TECHNICAL REPORT ON MIG-15 BIS

• Document Type: CREST
• Collection: General CIA Records
• Document Number (FOIA) /ESDN (CREST): CIA-RDP80T00246A027100330001-4
• Release Decision: RIPPUB
• Original Classification: S
• Document Page Count: 200
• Document Creation Date: December 22, 2016
• Document Release Date: June 28, 2010
• Sequence Number: 1
• Publication Date: August 1, 1957
• Content Type: REPORT

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6 4 25X1
SOURCE EVALUATIONS ARE DEFINITIVE. APPRAISAL OF CONTENT IS TFt~ITATIVF
SUBJECT Te"aca1 Report on MIG 15 BIB DATE DISTR. 1 August 1957
This material contain information affecting the National Defense of the United States within the meaning of the Espionage Laws, Title
18, U.S.C. Secs. 793 and 794, the transmission. or revelatlc of which In any manner to an unauthorized person is prohibited by law.
CENTRAL INTELLIGENCE AGENCY
Rek2a can 1 1
BIS No. 1
Part 0r,*
Tee Re ort on MIG 1
BIB No. 1
Part TWO
.
TMALW~L
on MIG 1
BIS No..
,
Part Thr"
Technical,
on MIG 1
BIB No.
Part Four
T+~?~nical Re
Oil MIG 1
BIS loo.
Part Five
lnwcal Report on NIG 1
BT6 ]
Pam S nt.
USAF, DIA, review
completed.
STATE ARMY
NAVY AIR I IFBI AEC
(Note: Washington distribution indicated by "X"; Field distribution by "#".)
Fleetrwtcik
ant
Electrioa]. 8 tem a
ufact l thuds k- 25X1
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TECHNICAL REPORT
ON
NO 1327
PART ONE
AIRFRAME
SECRET
SECRET
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PART ONE
AIRFRAf LE
S E .... E T
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.- N-- a . -
PART ONE
TABLE OF CONTENTS
SECTION I FORWORD
SECTION II AERODYNAMICS
A. General Description of Aircraft
B. Main Plane
E. Fuselage and Air Intake
J. General Comments
SECTION III STRUCTURES
A. General Statement on Structural Features
B. Wing
C. Fuselage
D. Horizontal Stabilizer
F. Landing Gear
Page
SECTION IV COCKPIT INSTALLATION AND EQUIPMENT 15
A. Introduction
B. Field of View
D. Control Column
E. Rudder Pedals
H. Undercarriage
I. Brakes
J. Radio Navigation Equipment
K. Instruments
M. Navigation Lights and Landing Light
0. Engine Controls and Instruments
SECTION V FLIGHT CONTROLS
B. Control Column and Rudder Pedals
SECTION VI PILOT.s SEAT 22
A. Description
B. Seat Ejection
SECTION VII COCKPIT CANOPY
C. Canopy Removal and Jettisoning
SECTION IX COCKPIT PRESSURIZATION AND AIR CONDITIONEL 25
SYSTEM
B. Cockpit Pressurization
C. Pressure Regulator
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Page
SECTION X OXYGEN SYSTEM
SECTION XI HYDRAULIC SYSTEMS
A. General Description
B. Main System
SECTION XIII INSTRUMENTS
SECTION XIV FUEL SYSTEM
SECTION XV PERSONNEL EQUIPMENT
SECTION XVIII ADDITIONAL PHOTOS
SECTION XIX DRAWINGS
60. Sketch showing Ronne Airport.
61. Outer Gimbal Caging Mechanism.
(Gyro Horizon)
62. Gyro Horizon, Wiring Diagram.
63. Turn and Bank Indicator, Wiring Diagram.
64. Turn and Bank Indicator (Scale Test).
65. Compass Transmitter Magnet Bearing.
66. Compass Transmitter, Wiring Diagram.
67. J.P.T. Relation between indication of
Jet Pipe Thermometer and ENI.
68. Relation between Thermocouple EMF
and Temperature.
69. Fuel Contents Transmitter Unit.
27
30
54
56
58
c.Pr Pr T
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1-1 "1 - \II a ~ L- Y
PART ONE
AIRFRAME
SECTION I
FOREWORD
The following report is based upon information gained from 25X1
examination of a Polish built aircraft (known as "FAGOT" designated
LIM-2 (Polish designation for the USSR MIG-15 Bis)
The full report has been divided into six separate publications
as follows:
Part One - Airframe
Part Two - Engines
Part Three - Radio and Navigation
Part Four - Armament
Part Five - Electrical
Part Six - Manufacturing Methods.
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Photo no. 2 - Three quarter rear view.
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3.
SECTION II.
AERODYNAMICS.
A. General Description of Aircraft.
Engine designations (frcm engine plate) LIS 2 (Polish designa-
tion of VK1).
2. Wing incidence was adjustable on the ground within certain limits
(see Section III Paragraph A.l.a.).
B. Mainplanes.
Wings.
Wing Body angle. Adjustable on the ground within small limits.
E. Fuselage and Air Intake.
1. Fuselage.
e. The position of the landing light has been transferred to
the underside of the port wing.
General comments.
13. The following photographic sequence and sketch are included at
this point to assist the analysis of emergency landing charac-
teristics. The aircraft came to rest approximately 475 meters
from point of touch down.
Conditions were as follows:
600 litres of fuel on board
flaps down
gear down but unlocked
full quantity of ammunition
soil was sandy and firm
air speed at touch down according to pilot was
260 kilometers per hour.
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,,W, -- ..+I %a~
Photo no. 3 - Point of touvh down
Photo no. 4 - 37 mm gun barrel
after nose gear collapse
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Photo no. 5 - 37 mm gun close-up
Photo no. 6 - Furrow caused by nose
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6.
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V`?\yI% k.- R,
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Photo no. 12 - Traces
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Photo no. 14 - Nose section damage
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SECTION III
STRUCTURE
A. General Statement on Structural Features
1. Wings
a. There were two other single pin attachments of the wing to the
fuselage. One at about 12% chord and the other at about 35%
chord. The center of the forward fitting was
2 cm below centerline and the center of the rear fitting was
0.7 cm below centerline, therefore both were load carrying.
Each was fitted with an eccentric cam which can be rotated for
making minor adjustments in the angle of incidence. The fit
of all pins was moderately loose. The diameter of the forward
pin was 18 mm and of the rear 12 mm. The adjusting arm of the
forward eccentric cam could be moved through an angle of 120
-600 or to +600 from the horizontal zero position. The arm could
be locked at the zero position and at plus or minus 200, 40
and 600. The eccentricity of the outside and the inside diame-
ters was 1.5 mm, giving a maximum vertical movement of 1.3 mm
at the 600 positions. The eccentricity of the rear cam was, of
course, in proportion to and less than that of the forward cam.
Note: Incorporation of the additional center pin would provide
greater torsional strength to the wing root section.
Note: The modification was reported by the pilot to be accomp-
lished at unit level. Close examination of the work confirmed
this and further indicated the use of modification kits.
Photo no. 15. Wing to fuselage fittings.
C T
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Photo no. 16. Forward Fitting.
Photo no. 17. Rear view of all Fittings.
Photo no. 18. Eccentrics and adjusting arms.
c,FC?FT
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b. Outboard of the main gear, wing construction consisted of
transverse ribs and tapering spanwise stringers having an
average pitch of 7 cm at mid span on the top surface and 8
cm on the bottom surface.
c. Double skin was used on both surfaces of the wing between the
two spars and extending from the root to the area of the wing
tank installation. The inner skin was extensively perforated
with lightening holes from the outer wing fence to the wing tip.
The outer skin was 1.7 mm thick and the inner skin was 1.4 mm
thick. Skin thicknesses as shown in Drawing no. 49 of the July
1953 report are identical to those measured on this aircraft
with these exceptions: the .064",.094", and .112" were measured
as .055", .067", and .067" respectively.
2. Fuselage
b. Intake duct skin thickness was .032". An additional wing fitting
has been installed at a position two feet and one inch behind
the forward wing fitting. This was bolted through the skin to
internal stiffness.
Photo no. 19. New Fitting.
Photo no. 20. New Fitting.
Photo no. 21. New and old Fittings.
SFC'PFT
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13.
3. Tail Unit
a. Dural plates about 10" long and .05" thick were attached to
the bottom and top forward surfaces at the center section.
b. The elevators had mass balance weights at the tips and a larg-
er mass balance weight at the center point. Elevator hinges
were at ribs one and five (measured from the tips) and a main
hinge was at the center point.
d. Skin thickness as shown on drawings no 48 and 49 of the July
1953 report are correct except for the upper tail which should
read .039" instead of .029".
4. Common Features
a. Delete the word "wings" and see paragraph A 1 a of this report.
B. Wings
8. See paragraph A la of this report.
C. Fuselage
2. See paragraph A la of this report.
Fuselage Serial No
(Plate in left hand
side of fuselage
at the wing attachments)
Photo no. 23.
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D. Horizontal Stabilizer
2. There were three stringers on the upper and three on the lower
surface of the stabilizer. Between the second and the third
stringers there was a forged main spar. A channel beam and a
concave web, which accomodated the leading edge of the elevator,
formed the rear spar.
3. See paragraph A 3c.
F. Landing Gear
1. Dimensional data
Main tire size
660 x 160 w
Right tire pressure
77 psi
Left pressure
80 psi
4. Nose Gear
c. No instruction plate installed.
Tire pressure 34 psi
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15.
SECTION IV
COCKPIT INSTALLATION AND EQUIPP~ENT
A. Introduction.
1. For alterations in the cockpit layout
following tex :
see photo 24 through 30 and the 25X1
Starting at the left rear and proceeding forward:
a. A VHF power pack has been installed.
b. Items 6 and 8 have been removed and replaced by the signal
flare discharger switch and firing buttons (items 7 and 9).
c. A VHF control box is now located where items 7 and 9 used
to be.
d. Item 85 has been removed and incorporated in VHF control
box.
e. Tripple toggle switch, item 8, has been replaced by a three
position rotary switch.
f. Item 17 has been labelled "Drop Tank warning switch".
g. Item 15 removed.
h. Item 113 removed and a combined pressure gauge-flow meter
was in place of item 114.
i. Undercarriage master switch has been installed to the left
of the undercarriage selector lever.
j. Item 119 was installed on the right side of the gun sight
and a turn and bank indicator installed in its former
position.
k. Item 44 removed.
1. Item 55 moved slightly right.
m. Item 93A should be called "Control Column bearing housing".
n. An extra brake lever has been installed on the control
column main bearing housing.
o. Under coaming, right of item 126, were three coupled toggle
switches for external undercarriage warning lights.
p. Extra gun sight lamp holder installed above item 67A.
q. Master switch for landing light installed in blank space
beside item 67.
r. A bank of three female electrical sockets installed to the
immediate right of pilot.s position. Top one was unused and
painted half white and half red. Middle one was marked "in-
specting lamp". The bottom one was also half white and half
red.
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16.
Photo no. 24 -- Cockpit, Left console
A. VHF Control box
B. Signal Flare discharger
C. Three position rotory switch
D. Undercarriage master switch
.E.-Combined oxygen pressure gauge-blinker
F. Extra brake lever.
Photo no. 25 - Undercarriage master switch.
T
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Photo no. 26 - Cockpit. Right hand console.
A. Master switch for landing light.
B. Switch for external undercarriage warning
lights.
Photo no. 27 - New cockpit pressurization control.
17.
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6.b. Provision was made in the parachute pack for emergency oxygen
system.
c. The emergency brake lever which was fitted on control column
bearing housing could, if the control lock linkage was so
designed, be used as a parking brake.
B. Field of view,
1. The field of view as described in the first report had been
decreased by the repositioning of the clock to the right of the
gun-sight.
D. Control Column.
l.f. Emergency Brake lever.
This lever had been fitted into the control column bearing
housing. There was no provision for keeping it in the "on"
position.
Photo no. 28
Forward vision
Photo no. 29
New parking lever
E. Rudder Pedals.
1. Each pedal was fitted with an adjustable rubber strap.
H. Undercarriage.
1. Although the selector lever appeared the same, it operated a
three position toggle switch since the undercarriage selector
was olectro-hydraulic. Just to the left of the selector was a
master switch (fitted with guard plates) for the undercarriage
circuit.
SEC RE' T
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19.
In addition to this safety switch the selector lever was still fitted
with a sliding mechanical lock.
3. The switches which controlled the external undercarriage down-lock 25X1
lights had been moved into the
cockpit. Operation of this switch, with the undercarriage locked down,
caused three green lights to go on: one under each wing and one in the
position which had previously been occupied by the landing light.
I. Brakes.
2. See paragraph D.l.f.
J. Radio Navigational Equipment.
1. a. VHF Command Transmitter/receiver - This new equipment is described
in detail in Part III, "Radio and Communication Equipment", of this
report.
K. Instruments.
1. g. (i) A new electrically operated gyro horizon was fitted which also
incorporated a wall-type side-slip indicator.
g.(ii) A separate turn and bank indicator was fitted.
2. a. The clock was fitted on the top of cockpit coaming.
Photo no. 30 - New gyro horizon
M. Navigation Lights and Landing light.
3. The landing light was installed under the port wing. It was control-
led by a two position switch fitted in the left hand side of the
cockpit. A master switch controlling the system was fitted in the
right hand console.
SECS
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0. Engine Controls and Instruments.
1. g. This switch, although relabelled Drop Tank warning switch,
still performs the same three functions.
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SECTION V.
FLYING CONTROLS.
B. Control Column and Rudder Pedals.
4. Adjustable rubber straps were provided.
Photo no. 31. Column and pedals.
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PILOT SEAT.
A. Description.
Although this was an entirely new seat
Canopy and
seat rejection levers were incorporated in each arm rest, per-
mitting ejection with either hand. Total weight of the components
listed in the July 1953 report has increased to 44,66 kg primari-
ly due to the increased weight of the back armor. With the addi-
tional weight of 3,5 kg due to the piston, firing pin mechanism
and yoke, the total weight of the seat as it is ejected was 48,16
kg.
B. Seat ejection.
1. 1. Remove ground safety pin.
2. Move either right or left canopy release handle forward.
This releases the canopy with in turn pulls a secondary safety
pin which is attached to the canopy. (See photo no. 36). The
firing pin mechanism is now armed.
2. The weight of the various components were as follows:
Head
armor with pad
8,98 kg
Back
armor
14,68 kg
Seat,
less armor
21,00 kg
Piston, firing pin mechanism
and yoke
3,50 kg
Total
48,16 kg
Photo no. 32 - Seat and dummy.
SEC RET
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23.
Photo no. 33 - Seat, rear.
Photo no. 34 - Seat, front.
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SECTION VII.
COCKPIT CANOPY.
2. When the canopy is either removed on the ground or jettisoned
in the air, the canopy pulls out a safety pin from the firing
pin mechanism. This pin is one end of a wire which is connec-
ted to the midframe of the canopy and coiled to form a long
spring. This spring imposed a load upon the secondary pin and
was therefore attached to the yoke by means of a safety wire.
The total length of this wire was 2 m.
If the canopy was jammed in the closed position, the pilot
could eject through the canopy by reaching back 4nd extracting
the secondary safety pin. This can be done without breaking
the safety wire.
Photo no. 35 - Canopy.
Photo no. 36 - Secondary safety pin cable.
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25.
SECTION IX
COCKPIT PRESSURIZATION AND AIR CONDITIONING SYSTEM.
B. Cockpit Pressurization.
2. The selector valve, located on the right hand side of the
cockpit, had four positions (see photo A-29). These posi-
tions were: - closed (ZAM), cold (Z), Mixed (C) and hot (G).
The cold, mixed and warm positions were marked by coloured
segments of blue, yellow and red respectively. The valve
could be placed in any intermediate position.
D. Cockpit Air Conditioning System.
1. The cool air nozzle in this aircraft had been moved slightly
to the right and turned through 900 to make room for the
clock, which had been repositioned.
Photo no. 37. Pressurization
selector valve.
Photo no. 38. Cool air
nozzle.
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26.
SECTION X.
OXYGEN SYSTEM.
1. All piping used in the installation was copper, painted blue.
Some piping in the nose :cction was chrome plated and paint-
ed blue. All high pressure pipes, after the main high pressure
valve, were marked by an lz" wide red band at each end of the
pipe; all low pressure pipes were marked by a l-2" wide, green
band.
2. Two plain steel, blue painted bottles were fitted. One of 4
litre capacity, the other of 2 litres. The weight of the bottles
was 7 kg and 4.7 kg respectively.
5. A pressure pipe line was taken forward from the reducer valve
to the pressure connection of a combined pressure gauge and
flow blinker mounted on the instrument panel.
The pressure gauge read from 0 - 150 kg/cm2 (0 - 2133 lbs/in2).
The gauge carried a danger mark in red just above the 150 kg/
cm2 mark. A pipe line from the low pressure side of the reduced
valve also went forward to the combined gauge. The flow-indica-
tor was of the blinker type.
Photo no. 39. Oxygen Gauge.
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27.
SECTION XI.
HYDRAULIC SYSTEMS. 25X1
A. General descri tion.
e tank pressurisation system was as
follows:
Pre s;.eure Reducing V~ alve
Air Filter
Relief Valves
lion return Valve
To Drain From To Drain
Engine Comp.
3. Air from the engine compressor was fed to the top of both
hydraulic tanks by way of a filter and a pressure reducing
valve. Two one-way relief valves were teed into the system
to limit the pressure in the tanks. These last two valves are
shown in photograph 15.2.88 in the July 1953 report. The filter
bore a label which said that it must be cleaned every 25 hours
of engine running.
Photo no. 40 - Air filter Photo no. 41
Filter, assembled
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A. 5. The main pressure lines were coded with three black bands
6. The pressure lines in the booster system were coded with three
black bands.
B. Main System.
2. A pressure line was taken from the accumulator to the electri-
cally operated undercarriage selector valve located on the left
hand side of the fuselage within the wing root. This valve had
two solenoids - one for "gear up" and one for "gear down".
Photo no. 42 - Undercarriage selector valve.
4.
A mechanical down lock was incorporated within the main under-
carriage actuator. Test and results showed that a hydraulic
pressure of 165 - 180 p.s.i. was required to unlock the gear.
When the inner door was open it was locked in this position
by a mechanical lock within the door actuator. A pressure of
90 - 100 p.s.i. was required to lock the door and a pressure
of 150 - 160 p.s.i. unlocked it.
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29.
7. There were self-sealing couplings in the air brake pressure and
return lines. These are located at the bottom of the fuselage just
aft of the main fuselage break joint.
Photo no. 43 - Coupling
Photo no. 44 - Coupling
Photo no. 45 - Coupling
^
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30.
SECTION XIII.
INSTRUMENTS.
Introduction.
The instrument equipment was the same as found previously, with
the following exceptions:
Gyro Horizon (new type).
Turn and Bank Indicator (added).
Gyro Magnetic Compass Transmitter (new type).
Oxygen Flow Indicator and Oxygen Pressure Gauge exchanged
by a combined instrument.
In the addition to tests and examination of the new type instru-
ments several tests were carried out on the other instruments.
CONTENTS
B. Flight instruments.
1. Airspeed Indicator.
2. Rate of Climb Indicator.
3. Altimeter.
4. Macbmeter and red warning light.
5. Gyro Horizon.
5 a. Turn and Bank Indicator.
6. Gyro Magnetic Compass.
7. Magnetic Compass.
8. Clock.
C. Engine Instruments.
1. Engine r.p.m. indicator system.
2. Jet Pipe Thermometer system.
3. Engine Gage Unit.
4. Fuel Burner Pressure Gage (low pressure).
D. Miscellaneous.
1. Cockpit differential pressure gauge and cob~pit altimeter.
2. Fuel Contents system.
3. Instrument Installation.
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31.
B. Flight Instruments.
1. Airspeed Indicator.
The airspeed indicator was exactly the same type as found pre-
viously.
The results of a scale test carried out are listed below:
km/h
Instrument
Error
km/h
108
-48
180
0
These errors are appreciab-
270
0
ly greater than found pro-
451
+19
viously.
631
+24
811
+29
992
+28
1082
+23
2. Rate of Climb Indicator.
The rate of Climb Indicator was exactly the same type as found
previously.
Scale tests carried out showed an accuracy comparable to that
found on the previous instruments.
3. Altimeter.
The altimeter was exactly the same type as found previously.
The following tests were carried out:
Leakage test: 15 p.s.i., no leakage.
Position error: max. 5 m.
Scale error test (Temp. 200 C):
Test point
Error
m
Friction error
m
asc.
desc.
m
0
-35
-35
1515
-15
+ 5
10
3030
-10
+30
25
6060
-20
+40
9090
-50
+40
35
12120
-60
-10
15150
-60
-20
18180
0
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32.
Scale error test (Temp - 40 C):
Test point
Error
m
m
asce
desc.
-21
-24
- 3
3027
+15
- 5
6087
+ 3
+50
9087
+ 8
+39
12035
- 5
-75
15241
+49
During the tests the pointer jumped about 300 m at approx.
6,5 km and 16,5 km.
4. Machmeter.
The machmeter was exactly the same type as found previously.
A scale test carried out at mach numbers 0,5 - 0,6 - 0,7 - 0,8
and 0,9 at heights of 0 - 10 - 20 - 30 - 40 and 50.000 ft showed
a max. error of 0,02.
The mach number warning switch was not installed in this aircraft
either.
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33.
5. GYRO HORIZON INDICATOR.
Introduction.
The Gyro Horizon Indicator is entirely different from the type
found earlier.
The instrument presentation (photo no. 46) is very similar to
that found on f. ex. the USAF Type J - 1 Attitude Gyro Indicator, but
the design is entirely different from the type J - 1 indicator.
The gyro rotor and housing has complete freedom of rotation
about the roll-and pitch axis.
The design is unusual as it, instead of one, contains two
gimbalrings, which are monitored by a small two-phase motor so they
will always be at a right angle to the rotor housing. Manual caging
of the instrument is possible only when the rotor is at rest, or run-
ning at low speed and only the outer gimbalring can be caged.
During flight tests carried out in a trainer aircraft the
performance of the instrument was excellent.
The instrument is not driven from the gyro-magnetic compass
inverter as earlier, but has its own inverter mounted behind the seat,
the two inverters are started simultaneously from the same switch.
Detailed description.
The instrument is composed of the following main subassemblies:
1.
Instrument housing incl. front panel and back plate.
2.
Outer
gimbalring.
3.
Inner
gimbalring.
4.
Rotor
and rotor housing.
1. Instrument housing.
The front panel is attached to the instrument housing by eight
screws.
It has a circular opening through which a portion of the indi-
cating sphere is visible.
Below this opening an ordinary ball-in-glass bank indicator is
mounted (photo no. 46 pos. C).
On the left side of the front panel is situated the trim indi-
cator adjusting knob (photo no. 46 pos. A).
The indicator is connected only in the left side (photo no. 48)
pos. A) to a sliding mechanism driven up or down by a crank shaft
mechanism (photo no. 48 pos. D) connected through a gear train to
the adjusting knob.
On the right side of the front panel is found a push button
(photo no. 46 pos. B), kept in its outer position by two small
leaf springs. When it is pressed, a ball bearing (photo no. 48 pos.
C) mounted on the shaft of the push button is pressed against the
front edge of the outer gimbalring (photo no. 47 pos. B), this edge
is curved with the "lowest point" in a position which corresponds
to the normal position of the gimbalring (see dwg. no. 61).
In this way only the outer gimbalring can be erected, there were no
provisions for caging the inner gimbalring or the rotor housing.
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I
..) L F L.
34.
The caging action of the mechanism is rather weak and inaccurate,
it is not possible to cage the ring when the gyro is running at normal
speed.
The instrument housing (photo no. 40a pos. A) is cast in light
alloy. On the side are four square holes which are normally hidden by
a light cover, which slides on to the housing from the rear, also co-
vering the back plate.
In the front part of the housing is pressed a steel ring (photo
no. 47 pos. D) which carries three small ball bearings supporting the
front part of the outer gimbalring.
The back plate is attached to the instrument housing by four mount-
ing screws. To the plate are attached the following compoments: (photo
no. 49)
Thermo switch (pos. A)
Two phase motor (pos. B)
4-pole supply plug (pos. C)
Brush assy. (6 brushes) (pos. E)
Two brushes (pos. F)
Resistor block (4 off) (pos. D)
The thermo switch is of the bimetallic type.
The two-phase motor is connected through a gear train, to a gear-
wheel mounted on the rear of the outer gimbalring.
Outer gimbalring. (photo no. 52)
The outer gimbalring is mounted inside the instrument housing.
The rear part is carried by a ball bearing in the back plate and the
protruding hollow rear axle is carrying six slip rings and a sliding cir-
cular contact plate corresponding to the abovementioned brushes.
To the rear part is furthermore attached a gearwheel driven by the
two-phase motor.
On the sides are mounted two ball bearings opposite each other,
carrying the inner gimbalring. At the left side ball bearing is mounted
the rotor of the torque motor effecting the bank erection of the instru-
ment, and two sliding contacts attached to the cover of this ball bearing.
A 4-brush assy. is attached to the right side.
Inner gimbalring.
This gimbalring is situated inside the outer gimbalring (photo no.
51 pos. B) carried by the abovementioned two ball bearings. To the left
side is attached the Stator windings of the abovementioned torque motor,
on the right side axle is found the slip ring assy, corresponding to the
abovementioned brushes.
The front part is circular carrying the rotor (photo no. 50 pos. A)
of the torque motor effecting the pitch erection of the instrument,. Across
this part is a bridge carrying the front bearing of the rotor housing, a
5-brush assy and contact plates (photo no. 51 pos. G), part of the outer
gimbalring controlling system.
The gimbalring is made slightly pendulous by a small balance weight
(photo no. 51 pos. C) attached to the ring.
Rotor Housing.
The rotor housing is carrying the following components.
A mercury switch (photo no. 52 pos. F), Stator windings of the
pitch torque motor (photo no. 51 pos. H).
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35.
The contact assy. of the front bearing axle, protruding through
the front bearing.
The indicating sphere is divided into two halves (photo no. 53
pos. A - B) each of which attached to two wheels (photo no. 52 pos. B)
mounted on the front and rear axle of the rotor housing.
The rotor is a normal 3-phase squirrel cage motor with the
stator windings attached to the inside of the rotor housing.
Electrical functioning.
(Refer to wiring diagram fig. no. 62).
The power for the instrument (36 volt, 400 cycle, 3-phase a - c)
is supplied through the 4-pole receotacle on the rear of the instrument.
The supply for the gyro rotor is fed through yellow, green and
brown leads through the sliding contacts to the rotor housing. The brown
lead is connected directly to the gyro rotor stator windings the green
lead is connected through the fixed-field windings of the pitch erecting
torque motor, and the yellow lead is connected through the fixed phase
windings of the bank erecting torque motor.
The erection to normal position is carried out by two normal two-
phase torque motors effecting bank and pitch erection respectively.
The pitch motor is considerably greater than the bank motor.
The mercury switch controlling the voltages to the motors, is
attached to the bottom of the rotor housing and is of a type similar to
the one found on the earlier instruments. It works in exactly the same
way (ref. dwg. no. 16 in first report).
When the gyro is in the normal position the white and black leads
are all connected to the brown central tap and consequently voltage is
fed to both signal windings of both motors. When the gyro is tilted to
be left f. ex., the connection between the left white and central brown
tap is disconnected, so only the one half of the signal-phase windings
is energized, thereby producing a torque in the proper direction and con-
sequently erecting the gyro. Voltage to the central tap of the signal-phase
winding of the pitch torque, motor is fed through a 300 resistor on the top
of the rotor housing from yellow phase.
Voltage to the central tap of the signal-phase windings of the bank
torque motor is controlled in the following way. When the instrument is
started the bimetallic thermoswitch on the back plate is closed. The green
phase is then connected directly to the blue lead feeding voltage to the
central tap, thus making the erection torque as big as possible.
The bimetallic leg of the thermo switch is continually heated by a
small coil; when the switch opens the signal voltage is reduced through
the 100_,'?-resistor on the resistor block on the back plate, thus reducing
the erection torque.
When the instrument is banked approx. 5 degrees the voltage is
further reduced as the two brushes of the circular sliding contact move into
the two isolated segments thus disconnecting the 100 JL resistor and leav-
ing only voltage to be fed through the 5,1 k f! resistor.
The purpose of this is to reduce errors when centrifugal forces act
on the mercury switch during turns.
The alignment of the outer and inner gimbalr.ing is accomplished in
the following way:
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36.
The two phase motor on the back plate rotates the outer gimbal-
ring when energized. The fixed-phase windin,";s are connected directly to
brown and yellow phase.
The two signal phase windings are connected through the two 3,3
k r'_ resistors to green phase.
The two violet leads are connected through the backplate slip
rings to the two brushes on the left side of outer gimbalring.
The contact plate is divided into two halves changing the connections
when the instrument is rotated 900 in pitch. From here the violet leads
continue and are finally connected to the two outer contact segments
of the contact plate (photo no. 51 pos. G) on the front part of the inner
gimbalring.
The corresponding brushes (photo no. 51 pos. F) are connected di-
rectly to green phase.
When the instrument is in its normal position the brushes are in
contact with the isolated segment in the middle, so no voltage is fed to
the violet leads. When the instrument is tilted the brushes come into
contact with one of the outer segments, thus connecting the green phase
directly to one of the signal-phase windings of the two-phase motor which
levels the outer (and inner) gimbalring again.
This system is introduced presumably because the outer gimbalring
is too heavy to be turned by the rotor alone. When the aircraft makes f.
ex. half a loop and continues in inverted flight the connections must be
reversed as the two-phase motor must turn in the opposite direction, this
being effected by the reversing switch on left side of the inner gimbal-
ring.
Contrary, when the aircraft is rolled into inverted flight the
two phase motor works normally, this being correct too.
Test.
The following tests were carried out on the gyro horizon:
Bench tests:
Starting voltage: Less than 5 volts.
Starting current: approx. 0,9 amp.
Normal consumption in the three phases:
0,35 amp; 0,35 amp; 0,54 amp.
Instrument in normal position 20 secs after start.
Thermo switch opens after approx. 12 min.
Erection tests:
Pitch
from
200
to
100
(down : 2 min 15 secs.
Pitch
from
290
to
100
(up) : 3 min 5 sees.
Bank
from
30
to
2210
(right): approx. 5 min.
Bank
from
30e
to
2220
(left) : approx. 7 min.
Flight tests:
The instrument was mounted in a training airplane and flight
tested.
The following manoeuvres were executed:
EGRET
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va..vi are.. R
37.
1. 180 degrees turns,
2. Roll,
3. Loop,
4. Stall Turn,
5. Spinn.
The following results were obtained:
180 degrees turns.
Several 180 degrees turns were made at different angular
velocities and bank angles (rate 1, bank angle approx. 15?; rate 4,
bank angle 25?).
After completion of each turn the aircraft was levelled and the
position of the indicating sphere was noted.
No errors could be read during these maneoeuvres.
Roll.
These were carried out at approx. 11 secs pr. revolution.
Up to four rolls were executed continuously; after levelling
the aircraft max. 2? errors were shown on the instrument.
Loop.
These were carried out at approx. 18 secs per revolution.
Up to four loops were carried out continously; after level-
ling the aircraft no error was shown 8n the instrument. During the loops
the instrument normally prucessed 180 at vertical climb on dive at-
titude, but in some cases this did not happen, the instrument only seem-
ed a little unstable at these attitudes.
The 180 0 prucession is quite normal for these type instruments (caused
by the so-called "gimbal-lock").
Stall Turn.
After stall turns (max. 3) up to 20 errors were shown on the
instrument.
Spinn.
A normal left spinn (2 turns) was carried out.
After levelling the aircraft 2? error in bank and turn was
shown on the instrument.
Trim Indicator Adjusting
Knob.
B: Outer Gimbal Caging Knob.
C: Bank Indicator.
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3g.
Photo no. 47
Gyro Horizon (Front
Panel removed)
A: Indicating Sphere.
B: Outer Gimbal Front Steel
Ring.
C: Outer Gimbal Front
Bearings.
D: Ball Bearing Support Ring.
Photo no. 48.
Front panel
Trim Indicator.
Bank Indicator.
Caging Ball Bearing and Leaf
Springs.
Adjusting Mechanism.
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39.
Photo no. 49
Back Plate.
A: Thermo Switch.
B: Two Phase Motor.
C: Receptacle.
D: Resistor Block.
E: Back Plate Brush
Block
F; Roll Cut-out Brushes.
G: Roll Cut-out Contact.
Photo no. 48 a
Instrument Housing.
A: Instrument Housing.
B: Outer Gimbal Ring.
C: Inner Gimbal Ring.
D: Rotor Housing.
E: Indicating Sphere.
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40.
Photo no. 50
Inner Gimbal Ring.
A: Bank Torque Motor
Rotor.
B: Bank Torque Motor
Stator.
Photo no. 51.
Gimbal Rings.
A: Outer Gimbal Ring.
B: Inner Gimbal Ring.
C: Balance Weight.
D: Inner Gimbal Brush
,,lock.
E: Rotor Housing Slip Rings.
F: Two Phase Motor Control
Brushes.
G: Two Phase Motor Contact
Plates.
H: Pitch Torque Motor
Stator Windings.
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a M%_- M c 1
. Photo no. 52.
Gimbal Ring and
Rotor HousinE.
A: Outer Gimbal
Gear Wheel.
B: Indicating
Sphere Moun-
ting Wheel.
C: Outer Gimbal
Rear Axle.
D: Inner Gimbal
Brush Block.
E. Rotor Housing
Slip Ring
Block.
F: Mercury Switch.
G: Rotor Housing.
Photo no. 53.
Indicating sphere.
A: Upper Half
Part (Brown).
B. Lower Half
Part (Blue).
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%J r%"
5. A. TURN AND BANK INDICATOR.
Introduction.
The turn and bank indicator is a normal electrically driven
type. The dial presentation is normal with three scale markings
for left and right turn respectively (photo no. 54). The bank
indicating mechanism is of the ball-in-glass type.
Description.
The instrument consists of the following two major sub-
assemblies:
1. Instrument Frame and Cover.
2. Rotor Housing.
The instrument frame (photo no. 55) is cast in light alloy,
with different components attached to it.
Inside the front part is found the pointer axle supported
in a bearing bushing (photo no. 55 pos. E). mounted on a Ushaped
bracket (photo no. 55 pos. F). The fork-shaped pointer axle arm
(photo no. 56 pos. C) is connected to a spigot attached to a cir-
cular plate mounted on the rotor housing.
The calibrating spring and the damping cylinder are attached
to the frame as shown on photo no. 56.
To the rear part of the frame are attached suppressor
condensers, suppressor coils and the two-pole receptacle (photo
no. 57 pos. A-B-C).
The rotor housing ball bearings are supported by spigots-at-
tached to the frame.
The rotor housing consists of two parts cast in light alloy.
To the front part is attached a circular disc on which is
mounted the pointer movement spigot, the connecting arm for the
damping unit, onto which the calibrating spring is anchored.
The rotor housing ball bearings are pressed into recesses
outside the housing.
The rotor (photo no. 59 pos. A) is bell-shaped with
windings and commutator situated inside the bell. Into the space
between the rim of the rotor and the windings is situated a light
alloy cylinder (photo no. 58 pos. C) with two cast-in pole cores
in connection with two semi-circular permanent magnets (photo
no. 58 pos. D). The rotor ball bearings are pressed into recesses
inside the rotor housing.
The instrument frame is covered by a presumably soft-iron
screen (photo no. 55 pos. B) and outside this the normal cover
(photo no. 55 pos. A) slides on the frame from behind.
Electrical Functioning (Wiring diagram dwg. no. 63).
Power for the instrument (27v d-c) is supplied through the two-
pin receptacle on the rear of the instrument. The two leads from
here are connected to a terminal block on the rear of the frame
through a suppressor system consisting of two coils and two
condensers as shown on the wiring diagram.
This terminal block is connected to a terminal block on the
rotor housing through two soft springs (photo no. 57 pos. D)
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and from here the leads continue directly to the two brushes. The
brushes protrude through the three openings in the magnets (photo
no. 58 pos. D), and slide on the vertical surface of the commutator
(photo no. 58, pos. B) and not on the sides as shown on the wiring
diagram.
No rotor speed regulator was included in the instrument.
Tests.
The following tests were carried out on the instrument:
Rotor starting voltage: 22 v.
Consumption at 27 v: 0,13 amp.
Rotor rpm: 1200.
Scale test: The results of this test are shown on dwg no. 63.
Full deflection (against stop) was obtained at:
300 sec clockwise
27 /sec counterclockwise.
0
Full deflection of bank indicator ball was obtained at 10 bank
angle. Weight of the instrument: 900 grs. 25X1
When the pointer, during turns, deflects to the outer marks-the
0
. At deflections to the
bank angle of the aircraft must not exceed 45
second or first mark, the bank angle must not exceed 300 or 150 respec-
tively.
Photo no. 54?
Front view of the Turn and Bank Indicator.
SE CET
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Photo no. 55-
A:
B:
C:
Cover.
Screen.
Instrument
Frame and Rotor Housing.
D:
Bank Indicator (from the rear side).
E:
Pointer Axle Bearing Bushing.
F:
Mounting Bracket.
44.
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Photo no. 56.
Front Part of Instrument Frame.
A: Damping Cylinder.
B: Calibrating Spring.
C: Pointer Arm Fork.
D: Rotor Housing Bearing Spigot.
P]:oto no. 57.
A: Suppressor Condensers.
B: Suppressor Coils.
C: Supply Receptacle.
D: Connecting Springs.
45.
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46.
Photo no. 58
A: Gyro Rotor.
B: Commutator.
C: Magnet Cylinder.
D: Permanent Magnet.
Photo no. 55-
A: Front Glass.
B: Instrument scale.
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0MLt w f
47.
GYRO MAGNETIC COMPASS.
Introduction.
The components of the gyro magnetic compass were the same as
found previously, with the exception of the transmitter (photo
no. 60), which was a new design, but working on the same prin-
ciples as the earlier type.
The transmitter was mounted in the starboard wing at the same
place as earlier with no vibrations damping outside the trans-
mitter.
Description.
The transmitter consisted of the following main parts:
Mounting Flange
Compass Transmitter Housing
Magnet Mounting Spider
Magnet and Damping Unit.
The compass transmitter housing is clamped to the mounting
flange on which is found scale ranging from -120 to +120, used
for A-coefficient compensation during compass swinging.
The compass housing is divided into an upper and lower part,
forming a hermetically sealed unit.
The lower part is mainly a bowl (photo no. 62) with eight
damping springs enclosed in rubber blocks, equally spaced along
the inside.
On the top of the upper part of the transmitter is situated the
compensator, which is of the normal type, consisting of two
pairs of permanent magnets at right angles to each other.
The compensator shafts are locked together as shown on photo no.
66.
Through a small circular aperture in the top, part of the compass
scale is visible.
Two small copper tubes on the top is used presumably for filling
the transmitter with an inert gas.
Inside the upper part is mounted a four legged spider (photo no.
63 pos. A) in eight non-magnetic springs connected to the inside
walls.
Furthermore the spider is connected through 16 smaller springs
to a ring attached inside to the top of the bowl (photo no. 63,
pus. B).
Through a universal joint (photo no. 63 pos. C) a bridge (photo
no. 63, pos. D) is connected to the damping ring (photo no. 64
pos. A), which also serves as magnet support.
The damping ring is made of an al.alloy and carries the magnet
bearing in its centre (photo no. 65 pos. D). This bearing is a
high quality ball bearing designed as shown on dwg, :.o 65
(Scale 5:1).
The magnet system (photo no. In. 21 pos. C) consists of a pair
of relatively big magnets attached to a circular compass scale
(photo no. 63 pos. E).
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Electrical functioning.
The two-pin and three-pin plugs are the same as those found
previously, but the connections inside the transmitter are
slightly different (dw>>g. no. 66).
The leads from the three-pin plug are connected directly to the
circular potentiometer whereas the leads from the two-pin plug
via two sliding contacts on the magnet system are connected to
two brushes sliding on the potentiometer 1800 apart.
The functioning is exactly as on the earlier type.
The magnets are damped by eddy currents generated in the damping
ring when the magnets are rotated.
Tests.
The gyro-magnetic compass was removed from the aircraft and the
interconnections between the components were made according to the
wiring diagram in the first report (dwg. no. 9).
When started, the compass functioned irreproachable.
When the magnets of the compass transmitter was deflected, it re-
turned to the original position within one degree (with vibration).
The friction error was approximately 20. The magnets were aperiodi-
cally damped and after a deflection of 900, the magnets returned to
the original position within 5 secs.
The weight of the transmitter was 2,5 kgs.
Photo no. 60.
Front view of the Gyro Magnetic Compass Transmitter.
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k. I I.4.""'
Photo no. 61 - Compass Bowl, upper part.
Photo no. 62 - Compass Bowl, lower part.
49.
C D "- t
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S t(., K t I
50.
Photo no. 63.
Compass Bowl, upper part.
A: Spider.
B: Spring Mounting Ring.
C: Universal.
D: Bridge.
E: Compass scale.
Photo no. 64.
Compass Bowl, upper part.
A: Damping Ring and
Magnet Support.
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Photo no. 65.
Magnets and Damping Ring.
A:
Damping Ring.
B:
Magnet
Bearing Bushing.
C:
Magnet
System.
D:
Magnet
Ball Bearing.
Photo no. 66.
Compensator unit.
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52.
C. 2. Jet Pipe Thermometer system.
The temperature indicator, the thermocouples and the leads
were of the type as found previously.
The indicator was tested as millivoltmeter with the results
shown on dwg. no. 67, showing practically linear relation
between m. V. and instrument indication between 400 and 900?
C. One of the thermocouples was tested in an electric oven at
temperatures between 90 and 600? C.
The thermo-electric force of the thermocouple was measured by
compensation, and the temperatures were measured by mercury
thermometers.
The values obtained are shown on dwg. no. 68.
These show two reversals of the thermoelectric force, the first
at approx 90? C, the second at approx 260? C. This explains why
no cold junction compensation is necessary as the normal varia-
tions of the cold junction temperature only produces minor varia-
tions in EMF.
As shown on photo no. 15-361 (first report the scale of the
instrument is compressed between 0 and 400
This is explained as the relation between hot junction tempera-
turS and voltage is not linear below a temperature of approx.
350 C.
An analysis (spectral and chemical)of the material of the thermo-
couple gave the following results:
1. 80,7% nickel, 15 % cobalt, 2 % manganese, 1,3 %
silicum, 0,1 % copper, 0,9 % aluminium.
11. 97 % nickel, 1,3 % manganese, 1,1 % silicum, 0,3 %
aluminium, 0,1 % copper.
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%3 CA... M C I
53?
Fuel Contents System.
2. The fuel contents system was the same as found previously,
with the exception that the fuel transmitter top and bottom
stops were positioned on the fixed transmitter arm (photo no.
67.)
The transmitter and indicator was connected according to the
wiring diagram shown in the first report dwg. mo. 5. and the
system was tested with the results shown on dwg. no. 69.
Photo no. 67.
Fuel Transmitter Stop.
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54.
SECTION XIV
FUEL SYSTEM
A. General Description.
No significant changes in tAnkage, layout or operation of
fuel system as described in the earlier reports were found, but
two new points of interest were:
(a) The fuel filler necks of both the main and rear tanks
had a small light-alloy 'cup' containing a paste ("Bluish" in
the main tank, "reddish" in the rear tanks) attached by four or
five inches of thin, flexible cable to the neck flange. It ap-
peared that this 'cup' was permitted to hang inside the gauze
filter of the filler neck when the filler cup was replaced. The
significance of these 'cups'and their paste was not readily ap-
parent since the paste did not appear to react to moisture, and
the 'cups' could only reach the surface of the fuel in full tanks,
but samples of the paste were obtained for chemical analysis.
Rubber anti-spill sock.
Quick release filler cap.
Filler neck flange.
ILight alloy cup
containing paste.
Photo no. 68. Main Tank Filling Point.
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a.) L-%.. 11. L.., i
Photo no. 69. Rear Tanks Filling Point.
Rubber anti-spill sock
Light alloy cup
containing paste
Quick release
access panel
55.
(b) An electrical modification, incorporating a new switch,
was found embodied in the drop-tank jettison circuit. This modi-
fication did not affect the fuel system operation and is fully
described in Part 5 (Electrical System) of this report.
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'JLAO. f F. I
56.
SECTION XV
PERSONAL EQUIPMENT
B. Emergency Oxygen in Parachute.
There was space provision for emergency oxygen pack at the base
of the parachute but no oxygen pack was carried.
C. Parachute.
a w eparacue o ccard.
SPADOCHRON
'I'ypa[{-1
The photos
Rodzaj 1- 64
Ned"
Photo no. 70. Parachute log card.
pb skhuih las umuvNw hips itf uw SN1.
6.Ob
-eDJss
N'. .7.. ss
tor.
11.033 IfS
.,,-,
26.08..
Photo no. 71. Parachute log card.
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Z) tC.Kt 1
D. Oxygen mask.
The photos show the mask.
Photo no. 72. Oxygen mask.
Photo no. 73. Oxygen mask.
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58.
Photo no. 74. Automatic harness release.
Photo no. 75. Landing lamp.
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JC.%.KI.C. 1
59.
Photo no. 76. External nose gear
down lock light.
Photo no. 77. Cockpit.
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60.
Photo no. 78. Cockpit.
Photo no. 79. Cockpit.
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Outer Gimbal Caging Mechanism
Schematic.
Steel Ring
Outer Gimbalrix
Instrument Housing
Dwg, no. 61.
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..j G~.F'% L.
Gyro Horizon Wiring Diagram.
Jri tek ws
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L- 1
Turn and Bank Indicator.
Wiring Diagram
TERMINAL BLOCK F
SUPPRESSOR I 0,5 MF 0,5 VF I SUPPRESSOR
COIL CONDENSERS COIL
4,5 sL
4,5 SL
GYRO
COJ&UTATOR
SOCKET
27 VOLT .DC.
0,13 Amp.
Dingo no. 63.
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Turn and Bank Indicator
Pointer
deflection 6
mm
1 2 3 4 5 6 7 8 pr, sek.
Angular velocity o/sec.
2 mm
2 mm-
Dwg* no. 64.
T
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Compass Transmitter
Magnet Bearing
Dwg, no. 65.
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Compass Transmitter
Wiring Diagram.
3?
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SIX.?.
m.Y.
A
of Jet Pipe Thermometer and EMP.
Indication M.V.
400 5,75
500 11,83
600 18,74
700 25,55
800 32,075
900 38,95
100
300
400
600
700
800 900
25X1
idioation
oC
500
ti_
7 ET
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Dwg* no. 67.
 A&f7.m.Y.
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5,0
4.0
3,0
2,0
1,0
` 0,6
0,2
0,2
140
168
189
216
240
280
302
320
332
342
364
378
390
398
405
415
426
439
449
462
469
480
491
517
530
541
556
579
592
608
Rela on a een erm oup
-0,084
-0,102
-0,100
-0,070
-0,020
+0,08
0,205
0,40
0,48
0,58
0,82
1,01
1922
1,34
1.45
1,64
1,88
2,12
2,31
2,55
2,71
2993
3,16
3,47
3,78
4,09
4,22
4,68
5,22
5,56
x,00
4 500 600
H 4t, otign temp. C.
Dwg. no. 68.
C'ik
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Fuel contents transmitter unit.
At this position
warning light came
on.
Dwg, no. 69.
cFrPFT
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Next 1 Page(s) In Document Denied
Iq
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TECHNICAL REPORT
ON
NO 1327
PART TWO
ENGINE
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PART TWO
C
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PART TWO
ENGINE
The following report is based upon information gained from
examination of a Polish built aircraft (known as "FAGOT"),
desi ated LIM-2 (Polish
Bis)
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SECTION I
TABLE OF CONTENTS
SECTION
II Introduction
III Engine Data
A. General
B. Weights
C. Centre of Gravity
D. Fuel Specification
E. Oil Specification
V Installation Notes
A. Dimensions
B. Engine Mounting Attachment Points
C. Jet Pipe Attachment and Support
D. Engine Slinging
E. Installation Connections Between
Engine and Airframe
1. General
2. Fuel System
3. Oil System
4. Starting System
5. Miscellaneous
6. Auxiliary Gearbox Connections
VI Engine Description
A. Wheelcase
B. Compressor Section
1. Air Intakes
2. Compressor Casing
3. Compressor Assembly
Page
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Page
C.
Combustion Assembly 12
1.
Combustion Chamber Cover
2.
(Expansion Chamber)
Air Casing
12
3.
Combustion Chamber Liner
13
Flame Tube)
D. Nozzle Box and Turbine Assembly
1. Nozzle Box
2. Nozzle Guide Vanes
3. Turbine Assembly
\)
E. Exhaust Duct
F. Engine Lubrication System
G. Engine Fuel System
1. High Pressure Fuel Pump
2. Throttle Valve and Idling Control Unit
3. Pressurizing Valve, High Pressure Cock
and Acceleration Control Unit
a. Pressurizing Valve and High
Pressure Cock
b. Acceleration Control
H. Starting System
1. General
2. Description
a. External Supply Socket
b. Starting Panel
c. Starter Motor
d. Booster Coils
e. Torch Igniter
f. Torch Igniter Feed Pump
3. Normal Starting Procedure
4. Relight in the Air
I. Centre Bearing Casing
J. Rear Bearing Housing
K. Cooling Fan and Casing
L. Central Coupling
VII Jet Pipe Assembly._
14
15
15
SCr -PF,
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VIII Auxiliary Gearbox
A. Generator
IX Engine Controls and Instruments
A. Engine Controls and Switches
B. Warning and Signal Lights
C. Engine Instruments
X Material Analyses
A. Compressor Casing
P age
15
15
16
16
17
19
B. Combustion Chamber Inner Liner 20
(Flame Tube
C. Turbine Disc 20
D. Turbine Blade 19
1. Spectrographic and Chemical Examination 20
2. Microscopic Examination 20
3. Hardness test 20
E. Nozzle Guide Vane 21
1. Spectrographic and Chemical Examination 21
2. Microscopic Examination 21
3. Hardness test 21
F. Acceleration Control Unit Filter
XI Miscellaneous 23
Fire Extinguisher System 23
XII Glossary of Photographs
Appendix of Drawings
25
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SECTION II
INTRODUCTION
The engine installed in this aircraft was, with few exception
identical to those reported in the previous) re-
ports.
? Photograph 1. LIS-2 Engine - General View.
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L- %-0 1 %&
Photograph 2. Engine Dataplate.
SECTION III
]ENGINE DATA
Designation of engine..... ..... e .........LIS-2
D. Fuel Specification
A sample of the fuel obtained from the aircraft was taken and
forwarded for analysis.
E. Oil Specification
Samples of the engine lubricating oil were submitted for analysis.
Photograph 3. LIS-2 Engine - Left Side View.
5.
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SECTION V
INSTALLATION NOTES
6.
A. Dimensions
The propelling nozzle diameter (in millimetres) was found etched on
the nozzle as fj 542.4.
B. Engine Mounting Attachment Points
The method of engine mounting was identical to the previous cases,
except for slight dimensional differences in lengths of supporting
members. the upper
supporting strut in this case was measured to be 31.25". The inter-
mediate support was 41.0" and the lower support 36.5" in length,
center to center.
The vertical distances between mounting points on both engine and
fuselage were found to agree with those previously measured i.e.
16.25" and 41.0" respectively.
Photograph 4. LIS-2 Engine - Attachment Points.
EGRET
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Photograph 5. LIS-2 Engine - Installation Details.
Photograph 6. LIS-2 Engine - Installation Details.
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u * ~~._ ~~ L Sir
y y/ 0
rbs Y
9 la
-6w ~ 6
d~41i~~y r Photograph 7 LIS-2 Engine Installation Details.
YS A
~ `` 'y yl ,~ ~< m
i ... 5~haan"w
Photograph 8 LIS-2 Engine - Installation Details.-
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9?
Photograph 9 Fuselage at Frame 13.
Photograph 10 Fuselage at Frame 13.
A. Auxiliary Gearbox Drain Valve.
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SECTION VI
ENGINE DESCRIPTION
A. Wheelcase
The breather on this engine was located in the central pad on the
front face of the wheelcase
The breather tube
on this aircraft.
was not present
Photograph 11. LIS-2 Engine - Front View
A. Engine Data Plate
B. Breather
C. Push-button Oil Level Valve.
mow/' I:w.. L'J
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v`. w.'^ %6v. .
B. Compressor Section
1. Air Intakes
were found, on
as being overall dimensions, 25X1
this engine, to be the dimensions of the open
mesh, i.e. not to include the sheet metal rims. This difference
is not thought to indicate a change in design but is considered
to be due to error in the production of those earlier drawings.
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C. Combustion Assembly
2. Air casing
Although the dimensions of the air casing remained the same
as previously, a different fabrication technique was employed.
Instead of the helical weld construction of the previous cans,
these air casings were made of two sections joined by a circum-
ferential seam. Each of the two sections was apparently formed
by deep drawing from a single piece. The seam joining the two
sections was overlapping, spot welded, and brazed to seal the
joint. The front flange on the casing was also attached by this
method, i.e. spot welded and brazed. The rear flange was attached
by means of a continous resistance weld.
Photograph 14. Air Casing and Flame Tube
A. Single-piece conical section of air casing.
B. Dimpled keyhole slots.
C. Reinforcing eyelets.
D. Flanged holes.
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3?
Flame Tube
The following differences were noted on flame tube construction
(see photograph 14):
a. The edges of the holes in the rear two rows were turned inwards
to form a stiffening flange.
b. Four rows of holes were eyelet reinforced similar to rows 3 and
4 of the previously observed engines.
c. The keyhole slots in the front conical section were dimpled
slightly inward at the rounded end of the slot.
Photograph 15.
Flame Tube Rear
End.
E. Exhaust Duct
The exhaust cone assembly was found to have slightly different
dimensions
The differences may be attributable to measuring techniques, but
a new drawing showing the currently measured dimensions is shown
in Dwg. no. 16.
Photograph 16.
Exhaust Duct Assembly.
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14.
F. Engine Lubrication System
A spring-loaded valve (push-to-open) was installed at the base
of the oil filler neck to permit easy draining of surplus oil.
(See photograph 11). This valve was also installed on the earlier
The pressure oil filters in this LIS-2 were identical to those 25X1
previous reported
Photograph 17.
Pressure Oil
Filter Element
Photograph 18.
Pressure Oil
Filter Element
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Stt..KI
H. Starting System
1. General
la redesigned
starter panel and an improved starter motor were fitted.
These two components are fully described in part 5 (Electrical
System) of this report.
SECTION VIII
AUXILIARY GEARBOX
A spring-loaded valve (push-to-open) was fitted in a pipeline
leading from the auxiliary gearbox to an overboard drain point.
Operation of this valve ensures against overfilling by draining 25X1
off any surplus.
A. Generator.
See part 5 (Electrical System) of this report.
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SECTION IX
ENGINE INSTRUMENTS AND CONTROLS
16.
A. Engine Controls and Switches
The only point of difference in engine controls and switches
between this aircraft and those previously reported on, was that
the throttle locking or friction lever had the reverse method of
operation to that described on page 95 in Section IX of the April
1953 report, (and shovin in photograph 15-79 on the same page).
In this aircraft the operation was:
"Down" for freedom.
"Up" for locked.
Photograph 19. LIM-2. Cockpit Layout
A. Throttle lever friction
control (In "free" position).
B. Engine starter button on throttle lever.
C. Throttle lever in closed position.
D. R.P.M. Indicator.
E. Jet Pipe Temperature gauge.
F. Engine gauge unit.
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17.
C. Engine Instruments
The only significant points of difference between the engine
instrument presentation in this aircraft and those previously
reported, lay in the colours painted on the instrument bezels.
The functioning of all instruments is fully described in Part
I of this report, and the position in the cockpit of engine
gauges is as described in earlier reports and as indicated in
photograph 19.
1. Engine Gauge Unit
This compound gauge has three indicating dials and their
ranges and colour bands are as follows:
a.
Top Main Scale - Pilot Fuel Nozzle Pressure
Reading - kg/cm2
Range - 0 to 100
Markings - 0 to 56 Blue
56 to 100 Red
b.
Left Bottom Scale - Oil Pressure
Reading -
Range -
kg/cm2
0 to 10
Markings -
0 to 2
Yellow
5 to 10
Red
c. Right Bottom Scale -Oil Temperature
Reading -
?Centigrade
Range -
-50 to +150
Markings -
-50 to +100
100 to 150
Blue
Red
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2. Jet Pipe Temparature Gauge
Reading - 0Centigrade
Range - 40 to 900
Markings - 0 to 680 Blue
680 to 700 Yellow
700 to 900 Red
3. RPM Indicator
Reading - RPM x 1000
Range - 0 to 15
Markings - 5 to 11.2 Blue
11.2 to 11.6 Yellow
11.6 to 15 Red
In addition, on this instrument, an inner ring of blue
extended from 11.4 to 11.7
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19.
SECTION X
MATERIAL ANALYSES
B. Combustion Chamber Inner Liners (Flame Tube
Approximate composition found:
Chromium ..... 2
Iron......... 0
Titanium.....a little
Nickel, .... remainder
Type of material: A Nimonic type, probably Nimonic 75-
C. Turbine Disc,
The turbine disc is made from a magnetic steel with the follo-
wing approximate composition:
Chromium.....3
Manganese....
1%
Vanadium.....
1%
Molybdenum...
-%
Carbon.......
0,24%
Sulphur......
0,024%
Iron.........
remainder
D. Turbine Blade.
Photograph 20.
Turbine Blade x 75
Diamond-polished, etched
with R1T03-HC1.
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%J r% I
Photograph 21.
Turbine Blade x 400
Polished and etched
as in photograph 20)
Photograph 22.
Turbine Blade x 750
Polished and etched
as in photograph 20)
1. Spectrographic and Chemical Examination
Approximate composition:
Nickel .............. 75%
Chromium .............23%
Iron ................ 1-Y
Titanium more than 1% (Similar to Nimonic BOA)
Manganese a little
Nickel .............. remainder
2. Microscopic examination
As it is to be seen from photograph 20 the grain size is very
irregular. From photographs 21 and 22 the second phase will
appear.
3. Hardness test: 402 Vickers, 500 grammes load.
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JG~r~G ,
E. Nozzle Guide Vane.
1. Spectrographic and Chemical Examination
Approximate Composition:
Iron............ 31%
Chromium........ 2c%
Tungsten........ W"
Silicon......... 115111
Titanium .....>... a little
Nickel.......... remainder
2. Microscopic examination
As is to be seen from photograph 23, the guide vanes show a
casting structure with coarse dendritic formations. By greater
enlargment, see photographs 24 and 25, it is to be seen that
the mixed carbides between the primary dendrites have been more
or less dissolved, as it seems that the material has been exposed
to a kind of solution heat-treatment with a following re-segre-
gation of the carbides into fine particles.
Therefore it is most probable that after casting the guide vanes
were heat-treated in order to achieve a better relative strength.
3. Hardness: 276 Vickers, 500 g load.
Photograph 23.
Nozzle Guide Vane x 75
Electro-polished and
etched in Knuth-electro-
lyte A2).
SEC R, F--- -r
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V L'_0 1 &.. I
Photograph 24 Nozzle Guide Vane x 7~0
(Polished and etched as in photograph 23)
Photograph 25 Nozzle Guide Vane x 1500
(Polished and etched as in photograph 23)
SFC P1-T
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23.
SECTION XI
MISCELLANEOUS
Fire Extinguisher System
The only difference in this system from those previously reported
lay in the steel C02 bottles which were marked in Polish, and whose
weights and capacities varied slightly from those marked in Russian
on the earlier MIG aircraft.
A translation of the Polish wording is:
Weight of cylinder
Capacity
Quantity of C02
Bottle 1 Bottle 2
4408 g. 4124 g.
2914 cm3 2955 cm3
1981 g. 2008 g.
Photograph 26. Engine Fire Extinguisher Bottles.
S F t" F T
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24..
25X1
the exploded view in photograph 27 is in-
cluded to clarify its build-up and operation.
Photograph 27. Fire Extinguisher Bottle Head.
A. Outlet elbow to spray pipes.
B. Circumferential groove.
C. Hollow diaphram-piercing knife.
D. Spring-loaded ball for restraining
free movement of knife.
E. Body of head fitting.
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E.J. fem.... 's %-
SECTION XII
GLOSSARY OF PHOTOGRAPHS - APPMrDIX OF DRAWINGS.
Photograph 28 LIS-2 Installation
Photograph 29 LIS-2 Installation
25.
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btLKt I
26.
Photograph 30
Installation Top View
Photograph 31
Nozzle Box Assembly
Photograph 32
Installation Top
Right View
SFc PFT
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S tL F I
%_ Photograph 33 LIS-2 Engine.
Photograph 34 LIS-2 Engine.
SE C R F ET
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V L..N-0 UN L.. I
Photograph 35
Wheelcase -
Rear View
Photograph 36
i?Theelcase -
Left side View.
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SF P T
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r... ,. ., ,..,_
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 I
TECHNICAL REPORT
ON
NO 1327
PART THREE
ELECTRONICS
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PART THREE
ELECTRONICS
S RET
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.7 fw.'., r% L. I
TABLE OF CONTENTS
SECTION
I. Foreword and List of Contents.
II. General Summary.
A. General
B. Electronic Installation.
C. Installation and Servicing.
D. Aircraft-Electronic Cabling.
III. Command Transmitter Receiver RSI-U-3M.
A. General Description.
B. Technical Data.
C. Circuitry and Cabling.
D. Operating Procedures.
E. Comments.
IV. Radio Altimeter RV2.
A. General Description.
B. Technical Data.
C. Circuitry.
D. Operating Procedure.
E. Comments.
V. Radio Compass ARK 5-
A. General Description.
B. Technical Data.
C. Circuitry.
D. Comments.
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Marker Beacon MRP-48-P.
A. General Description.
B. Technical Data.
C. Circuitry.
D. Operating Procedure.
E. Comments.
VII. Cabling and Connectors.
VIII. Aircraft-Radio Servicing.
A. General Description.
B. Servicing Problems.
C. Comments.
IX. Spare Utilization.
X. Pilots Interrogation.
A. Approach and Landing Systems.
B. Ground Radars.
C. Control Post S.D.
D. Servicability of Radio Equipment.
E. General.
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LIST OF DRA7INGS
III. No. 1 VHF Set RSI-U-3M, Circuit Diagram of VHF Trans-
mitter.
No. 2 VHF Set RSI-U-3M, Circuit Diagram of VHF Receiver.
No. 3 VHF Set RSI-U-3M, Interconnection Wiring Diagram.
No. 4 VHF Set RSI-U-3M, Circuit Diagram of Control Box.
No. 5 VHF Set RSI-U-3M, Cable Diagram.
No. 6 VHF Set RSI-U-3M, Circuit Diagram of VHF Power
Supply.
No. 7 VHF Set RSI-U-3M, Circuit Diagram of Inverter Unit
Type MA-100M.
No. 7A VHF Set RSI-U-3M, Tube Base Connections.
IV. No. 8 FM Radio Altimeter RV 2.
No. 9 FM Radio Altimeter Dynamotor Diagram.
No. 10 FM Radio Altimeter RV 2 Altitude Indicator.
V. No. 11 Radio Compass ARK-5, Inverter Type MA-250 M.
R, E 'IF
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PART THREE
ELECTRONICS
SECTION I
FOREWORD)
The following report is based upon information gained from examination
of a Polish built aircraft (known as "FAGOT" designated LIM-2
(Polish designation for the USSR IVIG-15 BIS
The full report has been divided into six seperate publications as
follows:
Part One - Airframe 25X1
Part Two - Engines
Part Three - Radio and Navigation
Part Four - Armament
Part Five - Electrical
Part Six - Manufacturing Methods
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SECTION II
GETNERAL SU! ARY
A. General
1. The radio installation of the MIG-15-BIS (Polish designation
LIM-2) No. 1327, consist of:
(a) VHF Command Transmitter Receiver Type RSI-U-3M.
(b) Radio Altimeter Type RV 2.
(c) Radio Compass Type ARK 5-
(d) Marker Beacon Receiver Type MRP-48-P.
2. All the electronic equipment was of Russian manufacture except
the Marker Beacon receiver, which was Polish. There is evidence
of the modification of some equipment e.g. Radio Altimeter,
since it was exploited in 1953. The equipment had differing dates
of manufacture, the latest being some time in August 1955-
3. Available space throughout the aircraft appeared to be
extensively utilized. However, more space could be made available
in the nose compartment, if the accumulator and oxygen bottle be
repositioned elsewhere, this could allow the installation of a
simple gun ranging radar in addition to the present equipment.
B. Electronic Installations.
1. Command Transmitter Receiver RSI-U-3M. This equipment is a four
channel crystal controlled VHF set, operating in the frequency
range 100 to 150 MCS. The installation comprises separate
transmitter and receiver units, and seperate power, inverter,
and control units. It uses a sword type antenna. The power
supply, obtained from the aircraft supply, is converted to 115
volt 400 cps AC in the inverter and is rectified by the power
unit to provide the various voltages required by the trans-
mitter and receiver. Channel selection is by the push button
method on a control box. The power output of the transmitter is
approximately 5 watts.
2. Radio Altimeter RV 2. This installation was similar to that
exploited in 1953 except for some minor modifications and
the inclusion of an additional filter unit in the LT supply.
3. Radio Compass ARK 5. This installation was similar to that
exploited in 1953 except for minor modifications including
the redesign of the filter units in the dynamotor.
4. Marker Beacon MRP-48-P. This installation was similar to
that exploited in 1953, but was of Polish manufacture.
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7.
C. Installation and Servicing.
1. The radio equipment in the MIG-15-BIS is installed in four
regions of the aircraft:
(a Underneath the rear fuel tank in the last section,
(b between the bulkheads at airframe stations 8 and 9 at the
rear of the gun bay,
(c) within the cockpit,
(d) in the nose compartment.
2. Apart from that installed in the nose compartment most of
the equipment was most inaccessible. Each equipment, including
sub-assemblies, was sealed with mounting fasteners locked in
position by wire; in most cases this had obviously been done
either at the factory or on the production line. In addition
the VHF receiver tuning and crystal access cover was wired in
position and all bulkhead plug and socket connections were
secured in positions by locking wire metal seals.
In one case, the VHF inverter, all instruments on the centre
instrument panel had to be removed before the unit could be
taken from the aircraft. It is unlikely that the installations
were intended to be removed for servicing.
D. Aircraft Electronic Cabling.
1. Connectors are clamped together using polythene straps and
a collar stud fastener, but no attempt was made to bond these
cables to the airframe. In many cases cables were much too long,
the slack being taken up by doubling the cable back on itself.
2. At least four different types of plug and socket are in use in
the electronic installations.
3. The negative low voltage power supply connection in all cases
was via the airframe, a single wire system being in use
throughout the aircraft except on armament installations, where
a two wire system was used.
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V L.V is I
SECTION III
COMMAND TRANSMITTER AND RECEIVER RSI-U-3M(PCi-Y-3M).
A. General Description.
1. The command transmitter and receiver set, RSI-U-3M was a
Soviet-built airborne VHF set, operating in the frequency
range of 100-150 mcs. It has four preset crystalcontrolled
channels. Channel selection was accomplished in a similar
manner to that in the American set, SCR-522, which probably
had been the model which inspired the RSI-U-311, although
the Soviet equipment differed completely in many respects
from the SCR-522.
2. The RSI-U-3M consisted of the following units:
(a) Transmitter RSI-U-3M(PC1 -Y-3M) (Fig.3-2), weight 9,5 kg
(b Receiver RSI-U-3M(PCl - Y-3M)(Fig. 3-1), weight 11,3 kg
(c Power Supply Type VM(BM) (Fig. 3-3), weight 5,8 kg
(d Inverter MA-100-M(MA-100-M) (Fig. 3-4)9 weight 7,1 kg
(e) Control Box Type P(TVtn - n ) (Fig. 3-5)
(f) VHF Stub Antenna (Fig. 3-6), weight 1,3 kg
(g) Mounting Tray for Transmitter and Receiver, weight 1,3 kg.
(Fig. 3-1)
Receiver RSI-U-3M.
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(Fig. 3-3)
Power Supply Type VM
9?
(Fig. 3-2)
Transmitter
RSI-U-3M.
(Fig. 3-4)
Inverter Type MA 100 M.
(Fig. 3-5)
Control Box
Type P
(Fig. 3-6)
VHF Stub Antenna.
SF'CPFT
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3. The receiver and the transmitter were each housed in separate
boxes, back to back, port to starboard, on a tray in the
nose compartment between the cockpit bulkhead and the battery.
The power supply was installed on the port side of the cockpit
bulkhead and the battery. The power supply was installed on
the port side of the cockpit, just aft and to the left of the
pilot's left arm.In previously investigated "Fagots" this area
was utilized for the RSI-6-M-1 receiver.
The inverter was installed behind the central instrument panel
and the central instrument complex leads had to be removed, one
by one, to permit removal of this inverter. The control box was
mounted on the port side of the cockpit, easily accessible to
the pilot's left hand. The VHF antenna, a quarter-wave sword,
was mounted just starboard of the top centerline of the aircraft,
immediately aft of the cockpit canopy.
4. The face of the control box we-s equipped with four push buttons
for the four channels marked 1,2,3 and 4; a volume control and
a switch for selecting the audio output of either the radio
compass or the VHF. A push button on the end of the box permit-
ted the disengaging of any channel selector button engaged at
the time. The channel selector windows were equipped with a
luminous slide, identifying the channel selected.
5. The VHF set RSI-U-3M was capable of transmitting and receiving
amplitude modulated signals in the frequency band 100-150 mcs.
No facility existed for transmitting MCW signals.
6. The RSI-U-3M appeared remarkably clean and new inside. Mar-
kings on resistors and capacitors were easily discernable and
Fig. 3-7.
RSI-U-3M Transmitter and Receiver Installation.
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JtI.KGI
11~
VHF Power Supply Port Side Cockpit.
VHF Inverter Behind Instrument Panel.
IV
VHF Controller Port Side Cockpit.
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V
showed no signs of being discoloured or burned through use.
The interiors were free of dust and chassis were not corroded.
B. Technical Data.
1. Transmitter RSI-U-3M (See Fig. 3-11 below).
Transmitter RSI-U-3M consisted of an oscillator (/1 -101,
6 1) 6C), which served as a frequency doubler as well and
fed a tripler (/1-102, 6 / 6C) which, in turn, fed another
tripler (f} -103, 1 Y -32), push-pull amplifier stage, making
a frequency multiplication of eighteen in all. At this point,
the signal was introduced into the push-pull power amplifier
( /)-104), pY -32). The RF output went to the antenna switching
relay P-101, which changed over the antenna between receiver
and transmitter. A small portion of the RF output was fed to a
pentode ( /) -105, 6 ( 311), wired as a diode detector, and the
detected signal was fed to the receiver audio amplifier to provide
sidetone in the pilot's headphones. The audio section of the
transmitter consisted of a speech amplifier (/1 -151, 6 r 2). which
amplified the microphone output and, in turn, fed the push-pull
modulators (I) -1529/7 -153; 6 /7 6 C). No facility existed for
using the audio part of the transmitter as an intercom system.
Channel switching was accomplished with.a ratched motor pushing
4 slides which turned 3 tuning shafts. The tuning shafts were
locked to their frequency settings by means of a locking knob
with an indicating dial calibrated from 100 to 150 mcs.
2. Receiver RSI-U-3M (See Fig. 3-12 overleaf).
Receiver RSI-U-3M was a 4 channel, 13 tube crystal controlled
superheterodyne receiver with one RF stage and a 12 megacycle
intermediate frequency (IF). The crystal frequency was deter-
mined as follows:
f - f -12
crystal - 18
mcs = 18 -0.6667) mcs
The oscillator functioned as a tripler and fed another tripler.
The RF signal and the 9th harmonic were mixed in the first mixer;
the second mixer combined the outputs of the first mixer and the
ninth harmonic of the crystal giving an output of:
(f-9 fcrystal ) -9 f crystal = 12 mcs
The signal was then amplified in a 3 stage IF amplifier and
was detected in a diode detector, after which it was
amplified in an audio amplifier, which, in turn, fed the output
tube. The receiver was equipped with a squelch circuit, making
the receiver "dead" when no signal was being received. The
AVC was amplified in an AVC amplifier, which was applied to
the RF, lst mixer, 1st IF and 2nd stages.
SEA:RFr
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L 4 r . t
? 1
to antenna
0 103
4 crystal
f-l8xf
crystal
oscillator
and double
6 J1 6 C
tripler
6J16C
tripler
rY-32
J1-101 Ji-102 J1-103
antenna
relay
4'105
to receiver
microphone_ 3- -
audio
amplifier
6r2
power am-
plifier
r Y-32
J1-104
modulator
2x6116 C
J1-151 J1-152
.1-153
monitor
6*3J1
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 \1/ 0 205 to antaasa
amplifier
2. IF
Detector
diode
r--
n -201
n -2o
6R4
4ztala
Oscillator
Tripler
1. miter
2. ^izer
Squelch
tf=12
and tripler
11-205
11-202
n -205
diode
:tali 18
n -204
Audio
eaplifisr
n-209
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t A .
AVC
diodn-;13
Audio
output
f1-210
Squelch
oscillator
551,10/8
n -211
Squelch
rectifying
diode
n-211
'wing
diode
AVC
amplifier
6r6G 612 6Y2 6*6C 6r2
t
t1:J--- ---------------------------------------------------------------------- J
(n
m
n
m
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15.
3. Inverter MA-100M.
Inverter MA-100M was a rotary inverter fed with 26 volts DC and
producing 115 volts AC, 400 cycles.
4. Power Supply BM.
Power supply BM was a rectifier producing high voltage for the
receiver and transmitter and negative bias for the transmitter.
5. VHF Antenna.
The antenna for the VHF set was of the wavelength ground-plane
type, mounted on the upper starboard side of the aircraft fuselage.
It was of the sword type, solid and swept back at an angle of 60
degrees with the horizontal.
It was composed of an aluminium alloy, type 24-S.Spectroscopic
test revealed it to be approximately 9c% aluminium, 5% copper,
2f magnesium, 1% manganese and its surface was anodized. It weighed
1.33 kg and was of sturdy construction. (See Fig. 3-13 for analysis
report and Fig. 3-13-2 to 5 for construction).
The antenna was removed and mounted on an insulator in a similar
manner to that on the aircraft, then mounted on a 1 m2 aluminium
plate, which was horizontally placed 2 metres above ground. The an-
tenna was then fed through a 7 meter RG-8 cable, and the standing
wave ratio of the entire system was measured with a 50 ohm reflec-
tion coefficient meter, type 136 B manufactured by the American
Sierra firm.
The tabular and graphical results of this test are included herein.
(See Figs. 3-14 and 3-15 overleaf). They reveal the VHF antenna to
be broadbanded from approximately 120 to 220 mcs with standing
wave ratio rising rapidly beyond the frequency limits stated. It
appears that the physical length of the antenna is a little too short
for the 100 to 150 me band. No means of tuning the antenna were in-
cluded on the aircraft.
V
Test of: Antenna.
for examination of: Type of Material and Surface Treatment.
Results of test:
Material: Aluminium Alloy.
Analysis of Alloy components:
Iron......... ca. 0.2%
Copper....... ca. 5%
Chromium...... Traces
Magnesium.... ca. 2l
Manganese.... ca. 1%
Titanium..... 0.3%
Zinc......... ca. 0.2%
(Fig. 3-13-1)
Analysis Report on VHF Antenna
SE RE -r
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Fig. 3-13-2
VHF Antenna
? Fig. 3-13-3
Base Mounting VHF Antenna
Fig. 3-13-4
Base Mounting VHF Antenna Removed
Fig. 3-13-5
VHF Antenna Central Connecting
Pin Removed
16.
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17.
DATA SHEET
MEASURE OF STANDING WAVE RATIO OF VHF ANTENNA.
RSI-U-3M VHF SET
it
FREQUENCY
SWR
FREQUENCY
SWR
(mcs)
(mcs)
75
20 (approx.)
165
299
80
14 (approx.)
170
3,2
85
15 (approx.)
175
3,0
9o
11
180
2,6
95
10
185
3,0
100
9
190
3,0
105
6
195
2,7
110
5
200
3,0
115
4,1
210
2,9
120
3,1
220
3,5
125
3,6
230
393
130
3,0
240
3,6
135
2,8
250
4,1
140
3,1
260
4,6
145
2,8
270
5,5
150
2,8
280
4,5
155
3,0
2 90
7
160
2,9
300
5
WEIGHT = 1,33 kg.
C
D
7
T ALL D EV NS IONS IN CM
UNLESS OTHERWISE NOTED
Fig. 3-14.
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I
VHF SET RSI-U-3M.
SWR. FOR VHF ANTENNA.
to N
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Standing Wave Ratio.
Fig. 3-15.
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19.
6. Crystals.
a. Findings.-
Both the transmitter and the receiver had sockets for 4
crystals each and external boxes for storage of 12 spare
crystals each. Crystals for each of the four operating
channels were included in the set but the spare crystal
storage boxes were empty except for a crystal inventory
list which contained a listing of 16 crystals (4 for the
operating channels and 12 spares) (See Fig. 3-16).
The 4 crystals in the transmitter
CnKCOK Kaapgea
Al
1
61
A16
2
616
A25
3
625
A61
4
661
A76
5
676
A172
6
6172
A301
7
6301
S
A313
S
9
A337
_
9
6x37
10
A349
10
i 6341l
lI
A361
11
12
A376
12
1i:17i
4
13
A460
13
G 1111
.
14
14
li181
15
A50S
1S
listls
16
A520
16
65211
1.%ww/nVP 0
(Fig. 3-16)
Crystal Inventory List
Transmitter
Crystal
Frequency(KCS)
Receiver Frequency(KCS)
Crystal
A
16
5625.2
B 1
6
4958.5
A
301
6944.8
B 30
1
6277.2
A
361
7222.6
B 36
1
6555.3
A
460
7681.1
B 46
0
7013.9
All crystals were mounted in identical heavy bakelite
holders, fitted with 2 banana pins. Measurements were
made under controlled temperature conditions over the
range from -20? to 70?C, with test points at every 3
degrees. The oscillator used was a Quartz Activity Test
Set QC57A (also marked type BW 670) built by the British
firm G.E.C... The instrument measured directly the parallel
impedance of the crystal with a capacitive load of 30 mmf.
The crystal frequency was measured with a Hewlett-Packard
Electronic Counter, Model 524 B. The counter was connected
to the test set QC 57A and was calibrated against the NPL
200 KC standard frequency from the Droitwich station.
In this way an accuracy of 2 parts per million was expected.
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20,
The results are given in the Graphs in Figs. 3 17 - 3 24
inclusive.
The Test Set QC 57A specification cards required a minimum
parallel impedance of:
16.5 Kohms for 5 mcs crystals
14.5 Kohms for 6 mcs crystals
12 Kohms for 7 mcs crystals
9.3 Kohms for 8 mcs crystals
b. Speculations:
The graph (Fig. 3-25) has the frequencies from 100 to
150 mcs plotted along the ordinate and the A numbers from
0 to 600 plotted along the abscissa. The four known A crystal
numbers are plotted in the system and lie along a straight
line which intercepts the 100 and 150 mcs points.
The resulting frequency distance between successive A
numbers is revealed to be 83.33 KC, providing 600 available
channels.
The crystal inventory card, Fig. 3-16, reveals that only
channel numbers which correspond to multiples of 250
KC (3 x 83,33 KC) were included in the spare crystal kit.
Fig. 3-16-1
Transmitter Crystals.
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VHF SET RSI-U-3M.
QUARTZ CRYSTAL NO.A.16.
Parallelimpedance in K
0
0
r-{ 0 0-11 x L - \O
N
LCD
U
0
o
0
0
N
N
LO
LCl
Frequency in kc.
Fig. 3-17-
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LCl
o
p.
VHF SET RS-I-U-3M.
QUARTZ CRYSTAL NO. A.301.
Parallelimpedance in K - L
N
ti
0
0
rn
0
0
0
6l
\_0
U
0
Frequency in kc.
Fig. 3-18-
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VHF SET RSI-U-3M.
QUARTZ CRYSTAL NO.A. 361.
Parallelimpedance in K JQ-
Frequency in kc.
Fig. 3-19.
r
I
j.r
It
j
+
,
I1
f
31;
+
r
;y
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7T~7
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.
77
1;z
-
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=FF
0
al
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J Gd, r-r L. I
VHF SET RSI-U-3M.
QUARTZ CRYSTAL NO. A.460.
Parallelimpedance in K _"!_
+t
}} 1 1.Y H }} t
Lx. ::I Fol. t!_ r a 1 1...- + ~*.
t- tt,
4~ --tit
.44
+. L
r-.
W4- T
.31111
N H
lCl
rn
Frequency in kc.
Fig. 3-20.
0
0
N rI 0
0
CD
c
CD
0
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.,L-%..I %%- I
VHF SET RSI-U-3M.
QUARTZ CRYSTAL NO. B.16.
Parallel impedance in K JL
O \10
N N
25-
U
0
T,
A
i
-
S
4- ~4
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r
It
i
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7
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7-
A.
.77
77:
0 0
N r i 0 Ol OD L- \D ~Il
n .~ w w . w w
00
LrN
Frequency in kc.
? Fig. 3-21.
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V`N.. I.%- I
VHF SET RSI-U-3M.
QUARTZ CRYSTAL NO. B.301.
Parallelimpedance in K _t-
`O d- N 0 cc d N O co ~O N
K1 K\ K\ K1 N N N N N r--?I r--I H H
^
.+
+
_.
1
,7
_Y
IF
4
I
f
r
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,141"
4
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+
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.fem.
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-JI
a
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4
1
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4 L
11
4 '
I
f_
r
:i
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n
i
i
!
11
9`
41
-4
-7
17 .
.
I
77
C--
C-
N
K\ N H
w
C-
N
Frequency in kc.
Fig. 3-22.
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26.
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VHF SET RS I--U-3P;i.
QUARTZ CRYSTAL NO. B.460.
Parallelimpedance in K L
rn
H
C='
7: d
777
V
ti
Frequency in kc.
Fig. 3-23.
0
0
\.o L-h zr rn
0
0
0
0
N
U
0
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V }R SET RSI-U-312.
QUARTZ CRYSTAL RD. B.361.
1 -4 4:
1: 7:
- .J.
41;. 44
t L.
Para.llelimpedance in K _(1-
..
i. ~.
r} i
N-H
. ..
- ---- ----- --
-777
0
0
o 3
LI\ n
Lfl
LIl
N ~7
O
0
LI)
LC\
L \
Frequency in kc.
Fig. 3-24.
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30.
C. Circuitry and Cabling.
1. Transmitter RSI-U-3M (see circuit diagram, drawing No.l)
a. Oscillator ( 17-101, 6/16C)
Tube 17-101 (6 )7 6c) was the oscillator tube in the crystal
oscillator circuit. The anode circuit was tuned to the
second harmonic of the crystal frequency. Tuning of the
anode circuit was accomplished with a variable condenser
which was ganged with the tuning condensor in the anode
circuit of tube 17-102, the first tripler stage. The cold
side of the grid resistor (33 K ohms) was grounded by the
antenna switching relay, P-101, when in the transmit po-
sition. In the receive position, negative grid bias was
applied to prevent oscillation.
b. First Tripler ( /7-102, 61/60)
Tube J] -102 functioned as a tripler. The cold side of the
grid resistor (56 K ohm) was grounded as was that of the
oscillator during transmitting and negative bias was applied
during receiving by the antenna switching relay, in order
to cut-off the tube.
c. Second Tripler (/j -103, /" Y32)
Tube 17-103 operated as a push-pull tripler and was the
Soviet type /")'` -32, similar to the American type 832 tube
used in the SCR-522. The anode circuit was tuned to the 18th
harmonic of th.e crystal frequency by means of the variable
condensor, C-118 which was set by means of its own tuning
shaft and tuning knob. Grid bias was applied to this tube
from the negative grid bias supply produced in the power
supply unit and applied to pin 3, plug 104 on the transmit-
ter.
d. Power Amplifier ( J7 -104, 1' Y-32)
Tube /7-104 functioned as a push-pull power amplifier and
was the Soviet /I Y-32, equivalent to the American type 832
tube as was the second tripler. Here, as in the second
tripler stage, the anode circuit was tuned with a variable,
condenser which was set by means of its own tuning shaft and
tuning knob. Coupling and the output was fed to the antenna
through the antenna switching relay, P-101.
e. Monitor ( /1 -105, 6 * 3)
The monitor tube /7-105 was a pentode, wired as a diode. A
sampling of the RF output voltage to the antenna was detect-
ed in this stage and the detected signal was used in tuning
the transmitter to the desired channel and for sidetone in
the pilot's headset.
f. Speech Amplifier ( 7-151. 6 [12)
The speech amplifier, A/ -151, employed a triode, type 6
/ 2 (Tube 6 /'"2 is a multipurpose tube a duodiode triode,
but the diode plates were connected externally to the ca-
thode and amplified the output from the microphone trans-
former B-13075-502. A carbon thrcatmicrophone was used. The
microphone was connected through the control box to pin 8
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Fig. 3-26. RSI-U-3M. Transmitter. Front.
Fig. 3-27. RSI-U-3M. Transmitter. Bottom.
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Fig. 3-28. RSI-U-3M. Transmitter. Back.
fta Fig. 3-29. RSI-U-3M. Transmitter. Top.
32.
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33.
on plug 102. Microphone voltage was applied to pin 4 on the microphone
transformer through pin 5 on plug 104 on the transmitter from whence it
had come through the receiver from pin 5 on plug 402 on the power
supply in which it was obtained from a voltage divider which provided
the microphone voltage from the 28 volt supply. Microphone voltage was
present only when transmitting. Grid bias to /7-151 was taken from the
negative grid bias supply, and the desired grid bias voltage was ob-
tained from a voltage divider consisting of an 82 k ohm and a 1 k ohm
resistor in the grid return circuit in the transmitter.
g. Modulator ( /'j 152, J7153, 6 )16C)
The audio transformer B - 13058 - 501 fed the push-pull modulator
tubesJi152 and/7153 (two 6/760 tetrodes). Grid bias was applied to
these two tubes from a voltage divider consisting of a 46 k ohm and
a 10 k ohm resistor which provided the proper voltage from the ne-
gative grid bias supply.
Transformer B - 13036 - 501 served as the modulation transformer and
the modulation was applied to the anodes and screen grids of the power
amplifier and the screens of the second tripler.
h. Plugs.
Transmitter RSI-U-3M was equipped with 4 plugs. Plug 101, fitted with
a metal friction grip cap, was used in all probability for connecting
a test meter to the transmitter to measure different voltages and cur-
rent to the tubes, thereby enabling the radio mechanic to tune the
channels and test the tube voltages and operation. Plug 106 was not in
use on this particular transmitter and was covered with a plastic cap.
Although the intended use of this plug was not evident, it was obvious
that channel selection could be accomplished from this plug with an
external control box, a microphone could be connected to pin 8, and the
push to transmit function was possible by connecting pin 2 and 8 through
a microphone. This plug was probably used for control and operation of
the transmitter during tuning and testing.
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34-
2. Receiver RSI-U-3M (See circuit Diagram)
a. RF Stage ( /1 -201, 6 4 3 /? )
The antenna signal entered the receiver through plug 205
by coaxial cable from the transmitter at which point the
coaxial transmission line from the sword antenna had en-
tered the set. The signal was applied the RF amplifier
Jf -201 where it was amplified and in turn, applied to
the grid of the first mixer. Tuning was accomplished by
means of a three section, ganged tuning condensor, two
sections of which tuned the grid anode circuit of the RF
amplifier, and the third anode of oscillator.
b. Oscillator ( 17 -204, 6 * 3 J7 )
Each of the four receiver crystals could be connected in
the grid circuit of the crystal oscillator, the anode of
which was tuned to the third harmonic of the crystal
frequency.
c. Tripler ( 1? -205, 6 ) 3 11 )
In this frequency tripling stage, the 9th harmonic of
the crystal frequency was produced. The anode circuit
of this stage was tuned by one section of the 3 sec-
tion ganged tuning condensor.
d. First Mixer ( f7 -202, 6 t 3 1? )
The 9th harmonic of the crystal frequency and the antenna
signal were mixed in this stage and the anode of this cir-
cuit was tuned to the difference between these two fre-
quencies. The 9th harmonic of the crystal frequency was
brought to the screen grid of this tube by means of a
tuned circuit, inductively coupled to the anode circuit
of the first Tripler stage ( 17-205).
e. Second Mixer. (17 -203, 6.r 3 J7 )
To the grid of the second mixer were brought two sig-
nals, the 9th harmonic of the crystal frequency and the
antenna frequency minus the 9th harmonic of the crystal
frequency i.e. 9f crystal and fantenna 9f crystal)' which
were capacitively coupled from the screen grid circuit
and the anode circuit, respectively, of the First Mixer
stage. The difference between the frequencies of these two
signals was the IF, 12 MCS (i.e.fantenna 18fcrystal - 12
mcs).
f. IF Amplifiers (J7-2O6,'7-2o7, 12-208) (type 6 H 4)
Three stages of IF amplification (12 MCS) were provided
utilizing tubes .17 2o6, /7207 and 17208; all were 6 H 4
type valves. The following inscriptions in relative po-
sitions shown were found on the four IF cans and may
give some insight into their content.
Results of selectivity measurements are included in sec-
tion III C 7, Laboratory measurements.
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JG\..nG I
35.
C 227-6 C 227-9
227
L 212 L 213
C 230
2031/
204
1 f
C 242 C 227-14
R 246
Tuning Knobs
The IF cans were hermetically sealed units and it was
not possible to investigate their interiors.
However, it was suspected that they were entirely con-
ventional double tuned networks with RC decoupling to
decouple the IF.
Audio Section; Detector and Audio amplifier ( /7209,
6 f 2), Audio Output (f7 210, 6 17 6 C).
The IF signal was detected in the right hand diode of
tube 1? -209, a duo-diode triode, Soviet type 6 /' 2
tube. The detector load consisted of 3 resistors (200
k ohm, 82 k ohm and 8.2 k ohm). The audio voltage was
applied to the grid of the audio amplifier (the triode
portion of tube 17-209). The sidetone from the trans-
mitter was introduced between the 82 k ohm and the 8.2
k ohm resistors and amplified in the audio amplifier.
The audio signal from the audio amplifier was applied
to the grid of the audio output tube, f7 -210 through a
coupling condensor.
The final audio was applied to the audio output trans-
former TP-201 and thence to the high impedance pilot.s
headphones. (14.5 k ohms at 800 cycles).
E C F'1 I T
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.a L-%,,
SIMPLIFIED DRAWING OF AVC CIRCUIT.
SECRET Fig. 3-30.
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 Sanitized Copy Approved for Release 2010/06/28: CIA-RDP80T00246AO27100330001-4
JGL.RG 3
37.
Fig. 3-31. Receiver RSI-U-3M. Top.
Fig. 3-32. Receiver RSI-U-3M. Front.
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 Sanitized Copy Approved for Release 2010/06/28: CIA-RDP80T00246AO27100330001-4
Fig. 3-33. Receiver RSI-U-3M, Bottom.
Fig. 3-34. Receiver RSI-U-3M. Back.
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. ..%W U .. e
39.
h. Squelch Circuit (11 211 and left diode of 11 209)
( 6 /" 2 )
The squelch circuit consisted of an oscillator operating
on 551 Kcs (triode position of 17-211). The 551 kc sig-
nal was rectified in the right hand diode of /1-211.
The load of this diode consisted of a 100 k ohm and 1 k
ohm resistors. When the oscillator was in oscillation?
a negative voltage was produced and applied to the grid
of the audio amplifier ( /7-209) cutting this tube off.
Oscillation was stopped when a signal was received and
rectified in the left hand diode of)? -209 and a negative
voltage was developed across a 390 k ohm resistor and
applied to the grid of the squelch oscillator. The squelch
circuit could be put out of operation, by a switch,re-
moving the high voltage from the squelch oscillator tube.
i. AVC Circuit ( fl 213 and 11212) (6* 6 C, 6 P 2).
The left hand diode of tube )7'212 was the AVC diode. The
IF output was applied to the cathode of this tube through
a 100 mmf condensor. The AVC voltage was amplified in
tube/1212 and AVC voltages were obtained from the anode
loads of this tube. Strong AVC was applied to the first and
second IF amplifier tubes. A weaker AVC voltage was applied
to the RF and first mixer tubes. The AVC could be varied b
means of a preset potentimeter marked y/ YBCTB (Sensitivity.
j. Plugs.
The receiver was equipped with 6 plugs. Plug 205 was a coax
plug where the antenna signal from the antenna switching re-
lay in the transmitter was introduced into the receiver.
Plug 201, covered by a metallic friction cap (see section
III9 C,1 (h)) was used for connecting a test meter to mea-
sure different voltages and currents in the set, enabling
the radio mechanic to test the receiver and tune it on the
various channels. Plug 202 was used to apply the various
voltages from the power supply to the receiver.
Plug 204 was connected through a cable to the transmitter
and brought the power supply voltages to that unit. Plug
203 was connected to the control box enabling the pilot to
select channels. Plug 206 was not in use in this particu-
lar receiver and was covered with a plastic cap. The func-
tion was evidently the same as that of plug 106 on the
transmitter, to enable local control of the receiver du-
ring testing and tuning (see section III, C,1 (h)).
SECRET
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 Sanitized Copy Approved for Release 2010/06/28: CIA-RDP80T00246AO27100330001-4
%J L_ %up UN L_ I
40.
3. Channel Selection and Push to Talk Circuit.
Channel selection was accomplished from the control box
(see photo and circuit diagram of control box DWG no. 4).
The control box (type P) was equipped with 4 push buttons,
one for each channel, marked 1.2.3. and 4. The control box
was fitted with a switch enabling the pilot to select either
the output of the VHF receiver or the radio compass, and a
volume control which permitted setting the audio output of
these equipments at a suitable value. The control box had 4
plugs. Plug 304 connected the pilot.s headset to the control
box. Plug 301 was connected to the push-to-talk button on
the throttle handle and through a Y splice to the output of
the radio-compass receiver. Plug 302 was connected to the
transmitter and plug 303 to the receiver. Channel selection
was accomplished by grounding the lead to the appropriate
channel contact on the ratchet motor housing. For example
for channel 1, pin 3 on plugs 302 and 303 was grounded, start-
ing the ratchet motor until the channel slide was pushed in
and its channel contact was broken by the channel contact
disengaging throw, stopping the motor. A button marked C6POC
was found on both the transmitter and the receiver. When de-
pressed ground was applied to the ratchet motor and it was
placed in operation. A similarly marked button on the control
box had no electrical function but mechanically disengaged
channel selector buttons. If this were done, the C6POC but-
ton on the receiver and transmitter could be used to control
the ratchet motor until slides were in a desired position e.
g. in position for a desired channel or completely disengaged
for manual, local tuning and locking.
These buttons started the motor and kept it running as long
as they were depressed.
The push-to-talk button was placed. on the throttle handle.
When it was depressed, antenna switch over relay P 101 in
the transmitter and the transmit voltage relay in the power
supply were actuated.
SEC RE 'r
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 Sanitized Copy Approved for Release 2010/06/28: CIA-RDP80T00246AO27100330001-4
Fig. 3-35. Control Box Type P. Front.
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