl with air must operate.
4. On the left-seat pilot circuit breaker panel switch. on the air contra
circuit breaker and press one of the push-button type lamps. On throttling don
one of the engines the blow-off band suet operate for closing and the shutterd
the undercowl.reaess air scavenging system must get closed.
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Fuel Shut-Off and Engine Fuel Cross-Feeds
1. On the right-seat pilot circuit breaker panel switch on two fuel shut-nee
cocteirmIt breakers, type A30-5, two fuel pump operation warning system circuit
tnikers, type 1311-2 , and the fuel cross-feed conk circuit breaker, type A3C-5.
2.0n the pilots' upper electric board set the fire shut-off cock switches
and the fuel cross-feed control switch to the OPEN ( ?TONTO ) position; In this
we the fuel shut-off cocks green warning lamps (OPEN) on the pilots' upper
electric panel must flash up. With the cock switches in the CLOSED (3AKPUT0)
position the warning lamps must go out.
The system operation should be checked during the engine operation by the
lummift technician or mechanic, who opens the fuel shut-off cocks prior to
mines starting and closes them after switching them Off.
Feasible Faults of Fire-Fighting System and
Their Elimination
Fault
Cause
Remedy
1
2
3
One of the push-button
(a)-Closing of contacts
Ghent the entire group
tria lamps is on when
In the fire-sensitive
of the push-button type
there is no fire
unit
lamp fire-sensitive unit.
On detecting a fault
inside the fire-emulative
unit replace the latter .
(b) Closing of contacts
Replace the push-button
inside the push-button
type lamp
type lamp
The push-button type
The change-over system
Replace the fuel Shut?
lump continues glowing
In the fuel shut-off
off cocks units.
aterpressing the '
cocks unit is out of
WI: The defects elini
hap when turning the
repair
mated, check the
tire-fighting system
firing mechanisms
snitch on and off
and if they have
operated, replace
the explosive
charges
Ime warning lamp is
(a) The lamp is burnt
Replace the burnt out
dead with the fuel
out
lamp
shut-off cocks open
(b) The adjustment of
Replace the fuel shut-
the limit switches In
the 1131-2 electric
mechanism is disturbed
off conk
(c) The pump operation
Switch on the circuit
warring system circuit
breaker.
breaker, type 1311-2 ,
is not switched on on.
Note: To determine the
fault of the foe)
the right-seat pilot
shut-off cocks
circuit breaker panel
warning system,
7-1.-01-lemees
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1
3
The fuel shut-off
cocks or the cross-feed
cock fall to open or to
close
The fuel shut-off
cocks or the cross-
feed cock let the fuel
through in the closed
position
(a) The fuel Shut-off
cocks circuit breaker,
type 130-5 or the cross?
feed cock circuit breaker,
type 130-5 on the right-
seat pilot circuit breaker
panel is OFF
(b) The connecting wires
are broken, or there is no
contact in the plug connec-
tor
(c) The electric mecha-
nism, type 1(311-2 is
out of repair
The adjupsent of the
limit switches of the
11311-2 electric mechanism
is disturbed
disconnect the Vac
connector from the
11811-2 electric
niam and connect
terminal 4 to the air.
craft frame by sewn
of a conductor; if
the warning lamp is
serviceable and the
wiring is in good
repair, the warning
lamp must be on
Switch on the circuit
breaker
Detect the places of
damaged wiring by ringing
out the wires and elhn.L.t
the damage. This dome,
check the operation
Replace the 1(3E-2
electric mechanism with
the cock
Replace the 0011-2
electric mechanism with
the cock
FUEL PURE'S CONTROL ARD FUEL GAUGE STWTRIC SISTER
The system is designed for:
(a) checking the total quantity of fuel In five tank groups;
(b) checking the quantity of fuel in each group of tanks separate/Y;
(o) automatic control of the fuel consumption sequence so an to preserve
the aircraft centring within permissible limits during flight and landing;
(d) manual control of the fuel pumps when the automatic control of fuel
consumptiqn fails;
(e) warning of fuel available for 30-minute and 15-minute flight;
(f) aircraft fuel filling control;
(g) measuring the amount of fuel consumed by each engine;
(h) signalling of the fuel pumps operation of each group of tanks :ferrate
-
LY.
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". 153 -
DB system includes the following instruments and units:
1, capacitive fuel level gauge, type CBTC-600.
Sis fuel level gauge set includes:
indicators 2 pieces.
weasuring device amplifiers, type YT-801( 2 pieces
automatic units, type 00-52-12 2 pieces
switches, type fl-7 2 pieces
main transmitters (with signalling) 10 pieces
additional transmitters (with signalling) 5 pieces
additional transmitters (without signalling) 2 pieces
2. Summing fuelmater, type FT0-16 (two sets): The fuelmeter set includes:
transmitter 1 piece
indicator 1 piece
twatron interrupter, type UT-51T 1 piece
5, Feel pump, type ang-T . with the MB-650A electric
'tor 12 pieces
a. Contactor, type 1-50 12
ri::::
5. Relay, type P0-2 10
6. Relay. type P11-61 piece
7. Earning lamps, type 04-51
with a red light filter
with a blue light filter
with a green light filter
6. Fuel-pressure warning unit, type
9. Resistor, type 110-10-5
4 pieces
4 pieces
12 pieces
CA-3 TY 12 pieces
8 pieces
Arrangement of Electric Mite Included
in the System
1. The fuel pumps operation warning lamps and control switches are located
cathe fuel supply electric board (Fig.99).
- 2. The arrangement of fuel pumps, fuses, switching contactors, 10040-5
nsietors and P11-3 relay designed for changing the fuel pumps over to the forced
smaitiona is indicated in Table 27.
3. The fuel gauge amplifiers (Fig.100) are located on the navigator-radar
elmmtor's Tack on the starboard side.
4. The automatic units (Fig.101) are mounted in the
sill on frame RO.22.
5. The 30- and 15-minute flight fuel remain warning lamps are located on the
left-seat pilot instrument panel and on the fuel supply panel.
8. The location of the fuel level gauge transmitters is 'Shown in Table 28.
7. lbs fuel-level gauge switches and indicators are installed on the right-
mat pilot instrument panel (Fig.102).
8. The fuel gauge A.C. supply fuses, type CA-1 , are mounted on the
mm4stor's upper electric board.
lemiei gear rose leg
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Table 27
Arrangement of Fuel Pumps and Knits and Their Switching by Groupe
Name ofP1aoe
tank
group
' of installation
Fuel
inge
tank
Pump supply fuse
Pump switching contactor
HO-10-5 resistor
Pump P11-2 function relay
1
2
34
5
6
left
at group
right
No.2 -
2 pieces
No.5 -
2 pieces
1111-75 fume in the
additional pump junction
box of tank No.2 (on
frame No.35)
1111-75 fuse in the
fuel pumps left junction
box (on frame No.33)
Two 1111-75 fumes in
the fuel pumps junction
box (on frame No.49)
K-501 contactor in the
additional pump junction
box of tank No.2
.K-50A contactor in the
fuel pumps left junction
box
Two 11-50A contactors in
the fuel pumps junction
box
,
In the additional pump
junction box of tank
No.2
In the fuel pumps left
junction box
Two resistors in the
fuel pumps junction. box
In the additional pump
junction box of tank
No.2
In the fuel pumps left
junction box
Two resistors in the
Mel pumpe junction box
left
2nd group
right
No.4 -
1 pieoe
No.3 -
1 piece
211-75 fuse in the
fuel pumps junction
box (on frame No.49)
1111-75 fuse in the
fuel pumps etarboard
Sunotion box (on frame
a...cc,
11-50Aoontactor in the
fuel pumps junction box
I-50Aoontactor in the
fuel pumps right junction
box
In the fuel pumps
junction box
In the fuel pumps
right junction box
In the fuel pumps junc-
tion box
In the fuel pumps right
junction box
1
V 2
5
a
5
6
. grd '
group
No.10 -
1 piece
RH-50 fuse in the
distribution panel of the
port and starboard engines
One E-50A oontactor in '
the Landing gear junction
box and one contactor in
the fuel pumps relay
junction box (in the port
and starboard landing gear
wells)
In the landireaaar
junction box and the
fuel pumps relay
Mn tba leading gear
junction box and the .
fuel pumps relay
4th .
group
No.16 -
1 piece
NA-50 fume in the
distribution panel of
the left and starboard
engines
.
One 11-50A 'contactor in
the landing gear junction
box and one contactor in
the fuel pump relay junc-
tion box (in the port and
starboard landing gear
wells)
Inthe landing gear
junction box end the fuel
pumps relay
left
.
5th group
right
No.6 -
1 piece
No.6 -
1 piece
11ilr50 fuse in the
fuel pumps left junc-
tion box (on frame
No.33)
1111-50 fuse in the
fuel pumps right juno-
tion box (on frame No.33)
KA-50A contactor in the
fuel pumps left junction
box
11-50A contactor in the
fuel pumps right junction
box
,
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. 156 -
'Table
Arrangement of Feel Level Gauge Transmitters
Name of tank group
let group
232,1 group
3rd group
4th group
5th group
left
right
left
right
Trensmftters
28
main transmitters
in tank No.
2
5
3
10
16
6
additional txanardttezi
in tank No,
3
12
6
9. The fuel level gauge D.C. supply circuit breaker, type A30-2
located on the right-seat pilot circuit breaker panel.
10. The automatic unit A.C. supply fuses. type CH-1 are mounted on lb
navigator's upper electric board.
11. The automatic unit D.C. supply circuit breaker, type 130-5 , is hmnh
on .the right-seat pilot circuit breaker panel.
12. Right- and left-hand circuit breakers, type A3C-2, of the 4th end5A
fuel pump groups are located on the right-seat pilot circuit breaker panel.
13. The fuel consumption control switch AUTOMAT-MANUAL (ABTOIAT-PRIMIN
relay is in the fuel-level gauge junction box on frame No.22.
124 The fuel-pressure warning units, type Car317 , cutting in warning het
are located near each fuel pump; they are connected to the fuel line.
15. The P20-16 fuel-level gauge indicators are mounted on the pilots' maid
instrument panel.
16. The P20-16 fuel-level gauge transmitters are mounted in the fuel lime
the engine lower part.
17. The thyratron interrupters, type 1lT-51, are located on the navigato*
radar operator right-band rack (Fig.103).
Checking Operation of ftelmeter System on Aircraft
Prior to checking the fuel pumps automatic control and fuelmeter sista
make certain that on the navigator's upper electric board CU-1 Pulses are mmte
in the fuel-level gauge A.C. supply circuits.
1. The aircraft mains must be supplied with 28 or 28.5 V D.C. from the
Apound power supply and with 115 V A.G. from the aircraft operating or starbl
inverter, type 110k-4500.
2. Switch on two FUEL-LEVER GAUGES (TOH)BBOUEN) circuit breakers, tYpe
130-2 , on the right-seat pilot circuit breaker panel.
eta
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Neat in two port and starboard engine fuel-level gauges supply ;witches,
SZ-250, on the right-seat pilot's instrument panel.
A. Bet the handle of the port fuel-level gauge switch to the I position.
Ioodom, in two or three minutes the pointer of the port fuel-level gauge
!Damao:twist indicate the amount of fuel filled in the let tank group with
poletdble error of =320 lit, by the lower scale marks. Press the button on the
oaceor case; the indicator pointer must stop at the zero scale mark, the
Amitmible error being =160 lit.; then release the button; in this case the
phter =St indicate the amoUnt of fuel filled in the 1st tank group.
After the fuel-level gauge bate been checked with the switch in the 1 posi-
04 perform checking with the switch In the 2 , 3 , 4 and 5 positions. The
beriont pointer must be in the same position as with the switch in the 1
ppition.
5. set the handle of the port fuel-level gauge switch to the TOTAL graL)
pinion; the instrument pointer must indicate the total amount of fuel filled
hdl the five tank groups with permissible error of =960 lit, by the upper
tele larks.
Emma the button on the indicator case; the indicator pointer must stop at
tee stale zero mark; then release the button; the pointer must read the total
moot of fuel in all the five tank groups.
A. Bet the handle of the port fuel-level gauge switch to the 1 position
midlumse the GROUP CHECKING (DPOBEPHA PPM ) button on the port group fuel-
leml gauge amplifier. The pointer of the fuel-level gauge left indicator must
Wheats the amount of fuel in the group (6000 lit.) with permissible error of
:320 lit. After the fuel-level gauge has been checked with the switch in the 1
*Men, perform checking with the switch in the 2 , 3 ? 4 and 5 posi-
dmm; the instrument pointer must be in the same position as with the switch in
the 1 position.
7. Set the port fuel. level gauge switch to the TOTAL (CYKKA.) position
'dimes the TOTAliCHECK (UPOBEPKA GYMMU ) button on the port group fuel-level
gattgo amplifier. The pointer of the fuel-level gauge port indicator must
Indicate 16,000 lit, with permissible Jeviation of =320 lit.
liztla 1. The starboard group fuel-level gauges are to be checked in the
same manner as the port group fuel-level gauges.
2. After the fuel-level gauges have been checked for proper opera-
tion, turn off the switches of the fuel-level gauges starboard
. and port groups on the right-seat pilot's instrument panel and
the 1:10-4500 inverter.
Checking Operation of Fuelmeter. Type P20-16
1. Prior to checking,set the indicating instrument pointer precisely to
the mammt of fuel filled in the fuel tanks starboard and port groups.
? 2. Prior to checking, make sure that the F1'C-16 fuelmeter A.C. supply fuses,
tgPe C11-1A , are mounted on the navigator-radar operator's electric board.
3. Make certain that the P20-16 fuelmeter D.C. supply circuit breakers,
k30-2 , on the right-seat pilot's circuit breaker panel are on.
4. Cheek the operation of the system with the engines switched on. The
hmummy of operation is determined by the fuel consumption per hour.
? 5. After the P20-16 fuelmeter has been checked for proper operation, switch
atlas 130-2 circuit breaker on the right-seat pilot's circuit breaker panel.
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grOtl" 41^'Tn.t1A71 Qt 1"1 I"" an-ikvtoTntic
Control System ard Their Camino: Svstam
I. Switch on the fuel pumps warning system circuit breaker, type A3C-2,
on the right-seat pilot's circuit breaker panel; the green and blue warning Imip,
must be dead.
2. Set the fuel flow control switch to the MANUAL ( MIME ) position andel
the pilots' upper electric board turn on the switch of the 1st front tank group
fuel flow. manual control 13C-5 circuit breaker. The green lamps of the lot
front tank group on the pilots' upper electric board must flash on.
3. Switch off the 1st front tank group fuel flow manual control A3G-5
circuit breaker (operating as a switch); the warning lamps must go out.
- 4. Switch on the 1st rear tank group 130-5 circuit breaker; the lamps of Us
let rear tank group must flash an. Then turn off the let rear tank group 130-5
circuit breaker switch; the warning lamps must go out.
5. Turn on the let rear and front tank group 130-5 circuit breakers; the
warning lamps must glow constantly without flickering.
6. Switch on the 2nd group 130-5 circuit breaker; the warning lanp ate)
2nd group on the upper electric board must flash on, while the let group pumps
suet change over from the nominal to the heavy duty.
P. Switch off the let front and rear tank group 130-5 circuit breaker; in
this case the let group warning lamps must go out.
8. Switch on the 3rd group 130-5 circuit breaker; the 3rd group warning
lamps must flash on, while the 2nd group pumps must change over from the nominal
to the heavy duty.
9. Switch on the 130-2 circuit breaker of the ath and 5th group fuel eske
supply on the right-seat pilot's circuit breaker panel.
10. Turn off the switch of the 2nd group 13C-5 circuit breaker; the 2nd
group pumps and warning lamps must get switched off.
11. Turn on the stand-by pumps switch, type 21-45, on the pilots upper
electric board; the 4th group pump warning lamps must flash on.
12. Turn on the 5th group switch, type 21-45; the 5th group warning lamps
swat flash on, the 3rd group pumps must Change ever from the nnsinsl to the
heavy duty and the 4th group pumps must change over from the stand-by to the
nominal condition.
13. Then checking the operation of the fuel pumps, pay attention to the
amount of current consumed by them which must be within the data given in the
Certificate of the fuel pump.
14. Set the fuel flow control switch on the pilots' upper electric board
to the ADTOZSTIC (ABTOCAT ) position and the fuel flow manual control 130-5
circuit breaker switch to the ma (BUKEVIEHO ) position.
15. Switch on the fuel automatic line 13C-2 circuit breaker on the right-
seat pilot circuit breaker panel.
16. On the pilots' upper electric board tura on the supply switch, type
21111-250 . of the starboard and port engines automatic control line amplifier;
nabs Bare that the fuel pumps automatic control operates correctly with the
given amount of fuel in the aircraft tanks. The flashing of the blue and green
warning lamps and the switching on of the fuel pumps depend an the amount of
fuel in each group separately. The operation of the automatic control versus
the amount of fuel filled is described in the Section "Operation of 111-311
Engines". Book ons? "Operating Instructions of TY-16 Aircraft".
I?. The fuel flow automatic control is performed through two channels
independent of each other. The left tank group fuel flow automatic control is
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dcalketed by the right tank group automatic control and vice vexes. Therefore,
then checking the operation of the automatic control, check the operation of
owbgmAlp separately, that is by switching the 2113-250 automatic control
witture on and off in turn. The flashing of the warming lamps on the pilots'
upper electric beard when the amplifiers are switched, on and off in turn
tce011es to serviceability of the amplifiers.
Notes: 1. After the operation of the automatic control has been checked up,
switch off the 1104500 inverter, if other units operati,T from
the A.G. power supply are inoperative.
2. Set all the switches of the automatic control system on the
pilots' upper electric board to the OFF ( BOAIREHO ) position.
3. In case faults of the automatic control are detected during the
check up, it is necessary to Cheek the system by means of the
YCA-53-60 installation as prescribed by special instructions
appended to the installation.
?p2cifications_of Fuel_Pumse_Control_and_Fuel
GaEgea Electric System
Fuel-Level Gauge
1. The fuel-level gauge set operates:
(a) within ambient air temperature range of -60 to +50?C;
(b) at A.C. voltage of 115 t 11.5 V, 400 - 28 c4.s. and D.C. voltage
of 27 I 2.7 V;
(c) with outside pressure changed from 760 to 90 mm of mercury, that is
at altitudes from 0 to 15,000 m.;
(d) in conditions of relative humidity from 30 to 98 per cent.
2. The error of the fuel-level gauge reading when bench tested under normal
conditions (at temperature of 20 t 5?C, pressure of 760 mm of mercury, relative
humidity of 30 to 98 per cent and voltage of 115 V, 400 c.p.s.) does not emceed
12 per cent at the zero mark and 14 per cent of the scale nominal value at the
other scale marks.
3. The error of the fuel-level gauge reading at -6000 does not exceed 16 per
cant at the zero mark and IS per cent at the other scale narks; at temperature
of +5000 the error does net exceed i3.5 per cent at the zero mark and t5.5 per cent
of the scale nominal value at the other scale marks.
4. The error of the signal unit operation checked by means of the bench
thes not exceed 10 mm of the float travel in the transmitter.
5. The additional error of the fuel level gauge reading at voltage Change
elle per cent does not exceed Il per cent; at frequency change of 1 per cent
it does not exceed Il per cent of the scale nominal value.
6. The insulance of transmitters and switches, type fl-7 , at normal tempera-
bse and relative humidity of 30 to 80 per cent is not less than 100 megohms and
at relative humidity of 95 to 98 per cent - not less than 20 megohms.
7. The insolence of the indicating instrument is net less than 20 megohms
at normal temperature and relative humidity of 30 to 80 per cent and not less
that 2 megohms at humidity from 95 to 98 per cent.
8. The similar elements of the set within one group are interchangeable.
9. The additional error is ti per cent of the fuel-level gauge scale nominal
/she (taking into consideration possible difference in the capacity of the: tanks
included in the groups)...
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10. The first earning signal 30-UTE FLIGHT-FUEL REMAINDER (00TILT0R
HA 30 MIH.HOXESA) is sent when the fuel left in one of the 4th tank groups
is equal to 60012 lit.
11. Tho second warning signal 15-LINUTE FLIGHT -FUEL REMAINDER (00TATOETO
HA 15 HIAH. nurno is sent when the fuel left in one or the 5th tank groups le
equal to 1600 =100 lit.
12. The electrical capacity of "dry" transmitters is given in Table 29.
Table 2y
Capacity of Dry Transmitters (Initial Capacity)
Nos ?
No. of tank
and
transmitter
Capacity vf.
transmitters,
PP
Nos
No. of tank
and
transmitter
Capacity of
transmitters,
PP
1
2
3300
7
6X
1000
2
3 ?
2300
8
7
1000
3.
Is
1000
9
10
2300 .
4
4
3300
10
12
1000
5
5
3300
11
16
2300
6
6
2300
PTQ-16_Fuelmeter_of Fuel_Consumed by_Each Engine
? 1. The summing fuelmeter, type PTC-16, operates within a range of 1200 to
16,000 lit, per hour.
2, The error of the fuelmeter set under normal conditions does not exceed
12.5 per cent.
.3. The fuelmeter set at temperatures of +50 and -60% does not exceed 4.5
per cent of the indicating instrument scale nominal value.
4, Pressure drop by the transmitter at fuel viscosity of 15 + 1 c.s.
(corresponding to fuel temperature of -40?C) and maximum fuel flow of 16,000 lit.
per. hour does not exceed 0.25 kg per sq.cm. with the impeller operating and
0.4 kg per sq.cm. with the impeller inoperative.
5. The inner chamber of the transmitter body, as well as the connections of
the branch pipe with the transmitter body, are gastight and withstand a testing
pressure of fluid (kerosene) of 9 ig per sq.cm,
6. Power consumed by the set is 40 W.
7. The thyratron fires with delay of 100 or 200 millfseconde.
Fuel_Pempe_Automatic _ and _____ Control
1. The pumps-switched-on signals of the subsequent groups - the transmittal
lower warming unit operate when 350 .= 150 lit, remain in one of the tank groups
of the same name.
2. The pumps-switched-off signals of the previous groups - the transmitters
apper warning unit - most operate when the following amount of fuel remain)] is
one of the tank groups of the same name:
2nd left group, tank No.4
2nd right group, tank No. 3
3rd group
4th group
2450 = 250 lit.
2250 = 250 lit.
5000 = 250 lit.
2300 = 250 lit.
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3. The reel pn.10, t7To 80-T , with the NB-650A electric motor:
(a) Electric motor power supply 27 = 2.7 V 'LC
(h) Current consumed by the electric motor:
maim duty
light duty
. (c) Fluid pressure drop produced by the unit at
iertem and voltage of 27 V-across the electric motor
main duty
light duty
(d) Pressure produced by the pump at light duty
deb the cock closed and voltage of 27 V across the
ihetric Bator terminals
not in excess of 31 A
not in excess of 19 A
the output of 14,000 lit.
terminals on the ground :
0.8 to 0.9 kg per sq.cm.
0.25 to 0.45 kg per sq.cm.
not on exceed of 0.8 kg
. per eq.om.
Note: Check these parameters at ambient air and pressure fluid tempera-
ture of 15 to 35?C.
(e) Period of continuous operation prolonged
(f) Permissible temperature of ambient air
berg the operation of the unit
(g) -Minimum permissible length of the electric
leer brushes
from +50 to -60?C
18 asa
lot,: 'Abe pump heavy duty continuous operation during 60 minutes
(15 minutes of them at zero output) is performed by connecting
a 5-ohm resistor to the main duty winding circuit.
CAUTION: Change-over to heavy duty can be performed only from
the main duty. It is not permitted to start the pump at heavy
duty.
4. The fuel pressure warning unit:
(a) At pressure Change from 0.35 ? 0.05 to 2 kg per sq.cm. the warning
hap flashes on.
(b) The device operates within the range of +50 to -60?C.
(c) The device warning lamp power is 3 W, its supply voltage is
27= 2.7 V.
(d) Errors of the warning unit operation:
at normal temperature =0.05 kg per sq.cm.
at temperature of +50 and -45Pb =0.075 kg per sq.cm.
(e) The gastightness of the device most meet the following requirements:
(1) at air pressure of 5 kg per sq.cm. no pressure drop as indicated
by the reference pressure gauge must take place in the warning
unit sensing element;
(2) the gastightness of the device body ensures that on delivering
air under a pressure of 300 mm of mercury simultaneously to
the static and dynamic systems pressure drop does not exceed 8 mm
of mercury during One minute.
Cr) The device can withstand pressure overload of 5 kg per sq.cm. during
Saint/tee.
(g) The device insulancr. at normal temperature and relative humidity of
30 to co per cent is not less than 20 megohms.
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Possible Faults of Fuel Nays Control Electric
System and Their Elimination
Fault
Cause
Remedy
2
The indicator indicator pointer is
(a) The transmitter
Detect the faulty
pressed to the left-hand
limiter
circuit is open
transmitter by sMt
on the transmittcm
group of transmitte
turn and check itec
tion In the plug col
Eliminate the fat*
(b) The outer connect-
Check the conant
ing wires of the trans-
wires and eliminate
mitter circuit are
broken
fault
(c) Break inside the
amplifier circuit running
to the transmitters
Replace the amplif
(d) There is no contact
on the P11-3 relay located
in the amplifier
Replace the amplif
(e) Short circuit of the
Replace the amphf
65-9 lamp grid wire to
frame in the amplifier
? (f) Shorting to frame of
Check the connecth
the circuit connecting
line and eliminates!
pin 11 of the switch
plug connector to pin 9
of the amplifier plug
connector to earth
circuit
(g) There is no contact
Check the contacts
between the transmitter
plates and the plug
connector pins
replace the tresseitt
The indicator pointer
(a) Shorting between
Detect the faultyt
s beyond the scale
axiom
the transmitter plates
sitter or the gni*,
transmitters by mdtc
them on in turn. BeM
the faulty transcd*
fran the tank. Mak
-
insulance between*
plates and betweenes
plate and frame. If.
insulance is lass*
100 megohms, wash*
faulty transmitter
clean fuel and *lit'
?
Then re-check the
to
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163 -
1
2
3
Van eeitched over to
fCCAL (CFNMA )
do, the indicator
?dater overshoots the
sale maximum
The indicator pointer
fails to move
Poor sensitivity
GD the indicating
instrument
-----QVC1174"011-
(b) Break of the self-
balancing bridge const-
ant capacity arm
(a) Break in the D.C.
supply circuit. The
A3C-2 circuit breaker
on the right-seat pilot
circuit breaker panel is
not switched on
(b) Tee relay, type
PH-6, fails to operate
(a) Break in the A.C.
supply circuit
(b) The amplifier lamps,
types 65-9 and 65-9, are
.71,t of repair
(a) The insulance
between the transmitter
plates and frame is
less than 100 megohms
insulance. If the
inaulance exceeds
100 megohms, mount the
transmitter in place.
In case the insulance is
less than 10 raegohas
replace the transmitter
Replace the amplifier
Switch on the 113C-2
circuit breaker on the
right-seat pilot's cir-
cuit breaker panel or
eliminate the break of
the outside connection
circuit wires
Replace the amplifier
Check the CH-1 fuse
on the navigator's upper
panel
Replace the lamps and
check the serviceability
of the set by pressing
the CHECK CP (HPOBEPEA)
'buttons on the amplifier
front panel. These buttons
pressed, the indicating
instrument pointer must
MVO towards the scale
maximum. The button re-
leased, the pointer must
return to the initial
position
Detect the faulty
transmitter or the group
of transmitters by
switching them on in turn
and remove them from the
tank. Check the trans-
mitter insulance between
the plates and frame: If
the, insulance is less than
100 megohms, dry the
transmitter and re-check
the insulance. If the .
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2
3
With the tanks empty,
the indicating instrument
pointer shoes the presence
of fuel in the tanks
The warning lamp fails
to flash on
The indicating instru..
Rent fails to operate
with fuel being consumed
and the transmitter
operating
(b) Lass of emiseion
of 68-9 and 68-8 lamps
(0) The turns of the
output transformer
primary winding are
Shorted
The insulance of the
line or transmitters in
too low
(a) The warning lamp
burnt out
(b) The D.C. supply
line is damaged
(o) Vas A.C. supply
line is damaged
(d) The relay. type
117-6 . is out of repair
(o) The special sensi-
tive relay is out of re-
pair
Ftelmeter, Type T1C-16
(a) fuses are not
fitted on the navigator..
radar operator electric
board
" (b) The A3C-2 circuit
breaker on the right-seat
pilot circuit breaker
panel is not switched on
(c) Open circuit
(d) Jamming of the
mechanism ir the indicat-
ing instrument
insolence is still 1m4
replace the trammitte
Cheek the fuel for
presence of moisture
Replace the lamps
Replace the amplifisz
The means of elisine.
tion of tlyt fault ieth
same as in the event d
poor sensitivity &Alm
indicating instrument,
Item A
Replace the lamp
Repair the line
Repair the line
Check the treoasmitter
coil by means of atedu
Replace the transmitter
Replace the automatic
unit
Check the presence of
the ell-1 fUS8 on the
navigator-radar operator
electric board
Switch on the A3C-2
circuit breaker on the
right-seat pilot circuit
breaker panel
Using a tester cheat
the connecting wires ad,
eliminate the fault
Replace the indicatied
instrument
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- 165 -
ve transmitter fails to
00sae with fuel being
=mod
Th indicating instrn-
watpointer fails to
rotate when the setting
Med 18 rotating
Great positive error of
the ,et (that is the indicat -
alamount of remaining fuel
Ii greater than the actual
'
mount)
Great negative error of
the oat (that is the indi-
cated amount of fuel re-
mising is smaller than
the actual amount)
The transmitter and the
Indicating instrument are
In good repair but the set
fails to operate with fuel
being consumed
(0) The thyratron In
the HT-51L thyratron
interrupter fails to
operate
(a) Clogging of the
transmitter bearings in
the guide mechanism -
the impeller fails to
rotate
(b) Clogging of the
contact mechanism - the
interrupter fails to
rotate when the impeller
La rotating
The stop spring is
damaged or deformed, the
stop ball drops out
(a) Clogging of the
transmitter
(b) Clogging in the
indicating instrument
mechanism
(a) Auer contact in
the connecting wires
(most often at the plug
connectors and lead-
ins)
(b) The indicating
instrument kinematic-
coupling is disturbed,
when the driving pawl
engages two teeth
during one cycle of
the relay operation
Failure of the
thyratron interrupter
elements
Replace the thyratron
Replace the transmitter
Replace the transmitter
Replace the transmitter
Replace the transmitter
Replace the indicating
instrument
Thoroughly check the
wiring and ensure reliable
contact
Replace the Judi ating
instrument-
Replace the thyratron
and, if after this the
set does not operate,
replace the thyratron
interrupter
PLAT CONTROL 2IECTR/C SYSTEM
The flap control electric system is designed for extending and retracting
the flaps, indicating the angle of their deflection and transmitting the horn
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signals in the front pressurised cabin when the engine throttle control is en
to the take-off rating while the flape are not in the take-off positiore.
The oysters includes the following electric unit's:
- electric mechanism, type 1613-3111;
- distant-reading electric flap position indicator. type 7811-47 ; the
instrument set includes one 7811-47 indicator transmitter and two 7811-47
indicators;
- limit switches mechanism, type 11EB-11;
- limit switches mechanism, type 1gB-2;
- relay, type P11-2;
- two contactors, type 1,250:
- switch, type. 21111-20;
- switch, type 3111111-45;
- fuses and circuit breakers.
The electrical units are located in the following places:
1. The flap control electric mechanism, type 11II3-3M , is mounted on the
centre plane between frames Nos 32 and 33.
2, The flap position indicators, type Y311-4?, are located on the right-gee:
and left-seat pilots, instrument panels.
3. The transmitter of the 7811-47 position indicators is installed on the
NKB-2 mechanism.
4. The limit switch mechanism, type 6E6-2, for switching on the warning
horn is located on the flap transmission shaft.
5. The limit switch mechanism, type 6KB-11, for switching off the electric
motors of the 1813-311 mechanism with the flaps in the extreme positions is on
the flap driving shaft.
6. The contactor. type 1-250, for switching on and off the supply of electric
motors Nos 1 and 2 of the 1013-31i mechanism and landing flaps junction boxiam
the bomb bay ceiling at frames Nos 34 and 35.
7. The fnse, type 1111-150 , for electric motor No.1 of the M113-3k socimWs
is in the double supply left-hand junction box and for electric motor No.2 lain
the right-hand junction box.
8. The flap control switches, type 314118-45 , of the left-seat pilot, and
type 2E-20 of the right-seat pilot are mounted on the angfina control panels
of the left-seat and right-seat pilots respectively.
9. The relay, type , for interlocking which prevents switching the
flags by one pilot for extension and by the other pilot for retraction is
installed on the left-seat pilot's engine control panel.
10. The limit switches for switching on sound signalling are mounted ones '
engine throttle controls on the right-seat pilot's console (Fig.104).
becifications of Electric Obits
1, Electric mechanism, type 1(113-31:
(a) mains nominal voltage
(b) range of mains operating voltage'
(c) loading moment:
nominal
maximum
(d) current with the mechanism operating with two
electric motors:
at nominal moment
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- 167 ????
et maximum moment
$00 with the mechanism operating with one electric
oter:
1 &twain?. moment
et maidaraw moment
(a) speed of rotation of the mechanism output Shaft at
animal voltage and nominal loading moment:
with the mechanism operating with two electric
setae
den the mechanism operating with one electric
motor
()epeed of rotation of the output shaft in!both direc-
tme of rotation at nominal voltage, simultaneous
operation of two electric motors and a moment of
214.m on the output shaft
(;)friction clutch slipping torque reduced to the
mechanism output abaft
item determining the direction of rotation of the
secheidem output shift from the side of the angle
trmeemission larger diameter the rotation to the
left corresponds to the flap extension and the
rotation to the right corresponds to the flap retrac-
tWn
(I)leclmnism operation duty
with two electric motors operating after the
extension or retraction of the flaps
with one electric Rotor operating after the
extension or retraction of the flaps.
i .
(1) the electric mechanism operates normally at ambient
27 V
air humidity of up to 58 per cent at temperature
24.3 to 29d'
ii' change from +50 to -6000 and at above-sea-level
i
/ altitudes of up to 5000 m.
.;
10 kgral I 2. 44 position inAleator, type 7311-47:
15441
(Odd= voltage
1.{. 04 the indicator operates at temperatures
not in gemese
Of 190 1 !
i
1
I
i
Wpm= consumed by the set
250 A
100 A
125 ?
not less than
240 r.p.m.
120 r.p.m.
not in excess
of 420 p.p.m.
18 to 25 kg-m
intermittent
5-minute inter-
val; complete
cooling of the
engines is neces-
sary atter
5 cycles
10-minute inter-
val; complete
cooling of the
engines is
necessary niter
2 cycles
27 + 2.7!
from +50 to -6000
not In excess
of 5 W
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(d) current communed by the transmitter
(s) set indications error
N. Malt suited mechanism, type EKB-2:
(a) nominal voltage
(ie) snactrain current at ohmic load
(c) meiranza load at inductive load
4. Unit switch mechanise, type WEBllt
(a) operating voltage range
(b) ilitEtana current at ohmic load
(o) maximum current at inductive load
9. Inn angle of flap extennien
6. The sound signnliing is switched off with the engine
thnettle control in the take-off patina position idiom
by the 731,47 flap position indicator the Minna=
extended by an angle
7. Blip extension time with both electric motors operating
simulteneensl,y and at current not in excess of 155 A
and voltage of 26
8. Blap retraction time with bah electric motors
operating simultaneously and at current not in excess
Of 160 A and voltage of 26 V
9. nep extension time with one engine operating at
current not in excess of 80 A and voltage of 26 V not in amen
of 50.see.
not in
of 100 n4.
not in oz,e4g
of 11? 1
24 V
15 A
A
23.4 to 307
15 A
8A
35 ? 1?
from 19 -
to 23..
not in etas
of 25 sin.
not In einem
of 25 sec.
10. nap retraction time with one engine operating at
current not in excess of 85 A and voltage of 26 V not in MENU
of 50 sec.
Ohackimmt.Flata Operation under Voltage
I. On the lett-seat pilot's circuit breaker panel switch on two LOME
mpg (ganonews BSTKL7 ) circuit preakers, type A3C-5 . and LIR= purse
AIR TB!! ATORB INDACATOR8 MURATORI 1100121091117 ZINGS id =PAVEL B8010
circuit breaker, type D0-2 I on the right-scat pilot's circuit breakoirpmml
switch on the BORN ( CUM ) circuit breaker, type A3C-2.
CSBTIctit 1. Prior to switching on the circuit breakers on the right ead
left pilote" tenable, check the position of the flsp,00ntrel esitchn
whichmnst be in the neutral position.
2. Prior to extnneing or retracting the flaps make sure thealle
nape and the flap driving gear are clear of. personnel and that the
ladders and the oases are removed.
.. 1
2. With the flaps in the retracted position, set the left-seae pilot *BP (
.to the BB:TRACERS (MIRO ) position. This done, by short pulses set the 110' ii,
I
seat pilot switch to the WITEBBION (BOYCE ) position; the flaps suet nee
extended.
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3. By =Was of the left-seat pilot switch extend and retract the flaps
agoetely (Fig.105). When extending the nape by operating the left-seat pilot
set the right-seat pilot switch to the RETRACTED (MIRO ) position by
oat pulses. The flaps must continue extending.
The flap control operation is checked iron the left-seat control switch
odiwithe right-seat control switch by means of both electric motors and by
*dielectric motor separately. Check the flap control operation from the right -
eepilot console after the flaps have been checked for complete extension and
retraction from the left-seat pilot console.
When checking the flap control from the right-seat pilot console, do not
*tend and retract the flaps completely (Fig.106).
5. When extending and retracting the flaps, set the engine throttle control
10 the take -off rating; the horn must hoot all the while. When the flaps are
deflected from 19 to 23? during extension and from 23 to 19? during retraction the
Waimea not hoot.
6. When checking the flap control operation check the operation of the flap
positlonindicators. During the extension and retraction of the flaps the pointer
of the 733-47 indicators must move without noticeable jerks and jamming. The
differnme in the flap position indicators reading of the right-seat (Fig.10n)
and left-seat pilots must not exceed '110.
TAIL SKID CONTROL AND LANDING GEAR WARNING
ELECTRIC SYSTEM
The tail skid control and landing gear warning electric system is designed
fa:
(a) sending out signals of the landing gear legs extended and retracted
pmitions separately;
(b) control of tbe tail skid extension and retraction;
(c) sending out sound signals In case the throttle control is in the off
position and the landing gear is not extended.
The system includes the following units:
I11-250 electric mechanism;
0IE1-51 warning lamp - 8 pieces (5 green lamps and 3 red lamps);
BK-44 limit switch - 6 pieces;
BK-2-14071 limit switch - 1 piece;
BK-2-14r limit switch - 2 pieces.
The electric units are located as follows:
1. The tail skid electric mechanism, type 0-250 , is mounted on frame
M465.
2. The landing gear extended and retracted positions warring lamps, type
C11I-51 are located on the pilots' central instrument panel.
3. The tail skid retracted position warming lamps, type CIEL-51 are
hmtalled in the rear cabin on the gunner-radio-operator's and rear gunner's
electric boards.
a. The blocking limit switches, type BK-2-142r designed for switching
en sound signalling in case the throttle control is in the off position with the
boding gear retracted ere mounted on the right-seat pilot engine control panel.
5. The BK-44 limit switches (Fig.108) designed for switching on the landing
narmaha legs extended position yarning lamps are located on the leedi,ig gear
shiboard and port legs struts.
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1
The 3K-44 limit switches designed for switching oaths main and nose leg
retracted position warning lamp are mounted on the landing gear legs up-locks.
The EX-2-1403 nose leg extended position limit switch is mounted on the nose 14
demo-lock. ?
6. Tho 3X-44 limit switch !carving to switch the tail skid for retraction
and extension is located on the nose leg up-lock.
7. The m11-5 fuse in the 181-250 electric mechanism supply circuit is in
the double supply left-hand junction box on frame No.17.
flpecifications of Electric Units
1. Tall aid control electric mechanism, type MU-250:
(a) eapeZy voltage 27 = 2.7 V
(b) rod load
nominal 250 kg
maximum 375 kg
(c) current ,
at nnmieel load not in excess,
of 3.4 A
at ;maximum load not in excess
of 3.8 A
180 1 ma
(d) rod travel
(a) rate of rod travel at voltage of 27 V and nominal
'load opposite to the rod travel 6.2 = 0.62 mm/soo.
(f) duty of operation at nominal data
(g) brushes A-12 measuring
(h) altitude
2. Unit switch, typo Maio:
(a) rod travel downward before the contacts are
changed over
(b) rod reserve travel downward after the contacts
Sr. changed over
intermittent,
consisting of
5 cycles follfolios.ed by an interval
of ems hour at
least
4x5x7
15,000 m.
.5 .1. 1.8 mm
(0) travel of the additional device downward atter
*blazing over
(d) full trr..-41 of the rad and the additional devicoa
button
(e) reverse travel of the rod upward after the contacts
Sr. changed over
(f) SIMMS travel of the rod upward after the contacts
are chanced over
(g) force applied to the rod to change over the
contests
not less than
1.5 mm
4 + 1.5 mm
from 10.5 to
15 mm
not in mess
of 4.5 2
not lees than
1.5 an
4 to 6 kg
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aj force applied to the rod at the beginning of compres-
sion of the additional device spring 5 to 7 kg
(i) difference between the forces not lose than
1 kg
(j) force applied to the rod at the end of compressional
the additional device spring 11 to 15 kg
(k) operating voltage 27 = 2.7 V
current flow-
ing through the
contacts 10 A
(1) the switch operates within the ambient air tempera,
tore range
(m) operating altitude
(n) range of the rod total length adjustment
Checking Operation of Tail Skid Control and
Lending Gear Warning System under Voltage
from -60 to
+50?0
from 0 to
15,000 a.
7.7 22
Checking the Operation of the
Warning System without the Land-
ing Gear Kinematics Adjustment
1. Switch on the L.G. legs position warning systom circuit breaker, type
13C.-2 , on the left-seat pilot's circuit breaker panel and the sound signalling
ciandt breaker, type A30-2 , on the right-seat pilot's circuit breaker panel.
The three green warning lamps mounted on the pilots' instrument panel must flash
MI.
2. Press the limit switches, typo BK-44, on the L.G. main legs up-locks and
the left-hand limit switch, type BK-44, on the L.G. nose leg up-lock. The three
Oman warning lamps on the pilots' central electric board must flash on.
Checking the Operation of the
Tail Skid Control and Warning
System with the Landing Gear Kinema-
tics Adjustment
1. Switch on the L.G. legs position warning system circuit breaker, typo
AW-2, on the left-seat pilot's circuit breaker panel and the seund signalling
Orcult breaker, type A3C-2 , on the right-seat circuit breaker paned; if the
lowing gear is extended the green warning lamps must flaah on.
2. As soon as the landing gear legs start rising, the three L.G. extended
Position green warning lamps must go out. The legs reaching the extreme retracted
Position, the three L.G. retracted position red warming lamps on the pilots'
mmtnd electric board must flash on. The L.G. nose leg reaching the extreme
reheated position, the tail skid control mechanism must get automatically on
homonaction.
With the tail skid completely retrocted, the electric mechanism aunt get
intonatically switched off; simultaneously the two green warning lampa of the
tal skid retracted position must flash on; one of the lamps is mounted on the
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rear gunner's electric board; the other lamp - on the gunner-radio operator's ;
electric board.
3. Pall up in turn both throttle controls on the right-seat pilot's coal% i
panel as far as they will go. In this case the born in the front pressuriud
cabin must hoot. Switch off the horn by pressing the button on the right-seat
pilot throttle controls designed for mechanical disconnection of the horn.
4. As soon as the L.G. legs start extending all the three L.G. retraeted
position red warning lamps slat go out and the tail skid electric mechanisat?
get switched on for extension; after the tail skid has been completely anew.,
the electric mechanism gets switched off and the two green lamps of the tall
skid retracted position go out.
5. Pull up both throttle controls as far as they will go. The horn must be
silent.
During the landing gear check make sure that the adjustment of the limit
switches is not disturbed. The adjustment of the lannin.- gear limit switches is
described in the Section "Landing Gear". Book one, "Service Manual of the Lim.
craft, Model TY-16",
TRIM TAB ELECTRIC CONTROL SYSTEM
The trim tab electric control system of the aircraft is used for remote
control of the aileron, elevator and rudder trim tabs, and at the same tine
as a system providing light indication of the neutral position of the aileron
and rudder trim tabs.
The system comprises the following units:
- two electric actuators, type. ME-100A-60;
- one electric actuator, type ME-100A-36;
- one electric actuator, type YT-11;
- aileron synchronization console;
- limit switches, change-over switches and circuit breakers;
- three tell-tale (warning) lights with 'white screens.
The electric units are located as follows:
1. The electric actuators, type ME-100A-60 ? of the aileron trim tabs-
between ribs 18 and 19 of the right and left wings; the actuators are accesado r
through the underwing access holes.
2. The electric actuator, type M11-100A-36, of the rudder trim tab - beams
ribs 2 and 3 of the fin; the actuator can be reached upon removal of the adjasoot
skin portion of the fin.
3. The electric actuator, type 7T-11 , of the elevator trim tab and its
DE-14118 limit switches of the up and down positions - at fuselage frame Na.69
the units are accessible upon removal of the stabilizer access hole panels.
4. The aileron trim tab control change-over switch, type 2UH-20 , amide
rudder trim tab control change-over switch, type UH-45M , - on the trim tab
control panels (stations) of the pilot and co-pilot.
5. The elevator trim tab control change-over switch, type EH-45M, - Oil
the control wheel spokes of the pilot and co-pilot (FiE.109).
6. The 8-45 switch used for emergency disconnection of the elevator trIn
tab electric control system - under the red cap on the overhead electric
board of the pilots.
7. The white Cl1-51 tell-tale lights indicating neutral position of the
aileron and rudder trim tabs - on the pilot's instrument panel (Fig.110).
e
8. The aileron trim tab synchronization console (Fig.111) carrying th till
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control
Change-over switChi type 118-45 ? white CMI.-51 light indicating
tab
the left aileron neutral position and auxillarY (blocking) contact, type E3-6, -
hetwean. fumes Nos 9 and 10, port side.
Technical Characteristics of_Electric Actuators
Ejectrid Actuator, Type ME-1001.
. Voltage requirement
1
2. =rent requirement:
at nominal rod load of 100 kg
at maximum rod load of 150 kg
5. Rod speed at 27 V voltage and nominal rod load
4. Tell-tale light glow duration with rod midposition
travel restricted by limit switch within iimm
5. Operating duty in nominal conditions
6. Brushes, mark 1-12, sizing
7. Rotor speed
8. Operating altitude
9.Vorking travel length of ME-100A-36 actuator rod
10. Working travel length of ME-100A-60 actuator rod
Electric Actuator,
1. Operating voltage range
27 ? 2.7 V
not over 1.35 A
not over 1.4A
1.65 mm/sec.
0.5 to 2 mm of
travel length
intermittent, consist-
ing of 6 cycles
followed by-obligatory
complete cool-down of
the actuator
4E5W mm
4100 = 410 r.p.m.
up to 15,000 m.
36 mm
60 um
Type 7T-11
2. Current requirement:
at nominal load of 180 kg/cm.
at sexism load of 260 kg/cm.
3. Output shaft speed
25.4 to 28.6 V
not over 2.8 A
not over 3.3
7 r.p.m. = 0.7%
Voltage Check of Trim Tab Electric Control System
1. Turn on the A3C-5 aileron, elevator and rudder trim tab control circuit
traOmms an the pilot's circuit breaker control board.
2. Prior to beginning the trim tab operation check, make sure that the
aileron and elevator covers are removed, and there are no obstacles under the
Msnraft to hinder the trim tab movement.
3. Engage the E-45 elevator trim tab electric control emergency disconnecting
mdtehonthe pilots' overhead electric control panel.
4. Hinge out the lock on the pilot's control wheel which secures the 1111-4581
slmmtor triaLtab change-over switch.
Operating the switch in pulses and engaging it for continuous operation,
4we the elevator trim tabs from one extreme position to the other. With the
1'1* tabs in motion, the elevator trim tab control handwheel will be rotating.
5. Operate the EM-45M elevator trim tab change-over switch on the scale
:Ertl* elevator trim tab control handwheel to set the trim tab neutral.
6. By operating the trim tab switch, type EH-45E , on the left-seat pilot's
414,111% *heel switch on the 7T-11 mechanism, then switch off the B,45 trim tab
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emergency cutout.. This done, the trim tab electric mechanism stops operatiagela
begins- to operate after the cutout is switched on.
7. Close the stop an the elevator tris tab switch, with a slight moveeezt
pull and push the switch; the electric mechanism must not operate.
8. By means of the Tedder trim tab switch, type UH-4511 , and the aileron
trim tab switch, type 20-20 , on the left-seat pilot trim tab control palud
(Fig.112) shift the trim tabs in both directions till they are completely.
deflected, then set the trim tabs to the neutral position. The trim tab neutral
position warming lamps must flash on.
9. Open the aileron synchronization panel cover. With the aileron trim bah
is the neutral position, the neutral position Warning lamp on the panel must
flash on. On pressing on the blocking contact, type KH-6, the warming lawn waft
go out.
10. Shift the B-45/1 switch on the trim tab synchronization panel to the
'right or to the left. This causes the L.H. wing aileron electric mechanisate
Operate, the R.H. wing aileron mechanism being inoperative.
11. The aileron trim tabs must be synchronised. For this turn on by pulses*
aileron tabs control switch on one of the pilotea consoles till the ailermatris
tab neutral position lamp flashes on-the left-seat pilot instrument panel, width
left aileron trim tab control switch on the synchronization panel till the Isle
on the synchronization panel flashes on. Synchronization is ensured if both lams
on the left-seat pilot's instrument panel and on the synchronization panel glow
simultaneously.
12. After the operation of the trim tab control from the left-seat pilot's
console has been checked, check the operation of the trim tabs control from Mm
right-seat pilot's console as prescribed in Items 4, 5, 6, 7 and g,
13. When checking the trim tabs operation, make sure that:
(a) the trim tab switches on the left-and right-seat pilots' console have
guards and that the stenciled markings are intact and not dirty (Fig.113);
(b) the elevator trim tabs are deflected upward when the elevator trim tabs
control switch is pushed forward and that they are deflected downward when the
elevator trim tabs control switch is pulled backward;
(c) the rudder trim tab is deflected to the left when the rudder trim tabs
control switch is shifted to the right and the trim tab is deflected to the right
when the control switch is shifted to the left;
(d) the right aileron trim tab is deflected downward and the left one 'Tara
when the aileron trim tab control switch is shifted to the right; the right
aileron trim tab is deflected upward and the left one downward when the aileron
trim tab control switch is shifted to the left.
CAUTION: It is prohibited to turn on the trim tab switches simultaneously
on the eonsoles and steering wheels of the right- and left-seat pilots.
14. The operation checked, set the trim tabs to the neutral position, fix
in position the trim tab switches on the steering wheels and close the synchroniza-
tion panel with the cover.
1.9
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Possible Faults of Electrical Peri of Trim
Tab Control System and Their Elimination
Fault
Cause
Remedy
The trim tab deflects in
one direction and fails to
deflect in the other direc-
tion
The neutral position lamp
is flickering
(a) Jamming of the mechan-
ism
(b) Failure of the elect-
ric motor
(a) Poor contact in the
plug connector for switch-
ing On the MePhanism
(b) Poor contact in the
mechanism warning lamp
switching on system
Replace electric
mechanism
Eliminate the defect
in the plug connector
Replace the electric
mechanism
BRAKE SYSTEM POMP CONTROL ELECTRIC SUMMER
The electric units mounted in the system regulate the pump operation thus
maintaining pressure in the brake hydraulic system within certain limits end
transmit signals at minimum permissible pressure.
The electric system includes the following main units:
- hydraulic pump 465 K with the electric motor, type A-4500K;
- presswe drop warning unit, type CBM-130;
- pressure Switch, type 111513-150;
- contactor, type I-4002; -
- relay, type 711-2;
- fuze, type 1111-250;
- imix1ng lamp, type 011-51 , with red light filter (2 pieces).
The electric units are located as follows:
1. The hydraulic pump 465 K, the 011-130 pressure drop warming unit and
the UMB-150 pressure switch are located in the hydraulic panel at frame lio,15.
2. The contactor, type K-400A , designed for stitching on the hydraulic
pimp electric motor, the intermediate relay, type 70-2 , for switching on the
hydraulic pump and the fuse, type 511-250 , are connected in the hydraulic pump
electric motor supply circuit and are mounted in the hydraulic panel junction
box at frame No.15,
3. The pressure drop warning lamps, type 0311-51 , of the normal and
emergency hydraulic systems are mounted on the pilots' central electric board.
Checking Operation of Hydraulic System Electric
Col,trol
1. On the left-seat pilot's circuit breaker panel switch off the two
hydraulic system control and warming circuit breakers, tyeeAaC-2, and release
to taro hydraillic pressure from the main and emergency hydraulic accumulatore
Of the brake hydraulic system. From the raie hydraulic accumulator pressure is
released by the Operation of the main brake system valves (by pressing the pedals)
or through the shut-off valve in the hydraulic panel on frame No.15; from the
emergency hydraulic accumulator pressure is released by the operation of the
emergency brake valve on the pilot's central panel.
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2. gedioh on the too hydraulic system control and warning circuit breabem,
type IDC-2 . The two red laepa warning of pressure drop in the normal and
emergency eysteme on the pilots' central electric board must flash on.
3. Turn on the hydraulic pvmp control switch on the pilots' central penal.
!be
hydraulic pump must start operating and increase pressure in the 47draulle
antes:.
With the hydraulic system pressure not exceeding 35 kg per sq.cm., releue
the switch handle; the hydraulic pump must continue operating. At a pressure of
100 t. 5 kg per sq.cm. the normal system pressure drop warning lamp must go out;
pressure reaching 13013 kg per sq.em.1 the emergency system pressure drop wain.
ins lamp goes out. At a pressure of 15013 kg per.sq.cm. the hydraulic pump gee,
automatically cut off.
4. By means of the main brake valve, release pressure in the normal hydranuc
system. With pressure dropping to 12013 kg per eq.cm., the hydraulic pump starts
operating; at a pressure of 15013 kg per sq.cm. the pump gets cut off.
5. By MEBUL9 of the emergency brake valve release pressure from the emergency
system. Pressure reaching 1301 kg per sq.cm. ,the emergency hydraulic system
pressure drop red warning lamp meet flash on.
, 1. The operation of the brake hydraulic system pressure control
electric system should be checked by the aircraft technician
together with an electrician.
2. When checking the operation of the hydraulic system, see that
proper operation duty of the hydraulic pump is maintained.
3. Daring the operation of the hydraulic pump make sure that the
current consumed by the pump electric meter is within the rated
Unite.
beelficetions_of Eystem Electric Units
Kl?ctric Pump 465% and Electric
Neter A450CE
1. Direction of rotation
2. Nominal voltage
3. Voltage operating range
4. Consumed current:
at operating pressure of 150 kg per sq.cm.
at maximum pressure of 180 kg per sq.cm.
5. Permissible peaks
6. Operation temperature range
7. Electric motor operating altitude
8. Break minimax length
9. Operation duty on the ground
at altitudes
left
27 V
24 to 30 V
not in excess of
180 A
not in excess of
260 A
not in excess of
300 A. up to
2 sec.
from .:70 to -600C
12,000 a.
14 mm
60-min, operation
followed by
complete cooling
(not less than
1 hour)
30-min. cperstion
followed by
Complete cooling
Pressure Drop Warning_ Unit. 'Flee ME-130
operation at normal temperature
ND* Dmtrumenc operates at 0.5 A
cexWmmvibretion overload
and 27 = 2.9 V
Pressure Switch. Type nms-no
0,,,mire operating range
2.1morof contact operation at normal temperature:
at points 30 and 100 kg per sq.cm. 5 kg per sq.cm.
at points 120 and 150 kg per sq.cm. 5 kg per Sq.CM
ib lest= vibration overload
error of operation:
at points 120 and 150 kg per sq.cm.
at pdate 30 and 100 kg per sq.cm.
4. The instrument operates at 27 t 2.7 V and 0.5 A.
130 kg per sq.ca.
not in excess of
+5 -_
_2 es per sq.cm.
not in excess of
2.5 g, with error
not exceeding
+6
_2 kg per sq.cm.
froze to 150 kg
per sq.cm.
not in OXCEISS
of 1.5 g
+5 kg per sq.cm.
-2
16 kg per sq.cm.
CABIN HEATING ELECTRIC SYSTEM
The cabin heating electric system is designed to prevent the glass pains
beadisaing, as well as for additional heating of the cabin by means of
electric heaters "Unit 107". In the front cabin the beater is installed at the
amtmad side near frame No.5; the switches are mounted on the pilots' upper
electric boards. In the rear cabin the heater is installed on the port side
0100 frame No.73 (Fig.114), the switches being mounted on the radio operator's
electric board (Fig.115).
The fumes, type 811-150 . of the electric beater circuits are located as
hEloas: for the front cabin on the starboard side at frame Ho.6 in the glass
lain heating system junction box, for the rear cabin on the pert side at frame
10.74intbe rear cabin junction box. 4
Decifications of the Heater "Unit 107"
1.Toltage
a. Current in the heating element circuit at V . 27 V
(adth 3 heating elements cut in)
5.0nnTentin the ventilator motor circuit at V . 27 V4.
%mting value:
(a) at altitudes from 0 to 7000 m.
ERT
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not in excess of
135 A
not in excese
of 30 A
'000+120
-510
per hour
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(b) at altitudes from 7000 to 15,000 at.
2oce80
-34olcal
per hour
5. Brushes, type VTC-7 j MiniMUM length 10 mm
Up to 15,0001.
Checking Operation of Cabin Heating Electric
System
1. Switch on the heater control circuit breaker, type A3C-30, on tsmit.
seat pilot's circuit breaker panel.
2. Turn on the HEATKR-VENTILLTOR (OBOTTEBATEIL-BEHTNIATM ) switch cc
pilots t upper electric board. The heater ventilator meet force air through,
switching on the 1st section warm air must itart coming out of the beaterk
some time; when the 2nd section is switched on in addition to the let stetim
still warmer air comes out of the heater. Male sure that air from the slotd
the pipe line nozzles of the navigator's, pilots' and blisters glass panel:es
at constant pressure.
Check the operation of the heater in the rear cabin in a similar nemere
the operation of the heater in the front cabin.
When checking the operation of the heater "Unit 107', measure the currati
consumed; normal current consumption testifies to the proper operation Mete
beater.
6. Operating altitude
Notes: 1. In case the electric motor, type A-400A , fails, it is prohIk
to switch on the heater.
2. Prior to switching on the power supply, make sure that theme
no foreign objects at the ventilator window and on the Wad
"Unit 107'. Remove foreign objects, if any.
3. Switch off the heater after its operation has been checked.
Possible Faults of Cabin Heating Electric
System and Their Elimination
Fault
Cause
Remedy
i
2
5
The heater body is over-
(a) The ventilator window
Remove the fore*
heated during operation
is closed by foreign
objects
.
objects
(b) The non-return valve
Check the operatio
operates with jamming
of the non-returner
in the tube common.:
the heater with the
line. If jam:tang it
detected, elinUmtet
'
(c) The thermoswitch
Remove "Unit 107'
-
fails to operate
the aircraft. atc"
operation of the dv,
switch. In the evedii
its improper opettO,
replace the theme0:0,
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-179-
1
2
3
. (d) The altitude relay
fails to operate
PREFLIGHT PREPARATION
Remove "Unit 107" from
the aircraft and check
the operation of the
altitude relay. In case
the latter fails to
operate properly, re-
place it.
systematic maintenance operations on the aircraft electrical equipment are
deautely necessary to ensure normal operation of the equipment; the main
dments of the maintenance procedure are the preflight preparation, postflight
impection and scheduled maintenance operations.
The scope of the preflight preparation depends on the scope and results of
Um previous postflight inspections and the thoroughness with which the troubles
&Meted in flight and during the ground check have been eliminated.
The preflight preparation and postflight inspection of the aircraft
electrical equipment consist in.inspecting the electric wiring and units for
condition and in voltage testing of the units.
It is advisable to adhere to the following ground check inspection procedure
(velleetround) during the preflight preparation and postflight inspections of the
electrical equipment:
(1) front cabin and fuselage between frames Nos 12-14;
(2) I.G. nosewheel well;
(5) I.G. left strut nacelle;
(4) navigation lights of left outer wing penal;
(5) stern cabin and tail skid;
(o2 :series compartment between frames Nos 56-69, fuselage belly section
ond b
(7) I.G. right strut nacelle;
(8) navigation lights of right outer wing panel;
(9) top sections of fuselage and icings;
(10) nacelles of right and left engines.
Preflight Preparation before Energizing Electrical Ectuinrent
Front Cabin and Fuselage between Frames Nos 12-14
1. flake sure that the storage battery switch on the radar operator's
olectric control board is OFF.
2, Carry out the following checks at the radar operator's station:
(a) check the ON-OFF and change-over switches, circuit breakers, rheostats
zoiceerating knobs of the cabin light and ultra-violet illumination system
improper functioning; the check is done by manually engaging and disengaging
thiabove-mentioned items; check for proper attachment;
(b) make sure that the glasses of the ammeters, voltmeters and lights are
haaet and that the instruments are securely attached in their mounting peti-
tions;
(a) see to it that the voltmeter and ammeter needles are zeroed and that
4* fuel ystem boosters are reliably fastened to their mounting platforms.
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180
3. The following checks should be carried out at the stations of the p0-lt4
and navigator:
(a) check (by engaging and disengaging with the hand) the ON-OFF amich200,
over switches, as well as the circuit breakers, operating knobs and rheostats
for sound operation;
(b) check the signalization and illumination equipment for condition and
secure attachment;
(c) maks sure that the cabin heater (Unit 107) and the AOC-81M automatic
glass- panel temperature controller are reliably attached and that their Shock
absorbers !Unction properly.
4. When through with the checks, place all the ON-OFF and change-over
switches and the circuit breakers (which serve as switches) to OFF (BWIED4o20)
Or NEUTRAL (ILEATFLIBHO ).
5. Maks sure that spare bulbs and fuses are available is the flight
maintenance kit.
6. See to it that the hydraulic control panel connections from the mitts
of the hydraulic system automatic control equipment are intact.
7. Inspect and make sure that the union nuts on plug connectors and fire
extinguisher discharge bonnets at frameproperl
y tightened up and
lockwired.
50.12 are
I. Check to see that the glasses of the la,Aiag, taxiing and well illanda.
tion lamps are intact and that the lamps are attached securely.
2. Check to see if the limit switches on the lock and brace strut of the
nosewheel are intact and reliably attached; inspect for secure wire
connections.
3. Check the fuel system boosters and 110,-4500 inverters for secure attach.
sent and see that the firing (discharge) mechanisms on the discharge bonnets of
the CO2 and inert gas bottles are properly locked.
Bight and beft LLLLL ?tr.q. Nacelles
1. Check the limit switches on the locks and shock absorbers of the oda
L.O. legs for secure attachment and sound operation.
2. Check the wires for proper attachment and connection to the limit
switches, towling. lamp and automatic brake control units, type YA-16
3. Check the bottom formation light and illumination equipment for scamd
Operation.
Navigation_Idghte of_Ieft and Right Outer
Wing Panels
1. Inspect the attachment fittings of the navigation light equipment and
make sure that the cover glasses of the lights are intact.
2. Make sure that there is no mator, ice or dirt under the light common
glass.
Stern Cabin and Tail Skid
?
1. Operating the switches, control Inas, circuit breakers and rheostats
manually, make sure that they function properly.
2. Melo sure that the cabin heater and the warning (signalization) equir
Lent are attached reliably.
3. Place all the switches and rheostats OFF.
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- 181 -
a. Check the voltmeter for condition and make sure that the instrument
udits cover glass are securely held in place.
5, Inspect the tail navigation light system and make sure that the glass
and the attachment fittings are intact.
6. Check the tail skid actuator and its electric wires for condition and
Attachment.
Acceesories Co2partment between Frames Nos 56 and 6? Fuselage
Belly Section and Bomb Bays .
1. Check to see if the MIKA-3A actuator, the de-icer junction box and the
chvWt breaker of the autopilot servo-unit heater system are attached securely.
2. Check electric wires for conditiqn and secure attachment.
3. Inspect the bottom formation light system.
e. In the bomb bays: check the electric wires for condition, and the
jection boxes and landing flap actuator, type 103-3M, for reliable attachment.
5. Check the 10-18000 ballast resistor for secure attachment and proper

ToR Section of Fuselage end Wings
Check the top formation light system for condition and reliable attachment.
Right and Left Engine-Nacelles
Check:the electric equipment of the engines for proper attachment and the
electric wires for condition; check to see if the MP-18000 generators, P3T-62
voltage regulators, TC-8 stability transformers and the overheat warning units
=attached securely.
Daring external inspections of the equipment in all the aircraft sections
sae sue that the fuses on the control panels and in boxes meet the Specifica-
tions Indicated on the respective nameplates and are reliably attached, that the
covers of the connector boxes are tight at their edges, and that the locks
mw lockwired and reliably retain the covers against vibration and falling out
in flight.
Autopilot, Type AH-5-2M
1. Carry Out condition and voltage checks of the autopilot units. Inspect
externally to check whether the autopilot units are free from moisture, dust
and breakdowns in connections to aircraft structural members. Remove the covers
bathe formation stick and directional stabilizer.
2. The autopilot preflight preparation procedure is obligatory before each
flight. If several flights take place during one day, it is sufficient to carry
Sit the preflight praperation before the first flight.
3. /f the ambient air temperature is below minus 2000, the autopilot
beaters should be engaged for one hour before the flight.
4. Turn on the A3C-15 circuit t mker of the torque MOUT assembly on the
nrigstor's circuit breaker control panel, the A30-5 circuit breaker on the
silot.s circuit breaker control panel, and the master switch on the autopilot
*nerd panel, and check the autopilot operation under voltage.
Check the clutch tension by hand, employing the following procedure:
(a) engage the bomb sight and autopilot clutches;
(b) turn the bomb sight so that the autopilot clutch lever would reach its
4q. In this position the autopilot clutch begins to slip on its drum; during
the bather rotation the clutch should not slip;
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() turn the switches on the autopilot control panel off.
'Voltage Test of Electrical Equipment
1. Carry out external inspection of the storage half-batteries, type
12-0171,55. If the batteries are operative, install them on the aircraft, euel
in place and close the container covers.
2. Place the storage battery change-over awitch to the =ma, (ROPROW
position, and check for loads by the ammeter on the radar operator's electric
control board. Engage the gyro horizon pets of the pilot and co-pilot smith.
interphone set which will correspond to a 10 to 12 A load on the battery, ma
check the battery voltage. The indicated voltage Should not be smaller the:n.241.
Disconnect the gyro horizon sets and the interphone system, and set the stomp
battery change-over switch to the neutral position.
3. Connect the storage battery in turn to the normal supply circuit and to
the triple supply busbar. To make certain that the storage battery energizes
' these circuits, engage the gyro horizon sets of the pilot. When the storage
battery is connected to the normal supply circuit, both gyro horizons shmile
operate. Mhen the. battery is connected to the triple supply busbar, it is maly
the stand-by gyro horizon which should operate. The operation of the gyro
horizons will be indicated by the noise of the inverter.
4. Disconnect the gyro horizon sets and the storage battery.
? 5. Connect the aircraft electric mains to a ground supply source.
6. Operating collectively with the aircraft technician or mechanic,
eheckthefollowing:
(a) operation of the control system of flaps, elevator and rudder trill:tea
and of ailerons. Synchronize the operation of the aileron trim tabs;
(b) operation of the tail unit de-icers;
(o) glass panel electric heating ;system;
(d) L.G. warming system: hand pressure upon the limit switches correspen?
ing to the L.G. retracted position should result in flashing up of the red mre.
ing lights; at the same time the green LA. position warning lights should.gc
on burning;
(e) operation of the main and stand-by inverters, type 110-4500 , with
reference to the aircraft 1.0. voltmeter;
(f) operation of the fuel automatic control system and of the feel flee
gauges;
(g) operation of the cabin ventilators and heaters.
7. Check the operation of the unit of fire-fighting system electromagnetic
valves; while checking, do not engage the 130-10 circuit breaker which opens
the CO2 bottles and the inert gas system switch, type B-45, on the overhead
electric control panel of the pilots since otherwise the discharge bonnets.
(firing mechanisms) will be actuated.
8. When testing the operation of the engines, check the operation of the
generators; if necessary, adjust the generator voltage and check the generator-
to-emergency supply circuit voltage supply.
POSTE-MOT II:MOTION
?
Gain information on the in-flight operation of the electrical equipment
from the crew members.
Do-energize the aircraft electric maim and disconnect the storage batter:0
This done, proceed to inspecting the system. The sequence of inspections is tia,
sane as that authorized for the preflight preparation.
egdect to inspection will be: the electrical equipment, the warning
(4eloption) system, the illumination equipment, the electric actuators, the
embed conductors and junction boxes. When inspecting, make sure that:
Lai. the equipment fittings, rheostats, switches, relays, bulbs, receptac -
eseenit breakers and other equipment itema are securely attached to their
outipe panels and boards.
2.111 the nameplates and instruction plates which concern the function or
operation of separate unite and switches are in good condition (are neither
armed nor fouled).
3. The clearance between the bunched conductors and moving parts is at
last 10 113.
4. Tbe union nuts of the plug connectors are adequately tightened up and
lodvired.
5, The mounting areas of the plug connectors and special 'wire adapters have
relameeelportions of cabin-sealing cement.
6. The gaps between the power contacts and the airframe members gauge at
heat 5 mm. Special attention shoula be paid to insulating the wires from the
Ian ('airframe") as aAy contact of a bare plus wire with the airframe results in
Mort-circuiting.
7. Reliable contact is ensured at the connections of power contacts.
8. In case of dirt, dust, oil or moisture on the electric wires or equipment
liens, wipe them with a clean cloth.
9. Carry out external inspection of the storage half-batteries and maks sure
that:
(a) the half-batteries are clean from the outside;
(b there are no cracks and breakdowns in the electric contacts and
:Marcell connections;
(e) the monoblock, cover and vent plugs are free from fouling and damage:.
dem fouled spots, if any.
If the storage battery is damaged, send it over for detailed inspec-
tion and correction of faults.
10. Check the condition of the storage battery containers:
(a) see to it that the felt is not moistened with electrolyte;
(b) check to see if the wires in the container are intact;
(c) see to it that the container cover locks are intact;
(d) make sure that the storage battery connectors connecting it to the
aircraft electric mains are sound.
The inspection over and the detected troubles eliminated, tura off all the
itches but for the interlock switch operating with the generator switch
connecting bar; place the storage battery change-over switch neutral, connect
Vie ground supply' source and carry out the voltage check of the electrical
kad1Pcmuta
Correct all the troubles detected during the voltage check. Troubles should
ha eliminated with the aircraft electric mains de-energized.
The inspection and trouble eliminating procedure over, report the electric
tIdNumt readiness for operation and termination of the operations to tee aio-
craft technician and the special equipment technician.
Checking Instruments for Serviceabilite
1. Turn on the 130-2 circuit breaker and the cabin air temperature regula-
fm'on the circuit breaker control panel of the co-pilot.
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2. Place the CABIN AIR SUM/ TEYPERATURR (TEUHEPATYPA 8A.13.11393A Naga)
'selector switch mounted on the co-pilot's instrument penal to the BET opi
position. In this position the MPT-1 actuator of the turbine-driven cooler
should close the cooler and open the cabin air temperature regulator.
5. With the selector switch thrown to COLD ( IOL. ),othe electric actuq4
should operate in the reverse direction.
4. Set the change-over switch to the AUTOMATION ( ABTOMAT ) positioa.
5. Set the cabin air temperature regulator thermostat scale of the firt
cabin to read 3 to 50C lower than the apbient air temperature. In this poai.
tion the MPT-1 electric actuator should cut out the cabin air supply teRpml.
ture regulator and engage the turbine-driven cooler.
6. Set the thermostat scale to read 3 to 50 above the ambient airt
tore. In this position the MP24 actuator should engage the cabin air-
temperature regulator and cut off the turbine-driven cooler.
7. The thermostat of the rear cabin will be checked with employmmt at
same procedure.
8. If the ambient air temperature does not permit to set the thermos*
scale at a temperature higher or lower than he original one, it is necessam
first to heat up or cool down the thermostat to a temperature of 19 - 23?,4,1
then to carry out the check according to steps 4 - 7 above.
'
Due, to the fact that the regulator check for meeting the Specificatimm
requires bulky fixtures which are not in quantity production, it proves *Da
sible to carry out the checks directly in the using unit. Therefore adequate
operation of the temperature regulator will be judged upon by its satisfados
functioning to maintain the pre-assigned cabin air temperature in the course
of the flight.
Automatic Cabin Air Temperature Regulator, Type FTBK-4
The regulator, type PTBE-45, is designed for automatically mainteining
the pre-assigned air temperature in the. pressurized aircraft cabin.
The regulator set includes:
- one thermostat, type TETRE-24;
- One electric actuator, type MPT-1.
Basic Characteristics
1. Nominal veltage requirement 2p.5 V
16.5 to 26.5%
3. Accuracy (no-response zone) ? =1?C
4?C
2. Temperature Control range
4. Degree of feedback irregularity
5. Current requirement by IET-1 actuator not over 1 A
.6. Nominal shaft load of MPT-1 actuator 120 kg/cm.
135o 30
7. Rotation angle of MPT-1 actuator output shaft
8. Time required for HI2-1 actuator output shaft to
turn through 135? I 3?
9. Operating duty of MPT-1 actuator
not longer than
45 sec.
intermittent
10. Resistance of MPT-1 actuator potentiometer 400 = 20 ohms
111 the units of P281-45 are interchangeable.
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PHOTOGRAPHIC EQUIPENT
GEMA.L
The photographic equipment carried by the aircraft includes:
-set of cameras A0A-33/5011 , L01-33/7511 and AU-33/100U intended for
roundgcametm argteAptsLl for
woo photography of the ground targets;
-set of camerae HA0A-3C/50 or RAA-6/5O for night time photography, of
Us
photographing the screen of the cathode-ray tube of
sderbomb sight p511-4;
-automatic tilting mount ABIOY-156H for all daytime cameras;
=era mount (frame HAOA ) for night time cameras;
-camera hatch;
-camera hatch and tilting mount control panel.
Arrangement of the photographic equipment on the aircraft is Shown in
11016.
The aircraft may carry only one of the aforementioned cameras (besides
=era IP-1 which is never removed) and one camera mount.
The camera mounts (tilting mount HAW and the frame) are installed on
spring-loaded shock absorbers selected according to the camera weight.Purnished
Ii th the aircraft are shock absorbers coming in three variants to fit cameras
1&t-33/1Oou . 10A-33/75M and RAWA-33/501f HAA-3C/5O and HAOA-6/50.
The automatic tilting camera mount ARA6Y-156H ensures two-strip vertical
*oblique photography. /n the case of two strip photography (AERIAL RECONNAIS-
MWM mode* operation), the camera mount departs from the vertical plane
though 6?30' to both sides when carrying camera A0A-33/100M and through 8?30'
teacamling camera A'BA-33/?5M.
112JW Camera LOA-33/50M is not employed on aircraft T146 with the
AERIAL RECONNAISSANCE mode of operation because only part of the
light rays of the camera vision field (34?) pass through the camera
batch hole.
During the oblique photography (BOMBING CONTROL
IlitOSatic tilting camera zu::d7F-156H deflects against the flight through
the esgles of 0; 10; 15;
a,Pending on the scale of aerial survey.
:ix:a:eras far daytime photography can be operated at various altitudes
Ilbsimum survey altitude depends on the flight speed and is calculated
mode
of operation).
the
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where: H -
N-
F
idnimum altitude of flight in km.;
exposure time in seconda;
focal length of aerial (miasma la:bug
speed of flight in kr://xe.
EMSIgoonSmo
Daytime Photography Caa_
1. Piettre size
2. Number of pictures
3. Size of film to be threaded up
4. Photography cycle times
at 15. to 25?S temperatures
at -60?C temperature
5. Power consumed
at 15 to 25?C temperatures
at -60?C temperature
6. Focal length:
? camera LOA-33/100E
camera 10A-33/7511
camera 101-33/5011
7. Interfreme space
8. Camera controller intervals
9. Thermoregulator:
engagement temperature
disengagement temperature
10. Camera controller ensures functioning of the
camera upon keeping electric bomb release button
SCEP pressed for 0.2 - 0.3 seconds
11. Exposure time (expressed in fractions of second):
cameras 101-33/507 and APA-33/794
camera ANA-33/100N
1. Focal length
2. Picture size
3. Number of pictures
4. Shutter
Camera 11041-30/50
5. Exposure time (expressed in fractions of second)
6. Power consumed:
at 10 to 3000 temperatures
at -60?C temperature
7. Photography cycle time
8. Shutter operation optical exposure
30:(30 cm.
190 to 195 pea
32x6000 cm.
not exceeding 2 se,
not exceeding 2.5 it
UP to 13.51
up to 16
100 cm.
75 cm.
50 cm.
10 to 25 mm
2 to 60 see.
3 to 13?C
20 to 30?C
1/75:
1/150;
1/300
1/75; 1/125; lra
50 cm.
18x24 cm.
approx.150p0
louvre type
1/25; 1/50; 1/1W
12 A
13.5 A
not exceeding 3,t.
2 to 15 luxes of
photocell
wel length
2. sietare Size
0
Ith81-6/50
50 cm.
louvre type
27crat.:760?Cdt'empersture
at 10.to 3000 temperatures 1/25;1/50; 1/100
4. weer*
time in fractions of second
laerceeding 15 A
6.1tetogdepby cycle time not exceeding 3 sec.
7.0Mmas operation temperature range +50?13 to -60?C
komer operation optical exposure 1 to 15 luxes of
photocell
Camera FAPA-1
Llama length 100 mm
2. Picture eine 13 coati dia.
(13x18 cm. frame)
3.111s4 perforated
width 19 cm.
length 28.5m.
I. Emilmmof pictures token without loading the file
assistne approx.200
5. Opals of camera operation alternative, depend-
ing on the antenna
revolutions or
sector scanning
angle
6. Power consumed:
with beater off
with beater an
5.3 A
15.6A
7. car.era operation temperature range +5000 to -6000
Technical and Adjustment Data of Automatic Tilting
Yount A1M-1565
I. Original position of the automatic tilting mount ALM in the vertical
imposition of the aerial camera AOA set within +0030 to -1? tolerance.
2. The tolerance for the tilting angle should stay within: +0?30. to -1?
for 6?30' and 8?30' tilting angles in the AERIAL RECONNALSSANCE mode of opera-
tion;
460' for 0; 10; 15; 20; 25? tilting angles in the BOMBING CONTROL mode of
iteration.
3. In the zero position, the play of the automatic tilting mount should be
oitidntO?30. (without taking into account the play in the reduction unit of
Um electric mechanism M10-2 ).
4. Time of changing the automatic tilting mount from one extreme position
to other:
in the AMA'. RBCONNAISSANCE mode of operation - 0.9 to 1.5 sec.
SP.erPTPT
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in the BOMBING CONTROL mode of operation when tilting
from 0 to.239 and from 25 to 0? - 1.9 to 3.5 sec.
5. Minimum permissible interval between exposures in the AERIAL aiscoaus.
SINCE mode of operation - 3 sec.
6. Current pulse sent by the contact-pulse mechanism of camera Ati slamu
not last longer than 0.5 sec,
7. The automatic tilting mount AKAW must reliably operate at tem/Ante:el
from +5000 to -60?0 and relative humidity up to 98%, withstanding vibratimm
of 10 to 80 cycles.
8. Service life of the automatic tilting mount UAW guaranteed coven
2 years including 21.000 cycles of operation (20,000 cycles in the AERIALREede
HAISSANOE mode of operation and 1000 cycles in the BOMBING CONTROL mode of
operation).
9. Current in the circuit of electric mechanisms MY-2 in the mike
BECONNLISSANCE and BOMBING CONTROL modes of operation with camera AtA installe'd
in the automatic tilting mount AKAN' should not exceed 10 A when the voltee
applied is within 27 ? 2.7 V.
10. During the BOMB/NG CONTROL mode of operation, reverse movement limit
switch must function at the moment when the frames =wing from the lower posh.
tion pass the zero by 1 to 1.5?.
11. The limit switch labelled STARTING FROM EITREME POSITIONS (TP0PAHMEI)
KPADHH1 nozommg ) must function in the zero position of the AERIAL REIM
MLISSINSE mode of operation, keeping OFF all the time the frame remains Lathe
extreme positions. .
12. Accuracy of opera-ion of the limit switches of all fixed positions for
the tilting angles - 10?30'
%clinical and Adeustment_Data of_Mount _ (Frame _ HAtAl
for Night Time _ Photography Cameras
?
1. The mount may accommodate either camera HAA-3c/50 or camera REM-6150.
2. The mount (frame HAA ) is intended to change the camera tilting mee
from 0 to 250 against the flight every 2030'.
3. The camera is set at the required tilting angle on the ground.
4. Femme Hifi is fixed in the lower attachment sleeves of camera mount
11881:1-15611 . The shock absorbers should be free of vertical play.
5. The inner frame of the camera mount (frame HAM ) must be fixed without
play at all tilting angle's of the camera.
6. The camera cables should not be in the way of the camera (frame WO
tilting irrespective of the angle.
Maim Technecal_and_Adjustment Data _ of Camera
Hatch
1. The camera hatch doors are opened inside the fuselage with the
the remote-controlled mechanism YP-7H.
2. Strain of band pulls - 8 to 12 kg.
3. Door opening and closing time - 40 sec.
4. The current consumed by mechanism YP-711 should not exceed 8 A wader
the rated voltage.
5. Coat all friction parts of the camera hatch actuator with lubricant 1511
State Standard ITICT 3276-54 ? There is no need to apply lubricant to the real
surface on which the doors and rod bearings move.
aid of
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189
PREFLIGHT PREPARATION
Preflight preparation of the daytime photography cameras includes:
(1) cheekine of the camera hatch;
(2) installation of the automatic camera tilting mount AKAOY;
(3) mounting of the camera and its preliminary checking;
(4) checking of the tilting mount operation;
(5) preparation of the cameras for surveying.
preflight preparation of the night time photography cameras includes:
(1) installation of frame HAtA;
?
(2) installation of the night time photography cameras and their prelimina-
preemeing;
(3) preparation of the cameras for surveying.
Preflight preparation of camera' OAPX-1 includes:
(1) installation of camera 0API-1;
(2) checking of the camera mechanism functioning;
(3) preparation of the camera for flight.
Preflight Preparation of Daytime Photography Cameras.
Checking of Camera Hatch
Check the camera hatch doors for proper closing and opening (Fig.117) by
eating the switch mounted on the control panel (Fig.118) 2 - 3 times on and off.
Whig made sure the camera hatch functions properly, proceed to installing the
mtmmtic camera tilting mount AKAtY or frame KAU.
Installation of the Tilting Mount AKAOY-150
Then doing survey jobs with the aid of camera ASA-33/1001I, install the
tlleing mount (Fig.119) with the shock absorbers, having on the cover marking
f-EXO on the upper row of the sleeves; in the case of cameras. Atk-33/75M
ad Ati-33/5011, install the tilting mount with the shock absorbers, having on
Uncover marking F-750 on ties lower row of sleeves. The tilting mount
Easing been installed, tighten the shock absorber sleeves as far as they will
govltekthe aid of union nuts 1.
Notes: 1. For installing the automatic tilting mount, remove the partition
separating the nose leg well from the camera bay.
2. Install the tilting mount horizontal accurate within +0030, an
-10 with the aircraft in the line-of-flight position.
Set the crank of the mount tilting mechanism with the aid of locking screw 11
at 6?30' when camera ASA-33/100m is to be installed and at 8030, when camera
MR-33/75 M is to be installed. Mount the bonding strips.
Camera_Installation
To install the camera:
1. Release hinged clamps 7 (See Fig.119).
2. Bring the camera trunnione in the Beats of the tilting mount AKAN and
Mx them with the aid of clamps.
Le_tee .The cbamner portion must be brought to the position shown by the
arrow marked on the film magazine (with the cardan shaft of the
driving unit set right of the aircraft fore-and-aft axis).
25X1
Approved For Release 2004/01/16 : CIA-RDP78-03066R000300070001-0
Approved For Release 2004/01/16 : CIA-RDP78-03066R000300070001-0
-190-
3. Arrange the driving and delivery unit on plate 6 and connect it tete
reducing gear on the chamber portion with the aid .of oardan abaft 5. 21, orot
bends in this event should not exceed 250 at the binge joints.
4. Use flexible hose 4 to connect the air blower volute chamber to the
chamber portion pipe connection.
5. Actuate screws 2 to zero the automatic tilting mount AKAN accurate
within +0?52! to -1?.
6. Connect all the emits with electric cables.
7. Mount camera controller .KDY-2 on the navigator's panel and join pism
connectors to it.
Checking_of Cameras 10A-31/100M _ _ A01-330511_ and
101-a3/5061 Functioning
1. Take the levers all the way out of the chamber portion and remove the
protective cover from the latter.
2. Set 5 - 7 sec. interval on the camera controller dial (118.120).
3. Dress the green button START ( nYCH ).
? As the chamber portion of the driving and delivery unit operates, chock
the air delivery to the chamber portion, functioning of the shutter and the
objective protective covers, and the illumination of the recording instruments
at the moment of the shutter operation.
4. Arrange the film magazine loaded with the exposed waste film on the
chamber portion and take the cover off the film magazine.
Press button START ( HYCK ) on camera controller K111-2 As camera MK
operates, check the film for proper rewinding watching the indicating lam
labelled REWINDING ( HEPEUOTKA), the pressure plate for proper rising and tom-
ing and the camera controller meter for proper operation. This done, deensrefte
the camera controller by pressing the STOP ( OCTAHOB. ) button.
5. Disconnect the card= dhaft from the chamber portion reducing gown('
the reducing gear driving unit.
6. Connect the hand drive to the input shaft of the chamber portion raison;
gear.
7. Slowly rotate the hand drive handle clockwise to check the functioning
of the shutter (accompanied by a click).
8. Beginning with the moment the shutter starts functioning count the
number of the hand drive handle revolutions lip to the closing and opening
moments of the protective covers.
9. Check the air pressure Lathe chamber Seeing that it is at least 13