SOVIET DEVELOPMENT OF ZENITH AND FALKE ROCKETS

• Document Type: CREST
• Collection: General CIA Records
• Document Number (FOIA) /ESDN (CREST): CIA-RDP80-00810A002000640010-8
• Release Decision: RIPPUB
• Original Classification: S
• Document Page Count: 10
• Document Creation Date: December 22, 2016
• Document Release Date: June 16, 2010
• Sequence Number: 10
• Publication Date: August 27, 1953
• Content Type: REPORT

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CENTRAL INTELLIGENCE AGENCY
INFORMATION REPORT
SBCRET
SECURITY INFORMATION
COUNTRY USSR (Moscow Oblast)
SUBJECT Soviet Development of Zenith
and Falke Rockets
DATE OF INFO.
PLACE ACQUIRED
This Document contains information aSeaUng the 3Ra-
tional Defense of the United States. within the mean-
ing of Title 18. Sections 703 and 794. of the ts.8. Code. as
amended. Its transmission or revelation gilts contents
to or receipt by an unauthorized perrson is prohibited
by law. The reproduction of this form Is prohibited.
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REPORT
DATE DISTR.
NO. OF PAGES
REQUIREMENT
REFERENCES
THE SOURCE EVALUATIONS IN THIS REPORT ARE DEFINITIVE.
THE APPRAISAL OF CONTENT IS TENTATIVE..
(FOR KEY SEE REVERSE)
PROJECT "ZENITH"
1. The Zenith mi4sile, by utilizing solid propellant (powder) rockets, was
designed to be used against flying targets for altitude;.3 up to 18 kilometers.
The maximum flying time was.20 seconds. The blast effect was supposed to be
sufficient to achieve total destruction or, at least, flight incapacitation,
when a hit was scored against heavy.. bomber-type aircraft. This weapon
was to be used primarily against flights of bomber-type aircraft.
2. To effectively combat mass targets, it was necessary to have a mass
weapon which could be produced cheaply and in great quantities. As a
consequence, certain basic requirements had'tb be kept'inmtind in
designing such a weapon. A remote control rocket with either
radio, ultraviolet, or similar control from its launching site,
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STA' ARMY #X I NAVY #x AIR FBl AEC
27 August 1953
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as well as a self-controlled rocket, was never taken into consideration
for the purpose of fighting bomber formations. Considering the fact that
a remote or nearby effect of the highly powerful
would demand an extremely large charge as compared to a direct 25X2 }
hit$ the utilization of proximity fuzes could be dispensed with by
using this type rocket. The effect at the target is caused exclusively
by a charge of the above-mentioned explosives, consisting of about 500
grams each, where the detonation occurs by means of a highly sensitive
delay moment involved. 25X1
Ballistic calculation proved to the Germans that the required flying
and its
specific impulse ispec is equal t
ubsequent ballistic exper men 3, ur-
thermore.,proved that especially favorable conditions for the two-stage
rocket were found when the initial stage, as well as the
Launching tlatform
3. Multiple instruments were provided for launching with a rate of fire
of five rounds per minute for the individual rockets. However, there
was an:arrangement whereby four rounds of six rockets each'and two
rounds of eight each within a common steel frame were constructively
investigated. These were mounted on a launching platform which was
adjustable by gun-laying radar in the usual manner for elevation and
traveise. As a variation, the utilization of the main stage with pay-
load 'tit without the first stage, for use against low-flying aircraft,
was also tested. Ground control with respect to fire control was to
have corresponded unchanged to the usual higher performance of ALL,
yet naturally omitting the flight time calculations for the timing
detonator.
PROJECT "X41KE"
Summary of Characteristics
4. The remote-controlled rocket Falke (air-to-air) with a long burning
powder propelling charge for combatting of flying targets was supposed
to have had, generally speaking, the following tactical characteristios:
a. To be used at altitudes of from zero to twelve kilometers.
b. Transverse acceleration in flying altitudes of fro six to eight
kilometers about ten g, equal to about 100M/second .
c. Maximum velocity of rocket at altitudes of from t3ix to eight
kilometers approximately 500m/second.
d. Combustion period of rocket at least ten seconds.
e. Velocity advantage of the carrier aircraft as opposed to the
target - about 50m/second.
f. Velocity of flying target up to 1250 kilometers/hoL-r.
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g. Manner off' utilization 'of rocket: pursuit flight.
h. Effbet on target: incendiary splinters and blast shockwave of
warhead.
i. Detonations electrical oriaoootiietioal (protimity fuze) minimum
detonator.
Controls VRP radio installation with pules+comman transmission.
Design CoMirnotion
5. attempted to reconstruct the overall data of the rocket :in 25X1
dimensions w oh are only approximate to those used.. at that time.fee
page 6? By means of prolonged calculations,. the Germans arrived
at the design values for the individual components. Hess it must be
mentianed that the wind tunnel results of profile and fuaelage models,
as .requested were not placed at the disposal of the German 25X1
mirk group, so that the necessary dimensions had to be caldulated on. a
theoretical 'basis through analysis of the,Rheintoohtir wind tunnel
measurements and development processes.
DESIGN DATA FOR THE ZENITH MISSILE
6. The design data for the Zenith missile are as followss
1. Initial Stage: Caliber approx 12 at
Charge Load approx 11.3 kg
Dimensionss
External Diameter approx. 107mm
Chamber approx 15oxi
Length approx 840Ilis
Combustion Time approx 3.8stc
Thrust approx 220bhg
Weight of Empty Rocket
Chamber including sup-
porting areas approx.14.5hg
Thickness of'Rooket
Chamber (internal surfaces are
protected against glowing
through by means of thermo-
insulating lacquer) approx 2.7am
Length of Rocket Chambers approx 1020mm
2. Main (Second) Stages Caliber approx 68mm
Weight of-Propelling
Charge (Tgl.RP) approx 2.lkg
Dimensions
External'Diameter approx 61mm
Chamber approx 8.5aw.
Length approx 475
Combustion Time approx 2.2 sec
'Weight of empty rocket
chamber (internal our- approx 2.35kg
face is insulated against
glowing through by asaas
of thereto-insulating
lacquer) approx 1.5mm
length of rocket chamber approx 575mm
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3. 'Payload: Weight of explosive charge approx 500 gr
Weight of hull inclusive-
detonator & stabilizer
(on rocket. chamber) approx 700 gr
4. Total Weapons Total, (over-all) weight approx 32 kg
Over-all length approx 1750mm
Nominal velocity increase
through rocket-base (initial
sta a approx 840m/s
Main (second) stage approx 62.6m/s
Total velocity increase approx 1660m/s
7. The most favorable moment for the separation of the initial rocket try
means of the powder gases of the initial stage) was variable under
various conditions, but in all instances it had to occur shcrtly after
completion of initial-stage propellant combustion. This result stands
in opposition to the utilization conditions of a multi-stage rocket for
which the achievement of maximum distance is desired. The main stage
and payload in this case were not to be separated during continuedflig}t.
AERODYNAMICS
8. The canard-type of construction, that is,with rudders arranged at the
.bow of the missiles, was chosens
a. To guarantee the effectiveness of the rudder in all flight
positions.
b. To make available at altitude the lift forces acting on the
rudder surfaces towards the total lift available.
9.
The pronounced sweepback of the wings was chosen, aside from the super-
sonic principles,,in order to obtain a sufficient do /o( . As lift-co-
efficient, based upon a Mach number of v/a about 1.3, a value of 0.05
for each degree of angle of incidence was accepted for incidence of up
.to seven degrees, whereas,,the resistance coefficient of the wings, within
the regions of from zero to seven degrees, was assumed to be within an
increase of from 0.12 to 0.17. The fuselage was, in assuming a nominal
diameter of bulkhead with 300 millimeters,.considered to have a resistance
coefficient of o of about 0.4 at zero degrees incidence, a value which
might rise to about 0,6 at an angle of, seven degrees. The calculation
of flight characteristics, during the-passing-through all Mach numbers,
was achieved according to a process of resistance, as ascertained within
the USSR according to SCHAPIRO, whereby a similar rule wag laid out, for
the progress of lift. The calculations proved that the demands made for
the proposed design would be met with a wide margin of success (so far as
the assumed aerodynamic prerequisites were attainable). During the fight-
ing against aircraft at the upper speed range, especially during low
flight altitudes, the fighting range fell naturally-to a marginal value
of 1200 meters and below, whereas, during the fighting-off of aircraft
with inherent velocities of about 900 kilometers/hour, the maximum
fighting separation of 1600 meters could, especially at altitudes of
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considerable value, be exceeded oonsiderabl . The lower fighting separa-
tion was determined in this casea minimum target pick-up time of five
seconds limitation.
CONTROL
10. The oQatrol of the flying missile was scheduled to be affected solely by
means of the rudders attaclied to its head in dispensation of the usual
traverse rudder. For this reason, it.was imperative Ab design the rudder
installation as a special vertical and horizontal combination, that is,
each rudder vane had to be adjustable individually. The steering knob
built .nto the pilot's compartment of the aircraft then interpolated the
stick movements into polar coordinates. Then, by means of a special
potentiometer circuit, these values were transmitted to the missile by
decimeter carrier wave common for both, the right as well as the left
vanes, through variations in the pulsing correlation of each individual
cf. a.rrl channel. Within the missile, after filtering of the low fre-
c:,.encies of the left-hand and right-hand channel, under admixture of the
in4tcated values of the directional gyro of the switching box, a change
to eleetro-mechanical commands for the Servo-unit see e 7 J.
Steering was' accomplished after a line of sight trajectory.
VARIATIONS
11. Taking into consideration dimensions, instruments of lesser acceleration
were likewise investigated, whereby the smaller wings led to smaller
overall dimensions. Also, variations for payloads were considered.
Furthermore, the utilization of a hinged wing unit was supposed to pe
investigated constructively but was never done due to centralization of
remote-controlled rockets within another Ministry in Putilovo (USSR).
A predecessor of the project development was represented by the Project
Moeve as elaborated on and worked on in the Technical Design Office of
Oberscboeneweide,.Berlin. However, this instrument was designed for
the current aircraft speeds and altitudes of up to eight kilometers.
Consequently, the total weight amounted to merely 140 kilograms.
PAGE 6 B Falke Rocket for Fighting Bombers with Legend.
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SERVO UPII'T:
KIM
, EPIA,rk
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LEGEND
The positions` marked are described as follows: Zs ee drawing7
1.? Dipole antenna as a combined receiving-transmitting antenna for the
electrical proximity detonator.
2. Supersonic propeller drive for the electronic installations in No. 3
and No. 5 position. The hollow propeller shaft has insulator bushings
under the above-mentioned dipole antenna (see 1.). Here it might be
mentioned that a variant was tested, using instead of the supersonic
propeller a ramjet turbine, but the installation did not work well
due to icing conditions which were not corrected. ;,.
DC-4C generator inverter for the generation of plate voltages for
Position No. 8 (250v DC) and the 24V DC for the magnetic coupling
of Position 5 as well as the three-phase AC inverter for the damped-
position-gyro of Position 7 at 500 cps.
4. Control rudder with plus or minus 70 variability. For a profile, an
acute symmetric rhombus was selected. The left and the right rutders
are independent of each other and controlled by the Servo-unit. Each
rudder has one needle and one ball-bearing support. The rudders are
sheet steel (approx. 4mm) and welded onto the rudder shaft which is
shaped like a conical thorn.
$. 'Eleotro-mechanical servo.-unit, whose kinetic system may be seen from
the sketch at the lower left Zee drawine. The control adjustment
is produced by the supersonic propeller, and is fed to the eleatro-
generator and then the central driving-gear wheel. This wheel its
geared to the left front and right rear drive wheels. Each of the
rudders meshes with a slightly smaller counter-wheel so that either
left-hand or right-hand turning is possible. The coupling shatEn,
with trapezoidal internal threads, are led out of the Servo-unit
housing through the hollow shafts of the front wheels. They have
the trapezoidal thread spindles whose yokes are coupled with a
joint lash to the adjusting lever of the rudder axle. Each of the
cogwheels is built as a pot magnstjand carries an excitation winding
that is fed by means of feed rings. The coupling shaft has an
axially-arranged displaceable friction disc, constructed as an
armature. Through excitation of the forward or rear pot magnet,
the clutch coupling disc is pressed into the corresponding cog-
wheel so that the coupling shaft may be switched from right to left
for rotation as desired. The frequency of shifting of this '
arrangement is given at fifty per pent' of the self-induotio4,of hg
exciter winding and of the-'mass inertia of the coupling disc. I:
amounts to about four seoondsa Rigidly combined ftth''tht? outer
rim is the Servo-unit. 4 , .
6. The housing for the electrical control instruments, capacitor, resis-
tance, amplifier tubes, and relay. The feed-back potentiometers,
(for reporting back of incidental rudder position) are built in
solidly within the Servo-unit. They are coupled together with the
threaded shafts.
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Centrally-flanged onto the Servo-unit is the damped-directional-gyro
whose interlock is released automatically at the moment of launching.
8. IIE' radio receiver, whose construction was copied from the Rh.#ntochter
instrument. By means of an antenna tuning device, it is connected to
the two wings which have metal sprayed on them in the shape of antenna-
strips. It was intended to use amplifier tubes with negative temipera-
ture coefficient cathodes (German patent - my lf_) in order to save on
the generator output. The activation of the complete installation
occurred through freeing, electrically9 of the supersonic propeller
from the observing site. In view of the low dynamic pressure during
initial acceleration for the actuation of the installation, the
electric generator was connected to the mother aircraft's supply,
thus funotionirn as a would-be rotary converter, generating the nec-
essary powero simultaneously with this9 the heating of the negative
temperature coefficient cathodes had to occur through the aircraft's
supply. After launching, the inertia of the generator's stared
eztergy was great enough to operate the installation until sufficient
ram. ressure of the rocket was attained to satisfy necessary power
requirements. The sloe cathodes had a suffioient heat-holding
capacity for a period of fifteen seconds operation.
The` front part of the fuselage was constructed of about four mili
meters thick aluminum sheet metal9 which also served as housing for
all control components. It was connected to the long-burning powder-
propulsion unit, which was arranged within the region of the center
of gravity by means of flanges.
10. Racket chamber for the powder-propulsions plant. The rear half of
the chamber had a coating of silicate thermo-insulation which was -01
sprayed on.
11. Glutinous-like thermo-insulating mass.
12. F:~ontal powder block (about 35 kilograms)', overall thermally-insulated,
except for a 20 milimeter center hole.
13. Bakelite-like thermal insulating mass on rear traverse of plane.
14. 8tee1 protective cover, to avoid mechanical damage to the rear thermal
insulation by hot powder gases.
15. As in 14, but for the protection of front traverse of the rear powder
body.
16. Rdar powder body (approx 30 kilograms). Front and rear traverse planes
thermally-insulated.' Whereas the. combustion process of the front
powder charge proceeds from tte center outwardq the combustion front of
the rear body progresses accordingly from the outside centrally towards
the central axis so that.during the period of combustion, a uniform
surface ratio is maintained. Since the front powder body experiences
a..aontinuously increasing loss of weight9 and the rear powder body
shown a continuous voluminous decrease, the center of gravity moves
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during combustion first forward and thong after combustion, returns to the
approximate point of origin. The coverlid of the rocket chamber arranged
in the rear, has six equi-distantly spaced jets on its circumference, it
was impossible to orientate the directivity of the jets to that of the
center of gravity, in this design, since that would have meant too great
a lose of propulsive thrust.' Nevertheless, eit..trientation of direction
up to about ten degrees would still be acceptable.
17. Electrical. firing installation for the explosive charge,
18. Wing (pair) with supersonic profile. Acute rhombus with d/t about 1/14
or eventually narrower yet.
19. Incendiary fragments, according to dimensions as with R100B8, about three
hundred pieces., that is, 18 kilograms over-all.
20. Blasting oharget about 30 kilograms of trinitrotoluene..
21. Aluminum container, of three milimeters heavy sheet metal, for blasting
charge. On its circumference, six pocket-like bays for jet stream.
22. Final discs, 2505 500 milimeters as damping fine for'vibration about
Y-axis.
23. Mounting angle brackets of wings welded to chamber cover.
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