THE SOVIET SURFACE-TO-SURFACE GUIDED MISSILE SCUD (SS-1B) (C)

• Document Type: CREST
• Collection: NGA Records (Formerly NIMA)
• Document Number (FOIA) /ESDN (CREST): CIA-RDP78T05439A000200370043-6
• Release Decision: RIPPUB
• Original Classification: S
• Document Page Count: 87
• Document Creation Date: December 28, 2016
• Document Release Date: October 11, 2002
• Sequence Number: 43
• Publication Date: May 1, 1963
• Content Type: REPORT

ILLEGIB Approved For Release 2002/11/04: CIA-RDP78TO5439A000200370043-6
Approved For Release 2002/11/04: CIA-RDP78TO5439A000200370043-6
 SEMET
AN ARMY INTILLIGENCE DOCUMENT
THE SOVIET SURFACE-TO-SURFACE GUIDED MISSILE
SCUD (SS-1B) (C)
MIS 24-63
MAYY 1963
PREPARED BY
4
U. S.ARMY ;MISSILE COMMAND
REDSYON#:ARSENAL, ALABAMA
THIS DOCUMENT WAS COMPILED AND PUBLISHED BY
THE ARMY MISSIL( COMMAND AFTER REVIEW AND
APPROVAL IN TNC P1CE OF THE ASSISTANT CHIEF OF
STAFF FOR INTELLI NCE AND THEREFORE CONTAINS
AGREED DEPART OF THE ARMY INTELLIGENCE.
I
WARNING NOTIC.E
SENSITIVE SOURCE "NI) METHODS INVOLVED
25X1 C
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25X1 C
(S) 1 S RD (U)
(S) This report is-an analysis of the SCUD SS-lb Surface-
to-Surface Guided Missile system, and represents an. evaluation
of intelligence information available through May 1963. This
document was prepared by the Directorate of Missile Intelligence,
U. S. Army Missile Command.
(U) Comments or queries regarding the material contained
in this report should be.submitted to the Commanding General,
U. S. Army Missile Cori. ALTN: AMSMI-Y, Redstone Arsenal, -
Alabama.
iii
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(S-NOFORN) TABLE OF CONTENTS (U)
I, eral System Description (U) ----------------------- 1
A. Introduction .(U)--------------------------------- 1
System Characteristics (U) -----------------------1
II. Missile Characteristics (U) -------------------------- 5
A. General (U)-------------------------------------- 5
B. Airframe (U)------------------------------------- 5
C. Guidance and Control (U) ------------------------- 7
1. General (U)---------------------------------- 7
2. Guidance System (U)-------------------------- 7
3. Control System (U)--------------------------- 8
4. Guidance and Control Checkout (U)------------ 10
5 Autopilot Erection and Alignment (U)--------- 10
D. Propulsion System (U)---------------------------- 11
E. Propellant System (U) ---------------------------- 12
F. Main Paver and Pneumatic Systems (U)------------- 13
G. rformance Analysis (U)----Z-------------------- 13
1. General (U) ---------------------------------- 13
2. tbi Propellant Weight (U)------------------------------- 14
3. Propulsion System (U)------------------------ 14
4. ku; ro ynamic Heating (U) ----------------------------- 15
5. (C) Airframe (U) ---------------------------------------- 15
6. Range (U)------------------------------------ 16
H. Warhead (U)-------------------------------------- 18
III. Ground Support Equipment (U) -------------------------19
A%. kaj neral (U) --------------------------------------------- 19
B. Missile Ground Transporter (U) ------------------- 19
C. Propulsion System Test Equipment (U) ------------- 24
D. Electrical System Test Equipment (U) ------------- 24
E. Air Compressor Unit (U) -------------------------- 24
F. Electrical Power Supply (U) ---------------------- 25
G. Fuel Transporter (U) ------------------------------ 25
H. Oxidizer Transporter (U) ------------------------- 25
I. Washdown and Neutralizing Vehicle (U) ------------ 25
J. Fire Fighting Vehicle (U) ------------------------ 28
K. Crane (U) ---------------------------------------- 28
L. Transporter-Erector-Launcher (U) ----------------- 28
M. Meteorological Equipment (U) ---------------------33
IV. Or anization and Field Operations (U) --------------- 35
A. General (U) -------------------------------------- 35
B. Line of Vehicle March (U)------------------------ -35
C. Processing the Missile (U) ----------------------- 35
D. Field Deployment (U) ----------------------------- 37
E. Launching Site (U) ------------------------------- 37
F. Deployment (U)----------------------------------- 37
G. Comments (U) ------------------------------------- 37
iv
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X1
Assembl
y and Checkout Procedure (U) ------------------
38
oducti
on (U)----------------------------------------
38
X1 B.
J
Tes
ting (U) --------------------------------------
38
General (U)----------------------------------
38
K1
38
Technical Positior (U) -----------------------
X1
Loa
ding and Assembly (U) -------------------------
38
1
Operations at the Fueling Point (U)--- -----
38
.
1
----
39
*
2.
Assembly and Joining Procedure (U) -------
ch (U) ---------------------------------------
39
x1
Missile Alignment
on Launcher
(U)____________
39
2.
Launch Area Tests
(U) --------
39
DO
Aim
ing Requirement (U)
------------
---------------
41
X1 APPENDIX I.
Supporting Analysi
s (U)-------
---------------
43
APPENDIX II.
(S) Heat T
ransfer Analysis (U
) -----------
---------------
60
APPENDIX III.
(S) Airfra
me Stress Analysis
(U) ---------
---------------
66
APPENDIX IV.
Transporter-Erecto
r-Launcher V
ehicle
nalysis U
Analysis-U
)------------------
) ------------------
------------
------------
---------------
---------------
71
APPENDIX V.
(U) Refere
nces (U) ----------
-
-------------
---------------
75
APPENDIX VI.
Glosasry of Terms
and Symbols
(U)------------
76
APPENDIX VII.
(U) Distr
ibution List (U)---
-------------
---------------
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X1.
Fissure
1.
2.
3.
4.
5.
6.
7.
8.
9.
10.
11.
12.
13.
14.
15.
16.
23.
24.
25.
26.
LIST OF GRAPHICS (U)
Paste
(CONFIDENTIAL) Moscow Parade Photographs, 1 May 1960 (U) -------- viii
(SECRET-
(SECRET-
(UNCLASSIFIED)
(UNCLASSIFIED)
(UNCLASSIFIED)
(CONFIDENTIAL)
(UNCLASSIFIED)
(UNCLASSIFIED)
(UNCLASSIFIED)
(CONFIDENTIAL)
(CONFIDENTIAL)
(CONFIDENTIAL)
(SECRET
(SECRET.
(SECRET)
(SECRET)
(SECRET)
(SECRET)
Parade, 1 May 1960 (U) ---------------------------
Inboard Profile (U)_____________________________
Block Diagram of SCUD Guidance and Control
System (C) --------------------------------------
ZIL-151 Standard Cargo Truck (U)-----------------
ZIL-l51 Van Body Truck (U)---.-------------------
ZIL-157 Standard Cargo Truck (U) ------------------
Missile on Transporter, Line Drawing (U) ---------
Electric Generator (Generator Trailer) (U) -------
Rotary Converter (Transformer Trailer) (U) -------
Fuel Transporter, Line Drawing (U) ---------------
A-Frame Crane, Photographs (U) -------------------
Transporter-Erector-Launcher, Photograph (U) -----
Transporter-Erector-Launcher with Missile in
Launch Position (U)______________________________
Typal Layout of Met Station (U)---=------------------
Organizational Structure of SS-1 (SCUD) Guided
Missile Brigade (C)-------------------------___
Typical Combat Formation of SCUD Missile 'K
Battalion (C) -----------------------------------
Grid Circle of Collimator (U) ------------------
Center of Pressure vs Mach Number Curve for SCUD
Missile (C) --------------------------------------------
Drag Coefficient vs Mach Number Curve for SCUD
Missile (C)--------------------------------------------
Performance, Parameters of SCUD Missile with 1550 lb
Warhead (C) --------------------------------------------
Performance Parameters of SCUD Missile with 2500 lb
Warhead (C) --------------------------------------------
(SECRET) Skin Temperature vs Time History for SCUD Missile (C)--
(CONFIDENTIAL) SCUD Missile on Carrier (C) ----------------------
(CONFIDENTIAL) Artist's Concept of Firing Table (U) -------------
(CONFIDENTIAL) Exploded View of Firing Table Assembly (U)-------
D
LIST OF TABLES (U)
6
20
21
22
23
26
26
27
29
30-
31
32
34
36
40.
54
55
58
59
64
70
72
72
1. (S Propulsion System Data (U) --------------------------- 15
2. (S Missile Characteristics (U)--------------=---- 16
3. (S Weight Breakdown (U)--------------------------------- 17
4. (S) ated Tank Volumes and Propellant Weights (U)---------- 44
5. (S Propulsion System Performance Characteristics (U) ---- 47
6. (S Estimated Tank Pressurization System Weight
7. (S) Estimated Main Power Supply Weight Breakdown (U)------------ 52
vi
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8.
(S)
Missile Skin Material Properties (U) ------------------------
60
9.
(S)
Nose Cone Skin Temperature (U)------------------------------
62
10.
(S)
Missile Body Skin Temperature (U) ---------------------------
63
11.
(S)
Fin Skin Temperature (U) ------------------------------------
65
12.
(S)
Estimated Vehicle Weight Distribution (U)-------------------
66
13.
(S)
Estimated Airframe Weight Breakdown (U) ---------------------
67
vii
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GENERAL SYSTEM DESCRIPTION (U)
Introduction (U)
The Soviet Union has developed a second generation 150 nm
missile system identified as the 8A61. This missile is designated the SS-lb
and has been nicknamed SCUD A by western intelligence. It is a replacement for
the Soviet V-2 type 150 nm missile.
The SCUD A was first seen in the 1957 May Day parade. This
probably a es the beginning of the development period in the 1953-1954 time
period.
I Intelligence information verifies that the SCUD A is heavily
deployed in the Soviet Union, lightly deployed in GSFG and possibly deployed in
other Bloc countries.
A%k
The SCUD missal ?is a highly mobile, extremely reliable, tactical
weapon. transported, erected and launched from a transporter-erector-
launcher.
System Characteristics (U)
The SS-lb (SCUD A) is a single-stage, surface-to-surface,
all-inertial, guided-ballistic fissile. It employs a storable propellant
combination of red fuming nitric acid (containing 181-22% N204) and kerosene.
In addition, a starting fuel, "Tonka", is used (507. ksilidin and 50% tri-
ethylamine). The trajectory for the all-inertial system is determined from
a preset program and cutoff is determined by a velocity seeker. The tankage
section is a full-monocoque welded construction, while the aft skirt and fins
are of semi-monocoque riveted construction.
1 -1 Three t
high-explosive, an nuclear. T
at the forward end of the tanks
explosive (HE) warhead, the its
nuclear warhead, the missile ha
nuclear warhead weighs approx
In future references within thi
warhead is in the missile, unle
to lift-off weight, burnout wei
I A brief s
data is listed e ow. The phot
parade were used, along with of
system description and perfo
Missile length
Diameter
Range (max)
Thrust (vac)
Specific impulse
of warheads exist for the SCUD: chemical,
warheads are attached to a structural member
section. When fired with the 1550 lb high-
ile.has a maximum range of 160 nm. With a
approximately half range capability. The
tely 2500 lbs and has a 10 to 15 kiloton yield.
study, it will be assumed that the 1550 HE
otherwise indicated. (i.e., with reference
etc.)
graphs (Figures 1 and 2) taken in the 1 May 1960
r intelligence information, in determining
(vac)
34.4 ft
2.85 ft
160 nm
21,500 lbs
256
1
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Propellant flow rate
84 lb/sec
Duration of thrust
92 secs
Lift-off weight
11,728 lbs
Burnout
weight
4000 lbs
Cutoff
acceleration
4.6 G's
A
41 urn
pogee
Flight
time
315 secs
Cutoff
velocity
4958 ft/sec
Cutoff
angle
36.5 degrees
3
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The SCUD airframe can be divided into the following sections:
(a) nose, (b) forward tank '(oxidizer), (c) instrument compartment, (d) rear tank
(fuel), and (e) tail section. (See Figure 3 for schematic.) The nose section
houses the warhead and fuse assembly.
SCUD A employs an all-inertial guidance system that remains
.systems.
inertially fixed to the local horizontal and azimuthal orientations of the
launch point. It has four carbon vanes, located in the jet exhaust, which are
mechanically deflected by steering motors, for control purposes. The missile
developes a vacuum thrust of 21,500 pounds and burns for a total of 92 seconds.
The carrier vehicle consists of several main subsystems; the airframe, guidance
and control, propulsion, propellant systems, missile power, and pneumatic
B. Airframe (U)
The center transition section contains an access door.
Part of the missile guidance and the hot gas pressurization system of'the fuel
pressure tank are contained in this transition section.
The tail section skirt is a semi-monocoque riveted structure
consisting of a structural frame and a thin sheetmetal covering. The loads
of this section are carried by the framework, and the skin acts mainly as
an aerodynamic cover. The thrust is transmitted from the engine through a
welded frame mount of steel tubing to the aft support ring which is bolted to
the aft skirt and to the tankage section. The tail section encloses the
propulsion unit and provides fin surfaces and trim tabs for the missile. In
addition, the tail section houses the electrical components of the control
system, part of the guidance system, and the main power supply unit.
The engine uses a liquid bi-propellant, has a sea level thrust
of 18,500 7us, and is assumed to be an uprated German Wasserfall type.
The forward transition section houses the hot gas pres-
General (U)
The forward and aft tank sections contain the missile
propellants. Stainless steel (0.081 inch thick) is used to withstand the 500
psi tank pressure of the hot gases from the. gas generator pressurization systems.
The tankage section is a full-monocoque structure consisting of a shell and
additional frame support members.- The missile skin serves as the tank wall.
The tanks are separated by a center transition section. All subsections of the
tankage are welded together to form a complete unit. An oxidizer line runs
through the fuel tank to the tail section. A conduit is located on the outside
of both tanks to.carry the electrical cables and pneumatic lines to center and
nose sections.
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rGUIDANCE SYSTEM:
S R T
YAWTRO
r
REQUENCY
AND( VOLTAGE
REULATOR
I 50 V
BATTERY
MISSILE
POWER
SUPPLY
0
i v
- -- PGIA
C O N T R O L
A M P L I F I E R
it i
32 V
BATTERY
4 ~ I
500', AC
INVERTER
L PITCH
ROGRAM
G G
8 B
PROPELLANT
- -a-- IV --
Figure 4. (SECRET
0
Tow
V.-
SUBSYSTEM
MECHANICAL
ELECTRICAL
BIAS CONTROL
2 POTENTIOMETERS
FEEDBACK
GAIN CONTROL
CARBON VANE
PENDULUM
ROTOR, GYRO
2 MOTORS, STEERING
SENSOR, ATTITUDE
TORQUE MOTOR
VALVE PROPELLANT
TIME CUTOFF SWITCH
VELOCITY CUTOFF
SWITCH
X YAW
Y ROLL
Z PITCH
PGIA PENDULOUS GYRO
INTEGRATING
ACCELEROMETER
Block Diagram of SCUD Guidance and Control System (C) _
6
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C. 1 -1 Guidance and Control (U)
1. General (U)
The SCUD employs an all-inertial guidance system
Guidance System (U)
The autopilot consists of a baseplate, two gyros, a
pitch gyro programmer, ana an integrating accelerometer. The baseplate is
mounted to the airframe in the instrument compartment between the propellant
tanks. This baseplate contains reference flanges for the mounting alignment
to the airframe, missile vertical erection, and inertial component alignment.
Gyros (U)
The two gyros of the SCUD autopilot are bas-
the German LEV-3 (V-2) autopilot; that is, they
y
are a two-degree-of-freedom, non-synchronous, ball bearing type.'
icall copies of those
The yaw-roll gyro is mounted on the base-
that remains inertially fixed to the local horizontal and azimuthal orienta-
tions of the launch point. During the thrust phase these orientations are
inertially maintained. All outside control is removed at launch. Attitude
deviations are detected by an autopilot, cutoff velocity is detected by an
integrating accelerometer, and attitude and trajectory control are maintained
by a control amplifier through steering motors and jet vanes. For a graphic
description of the guidance and control system see the block diagram (Figure
4).
plate with its spin axis perpendicular to the local vertical and to the firing
azimuth. (This spin axis orientation permits the missile to program in pitch
without affecting the yaw-roll reference.) An electromechanical pendulum and
torque motor align the outer.gimbal parallel to the local vertical. The spin
axis is aligned perpendicular to the outer gimbal by a torque motor and
bridged potentiometers. The inner gimbal potentiometer provides the roll.
attitude signal and the outer gimbal potentiometer provides the yaw attitude
signal.
(2)
Pitch Gyro (U)
The pitch gyro is mounted on the base-
plate with its spin axis perpen cu ar to the local vertical and parallel to
the firing azimuth. An electromechanical pendulum and a torque motor erect
the outer gimbal parallel to the local vertical. A torque motor and inner
-gimbal bridged potentiometers align the spin axis perpendicular to the outer
gimbal. (This alignment eliminates yaw cross-coupling while roll cross-
coupling can be tolerated to the mechanical limit of the gimbal freedom,
approximately ? 200.) The bridged potentiometer of the outer gimbal axis is
programmed around the pitch axis of the autopilot by a pulsed d-c motor.
7
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The pitch gyro programmer consists of a step
motor, sliprings, and bridged potentiometers. Pulses of d-c current from a
timer energize the step motor and the pitch attitude potentiometer body is
mechanically displaced around the pitch axis of the autopilot. (The potentiom-
eter wiper is attached to the outer gimbal of the gyro.) Programming of this
potentiometer does not affect the gyro's ability to detect missile pitch
attitude deviations, but this displacement introduces an attitude signal in
the pitch control circuits. This attitude signal will cause the missile to
pitch over until this sensor is nulled. The pitch program, during the thrust
phase, is controlled by this timer and the rate and duration of the program
is based upon the selected range.
The integrating accelerometer of the SCUD
autopilot is a pendulous gyro integrating type. This thrust cutoff device is
a velocity seeker that measures acceleration, integrates acceleration for
velocity information, and initiates cutoff. This device is so mounted on the
baseplate that longitudinal missile acceleration is sep,sed and, at a pre-
determined velocity, the cutoff circuits initiate thrust termination in a
single command.
The baseplate of the SCUD autopilot is somewhat
triangular in shape, has an adjustable capability on at least two of the three
points of support, and contains two machined-flanges that are used for plumbing
the missile after erection. The gyros and the accelerometer are initially
aligned to these flanges and these machined surfaces are used as a reference
for aligning the autopilot to the firing azimuth.
Integrating Accelerometer (U)
Control System (U)
The control system consists of the control amplifier,
four steering motors for the carbon jet vanes, four jet vanes, four feedback
potentiometers, the timer, the propellant cutoff valve and trim tabs. The
control system uses d-c attitude signals from the autopilot to control steering
rudder movement by a magnetic control amplifier.
Control Amplifier (U)
The magnetic control amplifier consists of a
preamplifier, mixer, and power amplifier for each of the three attitude
signals. Provisions are made to utilize rudder position signals (feedback).
The attitude and feedback signals are mixed and amplified by the preamplifiers
to provide sufficient amplitude for mixing with a constant phase source in the
mixer. Integrating networks of the preamplifier also provide rate and dis-
placement information for the mixer. The mixer determines the amplitude
8
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D5439A000200370043-6
by comparing the incoming signal with a constant phase source and the amplitude
is the result of rectification and filtering. The mixer output is amplified
by the power amplifier to energize the control relay of the steering motors.
Steering Motors (U)
The four steering motors are electrically-
driven oil pumps controlled by a polarized control valve. The control valve
is synchronized with the polarity of the attitude signals to provide vane
deflection needed for yaw, roll, and pitch. The steering motor consists of
a d-c motor, an oil pump, and a free-floating piston. The motor runs con-
tinuously and provides an oil pressure through the oil pump to the piston
cylinder. The piston's position, through linkage, determines the angular
deflection of the carbon vanes. The control valve is a polarized relay that
provides directional control of the oil flow to the cylinder. The control
amplifier output is applied to each of the four control valves and the polarity
of the controlling voltage determines the direction while the amplitude
determines the angular deflection of the carbon vanes.
The carbon vanes, located in the jet exhaust,
are mechanically deflected y the steering motors. The carbon vanes control
the missile in yaw, roll, and pitch; Vanes I and III in yaw, Vanes II and IV
in pitch, and all four in roll. Two potentiometers are spindle connected to
each of the four carbon vanes to provide a position signal for the control
amplifier and also for test purposes prior to launch.
The timer supplies the time reference for the
pitch program and the burning time limit for the engine. (This timer is
believed to be synchronized to the inverter frequency; that is, the 10th sub-
harmonic of 500 cps.)
Propellant Cutoff Valve (U)
The cutoff valve terminates propellant flow
on command of the velocity switch (PGIA) or the time switch (timer).
Normally the velocity switch terminates propellant flow after the missile
obtains cutoff velocity; however, should the PGIA fail to detect cutoff
velocity, the timer terminates propellant flow prior to depletion.
The use of trim tab control on the aerodynamic fins
has not been ascert , but their utilization would be desirable. A close
examination of the SCUD photographs indicates trim tabs on the rear portion
.of the fins; however, this is not positive because of the detail of the
photography. During the thrust phase, carbon vane erosion will introduce
false torques into the control system. The guidance and control system is
9
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 SECRET
capable of detecting and correcting these torques, but if trim tabs are used,
the correction torques needed for the carbon vanes would be reduced.
Guidance and Control Checkout (U)
Q The SCUD guidance and control system undergoes many tests,
alignments, and checkouts before launching.
I Components of the guidance system are tested
individually an as a sys em. The gyros and accelerometer are removed from the
missile during horizontal checkout, but replaced before the missile is erected
on the launch table. The guidance system functioning is tested with the control
system to insure proper polarization, gain, and synchronization. The integrating
accelerometer's capability to terminate thrust is also tested prior to and after
the missile is erected on the launch table. For further checkout procedures
see Section V.
I I The control system does not correct for lateral and
slant range errors during the thrust phase. Lateral errors due to wind shear
forces are determined from firing tables and met data and corrections in
azimuth are made prior to launch. Slant range errors are minimized by a nearly
constant engine thrust; therefore, providing a flight trajectory that closely
approximates the desired trajectory. A test sequence program is applied to
the overall control system to insure that:
(a) Th
e polarity of the attitude signals is correct.
(b) T
he synchronization of. the carbon vanes provides roll
(c) T
he pitch program is started, locked-in, and terminated
(d) T
he emergency shutdown circuits are actuated.
(e) T
he fuze-arming signals occur at the predetermined
(f) T
he thrust is terminated by an on-board timer (92
seconds after lift-off) if
cutoff velocity.
the integrating accelerometer has not detected
(g) The necessary final fuze-arming signals are applied
Autopilot Erection and Alignment (U)
When the missile is erected on the firing table with
Fin I aligned downrange, two levels are placed on the autopilot baseplate
reference flanges. The missile is leveled by a triangular leveling frame of
the firing table (see Figure 26). The final aiming procedure employs the
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reference flanges of the autopilot and a collimator (Section V, paragraph E).
A prism is employed to transfer (optically) the autopilot reference to the
carbon vanes.
I I The final aiming procedure requires that the auto-
pilot be energize cted; that is, the gyros running and the gimbals
aligned to their preset positions. The control amplifier must also be operating
to center-position the carbon vanes. The pendulums of the gyros detect the
local vertical and maintain the autopilot earth-fixed; therefore, the control
system will center-position the carbon vanes. This guidance and control system's
holding capability is remotely observed. A prism is mounted on and aligned
with Fins I and III and'the collimator is employed to monitor the fin's ref-
erence. After the autopilot compartment is closed, the autopilot's erection
capability is remotely observed by the collimator and the prism. This observa-
tion is conducted throughout pre-launch checkout and up to the moment of launch.
The prism mounted on Fins I and III may be removed before ignition; however,
the prism is believed to be expendable.
Propulsion System (U)
The power plant of the SCUD missile is a bi-propellant
liquid rocket engine. The thrust is generated for a period of 92 seconds.
The desired thrust is maintained by-a chamber pressure loop. Energy for this
thrust is provided by red fuming-nitric acid ('containing 18-227. N204) and
kerosene. Iodine is added to the acid to inhibit corrosion. A small concen-
tration of iodine additive results in a negligible decrease in"performance.
The combustion chamber cooling system consists of point
connections for distribution of the nitric acid within the chamber walls.
The chamber has a combined cooling system (circulation cooling and internal
cooling -- a vapor curtain). This cooling system permits the maintenance of
930-1110?F temperature.
The pressure in the compressed air tank located in the
engine compartment i re 3) is equal to 3000 psia. A reducer lowers this
pressure to 550 psia (this pressure is used to force the hot gas components
into the gas generator, and the starting fuel into the injector).
The propellants are pressurized and transferred from the
missile tanks to the engine combustion chamber by gases from two hot gas
pressurization systems. (See paragraph E below.)
1 -1 In summary, the propulsion system consists of the fol-
lowing components:
1. Combustion chamber.
2. Fuel tank.
11 1.'e1
reed a5 'i-7.
4. Oxidizer tank.
11
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 SECRET
5. Oxidizer feed assembly.
6. Distribution net.
7. Tank pressurization assemblies.
8. Engine control assembly (pyrovalves, diaphragms,
pressure relays, throttle unit, and the like).
9. -Engine assembly (piping system, delivery discs, and
feed and drainage valves).
10. Atomizers (ball-valve type, jet and centrifugal).
The orifice diameters are from 0.002 to 0.079 inches.
The propellant system transfers the fuel (kerosene) and
oxidizer (nitric acid). from the missile tanks to the thrust chamber under
pressure. The systems consist of two hot gas generator systems, two tanks,
two on-off control valves, a throttling control valve, piping, and the engine
passages.
I I The two tanks are made from high strength stainless
steel. The x tank is located above the fuel tank. The oxidizer passes
through the fuel tank (by way of a pipe) to the main on-off valve. After
passing through this valve, the oxidizer enters the coolant manifold. This
section directs the engine coolant fluid (the oxidizer in this case) around
the nozzle and up through the engine wall to the injector. Fuel flow is
direct to the injector head through the main on-off valve. The engine
temperature and chamber pressure variations operate the automatic throttle to
vary fuel comsumption.
A propellant tank pressurization system supplies a
pressure of about,500 psia for the oxidizer and fuel tanks. An investigation
of a simple compressed air pressurized propellant feed system was conducted
in an effort to determine the applicability of such a system. The study
revealed that a simple on-board pressurized feed system would impose severe
weight penalties on the SCUD missile. In addition it was found that the
missile does not have the necessary free volume to carry the required pressure
bottles. A minimum of 24.9 cubic feet is necessary as a gas volume require-
ment. This also imposes a weight penalty of approximately 500 pounds.
Therefore it is felt that two hot gas generator pressurization systems are
the best candidates for a propellant pressure feed system.
I It is assumed that a hydrogen peroxide gas generator-
is used for the aci tank. The steam and oxygen gases from this reaction are
compatible with the nitric acid at tank gas temperatures of 500-700?F. At
lower temperatures the steam would condense to water and dilute the acid. At
higher temperatures a prohibitive tank wall thickness would be required to
contain the pressure. Higher temperature would also cause decomposition of
prohibitive amounts of N204 in the acid.
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An ethylene oxide gas generator is assumed to be used
for the fuel tank. The methane and carbon monoxide gases from this reaction
are compatible with the kerosene fuel. A heating coil is attached to the head
of this gas generator, near the injector, for ignition of the ethylene oxide.
Nitrogen is used for pressurizing gas for the hydrogen
peroxide and ethylene oxide containers. The nitrogen gas is stored in a
sphere in the engine compartment.
It is also.a possibility that fuel could be used as a coolant.
There is enough fuel left as residuals for this purpose. However, nitric
acid was chosen as a coolant based on U.S..state-of-the-art and also on the
fact that the German.Wasserfall and Soviet R-113 engines use nitric acid to
cool the engines. (There is a great similarity between the Wasserfall, R-113
and SCUD). The missile parameters would be altered very little if fuel was
used.
Main Power and Pneumatic Systems (U)
A main missile power supply (MPS) is utilized in the SCUD
missile to supply electrical power for the guidance system and hydraulic power
for pneumaticovalves. Alternating current is supplied by a rotary inverter,
while hydraulic power is supplied by a d-c motor-pump combination. The MPS is
located in the engine compartment along with part of the guidance system and
pressurization system. Hydraulic and nitrogen pressure lines and cables are
routed to intertank and nose section through the outside cabling duct.
G. Performance Analysis (U)
General (U)
The initial phase of the missile performance analysis
consisted of calculating the design parameters of the SCUD A missile that
would be required for a trajectory analysis. The results of this trajectory
analysis were required to perform an aerodynamic heating analysis, to determine
the drag coefficient, and to provide "g" loadings for use in the airframe
stress analysis.
n For the trajectory, an estimate of the propellant weight
was made and thrU chamber parameters were determined. Using these estimated
values, the propellant tank pressure needed was calculated. Estimates were
then made. of the weight of the propulsion components, tank pressurization
system, main power supply, guidance system, and warhead. Calculations were
performed (see Appendix II) to determine the aerodynamic heating during re-entry
of the entire vehicle.
I I The results of this analysis, together with the
estimated component weights and thrust level, were used as a basis for the
first estimate of the airframe weight. Figures 1 and 2 were used to determine
the skin thickness based on wrinkles and dents in the skin, the locations of
riveted and welded sections of the skin, and the size of the welded joints.
The trajectory analysis (Appendix I) was performed utilizing the preliminary
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performance parameters, weights, and an estimated drag coefficient. Ranges
of 150 and 80 nm were obtained for the SS-lb (SCUD) using the conventional
and the nuclear warheads, respectively (see Figures 21 and 22).
Utilizing the propellant tank arrangement (see Figure 3),
the nitric acid weght in the forward tank and the kerosene weight in the aft
tank were determined to be 6200 lbs and 1750 lbs, respectively.
Propulsion System (U)
The thrust level used as the basis for the performance
calculations is considered as an approximation of the actual value. The
effect of not knowing the precise thrust level of the SCUD A is not significant
since missile range is not too sensitive to the exact thrust level. A thrust
level of 18,500 lbs has been calculated from propulsion parameters.
The only known engine available to the Soviets with a
thrust in this s is the World War II German Wasserfall engine, which has
a sea-level thrust of 17,600 lbs and a chamber pressure of 300 psia. However,
it is doubtful that the basic Wasserfall engine-is used because of the rather
large size of the engine in relation to its thrust level. It is possible,
however, that the Soviets have continued the-development of the Wasserfall
engine -- increasing the performance, and reducing the length of the engine
to_ make it more compact. On the other hand, it is also possible that the
Soviets have not made use of the Wasserfall engine and have developed a com-
pletely new engine for use in the SCUD A missile. However, it is felt#that
the SCUD A engine would utilize a low chamber pressure in order to use thinner
chamber walls and reduce the cooling requirements. Low chamber pressures were
used in the Wasserfall, the German V-2 and its Soviet counterpart, and the
Soviet K-102 engine which operated at 310 psis.
I For the SCUD A, the thrust chamber performance
calculations were base on the use of inhibited red fuming nitric acid and
kerosene at an injector mixture ratio of 3.4, a chamber pressure of 355 psia,
and an expansion ratio of 5.15. The corresponding nozzle exit pressure is
10.3 psia, which results in good performance, considering the burnout altitude
of about 100,000 feet. The nozzle throat area was chosen to provide a sea-
level thrust of about 18,500 lbs. The motor has both regenerative and film
cooling systems to maiptain a wall temperature of 930-1110?F.
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Sea-level thrust
lb
18,500
Vacuum thrust
lb
21,500
Propellant flow rate
lb/sec
84
Sea-level specific impulse
sec
220
Vacuum specific impulse
sec
256
Expansion ratio
--
5.15
Mixture ratio
lbs
3.4
Chamber pressure
psia
355
Nozzle exit pressure
psia
10.3
Duration of thrust
sec
92
A heat transfer analysis was performed (see Appendix II)
in order to helppTetermine the kind of material used and its required thick-
ness to withstand the temperatures encountered by aerodynamic heating. The
analysis revealed the maximum temperatures over a range of various skin
thicknesses for locations along the missile beginning at the tip of the nose
cone. A study of the SCUD A trajectory ahd its velocity revealed that
aerodynamic heating was not severe until re-entry. Re-entry velocity of
about 5,000 ft/sec results in an overall body skin temperature 9f approxi-
mately 700-900?F, and the nose skin temperature of about 1,000?F. These
temperatures are critical for aluminum; therefore, stainless steel is probably
used. -
L_J The width of the welds seen in the photography (Figures
I and 2) tends to support the use of stainless steel. The waviness of the
aft skirt skin is indicative of thinner materiel compared to the tank sections.
The heat transfer analysis (Appendix II) and the airframe stress analysis
(Appendix III) support the use of stainless steel for the SCUD missile airframe.
The airframe weight analysis was based on the assumption
that the SCUD a r rame is constructed of stainless steel of series type 300.
Flight loads were estimated by making use of the propellant and component
weights and their distribution, tank pressures, and the "g" loading curve
from the preliminary trajectory. The skin temperatures at various locations
along the missile were obtained from the heat transfer analysis.
A general description of the structure of the SCUD airframe
will help to clarify the results of-the airframe stress analysis. The structure
consists of three distinct sections; the aft skirt, the tankage section, and
the nose cone (Figure 3). The aft skirt is a semi-monocoque riveted structure
consisting of a structural frame and a thin sheetmetal covering. It is felt
that most of the loads on this section are carried by the framework, and that
the skin acts mainly as an aerodynamic cover. The thrust is transmitted
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through a thrust frame to the aft support ring which is riveted to the aft
skirt and to the tankage section. The tankage section is a full-monocoque
structure consisting of a shell and ring frames with no additional support
members. The section consists of the two integral propellant tanks (the
missile skin serves as the tank wall), an intertank transition, and a short
transition section on either end. The aft transition section connects the
fuel tank to the thrust unit, and the forward transition section connects
the oxidizer tank to the nose cone. All the subsections of the tankage
section are welded together to form the complete unit. The aft end of the
warhead is attached to a conical support which, in turn, is attached to a
support ring at the forward end of the tankage section. The nose cone is a
metal shield which provides the proper aerodynamic configuration and protects
the warhead from the heat generated during re-entry. The nose cone is
connected to the tankage section just ahead of the warhead support ring.
= For the stress analysis, the weights of the support
rings, the aft skirt, and the fins were estimated by using previous experience
with similar structures constructed of steel. The weights of the tankage
section and nose cone, however, were determined from the skin thickness
which provided adequate margins of safety for the most critical load on the
section. For the tankage sections, it was assumed that the entire section
was welded up from the same thickness of stainless steel. This facilitates
construction of the section, results in good welded joints, and provides
adequate although not excessive margins of safety for all points along the
missile. The resulting thickness for the skin and bulkheads of the tankage
section is 0.081 inch, while the nose cone thickness is 0.080 inch.
Range (U)
I I The computation of range is based on the estimated overall
missile weight breakdown as tabulated in table 2.
The zero-lift drag coefficient as a function of
Mach number was based upon preliminary trajectory analysis to obtain the skin
friction contribution. The final trajectory analysis was performed by making
use of the estimated propulsion system data, missile weight breakdown, and
drag coefficient. It is seen that for the 150 nm range, the most likely
warhead weight is 1550 lbs. The design characteristics of the SCUD are
presented in table 2.
TABLE 2
.
Length
34.4
ft
Diameter
2.85
ft
Missile dry weight
3,778
lbs
Nose cone
1,550
lbs
(Nuclear)
2,500
lbs
Range
150 ran
(Nuclear)
80 nm
Apogee
41 nm
Burning time
92 sec
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MISSILE CHARACTERISTICS (U) (CONT'D)
Flight time
315 sec
Cutoff velocity
4,958 ft/sec
Cutoff angle
36.5 degrees
Cutoff acceleration
4.6 G's
Lift-off weight
11,728 lbs*l
Fueled weight
11,775 lbs*2
Burnout weight
4000 lbs*3
Thrust .
18,500 lbs
(Vacuum)
21,500 lbs
Specific impulse
220 sec
(Vacuum)
256 sec
Oxidizer
6,200 lbs*4
Flow rate
66.3 lbs/sec
Residue
100 lbs
Fuel
1,750 lbs*5
Flow rate
17.7 lbs/sec
Residue
122 lbs
Propellant flow rate
84 lbs/sec*6
Starting fuel
47 lbs I-
TOTAL 11,775 lbs
Oxidizer flow rate: 66.3 lbs/sec
Total oxidizer burned in 92 sec
Residue
6,100 lbs
100 lbs
Airframe and nose cone weight
Oxidizer flight weight
Fuel flight weight
Lift-off weight
Starting fuel
Airframe and nose cone weight
Oxidizer and fuel residue
11,728 lbs
47 lbs
3,778 lbs
222 lbs
*See weight breakdown (Table 3).
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3,778 lbs
6,200 lbs
1,750 lbs
TOTAL 11,728 lbs
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WEIGHT BREAKDOWN (U) (CONT'D)
Fuel flow rate: 17.7 lbs/sec
Total fuel burned in 92 sec 1,628 lbs
Residue 122 lbs
TOTAL 1,750 lbs
Oxidizer burned during flight/sec 66.3 lbs
Fuel burned during flight/sec 17.7 lbs
TOTAL 84 sbl
The SCUD missile is designed with a non-separating
warhead. This war ead may be one of two existing types, i.e., a conventional
HE charge or an atomic charge designated as a special warhead 8K11. These
special warheads have 10 and 15 kiloton yields.
The HE warhead consists of the nose fuze, the detonator,
J
the explosive material, the base fuze, cables, and the fuze arming device.
L __
The explosive material consists of 60% trotyl (trini-
trotoluene), 25% hexogen (RDX) (trimethylenetrinitroamine), and 157; aluminum.
In addition to this 100% mixture, the explosive material has a 5% covering
of chloronapthalene. Overall weight of the explosive charge is 1175 lbs,
and the shell container of the warhead has three layers; a steel jacket -
(0.1 in.), an asbestos carton, and a steel jacket (0.04 in.). By means of this
protective arrangement of the warhead, the temperature at the surfaces of
the explosive material does not exceed 175?F.
The nose and base fuzes are installed and .the warhead
joined with the missile at the assembly point in the technical area.
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GROUND SUPPORT EQUIPMENT (U)
General (U)
Ground support equipment (GSE) includes all equipment used
Handle, test, service, and launch the missile.
The smallest self-sustaining operational unit for this
system is the battalion since the three launching batteries are dependent
on the headquarters battery, the missile transport battery,-and the technical
battery for support. For this reason, equipment is presented by area of
employment on battalion level rather than at battery level, and only vehicles
and equipment that have a technical function are included.
GSE for the SCUD missile system is both road and rail trans-
portable. addition, there is a capability for air resupply of missiles
by either helicopter or conventional cargo aircraft.
The basic vehicle used in ground support of the SCUD missile
is the ZIL-151. The standard cargo version of this vehicle is shown in
Figure 5 and the van body version in Figure 6. The ZIL-157 (standard cargo
version shown in Figure 7) is used interchangeably with the ZIL-151.
B. Missile Ground Transporter (U)
The purpose of this vehicle cis to transport the SCUD
A missile from the epot to the technical position; to serve as a dolly
through test, checkout, fueling, and warhead installation; and to transport
the ready missile to the storage area. The vehicle is a two-axle trailer
consisting of chassis, front and rear cradles with rubber liners and straps,
draw bar, guard rails, undercarriage, and brake mechanism. This vehicle
has the following characteristics:.
Load capacity 6 tons
Weight 3.2 tons
Length 32.2 ft
Width 8.9 ft
Height 5.9 ft
Ground clearance 9.8 in
A variety of prime movers could be used to tow this trailer;
however, tost likely candidate is the ZIL-151 shown in Figure 5.
Figure 8, which shows a trailer transporter carrying a
SCUD missile, s based on a photograph that was taken.during a transport
helicopter demonstration at the Tushino air show in_1961. The ground
transporter described above is of the same general configuration but of
more rugged construction.
I
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0 ZIL-151?STANDARD CARGO TRUCK (U)
CHARACTERISTICS (U)
1.
Weight--------------------------------
6 short tons
2.
Wheel base----------------------------
191 inches
3.
Overall length------------------------
22 ft, 9 in.
4.
Width---------------------------------
7 ft, 7 in.
S.
Engine--------------------------------
92 HP 6 cyl
gasoline
6.
Speed---------------------------------
41 mph
7.
Cruising range------------------------
413 miles
8.
Payload-------------------------------
5.2 short tons
9.
Towed.load----------------------------
4 short tons
Figure 5. (UNCLASSIFIED) ZIL-151 Standard Cargo Truck (U)
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This equipment is mounted in the van body of a ZIL-151
truck similar to that shown in Figure 6. The purpose of the equipment is to
provide a capability for a complete pressure testing of the propulsion system
and its components. The equipment consists of manometers, a voltmeter, and
an ammeter. Also included is equipment for filling the spherical tanks in
the missile with compressed air to check the correct setting of the reducer,
stands, piping, valves, etc.
The electrical system test equipment, like that for
the propulsion system test, is mounted in the van body of a ZIL-151 truck
similar to that shown in Figure 6. Included in this van is a console and a
mock-up of the missile electrical system for checkout of guidance system
components, and a console for checking insulation of the cable net. Guidance
system checkout includes check of pickup current of steering motors, check of
functioning of the autopilot, and check of accelerometers.
This unit is used for filling the missile propellant
tanks with compressed air and for pressurizing the air bottles on the trans-
porter-erector-launcher (TEL) (230 to 350 atmospheres). The unit. includes
the power plant, compressor, air system, cooling system, regenerating unit,
oxygen-water condenser, moisture absorber, and control panel. The unit Is
mounted on a ZIL-151 truck and has the following.characteristics:
A photoelectric automatic moisture indicator is used to
control the humidity of the air delivered by the compressor or from the
bottles. It consists of a measuring head with cooling system, an electronic
unit, a power unit, and a front cover. The measuring head consists of a
measuring mirror, illuminating bulbs, two photoelements, two objectives,
and diaphragms. The cooling system consists of a throttle, a coil, a pre-
heater, and a heat transfer unit. The power unit is fed with a 220 volt
source.
Propulsion System Test Equipment (U)
Electrical System Test Equipment (U)
Air Compressor Unit (U)
Also included as part of the air compressor unit is an
air preheater consisting of an electric motor driven blower, a gasoline burner,
and grids through which the air is blown. The exhaust air temperature is_
250?F.
Overall weight
10
tons
Weight w/o ZIL-151
3.5
tons
Air delivery
53
cu ft/min
Engine
55
to 60 HP
Compressor speed
1250 to 1800 rpm
Maximum speed of vehicle
25 mph
Range (governed by fuel)
400 miles
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6X1
F.
The electrical power supply for the SCUD requires two
X1 -
pieces of equipment in the test area. A gasoline electric generator similar
to that shown in Figure 9 is used as the basic power source of 220 volt
alternating current and probably has a capacity rating of 25-30 kw. The
rotary converter or transformer trailer similar to the one shown in Figure
10 is used in conjunction with the generator as a source of direct current
supply.
The purpose of this vehicle is to transport kerosene
is shown in Figure 11 and has the following characteristics:
for the SCUD missile to the fueling site and there transfer it to the missile
tank by means of the integral pumping system mounted in a compartment on the
vehicle. Kerosene tanks are constructed of steel with a zinc coating. The
tank and pumping system are mounted on a ZIL-151 truck chassis. This vehicle
Weight loaded
Capacity
Working capacity
Weight of contents
Method of transfer
Maximum working capacity
10 tons
800 gallons
780 gallons
3 tons
Pump
95 gpm
Oxidizer Transporter (U)
The purpose of this vehicle is to transport nitric
acid for the SCUD missile to the fueling site and there to transfer it to
the missile oxidizer tank by means of the integral pumping system mounted
in a compartment on the vehicle. Nitric acid tanks are constructed of
aluminum and aluminum alloy. The tank and pumping system are mounted on
a ZIL-151 truck chassis and in general appearance would be similar to the
kerosene transporter shown in Figure 11. This vehicle has the following
characteristics:
Weight loaded
10.6
tons
Capacity
810
gallons
Working capacity
740
gallons
Weight of contents
3.4
tons
Method of transfer
Pump
Maximum working capacity
95 gpm
1 Washdown and Neutralizing Vehicle (U)
The purpose of this vehicle is to provide equipment to
remove and/or ren er harmless any propellants spilled during transfer to the
missile tanks. The unit consists of a tank and integral pumping system with
hoses mounted on a ZIL-151 truck chassis. Its general appearance would not be
too unlike that of a fuel transporter, and it has the following characteristics:
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Switch panel with cable plugs
Figure 9. (UNCLASSIFIED) Electric Generator (Generator Trailer) (U)
Automatic voltage regulator
Figure 10. (UNCLASSIFIED) Rotary Converter (Transformer Trailer) (U)
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Weight loaded
10.1
tons
Capacity
500
gallons
Working capacity
500
gallons
Weight of contents
4200
pounds
Pumping rate
250
gpm
Fire Fighting Vehicle (U)
The purpose of this vehicle is to provide equipment for
fire protection for the technical area. The unit consists of a tank, integral
pumping system for filling and emptying the tank, and suction and discharge
hoses mounted on a ZIL-151 truck chassis. It has the following characteristics:
Weight loaded
Capacity
Working capacity
Weight of contents
Pumping rate
Crane (U)
The crane is a truck-tractor and semi-trailer combina-
tion as shown in Figure 12. It has a..30 kw alternating current
generator and two electrically-driven motors. One motor is used to power the
winch that controls the boom, and the other is.used to power the winch for
the hook.
The crane has the following characteristics:
Weight of semi-trailer
7.8
tons
Weight of truck-tractor
14.9
tons
Length of combination
49.5
ft
Height in mobile condition
10.5
ft
Height in working. position
28.2
ft
Width
11.5
ft
Maximum lifting capacity
13.7
tons
Hoisting speed
3.95
ft/min
Maximum elevation of hook
23 ft
This crane is used at the assembly point in mating the
warhead with the missile, and is also used to transload the missile from
the ground transporter to the TEL.
The purpose of this vehicle is to transport the SCUD
A missile from the ready missile storage area to the launch site, erect it
on the launch pad, and provide facilities for final preparation and launching.
It consists of: the body; engine mount; transmission; chain drive; suspension;
missile booms; launch pad; test-launch equipment; aiming instruments complex;
equipment for loading the missile with compressed air and starting fuel; the
hoisting mechanism of the boom; general electric equipment; radio-equipment,
with a range of approximately 25 miles; and fire fighting equipment. The TEL
9.8 tons
515 gallons
515 gallons
4300 pounds
250 gpm
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is shown in figure 13 as it appeared in a Moscow parade, and has the following
characteristics:
Weight without missile
Weight with missile
Chassis length
Length with boom
Width
Height in working position
Maximum speed
Cruising range
Average specific pressure
Maximum ascent and descent
Lateral angle of tilt without
missile
Lateral angle of tilt with missile
Capacity of tanks - fuel
Capacity of tanks - oil
Capacity of tanks - water
Engine - 12 cyl tank diesel
Fording depth
Figure 14. (CONFIDENTIAL) Transporter-Erector-Launcher with Missile in. Launch
Position (U)
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Figure 14 shows the TEL with the
35 tons
41 tons
23 ft
94 ft
.8 ft
9.4 ft
6 mph
187 miles
0.046 lbs/sq
25 deg
20 deg
16 deg
234 gals
34 gals
22 gals
520 HP
4.6 ft
launching position, with the boom in the vertical
position.
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NM4V;) 6 ti
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5X1
Heteorological Equipment (U)
t rolo Leal station of the SCUD missile system
h
eo
e me
g
I I Z
is part of the equipment operated by a platoon of the battalion headquarters
battery. The basic equipment for this station is carried in a ZIL-151, with
a van body similar to the one shown in Figure 6, and includes standard
thermometers, psychrometers, aneroid barometers and wind measuring devices.
This vehicle serves as a processing point and is designated as a mobile met
office. Vehicles for a meteorological platoon probably include one mobile
met office; two cargo trucks, each towing an electric generator; and a radar
with prime mover. An assumed typical station layout is shown in Figure 15.
33
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 Approved For Release 2002/11/04: CIA-RDP78TO5439A000200370043-6
SECRET
1 IV.
1
ORGANIZATION AND FIELD OPERATIONS (U)
The tactical operational unit for the SCUD (8A61) system
is the missile battalion. It is composed of a battalion command group, a a
headquarters battery, three launching batteries, a technical battery, and .missile transportation battery. The headquarters battery consists of a
surveying platoon, two signal platoons, and a mobile meteorological artillery
station. The launching battery consists of a 20-man launching section,
one 12-man electrical firing section, and an-8-man compressor and battery
charging station unit. The technical battery is composed of three missile
testing stations, a section for handling special fuel, and one assembly section.
.The transport battery has two missile supply sections, one missile r.gging
platoon, one section for supply of special fuels, and a dry missile depot.
.By utilizing the one launch facility in each battery, the battalion has a capa-
bility of launching three missiles simultaneously every 2 hours. The fueled
missile normally moves by road on its JSU tracked vehicle from the forward
missile storage depot to the launch area. It has an off-road capability and
is air transportable. It is integrated into all Soviet tactical forces and
is employed in troop maneuvers and other training exercises for the achieve-
ment of combat effectiveness. Its relative short-range capabilities limit
its use to front line objectives.
B. Line of Vehicle March (U)
al convoy of a launch section moving within rear
i
c
I I A typ
areas would include a command car followed by an instrument vehicle, a
launching unit equipment vehicle, and the transporter-erector-launcher
vehicles -- including a test launch set, a dry missile transporter;, a neutral-
ization and washdown vehicle, and personnel carriers. Only the amount of
ground support equipment required-to launch the missile would be moved beyond
the forward dry missile depot storage area, which comprises part of an area
known.as the Technical Site.
C. Processing the Missile (U)
s are loaded in containers and moved by rail
il
i
e
ss
I I The m
to a forward area supply depot -- 3 missiles per special car and up to 60
missiles on one train. Nose. cones are transported separately -- five in
2-axle cars, and twelve in 4-axle cars. From the supply depot missiles are
moved by vehicle to the Technical Site, approximately 50 km closer to the
forward edge of the battle area (FEBA). Here the dry missiles are held
temporarily in a storage depot until.preparation of the missile for firing
takes place, in which event it is processed through a serieshofftesting
points. These tests include the ground control equipment, fuzing system,
the propulsion unit, and the guidance system. These.and other technical
checkouts and services discussed elsewhere herein are made at the test points,
the fuel loading point, and two points for the assembly of components (all
within the Technical Site).-- including the warhead. The missile is then
held at the ready missile storage point to be picked up by the using unit.
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Field Deployment (U)
Current doctrine for deployment of the SCUD missile
battalion prescr es dispersion of two launch batteries along a front of 12
to 14 km. The Command Post, third launch battery, meteorological sites, and
service units are echeloned in depth up to 8 to 10 km (Figure 17). This
? battalion area is normally located not closer than 30 to 40 km from the FEBA.
Criteria for selection of the launch area is that it should
provide for coverage of all assigned targets, deployment of all combat
elements of the battalion, camouflage of all personnel and combat equipment,
ease of control of fire, and convenient and concealed roads or trails.
E. Launching Site (U)
Batteries are normally held 15 to 20 km to the rear in
a waiting a a pen g movement to the launch area.
At this pre-surveyed position, little remains to be
installed in the missile. Final operations include an inspection of the engine
assembly and the guidance system, and a check of the insulation of the cable
network. The missile is mounted vertically on the launch platform and the
battery is installed. A collimator is fixed on a designated point and the
missile is layed. Storm moorings are affixed when the wind velocity exceeds
20 mph, a gasoline generator and a quartz generator (a frequency control
device) are started, and the missile is launched. Fu0 1 geodetic preparation
for firing the missile requires 3 days far completion.
Missile troops are deployed with the Group of Soviet
Forces in Germany (GSFG), the Southern Group of Forces (SGF), the Turkestan
Military District (Turk VO), the Transcaucasian Military District (Zak VO),
and the Far Eastern Military District (DVO). It.is likely that the SCUD is
operational throughout all of these areas. The USSR has probably supplied
the conventionally armed SCUD missile system in small quantities to both East
Germany and Communist China. it has been produced in the USSR in large
quantities, is an efficient weapon, and is well suited for training of troops
with little or no missile experience.
c_
mumm (RI
Tactical exercises involving Soviet missile troops
during 1960 and 1961 revealed a weakness in combat readiness of missile
technical units (Mobile Repair Technical Base - PRTB). At no time during
these exercises were missiles delivered by the PRTB to the missile units in
condition to be fired without undue delay. Soviet military leaders realize
the necessity for developing close coordination and cooperation between
missile units, missile technical units, and rear area service type units --
and they can be expected in the future to place extreme emphasis on correct-
ing or at least greatly improving these shortcomings.
I
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 SECRET
-ASSEMBLY AND CHECKOUT PROCEDURE (U)
Possible deployment concepts of the SCUD have been presented
in numerous publications since the missile's first public appearance.
Although these various concepts cannot be verified, it is certain that the
missile and its components must be thoroughly tested before launch. The
actual deployment and checkout by troops in the field may be done in the
following manner. The phases through which a missile must pass after
manufacture are:
Storage. All components are held.in a ready state at a
for the troops.
Testing. After the troops are issued a missile system it
to test all components together.
Loading and Assembly. The missile is moved to the fueling
area and the assembly area.
Testing (U)
After being issued at the supply depot, the
associated components are transported to the technical
position, an area some distance away from the launch area, where testing
and loading and assembly are accomplished. The equipment is disassembled
and prepared for testing. Individual tests are made on every unit of the
guidance system and fuzing equipment as well as the leak test on the rocket
motor, fuel tanks and associated plumbing and valves. The missile remains
in a horizontal position for these detailed tests or'checks.
The testing position is divided into two work
areas, work area one for propulsion system testing and work area two for
electrical system testing, where special checkout consoles are erected for
sequentially operating flight components of the missile in individually
controlled simulated flight tests.
Loading and Assembly (U)
Operations at the Fueling Point (U)
Two working areas are organized at this point:
one for fueling the missile with kerosene, the other for adding the oxidizer.
The maximum loading speed is 250 liters per minute (66.25 gal/min).
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 SECRET
and joining point on a
Assembly and Joining Procedure (U)
After fueling, the missile proceeds to the assembly
carrier. The warhead is brought up in a special con-
tainer. The fuzes are also delivered to this point in special packaging.
missile. A crane is used to-assist in joining the warhead to the missile. All
connecting surfaces have been smeared with a fireproof substance. The entire
missile is then placed on the tracked erector-launcher and moved to the launch
area and turned over to the launching section of the firing battery.
After the missile arrives at the launching area there
is a sequence of operations which is conducted to assure a successful launch.
First, a thorough inspection is made of the engine assembly, the guidance
system, and the gyroscopic instruments which are already mounted in the missile.
Missile Alignment on Launcher (U)
The n the previously surveyed point wi
th a collimator,
which has been placed upon a point designated by the geodetic personnel. The
missile is aligned,. with fin I in the direction of fire, and a vertical check
is made. Leveling is probably accomplished by adjustments on the leveling
frame on the launcher.
Aiming the missile at the target requires more than
just orienting the body of the missile. The gyro plate must be leveled and
properly balanced and the missile must be perfectly vertical. Balancing the
gyro plate implies the orientation of the gyro plate, together with the
gyroscopes mounted thereon, with respect to the stabilizer fin surfaces.
measuring the angle
A rod and an optical quadrant which are designed for
of inclination of the stabilizer surfaces to the horizontal
are used to balance the gyro .plate. The rod is fastened to the adapter of the
gas-jet vanes. The optical quadrant, which consists of the body, a dial, and
the traversing mechanism of the dial, is fastened on the control plate of the
rod. A scale in degrees is painted on the dial. There is also a vernier of
the. dial which has a grid with 1-minute angular graduations. The optical
quadrant provides a measurement of the angles +120?. In aiming the missile
at the target, an instrument set is used. The whole set is comprised of the
collimator (as one unit of the set), a horizontal sight, two illuminated aiming
stakes and two magnetic levels. These magnetic levels are for determination
of the magnetic deviation peculiar to the particular launching site and will
indicate the magnetic dip.
Continuity checks are made of the cable networks
by a megohmmeter. The main missile batteries are installed and the grounding
39
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The final launch operations include a number of
checks of the missile that cover the pyrocartridge circuit, the guidance
system and a first and second general check. These checks are made with auto-
matic test equipment and consist of a series of sequential events that are
initiated by the test equipment; test results are indicated in a go-no-go
display on the launching consoles.
connection is made. Storm moorings are attached if the wind velocity exceeds
10 meters/sec.
I J It would appear that no problems should exist for
the launching crews after the missile is erected and aligned and the plug
connections are made. The high-pressure feed connections appear to be an
integral part of the erector launcher and are most likely connected at the
time the missile is placed on the launcher. Otherwise the actual firing is
automatic and controlled by a console.
Aiming Requirement (U)
The geodetic survey of the launch point requires about
3 days which indicates that a first order survey is mandatory in order to
achieve the target accuracy required of the guidance system. The use. of a
collimator, an instrument to measure horizontal and vertical angles and to
determine the aiming point, in the survey supports the conclusign that this
high-order geodetic survey is a necessary measure.
The collimator consists of a tube, a sight, au upper
plate, a dial (or angle-measuring ring), a lower plate with a fixed section,
together with adjusting screws and auxiliaries. There is the usual tripod,
small battery (2.4V), and a cover. The optical system of the collimator
consists of an objective and a grid.
I Objective. The objective, consisting of three sections,
is a cone at one of whose focal points is placed a small source of
light. Rays diverging from this focal point emerge from the objective lens
in a parallel beam.
Grid. The grid is painted on the surface of the lens
of the third objective section and is placed in the focal plane of the col-
limator objective. The grid circle is divided into 76 vertical stripes. On
the left half of the grid, numbers are inscribed in a vertical and horizontal
arrangement from 1 through 18. The right half is filled with 18 very narrow
letters which are also arranged in the same manner as the numbers,- (See
Figure 18.) A ground glass provides uniform illumination of the grid. The
grid is illuminated by :a bulb when working under night conditions and is
illuminated' during the;'day by means of a reflecting mirror.
I ILevels. A cylindrical level and two spherical- levels
are used to level a collimator horizontally. The cylindrical level is
graduated to 1 minute;lthe spherical levels are graduated to 20 minutes.
41
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1 -1 Sight. A panoramic sight, consisting of an objective,
a rotating prism, grid and eyepiece, is fastened to the body of the collimator
tube and is used to aim the missile at the target. The angle of the field
of vision is-10 degrees. The limit of the horizontal measurement is 60-00;
the vertical angle is +6-00. The grid has an upper raw of numbers used for
aiming at the target and a lower row of numbers for balancing of the horizontal.
The horizontal sight is balanced when its optical axis, in the zero settings
of the angular scale, is in the plane with the perpendicular control stripe -
of the table. For a horizontal balance, the collimator must be fastened to
the control table so that,. markings on the collimator are superimposed on
identical markings of the horizontal plate. The horizontal plate is turned
with its head down (to 1800). By means of the angular scale of the horizontal,
the identical markings of the collimator and the horizontal plate are super-
imposed. The reading of the angle is taken from the horizontal angular scale.
Screw clamp. A screw clathp, made up of a body with
supports, a swinging mechanism, a telescope socket, two cylindrical levels
and a control table, makes it convenient to balance the horizontal plate and
install the collimator on either a tripod or the launcher.
The above aiming procedure is based primarily on the V-2 type
procedure. (Reference: Operation Backfire.)
42
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SUPPORTING ANALYSIS (U)
Introduction (U)
This appendix includes all calculations used in the evalua-
tion of t e soviet SCUD missile with the exception of the Heat Transfer
Analysis and the Airframe Stress Analysis.
Fr-6 Conclusions based upon these calculations are found in the
body of t y.
II
Discussion (U)
Airframe Dimensions (U)
re 1 shows a side view of the SCUD missile taken
F fi
gu
~~
at the 1960 Maguey' parade. This photograph vas utilized to determine the
distances between the major circumferential welded joints, the length of
the nose cone, and other pertinent data needed for this analysis. The
overall missile length and diameter were obtained from the photo intelligence
study of the SCUD A missile seen at the 7 November 1957 Moscow parade.
Tankage (U)
The tank volume calculations are based upon the end
closure locations and missile diameter, as shown in figure 3.` Two end
closures, one on each tank, are assumed to be hemispheres. The remaining
two closures are assumed to be of the 2 to 12 semi-elliptical typse. The
volume of the fuel tank (V fuel) is :
D2L + f r' + D3
Ir'
V fuel =
where:
D = Diameter, ft
L = Length, ft
Vfuel = 3.14 (2.8)2 (4.4) + 3.,714 (2.8)3 + 3_.14+ (2.8)3
Vfuel = 6.16 x 4.4 5.7 = 27.1 + 5.7 + 4.2 = 37.0 ft3
75 for feed line volume through
btract
t
.
su
From the 37 cu ft, one mus
fuel tank, and 1.85 cu ft for unusable residue. Therefore, 34.4 cu ft
is the usable fuel volume. .
The volume of the nitric acid tank, V RFNA' -
VIRFNA = 6.16 x 9.0 + 5.7 + 4.2 =.65.3 ft3
+ 0.75 for feed line volume = 66.05 ft3
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 SECRET
where
Using a 1+;'% ullage,the total weight of fuel is:
W = p x v x 0.96 = 53 x 34+ . 4 x 0.96 = 1750 lb
W = Weight, lb
= Density, lb/ft3
Using an 7% propellant-tank-volume outage, the usable weight of fuel is:
W = 1750 x 0.93 = 1628 lbs
is calculated L
By the same procedure, the total weight of nitric acid
W =.P x V = 93.9 x 66.05 = 6200 lb (ullage included)
Using approximately 1.6% propellant tank volume outage, the usable weight
of nitric acid is :
11 = 6200 x o.984 = 6100 lb
It is felt that the Soviets could easily attain the ullage and outage
figures quoted above; therefore, the analysis is conservative from a
propellant utilization point of view.
Since approximately 9 % of the nitric acid flow rate is
used for filing of the thrust chamber, the amount entering the injectors
is:
W = 6100 x .91 = 5551 lbs
The mixture ratio at the injectors is:
1,..R. =5551=3.4+
162 _
A s:ary of the propellant weights and volumes is
-
-
-
provided in tab
l
e
PT.
TABLE 4.
CALCULATED TAUK
VOLU1.ES
A:-JD PROPELL4 TT ? TEIGHTS
(U)
OXIDIZER
FUEL
U1,TITS
(F',i, ITA
(Iff '.OSE;1E)
TOTAL
Propellant tank volume
ft3
66.05
34.11.
100.11-5
Total propellant weight
lb
6200
1750
7950
Usable propellant weight
lb
6100
1628
7728
Propellant flow through injectors
lb
5551
1628
7179
?ilri cooling
lb
549
0
549
Outage
lb
100
122
222
44
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The thrust chamber performance calculations are
Performance Calculations (U)
Thrust Chamber (U)
based upon the following values:
Chamber pressure (Pc) = 355 Asia
nitric acid/kerosene.
For a mixture ratio of 3,4 and a chamber pressure of 355 psia, the theoretical
value of characteristic exhaust velocity (C-) is:
C=am- = 5415 ft/sec (C:: = PCAt = PcAtg = 355 (39.8) (32.2) = 51+15 ft/sec)
For the ratio of _peclfic heats (k) of 1.18, and the expansion ratio of
5.15, the ratio of nozzle exit pressure (Pe) to chamber pressure (=c; is:
Pe = 0.02901
Pc-
Thus, the nozzle e:cit pressure is:
Pe = 0.02901 x 355 = l0.3 psis
For k = 1.18 and E = 5.15, the theoretical value of vacuum thrust coefficient
(Cf ) for a 15-degree, half-angle, conical nozzle is:
VAC
C = 1.522
fVAC
The theoretical sea-level thrust coefficient (Ci is:
SL
P..
CfSL CfV:~C PC E
Propellant combination = inhibited red fuming
Mixture ratio (bi.R.) = 3.1+ (at injectors)
Throat diameter (Dt) = 7.118 in = 39.82 in (At)
Exit diameter (De) = 16.2 in = 2052 in
Expansion ratio (.E) = 5.15
45
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 SECRET
where:
Pa = Ambient pressure, psia
Cf = 1.522 - 14.7 x 5.15 - 1.522 - 0.213 - 1.309
The sea-level and vacuum. specific values (Isp)Gcan be calculated as:
Cf C*
Isp = SL = 1.309 x 5415 - 220 sec
SL g 32.2
Cf C*
Isp VAC = 1.522 x 5415 = 256 sec
VAC g 32.2
For the throat area (AT) of 39.8 in2, the thrust can be calculated from:
T = CfPcAt
T = 1.309 x 355 x 39.8 = 18,500 lb
SL
T = 1.522 x 355 x 39.8 - 21,500 lb
VAC
The weight flow rate (w) can be calculated as:
T = 18,500 - 84 lb/sec
Isp 220
From the total propellant flow rate, mixture ratio at the injectors, and
percentage of the nitric acid used to film-cool the thrust chamber, the
oxidizer and fuel flow rates are calculated as:
wIRFNA = 66.3 lb/sec
*fuel = 17.7 lb/sec
TOTAL = 84 lb/sec
b. ^ Performance Characteristics
(1)
T = 18,500 lb -- T = 21,500 lb
SL VAC
(2) 1 For the usable propellant weight of 7728 lbs, the
t = 7728 lb = 92 sec
84 lb/sec
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(3) The overall specific impulse is the net thrust
divided by the tot ropellant flow rate:
IS= 18,500 = 220 sec
pSL 84
(4) 'sr' A summary of the performance characteristics of the
SCUD propulsion sy~is shown in table V.
TABLE 5. PROPULSION SYSTEM PERFORMANCE CHARACTERISTICS (U)
Oxidizer
Red fuming nitric acid
Fuel
Kerosene
Mixture ratio
at the injectors
3.4
Chamber press
ure psi
355
Total propellant flow rate
lb/
ec
84
Sea-level specific impulse
sec
220
Vacuum specific impulse
sec
256
Sea-level net thrust
lb
18,500
Vacuum net thrust
lb
21,500
Duration of thrust -
sec
92
4.. Propulsion System Weight (U)
The propulsion system weight can be calculated based upon
current Unite tea state-of-the-art and multiplying this value by a factor
which takes into account the differences between current United States and
early 1950 Soviet state-of-the-art. The factor can be determined by comparing
the calculated weight to the actual weight of a known Soviet propulsion system;
for this comparison, the Soviet K-102 engine was chosen. It is believed that
this engine is essentially a V-2 engine with the thrust increased from 56,000
lbs to 80,000 lbs at sea level. The increase in thrust was obtained by in-
creasing the chamber pressure, maintaining essentially the structure of the
original V-2 engine. Since the known Soviet engines produced during the
development'of the SCUD missile were of the same state-of-the-art as the K-102
engine, the calculated weight of the SCUD propulsion system could be approxi-
mately correct.
Based on current United States state-of-the-art, the weight
of the K-102 propulsion system, less tutbopump, would be about 1140 lbs, while
the actual weight (assuming that it is a V-2 engine) is 1360 lbs. Thus, the
ratio utilized to convert current United States state-of-the-art propulsion
system weights, less turbopump, to early 1950 Soviet state-of-the-art is:
ratio = 1360 = 1.2
1140
Isp = 21,500 = 256 sec
VAC 84
47
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Based upon current United States state-of-the-art, the
weight of the S propulsion system, less turbopump, is 460 lbs. Increasing
this weight by 207 to convert to early 1950 Soviet state-of-the-art, the
weight is:
However, this weight is based upon an expansion ratio of 3.4, as used on the
V-2 engine. For a more probable expansion ratio of 5.15, the weight would be
increased by about 30 lbs. Thus, the estimated total propulsion system weight
is approximately 580 lbs.
5. + 1 Propulsion System Vol$me (U)
From the tankage location (Fissures 1 and 3) the length of
the propulsion compartment is 7.2 feet. The forward 1.4 feet of this length
is occupied by the rear bulkhead of the fuel tank. The apparent nozzle exit
is about 0.4 foot within the aft end of the missile. Thus, the usable length
of the propulsion-compartment is about 5.4 feet.
L With a sea-level thrust of 18,500 lbs, a chamber pressure
of 355 psia, and-a] characteristic chamber length of about 100 inches (the
same as that of the German Wasserfall engine), the approximate length of the
thrust chamber is 4.0 feet. Subtracting the 4.0 foot thrust-chamber length
from the 5.4 foot usable length of the propulsion compartment leaves 1,4 feet
between the thrust chamber and tankage. There. is adequate room between the
thrust chamber and tankage to house the pressurization bottle, the MPS, and
the autopilot amplifier.
It is concluded that the available propulsion compartment
is of the prope ze to house a conventional pressure-fed thrust chamber
with a sea-level thrust in the range of 18,000 lbs.
Nitric Acid Tank Pressurization System (S)
The nitric acid tank is assumed to be pressurized
by hot gases from a small hydrogen peroxide gas generator. The water and oxygen
of these gases are compatible with the acid. These gases are produced in ac-
cordance with the following equation:
2 H202 f
2 H2O + 02
Therefore, two moles of liquid are required to produce three moles of gas.
The resulting mole ratio is:
The number of moles of gas required to pressurize the 66.05 cu ft acid tank
to 500 psia can be determined from the following equation:
n = P V
R T
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where:
P = Tank pressure, psis
V = Tank volume, cu ft
Tank gas temperature in OR
A 937. hydrogen peroxide (H202) monopropellant is used in these calculations.
It has a density of 88.7 lb/ft3 and a chamber temperature of 1483?F. A tank
gas temperature of 540?F (1000?R) is assumed.
n = 500 x 66.05 = 3.08 lb moles
10.73 x 1000
Pound moles of R202 = 2/3 (3.17) = 2.05 lb moles
The weight of hydrogen peroxide required is:
= 3.05 (35) = 71.8 lbs
Wg202
The volume of hydrogen peroxide required is:
Vg = 71.8 = 0.81 cu ft
202 88.7
The diameter of two spheres containing this volume is determined by the fol-
lowing expression:
D= 6 V 2 = 6 x 0.81 = 0.918 ft
3.14 (2)
Using a material similar to 321 or 347 austenitic stainless steel, with a
yield strength of 180,000 psi and a factor of safety of 3.0 the design stress
(s) or 60,000 psi is used. The thickness of the sphere is calculated as:
t = PD = 710 x .918 x 12 = 0.0326 in
We 4 x 60,000
The total weight of the two spheres is:
W = 21rD2t1 = 2 x 3.14 x ?(0.918)2 x .0326 x 489 = 7.03 lbs
12
The weight of the hydrogen peroxide spheres is therefore estimated to be 10
lbs. The weight of the gas generator, which is 3 inches long and 3 inches
in diameter, and the piping and control valve is estimated to be 5 lbs.
49
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Kervseue Task ~essurl I Ic,a 0-vAtaa %)
The kerosene tank is assumed to be'pressurized by hot
gases from a small et lene oxide (C2H40 gas generator. The methane and
carbon monoxide from this reaction are compatible with the kerosene. These
gases are produced in accordance with the following equation:
C2H40 C H4 + CO
Therefore one mole of liquid is required to produce two moles of gas. The
mole ratio is:
mole ratio = 1/2
The number of moles of gas required to pressurize a 34.4 cu ft fuel tank to
500 psis is determined by the following expression:
n = 500 x 34.4 ? = 1.603
10.73 (1000)
This calculation is based upon 1900?F chamber temperature for C2H40 and a
tank temperature of 540?F, or 1000?R. The C2H40 has a density of 54.3 lb/ft3
at 60?F.
Pound.moles-of C2H40 = 1/2 (1.603) - .802 lb mole
Weight of C2R40 = .802 (4.5) = 36.1 lbs
Volume of C2H40 = 36.1 - .665 cu ft
54.3
The diameter of two spheres containing this vol- is determined by the
following:
D = 3 x .665.= 0.866 ft
The total weight of these two spheres, the gas generator, control valves, and
piping is estimated to be 15 lbs.
Summary (U)
The total pressurization system weight is 140 lbs. A
detailed breakdownhis figure is given in table 6.
TABLE 6. ESTIMATED TANK PRESSURIZATION SYSTEM WEIGHT BREAKDOWN
ITEM NITRIC ACID TANK FUEL TANK TOTAL
Monopropellant 71 lb's 39 lbs 110 lbs
Propellant bottles, valves, 15 lbs 15 lbs 30 lbs
and gas.generator
TOTAL 86 lbs 54 lbs 140 lbs
50
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 SECRET
The tank wall (also the missile skin) is assumed to be
0
Propellant Tank Weights (U)
Nitric Acid Tank (U)
constructed of a marial similar to heat-treated 4130 austenitic stainless
steel. This steel has excellent weldability; its joints and weldments are
exceptionally tough and ductile. Heat-treated 4130 steel has a yield strength
of 160,000 psi at 600?F: A factor of safety of 1.5 is used to get a design
stress(s) of 106,600 psi.
The thickness of the steel cylindrical wall is calculated as:
t = PD = 500 x 2.8 x 12 = 0.0788 in
2s 2 x 106,600
The thickness of the steel domes are calculated as:
t = PD = 0.0812 = 0.0394 in
The total weight of the acid tank is:
W = 11D2tA0 +fDLtf
3.14 x 2.132 x 0.0394 x 489 + 3.14 x 2.8 x 9 x 0.1288 x {+89
12
W = 40 + 254 = 294 lbs
thickness of this tank is the same as the acid
Th
I I
e
tank since it has Lu same diameter, is made from the same material, and
contains the same pressure. The total weight of the fuel tank is:.
W = 3.14 x 2.82 x 0.0394 x 489 + 3.14 x 2.8 x 4.4 x 010788 x 489
2
12
W = 40 + 124 = 164 lbs
8. II Missile Power Supply (U)
The missile power supply (MPS) is assumed to be a battery-
IJ
powered system comparable to the system on the V-2 missile; that is, alternat-
ing current is supplied by a rotary inverter.
For the guidance system, it is estimated that approximately
65 watts of power are required, 907. of which is supplied by an alternating
current power supply. Thus, a-c power supplied by the rotary inverter is 65
x 0.90 = 58.5 watts. Rotary inverter efficiency is about 607.; therefore, the
power input is (58.5 x 0.60 =) 75 watts. A separate 50 VDC command voltage
supply is used for attitude sensor voltage, actuator feedback voltage and
51
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velocity cutoff circuits. The total current required will be 2 amperes or
less. Power required for this supply is provided by the inverter (rectified
and filtered). The power requirement for the missile command voltage is 100
watts. DC actuators for a missile of this range during this time period re-
quire peak currents of 100 amperes. An average current drain of 30 amperes
is considered conservative; therefore, power needed for the actuators would
be (32 volts x 30 amperes =) 960 watts.
The total battery output power is:
Power for autopilot 75 watts
Power for actuators 960 watts
6
Power for command voltage 100 watts
TOTAL 1,135 watts
Using a 50% margin of safety for the battery, the total battery capacity should
be at least (1,x35 watts/32 volts =) 35 amperes for about 2 minutes. Batteries
for this capacity during this time period are known to have weighed about 20
lbs. Two such batteries could'supply more than sufficient power for the rotary
inverter and actuators. Battery weight is estimated to be:
Rotary inverter (100 watt class) during this time period are estimated to have
weighed 15 lbs. The weight of the four servo actuator assemblies to control
the jet vanes is estimated to be 80 lbs.
The MPS weight summary is shown in table VII.
TABLE 7. ESTIMATED MAIN POWER SUPPLY WEIGHT BREAKDOWN (U)
ITEM
Batteries.
Rotary inverter
Frequency regulator
Miscellaneous
5
70
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Figure 19 shows the center of pressure vs Mach number
for the SCUD missiT-eJ The center of pressure is considered as that point
along the axis of the missile where the total lift force is acting.
The missile has two major components contributing to
the lift of the vehicle near the zero angle of attack. These components are
the nose cone (with its attendant carry-over on the body) and the fins.
The lift of the nose cylinder at supersonic speeds was
estimated by interpo acing curves in reference 1, Appendix V. These curves
show good correlation with experimental data for this application. For sub-
sonic and transonic speeds,_a slender-body theory was used for the nose con-
tribution. The carry-over of lift from nose to cylinder was estimated by
extrapolating experimental data of reference 2, Appendix V.
The supersonic lift of the fins and the fin centers of
pressure were determined by linear theory and the design charts and methods of
reference 3, Appendix V. Lift of the fin at Mach 6 was determined by the ex-
pression derived in reference 4, Appendix V. The subsonic fin conditions were
determined from wind-tunnel data presented in reference 5, Appendix V. These
data were for a fin with the same planform, but with an NACA 63A series, 4%-
thick airfoil instead of the 4%-thick, single-wedge fin of the vehicle. This
discrepancy is of small effect, however, since the fin planform is much more
important than fin profile in this case.
The cabling tunnel was ignored in the aerodynamic analy-
sis. Its contribu n to the total lift is very small and, because of its
length, it would have little effect upon the center of pressure. It may cause
some-turbulent flow over the fins, but this would be minor as its height is
small, the area of the fins near the body is in turbulent flow, and the cabling
tunnel is interdigitated about -5 degrees with respect to the fins.
The methods used in determining the center of pressure
of the total vehicle give an accuracy of ?5% of the body length.
U Figure 20 shows the zero lift drag coefficient versus
Mach number curve used for the SCUD missile.
c. = Range Optimization (U)
(1) 1 Method of Analysis (U)
The trajectory analysis was based on the following
equations. The rate o change of missile velocity is given as:
53
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 SECRET
is
2 3 4 5
MACH NUMBER
Figure 19. (SECRET) Center of Pressure vs Mach Number Curve for SCUD Missile (C)
54
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 SECRET
Cd
2 3
MACH NUMBER
Figure 20. (SECRET) Drag Coefficient vs Mach Number Curve for SCUD Missile (C)
55
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V = T (cos,* )-D - g sin
(v = rate of change of velocity
g = go (r /r)2
go = acceleration of gravity at sea level
ro = radius of earth
r = ro + altitude
40= Angle between missile velocity vector and local horizontal.
T = Thrust (sea level). + (Po-Pa) Ae
Po = Sea-level pressure
Pa = Pressure at altitude
Ae = Engine exhaust exit area
Angle between thrust vector and missile velocity vector.
D = i V2 Cd A
P = Atmospheric mass density at altitude
m = Missile mass
V = Velocity
Cd = Drag Coefficient
A = Cross-section area of missile
The rate of change of a is:
9 = V a cos 9 + $ + T (sin,B )
im
- ; --.( V
The program allows the missile to proceed vertically
for 6 seconds at which time B is changed at the rate of 0.875 degrees/second
56
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X = ro V cos 6
 5rprp oved For Release 2002/11/04: CIA-RDP78TO5439A000200370043-6
SECRET
(56?/64 sec). The thrust misalignment remains constant for 10-20 seconds to
cause the missile to start to pitch over. Then O is reduced to zero and the
missile undergoes zero angle of attack until cutoff (92 seconds). The value
of,& was chosen to produce a pitch program that permits maximum range.
for trajectory analysis.
Results of Analysis (U)
I I A detailed trajectory analysis utilizing warhead
weights of 1550 and 2 s conducted to provide performance parameters shown
in figures 21 and 22. Since the range of this vehicle would be affected only
slightly by the rotation of the earth, non-rotating earth programs were used
SEERET
 SECRET
ACCELERATION
VELOCITY
ALTITUDE
2
O
aN
NOSE CONE WT. 1550 LBS.
C.30 a
O O
100 200 30,10
TIME (SEC)
Figure 21. (SECRET) Performance Parameters of SCUD Missile withl1550 lb
Warhead (C) 58
SECRET
 Approved For Release 200SECRCl RDP78TO5439A000200370043-6
80 3000
20 I
60
4O
ACCELERATION
E-
U.
I-
V
0
J
TIME (SEC)
Figure 22. (SECRET)
Performance Parameters of SCUD Missile with
2500 lb Warhead (C)
T
300
59
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 SECRET
HEAT TRANSFER ANALYSIS (U)
The purpose of this analysis is to obtain heat transfer data
needed to perform the stress analysis of the SCUD missile. Specifically,
maximum skin temperatures vs thickness at various sections along the missile
are desired. Temperatures along the nose cone and cylindrical section of
the missile were obtained directly by the use of an IBM 704 computer. The
computer printouts were also utilized to estimate=the heating which occurs
at the tip of the nose cone'and the skin temperature of the fins.
The basis for this analysis is the missile configuration (Figure
3) and the trajectory (Future 21). It was assumed that the missile skin was
AISI 347 stainless steel with physical properties as shown in table 8.
TABLE S.
ITEM
MISSILE SKIN MATERIAL PROPERTIES (U)
Specific heat
Thermal conductivity
Surface emissivity
Density
I Discussion
UNITS VALUE
Btu/. lb ?F 0.13
Btu/hr ft ?F 10.2
0.30
Lb/ft3 494
Method of Analysis (U)
The basic equations of the IBM 704 computer program are
as follows:
The heat transferred from the missile skin by radiation is:
qr = e 0'(T.)4
where:
qr = Heat transferred by radiation, Btu/ft2 hr
e = Vehicle surface emissivity
6 = 8
Stefan-Boltzmann constant, 0.173 x 10 Btu/ft2 hr?R4
TV.= Wall temperature,
The heat transferred to the skin by convection is:
qc = h (Tr - TO
60
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 SECRET
qc = Heat transferred by convection, Btu/ft
h = Heat transfer coefficient, Btu/ft2 hr OR
Tr = Recovery temperature, OR
The heat transfer coefficient, h, is:
1/3
?
0.5
h = 0.332 k*
x
(Pr*)
(Re*)
(for laminar flow)
. 1/3
0.8
h = 0.0296 k*
(Pr*)
(Re*)
(for turbulent flow)
x
where:
k* = Thermal conductivity at T*, Btu/hr ft ?R
T* = Reference temperature, ?R
x = Distance from leading edge, ft
Re= Reynolds number at T*
The recovery temperature is:
02
Tt = Too + R 0
V 2g JCp
To = Ambient temperature, ?R
R = Recovery factor
Voo = Free stream velocity, ft/sec
g = Acceleration of gravity, ft/sec2
J = Mechanical equivalent of heat, ft-lb/Btu
Cp = Specific heat of air, Btu/lb OR
The reference temperature at which fluid properties are evaluated is:
T* = Too + 0.5 (Tw = -Too) + 0.22 (Tr - Too)
SECRET
 SECRET
Results of Analysis .(U)
I I 1S[71ca +, 's
function of s in hickness for various locations on the nose cone, missile
body, and fins, respectively.
Temperatures on the nose cone and missile body were obtained
from the comp printouts, while fintemperatures were estimated by hand
calculations based upon the computer printouts. gu
-skin temperature time history, in this case for an.081 inch wall and 8.2 feet
from the forward end of the missile.
n It is seen that the skin temperature at this location does
not exceed 20v?~ during take-off, but that it rises very rapidly upon re-entry
to a value of about 820oF and then decreases to about 770?F upon impact.
II
MISSILE BODY SKIN TEMPERATURE (U)
DISTANCE FROM SKIN MAXIMUM
SSILE THICKNESS TEMPERATURE
TIP OF MI
(IN) (?F)
(FT)
8.2
0.081
820
0.100
750
0.125
675
13.2
0.060
855
0.070
800
0.090
725
18.2
0.060
835
0.070
775.
0.090
700
23.2
0.060
810
0.070
760
0.090
690
28.2
0.060
920
0.070
750
0.090
670
33.2
0.030
1020
0.050
850
0.070
745
63
SECRET.
 SECRET
zoo
1
100
1
200
1
300
Figure 23. (SECRET) Skin Temperature vs Time History for SCUD Missile (C)
64
SECRET
 SECRET
DISTANCE FROM SKIN
LEADING EDGE THICKNESS
(FT) (IN)
MAXIMUM
TEMPERATURE
(?F)
0.050
963
0.100
581
0.050
1280
0.100
738
0.050
1590
0.100
894
The most forward station for machine calculations was at a
1-foot distance from the tip of the nose cone. The heating rate at this point
was extrapolated to obtain the heat transfer rates from the 1-foot distance to
the tip of the nose cone by the method of reference 6, Appendix V. This heat-
ing rate is maximum at the leading edge and decreases with increasing distance
from the leading edge. The ability of a solid cone to absorb heat per unit
surface area, however, increases with the distance from the leading edge. At
the tip of the cone, the metal has a small heat capacity, and the temperature
approaches the melting point. Further aft, the metal has adequate heat capacity
to absorb all of the transferred heat and remain well below its melting point.-
At a distance of 0.1 foot from the tip of the nose cone, the metal reaches a
maximum temperature of about 1000?F. Thus, a solid tip about 0.1-foot long
has adequate capacity to absorb the heat transferred during re-entry and main-
tain structural integrity.
Because of the rounded leading edge of the fin, the tempera-
ture does no h the high values of the nose cone tip. The temperature
0.3 feet from the leading edge of the fin can be considered about equal to the
temperature of the leading edge (table 11).
65
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 SECRET
APPENDIX III AIRFRAME STRESS ANALYSIS (U)
^ This analysis provides the estimated airframe weight of the SCUD
missile. It takes into consideration loads imparted to the vehicle
and flight
erection
carrier during transportation and loads occurring during
of the missile. Tankage and transition sections are considered to be of
full-monocoque welded construction, while the aft skirt and fins are of semi-
monocoque construction.
The airframe is assumed to be constructed of a material similar to
321 7 austenitic stainless steel. Paragraph 9, Appendix I, assumes that
the propellant tanks are heat treated. No heat-treatment is needed to improve
physical properties after welding for the thinner sections encountered in the
SCUD airframe.
n The SCUD configuration is shown in figure 3, and the overall weight
theweight
breakdowncomponent
re12 shows the sulting weightmissile
break is shown in table showsTable
dis
air-
dstribribut ution, and and table 13
frame components.
Tank pressures used as the basis for this analysis were 500 psia for
the oxidizer tank and 500 psia for the fuel tank. The acceleration of the
missile (g's) was obtained from the trajectory (Figure 21). The missile skin
temperature as a function of thickness was obtained from tables 9. 10, 11.
TABLE 12. ESTIMATED VEHICLE WEIGHT DISTRIBUTION (U)
INITIAL WEIGHT (LBS) BURNOUT WEIGHT (LBS)
PROPULSION COMPARTMENT 168
Fins 168
Jet vanes 90 90
Jet vane actuators _ 80 80
Trim motors 30 30
Engine 580 580
Inverter/batteries (MPS) 80 80
Autopilot amplifier 30 30
Nitrogen pressurization system 12 12
Ignition pressurization system 30 30
-Engine support frame 30 30
Aft transition 30 30
Aft support ring 9 9
Aft cylinder structure 122 122
Misc & power distribution 73 73
TOTAL 1,364 1,364
FUEL COMPARTMENT 164 164
Tankage 122
Fuel (kerosene) 1,750
TOTAL 1,914 286
66
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 66
60
54
SECRET
ESTIMATED VEHICLE WEIGHT DISTRIBUTION (CONT'D) (U)
INITIAL WEIGHT (LBS) BURNOUT WEIGHT (LBS)
INTERTANK TRANSITION
Intertank transition 66
Autopilot 60
Fuel pressurization system 54
OXIDIZER COMPARTMENT
Tankage 294 294
Oxidizer (nitric acid) 6200 100
TOTAL 6494 394
FORWARD TRANSITION
Oxidizer pressurization system 86 86
Ballast 63 63
Forward cylinder structure 43 43
Forward support ring 9 9
Warhead support structure 25 25
TOTAL 226 226
NOSE CONE
Warhead (HE/nuclear) 1175 2125
Structure (shell) 375 375
TOTAL 1550 2500
LIFTOFF WEIGHT . . . . . . . . . . 11,728/12,678 lbs.
BURNOUT WEIGHT . . . . . . . . . . 4000/4950 lbs
TABLE 13. ESTIMATED AIRFRAME WEIGHT BREAKDOWN (U)
ITEM WEIGHT
Warhead-support _ 25 lbs
Forward support ring 9 lbs
Forward transition 43 lbs
Oxidizer tank 294 lbs
Intertank transition 66 lbs
Fuel tank 164 lbs
Aft transition 30 lbs
Aft support ring 9 lbs
Tail section structure 122 lbs
Fins 168 lbs
Miscellaneous 73 lbs
TOTAL 1003 lbs
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 SECRET
The weights shown in table 13 for support members, aft cylindrical
structure, and fins were estimated using previous experience with similar
0
. The nose cone must withstand the aerodynamic heating and pressure
differential across the skin created during re-entry of theovehicle. For a
skin thickness of 0.080 in.l
maximum pressure different
fiumption that the static p
pressure.- This assumption
ence gained with the Aerobe
The compressive stress (fc)
f = pr = 7.5
, the maximum skin temperature is 940?F and the
1 is 7.5 psia. This figure is based upon the as-
ssure within the cone is equal to the local ambient
s essentially been substantiated through experi-
c
t cos a 0.0180 x 0.981
where:
p = Pressure, psi
r = Radius, in
t = Thickness, in
Q = Cone half angle, degrees
Conservatively, by assuming a cylinder 72 inches long and 33.6 inches in
diameter, the critical buckling stress is computed as follows:
ZL = L2 (1-1I2) 1/2
where:
Z = General length range parameter
L = Length, in
/4 = Poisson's ratio
ZL= (72)2 (0.954) = 3680 psi
16.8 x 0.080 in
The ratio of radius to thickness is:
r/t = 16.8/0.080 = 210
From a curve of K. (buckling coefficient) vs ZL for r/t equals 210,
Kc = 70
68
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 SECRET
critical buckling stress, Fcr, is calculated from:
Fcr = Kc 112E (Lt )2
12 (1~ Z) L
where:
E = Modulus of elasticity (23 x 106 psi at 940 ? FF
Fcr = 70 (3.14)2 (23 x 106) (_g.080 52 = 1800 psi
The margin of safety, M.S. = Maximum allowable stress
Actual stress
M.S. = Fcr = 1.12
fc
The actual margin of safety is higher because a cone is more resistant to
buckling than a cylinder; therefore, the skin thickness of the nose cone is
sufficient.
SECRET
 SECRET
@eneral (U)
This appendix contains a detailed description of the carrier
vehicle kndLaunching equipment used with the SCUD missile system. This analy-
sis is based upon the 1960 and 1961 Moscow parade photographs of the vehicle
and missile, and is supported by other intelligence data.
Basic Vehicle (U)
The basic chassis used for this vehicle is the same as that used
for the Joseph Stalin tank and also for the JSU-122 assault gun.
C. Erector-Launcher Assembly
i 1. Operation (U)
The mechanical operation of the erector-launcher assemblies
depends on use -o=:he two hydraulic pistons for erection, and the winch cable
and pulley arrangement for return of the firing table to the transporting po-
sition. The. missile is supported on the transporter-erector carriage by a
retaining strap assembly near the nose cone, and by alignment fixtures at the
aft skirt section. The carriage is held on a 3-point support, the rear pivot
points, and a lock forward on the hutment.
I The operation necessary to erect the missile and
complete the mission is as follows (refer to Figure 24 for items in parentheses
below):
Position erection clamps (6) around missile and lock
in place on lowering jacks.
Release carriage lock (38).
c. I I Activate lifting piston (11) -- rotating missile,
carriage and firing to le assembly about pivot point (39).
Disengage and store carriage locking arms (16).
f. II Position missile in cross level for alignment with
firing table (17), using alignment fixture (13).
i.
clamp (6) assembly.
Release retaining strap (1).
Raise support feet (18) to contact lugs on missile
Transfer missile to firing table with the erecting
SECRET
 SECRET
Figure 25. (CONFIDENTIAL) Artist'
HORSESHOE WELDED FRAME
FLAME SHIELD (FOR ADJUSTMENT SCREW THRUST BEARING)
THRUST BEARING HOUSING
LEVELING SCREW ADJUST AND BEARING ASSEMBLY
FRAME RETURN FULLY
TRIANGULAR LEVELING FRAME
INNER AZIMITIH RING
OUTER AZIMUTH RING
Exploded View
TRANSFER RING
SUPPORT FOOT
SUPPORT FOOT LOCK
CABLE TIE DOWN RING
CABLE TIE DOWN RING LOCK
:MISSILE TIE DOWN CLAMS
LEVELING SCREW HANDLE
RING SPACER
of Firing Table Assembly (U)
72
SECRET
 SECRET
Disassemble erecting clamps and stow.
Fine level missile.
Lower carriage to horizontal.
M. After launching, the baseplate assembly is. winched
into the transporting position.
Althrough the lowering jacks are not visible in the
I _J
photographs, the presence of the two erecting clamp assemblies on the trans-
porter indicates that the jacks are probably on the inside of the carriage
beams. In addition, this is the only method available in the system to position
the missile on the launch table. It cannot definitely be determined whether
the alignment fixtures are used in the role indicated above, however. It is
possible that the table may not be leveled until after the load of the missile
is placed on it. It is possible to lower and adjust the firing table assembly
separate from the carriage using the win2h. Since the load of the missile is
held on the carriage, however, the firing table can as easily be lowered and
adjusted at the time of missile erection (see Figure 25 for artist's concept
of firing table).
Firing Table Assembly (U)
The horseshoe-shaped pipe frame is the basic
i
structural member (Figure 26). To this, a triangular, welded member is
attached by adjustable collars. Two of these are screw adjustments for cross
leveling while the third collar provides leveling in the fore/aft direction,
a 3-point support system. The next major component is a set of two circular
rings separated by spacers. It is believed that the first ring contains the
azimuth rotating mechanism while the second ring with the four support feet
is probably only a load transfer plate. Azimuth adjustment is verified by-
mis-
examination of photographs. The blast deflector, hold-down clamp,
cellaneous support pieces complete the baseplate assembly.
Both-the baseplate assembly and the launching
carriage are supported an hinged at the same point at the rear of the vehicle.
Mechanical stops prevent the-angle between baseplate and launching carriage
from becoming less than 180 degrees.
3. Transport Carriage (U)
U t ans ort carriage is made up of two major
r
p
I I T e
structural members; the side pieces. The-supporting structure includes those
members necessary to tie the major structural-members-rigidly together and the
tubular steel framework forward of the retaining strap. A ladder built on the
carriage beams is provided for servicing the missile when in the vertical
position.
ort carriage is supported by the aft
trans
Th
p
e
carriage supports' t e y raulic lifting arms and, when in the horizontal
L_ I
73
SECRET
 SECRET
position, the hutment of the vehicle. Just forward of the hutment the
launching carriage release provides a position lock, holding the launching
carriage to the vehicle.
4. I IHydraulic Lifting Mechanism (U)
The hydraulic system which provides erecting service
to the system is compose of two large extensible hydraulic pistons, a reservoir,
a pump, and associated fill and bleed lines. The extension pistons permit
erection to the vertical position and are well within design limits for this
operation.- Hydraulic exit orifices are controlled so that these same lifting
pistons slow the descent of the launching carriage when it is lowered into
the carrying position. The power for the hydraulic system is controlled by
the driver who activates the auxiliary engine.
Miscellaneous Launching Equipment (U)
Winch (U)
I The winch is located on the top of the hutment.
This item is used to manipu ate the baseplate assembly from the carrying posi-
tion and return. There is a bar lock for securing the baseplate on the right
side of the vehicle in the carrying position. Power for the winch is taken
from either the main engine generator or the auxiliary engine generator.
Cylindrical Tanks (U)
The purpose of the cylindrical tanks along either side
c. I I.Remote Checkout Gear (U)
Due to the nature of the missile,. the vehicle, and the
of the hutment is not specifically indicated. Most probable uses, however,
based on the needs of the missile are for storage of cold gas for use during
pre-launching checkout and for pressurization. Inert cold gas is necessary
for tests during the checkout period. If these external tanks carry nitrogen
at an assumed pressure of 2940 psi, the size of the tanks is such that pres-
surization and checkout requirements are met and exceeded by a small amount
(possibly an allowance for leakage due to field operation).
launching equipment,'mr'is deemed unsafe to fire the missile while the crew is
on the vehicle. Thus, remote equipment, both for radio operation and last
minute checkout and firing, is essential. The opening on the right side of
the hutment, covered by a hinged access plate, could be used to house this
remote gear. It is believed that this remote equipment is-used beginning
approximately 5 minutes prior to launching time for final go-no-go circuit
checks and for initiating tank pressurization.
74
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 SECRET
APPENDIX V (U) REFERENCES (U)
1. (U) Syvertson, C. A., and Dennis, D. A., "A Second-Order
Shock-Expansion Method Applicable to Bodies of Revolution Near Zero
Lift" (U). NACA: Report No 1328, 1957. (UNCLASSIFIED).
2. (U) Purser, P. E., and Fields, E. M., "Some Research on the Lift
and Stability of Wing-Body Combinations" (U). NACA: Report No RML 55GO6a,
2 July 1957. (CONFIDENTIAL).
3. (U) Pitts, W. C., Nielsen, J. N., and Raattari, G. E., "Lift and
Center of Pressure of,Wind-Body-Tail Combinations of Subsonic, Transonic,
and Supersonic Speeds" (U). NACA: Report No 1307, 1959. (UNCLASSIFIED).
4. *(U) Dorrance, W. H., "Two Dimensional Airfoils at Moderate
Hypersonic Velocities" (U). Journal of Aeronautical Sciences. Vol 19,
No 9, 1952. (UNCLASSIFIED).
0
5. (U) Emerson, H. F., "Wind Tunnel Investigation of the Effect of
Clipping the Tips of Triangular Wings of Different Thickness, Camber and
Aspect Ratio-Transonic Bump Method" (U). NACA: Report No RNA 53L03,
3 February 1954. (CONFIDENTIAL)
6. (U) Knapp, R. L., "Analytical Procedure for Estimating Aerody-
namic Heating" (U). Aerojet-General Corporation: Technical Memorandum
112 SRP, 19 February 1959. (UNCLASSIFIED).
75
SECRET
 DVO
GRAU
GSFG
KP
PAMS Station
Panoramic Sight
PRTB -
Quartz generator
RV
SGF
Siting Area
SP
S/S
Technical Site
Turk VO
Waiting Area
Zak Vo
SECRET
GLOSSARY OF TERMS AMID SYMBOLS (U)
A component of an instrument set that is designed
to measure horizontal and vertical angles, and also
is used as the aiming point.
Far Eastern Military District
Chief Missile Artillery Directorate
Group of Soviet Forces in Germany
Command Post
Mobile Meteorological Artillery Statioll
A component of an instrument set used to aim the
missile at the target.
First letter of the four Russian words meaning Mobile
Repair Technical-Base.
Frequency control device
Missile troops
So?athern Group Forces
Launch position area
Siting position or launch point
Supply depot
Forward area missile storage and checkout area
Turkestan Military District
Forward holding area occupied temporarily by units
prior to moving forward to siting area.
Transcaucasian Military District
Another Soviet designation for the SCUD missile
76
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 SECRET
APPENDIX VII' (U) DISTRIBUTION LIST (U)
, 25X1 C
25X1 C
State Department----------------------------------------------- 1
U. S. Intelligence Board--------------------------------------- 24
(8)
(8)
Central Intelligence Agency------------------------------------ (8)
National Security Agency--------------------------------------- 5
Secretariat State Def Mil Info Control Committee--------------- 1
Secretary of Defense------------------------------------------- 1
Assistant Secretary, Research and Engineering------------------ 1
Joint Chiefs of Staff, J-2------------------------------------- 1
Defense Intelligence Agency--------------------- 24
Weapon System Evaluation Group-----.---------------------------- 2
Defense Atomic Support Agency--------- ------------------------- 1
Commander-in-Chief, Europe------------------------------------- 1
Commander-in-Chief, Pacific----=------------------------------- 1
Commander-in-Chief, Atlantic-------=-----=--------------------- 5
Commander-in-Chief, Strategic Air Command---------------------- 2
Commander-in-Chief, Strike Command----------------------------- 1
Commander.US Forces, Rorea------------------------------------- 1
National War College------------------------------------------- 1
Disarmament Command Staff-------------------------------------- 1
Industrial College Armed Forces-------------------------------- 1
a
National Photographic Intelligence Center---------------------- 7
(Central Registry) (5)
(Army Group) (2)
Air University (Army Member) ----------------------------------- 1
Air War College (Army M,ember)---------------------------------- 1
Air Force Intelligence----------------------------------------- 10
U. S. Marine Corps--------------------------------------------- 2
U.. S. Marine Corps School--------------------------------------
U. S. Forces Japan---------------------------------------------
~Ink~r*
Undersecretary of the Army-------------------------------------
Assistant Secretary, Research and Development------------------
Office of Chief Research and Development-----------------------
Deputy Chief Staff, Military Operations------------------------
Deputy Chief Staff, Logistics----------------------------------
77
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 Approved For Release 2002/11/04: CIA-RDP78TO5439A000200370043-6
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Office of Assistant Chief of Staff, Intelligence--------------
26
(US Army Attaches)
(11)
Headquarters, USCONARC----------------------------------------
2
U. S. Army Caribbean-G-2--------------------------------------
3
U. S. Army Europe---------------------------------------------
1
U. S. Army Pacific--------------------------------------------
1
U. S. Army Alaska-G-2-----------------------------------------
1
U. S. Army Japan----------------------------------------------
1
Eighth U. S. Army---------------------------------------------
1
CG, XVIII Airborne Corps--------------------------------------
1.
U. S. Army Materiel Command-----------------------------------
9
(FSTC)
(8)
(R&D).
(1)
U. S. Army Mobility Command--=--------------------------------
1
U: S. Army Munitions Command----------------------------------
1
U. S. Army Weapons Cc -nd-------------- ----------------____--
1
U. S. Army Test and kvaluation Command------------------------
2
U. S. Army Electronics Command--------------------------------
2
U. S. Army Combat Development Command-------------------------
1
U. S. Army Combat Development Command Experimental Center----
1
CG, Fort Bragg, N. C., Staff LNO, Sec 9-----------------------
1
CG, Fort Ritchie,.Md., Det F, Sec 9---------------------------
1
Army Photo Interpretation Center------------------------------
3
U. S. Army Security Agency --------------------- ---0-----------
I
U. S. Army Map Service----------------------------------------
1
U. S. Army War College----------------------------------------
1
U. S. Army Military Academy-----------------------------------
1
U. S. Army Command and General Staff College------------------
1
U. S. Army Artillery and Missile School-----------------------
1
U. S. Army Special Warfare Center-----------------------------
1
U. S. Army Signal School--------------------------------------
1
U. S. Army Ordnance School------------------------------------
1
U. S. Army Artillery'Board------------------------------------
1
U. S. Army Armor Board----------------------------------------
1
U. S. Army Infantry Board-------------------------------------
I
U. S. Army Airborne and Electronics Board---------------------
1
U. S. Army Missile Command------------------------------------
30
(AM:4II -R)
(10)
(AMSMI-I)
(1)
(AMSPII-s)
(1)
(AMSMI-X)
(1)
(AMSAII-XB)
(2)
(AMSIII-XG)
=
(3)
(ANcPM
MA)
(1)
(AMCPM-MB)
(AMCPM-z8)
(1)
(9)
78
SECRET
 WARNING NOTIC
THIS DOCUMENT CONTAINS INFORMATION AFFECTING
THE NATIONAL DEFENSE OF THE UNITED STATES WITHIN
THE MEANING OF THE ESPIONAGFr LAWS, (18 USC 793,794),
THE TRANSMISSION OR REVELATION OF WHICH IN ANY
MANNER TO AN UNAUTHORIZED PERSON IS PROHIBITED
BY LAW.