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SECTION
TION VI BRT ee

firet 12 hours after launch. Dual payload configuration restricts initial checks
of visual payload function to qualitative evaluations only. These limitations
are imposed by the joint usage of the Sequence programmer, the testricted
view of the camera in looking through the ferret Payload antenna horn, and

the obscuring of the horizon optics field-of-view by the ferret Payload struc-
ture. Quantitative testing of the ferret load and qualitative evaluation of
the visual payload are completed in the Feat four days of orbital operation,
The ferret payload is then ejected within range of the Vandenberg Tracking

Station, quantitative testing of the visual payload occupies the remaining
Gatellite life,

7-76, Paragraphs 7-77 through 7-87 indicate the nominal sequence of system
events during the first 10 orbits, Figure 7-2 supplements the information in

these paragraphs by showing a schedule of the orbits during the first 24 hours
of orbital operation.

OO

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TH16 —SECRET—

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7)


SECTION Iv

LMSD-437237
CONTROL GAS
SUPPLY SPHERE

GAS REACTION
JET (6 PLACES):

— =~

, SSS
PRessune
CONTROL Gas ! nEcucATiNo |
SUPPLY SPHERE | PACKAGE |
. 1 {S|rton pressure
| IS} revier vatve
=o :
eiren | PRESSURE |
| REGULATOR |
| Low pressure |
| RELIEF VALVE | ROLL GAS ROLL Gas
HORIZON L J CONTROL VALVE jICONTROL VALVE
SCANNED ————— — —
ROLL ROLL GAS ROLL GAS
GYRO REACTION JET REACTION JET

ROLL AC > DEMODULATOR ROLL DC
AMPLIFIER AND RATE > AMPLIFIER
NETWORK

-—-——

— — — — —SSHT conTROE ELECTROMes unr

Figue 4-18. _ Pneumotic Contre! System, Functional Schematic
. ——_SEeRET— hea?



Post-Nosedown Orbiting 7-14
Description 7-14
Pre-Nosedown Orbiting 7.14
Description 7-14
Reconnaissance 7214
Paytoad Testing 7-14
Read-in and Readout 7.14

Orbit 1 Sequence of Events 7.16
Orbit 2 Sequence of Events 7.16
Orbit 3 Sequence of Events 7-16
Orbit 4 Sequence of Events 7.16
Orbit 5 Sequence of Events 7.16
Orbit 6 Sequence of Events 7.16
rbit 7 Sequence of Events 7-17
rbit 8 Sequence of Events 17
Orbit a Sequence of Events 7.17

7-17
Orbie 10 and Subsequent Orbit

Events 7.17

PRELAUNCH OPERATIONS AND
COUNTDOWN 7-7
Cuntdown 7.19

General 7.7

Ini

int

et
c

cRrated Fricitity P.

Satellite Interface
te 7.7

Ground Space Communications

Testing 7.7
¢rsonne! Readiness
ehearsal 7

Feparations 7.,

CHECKour 7-6
* + Description 7-6

“BEEP METAL Daa

Samos Vehicte Checkout, See; Samos
System Integrated Operations,
ra Bases, Centers, and
Stations, and Samos Satellite

Samos Satellite

SATELLITE resp CENTER (stc), °
8 ting B, Centers,
and sae ing Bases, Con ere

SECTIONS OF THE MANUAL, See:
round Space Communications and
ontr: .
Cperating Bases, Center, and
Pavigae Data Pp. ing S$
‘ayload Dats Process 18 System
Samos Booeter
8 Satellite
mos tem General Description
Samos System, Integrated Operation

SERVICE REQUIREMENTS, seq,
Semon SegUIn ue

SATELLITE-2oRNE COMMUNICATIONS
FQUIPMENT, See: ;



TON IV ——— SECRET
SECT LMSD-437237

midbody structure so that the tank support fitting engages inner midbody
structure. Tiedown fittings retain the assembly in thi position,

4-53. PROPELLANT FEED-AND-LOAD AND PRESSURIZATION SYSTEMs,

syetem provides the equipm:
Propellant tanks and cons:
load-and-dump, nd

tings to the tanks, Plus the propellant feed lines that Connect the tanke to the
engine Pumps, . .

Spheres on the aft equipment rack, the Pressurizing lines that interconnect
the spheres and the propellant tanks, the lines for tank and sphere Mpeone
the line for Preseurizing the oxidizser-pump, double-lip seal, and the
valves and controls ‘=<: stled in the system lines, .

4-56. For Propellant loading, external Propellant-loading lines and external
vent lines are connected to the appropriate quick-disconnect fittings of the
Satellite. One loading line and one vent line is for the fuel tank. -This pro.
vides a clos j-loop loading system between each Propellant tank and its Tee
spective ground-support Propellant weighing tank, In this manner, the
Propellant tanks Pressurized during the Pumping of Propellants into
the satellite, By me. ground-equipment controls, the 3 can
be forced back into the weighing tanks or, pe senavely, into the external
dump tanks.

applying pneumati ‘sure to release actuators in the quick-disconnect fit.
tings. The relea: stem is backed up by an independent lanyard release

4-57. The two helium Preseurization spheres can be loaded when the external
helium source is connected to the helium-fiil, ‘quick-disconnect fittin Of the
Satellite. Each Sphere has a sapacity of 2200 cubic inches and is loaded with
dry helium to a pressure of 3000 psi and at a maximum temperature of 120°F,

4-58. Usii ig the closed-loop Propellant loading Setup, the propellant tanke

‘Surized to full Operating pressure either by use of the external

vent lines during Propellant loading (par. raph 4-56) or through the helium
em during the process of loading the helium Spheres, With the latter

flow.to the tanks but prevent backflow of Propellants to the regulators, -A
normally-closed, squib-cperated valve prevents main-regulator 88 flow to

4-20 ~——SEGRET__


SECTION va a

7-62. SAMOS VEHICLE ACTIVITIES,

7-63. LAUNCH AND BOOST PHASE. (See figure 7-1.) After release of . C
the Samos Vehicle from the pad, and continuing through retro- rocket ignition,

fll Samos Booster sequences are controlled by the booster flight program-

mer and by guidance commands and discrete commands, which are gated by

the programmer. After 42 inches of vertical motion, the booster autopilot

is activated, Providing attitude control. The Samos Vehicle ascends verti-

cally for the first 15 seconds of flight. Starting at 3 second: continuing

through 14 seconds, the vehicle executes a programmed 35-degree roll. At

the end of 15 seconds, a programmed pitchover is initiated and continues

until cutoff of Samos Booster power, : .

7-64. The Samos Satellite Performs no guidance and control functions of

its own until immediately before vehicle separation. The booster guidance
system then disarms the command destruct system, tellite
guidance timer, and uncages the attitude gyros. The satellite guidance
timer applies backup signals to disarm the command destruct system, turn
on the flight control power relay, activate command separation, and open the
control gas supply valve. .

7-65. SEPARATION AND COAST PHASE, (See figure 7-1.) Following the
command for separation and the firing of the retro-rockets, the Samos

P
Pitch gyro is torqued to give the satellite the desired Pitch rate. Error
signals from this Byro cause the flight control pneumatic system to pitch the
satellite over until the roll axis is nominally horizontal, at which time the
command pitch rate is reduced to maintain this satellite attitude relative to
the earth. The satellite guidance timer then connects the horizon scanner,
optically establishes the local vertical, and then resolves the angle between
the satellite yaw axis and the local vertical. The angle is resolved into pitch
and roll component signals and transmitted to the inertial reference package.

7-66. ORBITAL BOOST PHASE, (See figure 7-1.) The flight control hy-

draulic system for engine gimbaling is activated prior to ignition of the sat-

ellite engine, However, satellite attitude is controlled by the. pneumatic -

system until this time, at which, Pitch and yaw pneumatics are turned off.

Prior to engine ignition, the ullage rockets fire. Then, during engine firing,

engine gimbaling provides Pitch and yaw control, and roll control is by the

Pneumatic system, The longitudinal accelerometer output is connected to

the integrator to determine velocity gained and to maintain the satellite -
attitude relative to the earth, Engine shutdown is commanded by the inte-

grator at a predetermined vehicle velocity. The pitch and yaw pneumatic

Control is turned on again, the hydraulic system is turred off, and the

longitudinal accelerometer is disconnected from the integrator, Engine

shutdown occurs at an altitude o approximately 300 miles, The vehicle

then enters the orbit phase of operation, .

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EC Iv
LMSD-437237 SECTION

nose-down attitude. The orbit phase begins after orbit reorientation, it in-
cludes the ejection of the ferret paylosd, and continues until the electrical
power of the satellite is exhausted. There are also several significant differ-
ences between the guidance and control equipment scheduled -for the initial
Samo~ flights and that scheduled for later Samos fights. . The equipment
sch. uted for the initial flights ts discussed in this section, and the changes
to be made for the later flights are indicated in the appropriate discussion.

rs

The operations performed by the guidance and cont jyeteme are the same
for all flights. .

4-116, GUIDANCE SYSTEM. The guidance system provides three types of
guidance signals: attitude error signals, pitch and roll attitude reference
signal nd pitch attitude program command signals, The attitude error
signals correspond to the amount that the Samos Satellite has deviated in
pitch, roll, and yaw about the pitch, roll, and yaw axes that are established
by the guidance tem. The pitch, roll, and yaw axes are established by
the guidance sys! to coincide with the reference pitch, roll, and yaw axes
described in par: h 4-117. The attitude error signals continually com-
mand the control em to correct for random deviations in satellite attitude.
The pitch and roll attitude reference signals correspond to the amount that
the satellite is being consistently held off course due to a constant error be-
tween the pitch and roll axes established by the guidance system and the ref-
erence pitch and roll axes. The pitch and roll attitude reference signals
continually correct the guidance system to eliminate any euch constant error.
The pitch attitude program command signals correspond to the amount that
the attitude of the satellite must be changed in pitch for different phases of
Might. The pitch attitude program command signals are applied to the guid-
ance system at scheduled times to rotate the nose of the satellite downward
with respect to the reference pitch axis. The guidance system aleo measures
the velocity gained by the satellite during the operation of the satellite rocket
engine and cuts off the engine. The equipment that forms the guidance sys-
tem and the functions they perform during the ascent and orbit phases are
described in paragraphs 4-120 and 4-133. 7

4-117, The reference pitch, roll, and yaw axes for the Samos Satellite are
defined with reapect to the earth as shown in figure 4-14. These reference
axes are independent of the satellite and they are in no way affected by its
attitude. The reference yaw axis is always along a line draws from the center
of the earth to the satellite. The reference roll axis is always perpendicular
to the reference yaw axis and pointed in the direction that the satellite is mov-
ing. The reference pitch axis is always perpendicular to both the reference
yaw axis and the reference roll axis, When the satellite ie changed from its
horizontal attitude to a vertical, nose-down attitude, the reference Pitch, roll,
and yaw axes remain unchanged with respect to the earth. Notice in

figures 4-14 and 4-15 that the X-X, Y-Y¥, and Z-Z are determined with
respect to the satellite and remain fixed within the satellite.

4-118, CONTROL SYSTEM. The control system actually consists of three
control systems: the pneumatic control « » the hydraulic control sys-
tem, and the orbital attitude damping syst: "These three systems contro!
the pitch, roll. and yaw movemente of the Samos Satellite about the refer-

ence pitch, roll, and yaw axes during the
pitch, roll, and yaw moven.ents for both are illustrated in .
figure 4-15, The pneumatic control system jocated at the rear of the

satellite ond the six gas reaction jets of th system are mounted on the aft

——SEGRET- "4-37



1 eee TR,

AR smonow vm

System tests and test activities as they may affect the achievement of ulti-
mate program goals and objectives. “Kaajor emphasis, however, is devoted
to the rapid evaluation of System flights as required to properly redirect
the program. As prime contractor, Lockheed is also responsible for oper-
ations evaluation of all satellite ‘stems, with particular emphasis placed
00 operations evaluations of the ‘ret and visual payloads.

7-114, Payload, telemetry, and tracking data collected by the tracking
stations are forwarded by communication links to the Satellite Test Center.
Payload data are analyzed for engineering evaluation, and copies are for-
gencies. Telemetry and tracking data are
tablish system performance. Computers at
the tracking stations and the Satellite Test Center are utilized to refine the
orbit ephemer: tablish new tracking stations, and generate updated
acquisition messages and realtime and stored program commands forthe | _ .

launch and test must be accomplished within the time span occurring between
launches, Because of the scope of Samos teat operations, & most stringent
time factor is placed on the tasks associated with post-test handling and
evaluation,

7-116. DATA REDUCTION.

7-117, GENERAL, Satellite telemetry data requiring reduction to suitable
form are processed by Lockheed data. services. Certain metric data (optics)
and FPS-16 launch data are reduced by Pacific Missile Range. Engineering
surveillance and documentary launch films are pr ed by the Ain Force
Lookout Mountain Facility, Payload visual data are reduced to the form of
Primary records by Vandenberg Tracking Station and are reassembled and
analyzed by Lockheed at the Satellite Zest Center and by the associate
contractor for data Processing. Raw ferret data tapes are delivered tothe
Satellite Test Center and to the data processing contractor, In order to
Permit an early evaluation of Samos test results, the satellite data {excluding
payload data) reduction Process is accomplished in two perts, launch quick.
look data and final data, Following the quick-look and final data aspects of
data reduction, a Samoa Subsystem Critique Meeting is conducted wherein
the Samos Satellite systems undergo detailed review. and qualification of
available final data, Following this meeting, detailed analyses of satellite
flight trajectory and vehicle System are conducted, All factors {avolving

nt or refinement of satellite hardware are thoroughly oan
investigated. Results of these analyses and i; tigations are compiled and
Published in specified reports (refer to Paragraph 7-131),

7-118, LAUNCH QUICK-LOOK DATA, Data reduction required to support
quick-look launch evaluation activity is accomplished on a first priority
basis within 24 to 36 hours after rec. pt of metric data and: telemetry tapes.
Copies of launch data are distributed to participating Lockheed organizations
12 hours in advance of a quick-look meeting. The Meeting is called to
initiate a preliminary evaluation of Samos Satellite flight results and is at-
tended by representatives of all contractors and Air Force organizations

——Szerer— . 7-23

ne



om nomen net

—seeretT—
LMSD-437237

GLOSSARY (Contioved)

LAUNCH PapD-a hard surface area on which a missile launcher is positioned.

MISSILE LAUNCHER-A structural device which cupports and holds a missile
in position for launch.

MONOCOQUE-A type of airframe construction characterized by stressed skin
and the absence of stiffening framework,

MONITOR—To continuously receive and analyze information to confirm that a
Specific condition is being maintained or that a Particular action is performed.

MULTIPLEXER-A mechanical or electronic component which converts a
number of voltage signals into a sing) ouput voltage signal that consists of
a successive series of voltage samp! Inder certain circumstances of
pulse inputs and appropriate timing, single output voltage signal can con-
tain all the input voltage signals applied to the component,

PITCH—Motion of the satellite about the reference pitch axis,

PROGRAMA predetermined sequence of events, provided as a series of
inatructions, that ie executed by an automatic device or computer,

RATE GYRO-A Syro that provides an electrical signal proportional to the
Fate of change in angular displacement of the syro from a reference Point.

RF-~Radio frequency, Electromagnetic fadiation in the frequency range
from 15 ke to $0,000 me, .

ROLL—Mation of the satellite about the reference roll axis,

REFERENCE PITCH AXIS=Refer to Paragraph 4-117 regarding Samos
Satellite, .

REFERENCE ROLL AXIS-Refer to Paragraph 4-117 regarding Samos
Satellite,

REFERENCE Yaw AXIS-Refer to Paragraph 4-117 regarding Samos Satellite.

REORIENTATION-The action of pitching the Samos Satettice from a horizon.al
attitude to a vertical, nosedown attitude,

REAL -TIME COMMAND~A Command that te tranamitted to the Samos Vehicle >
and executed immediately, Used to initiate an equipment function waen the
vehicle ie within range of the ground command station,

RECONNAISSANCE SATELLITE~An earth satellite launched from earth;

dreigned ty obtain airategic information through such means as photography
and detection of elec tromegnetic emissions.

———"-srerer— : G.3


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SECTION IV 1 Man isioa7

to operate as a rate gyro. The input axis of the pitch gyro is parallel with the
reference pitch axis when the vehicle is horizontal, and, because of its orien-
tation in the Samos Satellite, it remains parallel with the reference pitch axis
when the satellite is turned to a vertical, nose-down attitude. The attitude
error signals from the pitch gyro are applied, as shown in the diagram of
figure 2, through the time-lag integrator and the Servo amplifiers to the
motor that drives the pitch inertial reaction wheel. The apin axis of the wheel
is also parallel with the reference pitch axis. When pitch damping is com.
manded by the attitude error signal, the pitch inertial reaction wheel is accel.
erated in the direction of the disturbing torque. The wheel housing reacts to
the accelerating torque on the wheel and produces an equal and opposite torque

4-150. The roll/yaw attitude damping package is controlled by an attitude
error signal from the yaw gyro which is operated

input axis of the yaw gyro is parallel with the re:
Samos Satellite is horizontal, but, because of i entation in the satellite,
it is parallel with the reference roll axis when the satellite is turned toa
vertical, nose-down attitude, The yaw gyro will generate an attitude error
signal if a disturbing torque cau: n irregular motion of the satellite in
either roll or yaw. The attitude error signal is applied through the servo
amplifiers to the motor that drives the ro! /yaw inertial reaction wheel. The
spin axis of the wheel is parallel with the reference yaw axis. The principles
of operation of the roll/yaw attitude damping package to provide damping in
yaw and roll are illustrated in figure 4-23, . ot

4-151, When yaw damping is commanded by the attitude error si, + the roll/
yaw inertial reaction whee! is accelerated in the direction of the sturbing
torque. The wheel housing reacts to the acce! ‘ating torque on wheel and
produces an equal and Opposite torque on the lite that opposes the disturb-
ing torque in yaw. This action is identical to that used to damp disturbances
about the reference pitch axis. .

4-152. When roll damping is commanded by the attitude error si
angular velocity of the roll/yaw inertial reaction wheel is changed to produce

a restoring torque. The restoring torque results from gyroscopic prec

of the wheel as it reacts to the constant orbital pitching of the satellite. Wh,

@ gyro is twisted, it reacts by twisting about an axis which is perpendicular to
the axis of the original motion. So, as the Constant orbital pitching of the satel-
lite tends to twiet the roll/yaw inertial reaction wheel about the refere pitch
axis, the wheel reacts by twisting about the reference roll axis. Since the
wheel is not free to twist about the reference roll axis, it exerts a restoring
torque on the satellite that Opposes the disturbing torque in roll. The amount

of twist, and consequently the restoring torque, depends on the angular velocity
of the wheel. . . .

4-153. SEQUENCE oF GUIDANCE AND CONTROL EVENTS,

4-154. GENERAL, The sequence of guidance and contro! events for the Samos
Satellite includes most of the events that are necessary to place the satellite in
orbit. Since the guidance and control events are so important, and
cussion of them summarizes the functions ofthe Satellite-borneguidance and

control equipment, the sequence is described in Paragraphs 4-155 through 4-172,

4-54 ——SECRET__



SECTION Vv

LMSD-437237

5-38. The Up-Converter Transmitter, which is’ modulated by the output of

the receiver, is tuned ¢

0 the

frequency of the Narrow-Band Data Receiver

and supplies its output to a small antenna located on the Data Receiving An-

tenna Boresight Tower,
ward the Data Re
tower, the Range
repla the equipment
compensatory adjustme

By pointing the Command T;

in the vehicle) and it is possi!
nt.

Transmitter Antenna to-

ving Antenna and pointing the latter toward its boresight
‘easurement loop is closed (the Up-Converter in effect

ble to make the necessary

~


ECTION
LMSD-437237 SECTION IV

ground separation command does not successfully initiate separation, a back-
up signal to initiate separation is provided by the sequence timer 25 seconds
after its brake is released. Also, the Pneumatic control system is activated
at separation by a mechanical switch that is tripped as the Samos Satellite
moves out of the Samos Satellite and Samos Booster adapter,

4-159, FROM SEPARATION TO START OF SATELLITE ROCKET ENGINE,
From separation to the start of the satellite rocket engine, the Samos Satellite
coasts through space.for approximately four minutes. During this time its
attitude is controlled in pitch, roll, and yaw by the gas reaction jets of the
pneumatic control system. At separation, the nose of the satellit: atan
angle of 25.6 degrees above the horizontal. Within fifteen after separ-
ation, the gas reaction jets damp any satellite oscillations about thia attitude
in response to attitude error signals from the pitch, roll, and yaw gyros in
the inertial reference package. The Samos Satellite will be maintained with-
in 0.3 degree in pitch and yaw, and within 0.6 degree in roll about the stable
attitude of 25.6 degrees above the horizontal,

4-160. Immediately after the Samos Satellite has been stabilized following
separation, a pitch attitude program command of 40 degrees per minute: from
the sequence timer is applied to the inertial reference package. The satellite
is then turned from its attitude 25.6 degrees above the horizontal to a hori-
zontal attitude. When the satellite is horizontal the pitch attitude program
command sigaal of 40 degrees Per minute is replaced by a pitch attitude pro-
gram command signal of 2.8 degrees per minute from the integrator in the .
computer. The gas reaction jets respond to this signal and pitch the orbiting
satellite at the rate of 2.8 degrees per minute so that it remains horizontal.
After the satellite rocket engine is started, the pitch attitude program com-
mand signal from ths integrator is increased by the integrator from 2.8 to

3. 82 degrees per minute as the velocity of the satellite increases.

4-161. The horizon scanner fairing and the nose cap are ejected while the
Samos Satellite is being turned to a horizontal attitude. After the satellite
has been stabilized in the horizontal attitude, the horizon scanner establishes
‘ence yaw axis within 0.6 degree and detects any constant error in
horizontal attitude about the reference pitch and roll axis. The horizon
scanner provides the pitch and roll attitude reference signals that are sent

to the inertial reference package to adjust the pitch and roll gyros and correct
for the constant error,

4-162. Approximately eight minutes and twenty seconds after launch, which
includes four minutes and twenty ‘onds of Samos Booster boost and four
minutes of coast, the sequence tim activates the hydraulic control system
and fires the ullage rockets. The ullage rockets propel the Samos Satellite
forward to ensure that the oxidiser and fuel are at the inlets to the satellite
rocket engine. This is the first step in starting the satellite rocket engine.
The hydraulic control system does not perform a control function until thir-
teen seconds later when the satellite rocket engine is started.

4-163. FROM SATELLITE ROCKET ENGINE START TO CUTOFF. From
satellite rocket engine start to cutoft, the horizontal attitude of the Samos
Satellite is controlled in pitch and yaw by the hydraulic control system and in
roll by the gas reaction jets of the pneumatic control system. Eight and one-
half minutes after launch, thirteen seconds after the ullage rockets are fired,
the sequence timer starts the satellite rocket engine, and changes the pitch

—— SEGRET—— . + 6-62



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1
{

LMSD-437237 ”
‘
General 7.18 oy Program Evaluation and
+ Postreadout 7-18 Redirection 7-23
+ + Preparation, Readout, and :,.° General 7.23
Command 7.18 Test Reporte 7-23
Post-Pags Analysis 7-20 « +, General 7-23
: General 7.29 . Peoster Costractor ‘Reports 7-23
Preacquisition Preparation 7-18 + Lausch Reports 7.23
z.: Description 7-18 « Samos Reports 3 pee.
Satellite Control TH19 Tracking Station eportes 7.24
+ + Prereadout 719 INTIAL PROGRAM NARS I+
+ + Satellite Antenna General 74. 4
Orientation 7-19 eo . 8 of ramming 7.
SPC Transmiesion TH19 Samos Laid 75
:_: Description 719 o, General’ 7.5

Tracking 7219 + Communications. rations 7.6
.. Description 7-19 « Countdown Mansa aat
COORDINATION AND CONTROL Tod + Detailed Teet Objective 7-5

AF Ballietic Missile Division 7.2 .

+ ° Program Coordination 7-2 .

Control of Samos Operations 7-2 .

+ Central Control by STC 7.2 + Syetem Test Directive 7-8

. Saree tem Test Working Gr 725

+ Launch Base 7-3 Sateliite : #6 Group

+ «+ Launch Conductor 7-3 « Inflight 7.6

+ + Launch Conductor/Booster 704 s Pre-Launch 7.6

: + Launch Controtier 7-3 INTRODUCTION AND SECTION

* + Saunch Operations ARRANGEMENT TH2
Coordinator 7-3 * » Scope of Section 7.2

+ + Pad Safety 7.4 LAUNCH AND PREORBIT

+ 2 Range Safety 7.4 PHASES 7.1)

+ STC and Launch Base Base, Center, and Station

Coordination 7-4 . Activities 7.13 :

+ + Command and Control 7-4 + « Ascent Tracking 7-22

DATA EVALUATION 7-20 s+ Range Safety Tracking 7-11

Booster and Satellite 7922 eo. Vandenberg Tracking TU

s+ General 7.23 General 7.3]

Data Reduction 7-21 s+ Start of Launch Po

+ Final Data | 7-22 Samos Vehicle Activities 7-12

+ Generai 7.24 + SFlight Promise Tel3

+ Launch Quick -Look Daw 7.23 + Launch and Boost 7.22

General 7.29 + Orbital Boost Fol2

+ + Payload, Telemetry, and oe Preseparation Tol2
Tracking 7.23 . Jeparation and Coaat Fol2

+ + Samos Booster 7-21 ORDIT AND RECONNAISSANCE

; + System 7.29 PHASES Tole ‘

Launch Base 7-22 Ferret Payload Ejection 7.37

s+ General 7.22 oe Attitude Control After

Overall System 7.22 Ejection 7.18

:,.: General 7.22 s+ Conditions of Ejection 7.17

Payload 7.22 General 7.14

++ Evaluation Areas 7.22
w+ General 7-22

o. Possible Orbit Variances 7-14



SECTION IV DES

A simplified listing of the guidance and control events is Provided in es
figure 4-24. The times given in figure 4-24 are typical and given for example . oO
only, The events and times may be changed at any time before a Samos flight

to ensure optimum performance of the satellite equipment. The timing of each
event is adjusted to ensure Compatibility with the characteristics of the partic.
ular booster used for the flight. The periods before launch and from launch to
separation are discussed separately, The ascent phase is divided for discussion
into three periods: from separation to start of the satellite rocket engine, from
satellite rocket engine start to cutoff, and from satellite rocket engine cutoff
through orbit reorientation. The orbit phase and the ejection of the ferret pay-
load are discussed separately. . : .

4-155. BEFORE LAUNCH. Before launch, power is applied to the Samos satel-
lite guidance and control systems but the equipment does not perform any guid-
ance or control functions until after separation of the Samos Keellite from the
Samos Booster. Approximately two minutes before launch, power is a; lied
to the inertial reference package, the computer, and the flight control elec-
tronic unit. A 28-volt "keep alive" voltage is applied to the orbital attitude
damping system. The Gyros are operated in the caged mode, and the sequence
timer {s held by ite brake until just before separation. The horizon scanner,
the horizon sensor, the pneumatic control system (the 6 reaction jets), and.
the hydraulic control system are not connected to perform their guidance and
control functions. The accelerometer in the inertial reference package is not
connected to the integrator in the computer,

4-156. FROM LAUNCH TO SEPARATION, From launch to separation, power .
remains applied to the Samos Satellite guidance and control systems and they O
remain ready for operation. After launch, the Samos Booste sustainer and

boo. ingines operate for approximately four minutes. The vernier rocket
engi: Operate for an additional twenty seconds. During this time the Samos
Booster guidance and control systems perform the guidance and control func-
tions that are required. After the Samos Booster engines have operated for
four minutes and twenty seconds, a ground eparation command cuts off the
vernier rocket engines and init a the separation of the Samos Satellite from
the Samos Booster. Earlier, jeconds before cutoff of the vernier rocket
engines, a ground command released the brake on the sequence timer and
started the sequence of Samos Satellite guidance and control events.

# between the release of the sequence timer
ration, the sequence timer performe two back-
eps performed earlier to ensure that the steps

have been done. The two back-up Operations consist of rming the destruct

4-158. The following back-up signals are provided at Separation, Approxi-

mately 4 minutes and 10 seconds after launch, a ground command is sent to

the Samos Satellite to release the brake on the sequence timer, and ten sec-

onds later a back-up signal to release the brake is generated when the pull. (
away plug opens during separation. The ground separation command uncages

the gyros in the inertial reference package and fifteen seconds later a back-up.

signal to uncage them is provided by the sequence timer. Ta the event the :

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SECRET
LMSD-437237

SECTION VI

SECTION v1

INTRODUCTION and SECTION ARRANGEMENT 6-2

PAYLOAD DATA PROCESSING SYSTEM

“SSENERAL
OPERATING GROUP
SUPPORT GROUP
MANAGEMENT GROUP
SYSTEM INPUTS
SYSTEM OUTPUTS

6-2
62
62
6-2
63
63
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SECTION IV

test equipmert is used to control the
payloads of the satellite during a complete, simulated flight,

monitors, measures, and records all signals and responses
generated during the procedure. The power required during the Procedure
ia provided by the system checkout complex 2A, Items of special subsystem
test equipment may be used to isolate malfunctions detected during system
checkout.. After system checkout is completed satisfactorily, the satellite
and the payloads are transported to the launch pad complex.

4-306, The Special system test equipment used for system checkout is tabu-
lated below: . .

SPECIAL TEST EQUIPMENT FUNCTION
SYSTEM CHECKOUT COMPLEX 2A

Signal Distribution and Provides test Points for each umbilical

Monitor Console signal and functions a central distri-
bution point of signals and control func.
tions for system checkout equipment and
satellite. Monitors and simulates actual
squib firing and records measurements,
Enables LACE to select various um-
bilical and test plug signals during sys-

tem checkout,
Control and Indicator Provides for control and metering of
Console Operating time of guidance and control

equipment and satellite-borne commu-
nications equipment during system
checkout by using two Funning time
indicators which correspond to the two
timers in the satellite Measures
power levels of various atellite power
supplies, Remotely controls power
distribution console, Controls and
monitors tests of guidance and control

equipment.
Command Decoder Sequence Checks command decoder and sequence
Programmer Console programmer by generating 47-bit com-

mand words at a rate of 1000 bite/
second, Compares response to com-
mands with specified standard, Gives.
"GO" or "NO GO" indication,

UHF Telemetry and Data "Provides Circuitry for checkout of com-
Link Console pletely mbled satellite-borne com-
pment. Selects and
elemetry signals and
0 and payload signals to

recorders,

—_SECREF— 4-109



SECTION Iv LMSD.437237

a -28-volt power supply, und two 400-cycle power amplifiers. The power
control equipment inclu: such items as Telays, switches, umbilical con-
nector, wiring harness, and protection devices, (Refer to paragraph 4-89
for major equipment item descriptions, )

electrical load requirements of the vehicle in figure 4-12, The system sup.
plies both + and -28-volt direct current power, 2000-cycle, single-phase,
115-volt power; 400-cycle, 3-phase, 115-volt power; 400-cycle, single-phase,
115-volt, precision regulated power; and upregulated battery power, The un-
regulated battery power is directly furnished OF satellite equipment, such as
the hydraulic Pump drive motor, that can utilize the 22 to 29,8 battery voltage
range. For equipment that requires more closely controlled voltage, the in-
ternal electrical Power system utilizes the two +28-volt Fegulators which oper.
ate in series with the unregulated battery source and which also draw power
from the 2000-cycle inverters to Provide regulated ¢28-volt direct current,

In addition, the - 28. volt power Supply unit operates from the 2000-cycle sys-
tem and is arranged to track, in voltage Tegulation, the +28- vole regulator,
Positive and Regative direct current from the power supply unit and the Fegu-
lator are required for the flight control and guidance systeme and the Payload.
The 100-cycie, 3-phase inverter furnishes the basic 8-C power to the satellite
guidance and control system, the rec Fe, and payload; the 400-cycle power
amplifiers, synchronized with the Fence phase from the inverter, furnish
the required guidance-and-control system, single-phase, 200-syele. refer-
ence power. During countdown, the Outputs of the + and - it regulators
and the a-¢ power inve rter are monitored from the launch operations building, .

4-87. A power Junction box serves to interconnect the system componente,
The box includes the nece

connected components,
junction box with » Broun Connection to the vehicle frame. Power dis-
tribution leads are carried from the junction box to the power distribution

box From the Power distribution box, Power is distributed to the various
subsystems.

4-88, Ground- checkout
Power junction box to th
system components
internal power trans. let ewitch is mechanically dovigned as an integral part

of the umbilical connector, It incorporates a8 external solenoid which actuates
the awitch to permit ewitching from @xternal to internal power during count.
down prior to launch, Provisions are also made ia the power transfer switch
for lanyard Operation in the event of failure of the solencid mechasiom or of
the pneumatic, umbilical power dieconnect, The switch also apcorporates
contractors to furnish a Power-transfer signal to the system-chechout, and

to the monitor-console Panel in the launch operstions building.

4-89. MAJOR EQUIPMENT ITEMS DESCRIPTION,

4-90. GENERAL, (See figure 4-10.) The major equipment items of the in.
ternal electrical power syatem that are discussed ‘athe following Paragraphs
include the batteries, inverters, regulators, power amplifiers, and Regative
Power supply. With the except: of the batteries which are installed on the

ait equipment rack, the major items are located on the power panal of the
forward equipment rack,

4-30 SE GR ET__



——S8EeRreT— SECTION Iv
LMSD-437237

4-72, THRUST CHAMBER ASSEMBLY. The thrust chamber assembly pro.
vides the medium wherein the Propellants, as received from the turbopump
feed assembly, unite to Produce the desired forward thrust, The assembly is
an integral unit Consisting of a regeneratively-cooled combustion chamber, a
Rozzle throat section, and a divergent nossle section, and contains c
Passages within the walls, The oxidizer, used as a thrust chamber coolant,
enters the chamber cooling passages through a manifold Ping located slightly

4-73. The thrust-chamber Propellant injector is « “one-on-one” impinging
type with an additional circle of fuel o: es fed by a Ceparate manifold for
film-cooling the combustion chamcer walls, The injector di iea cavity
in the fuel section into which the fus! valve fits, thus allowing fuel valve
Poppet to be located close to the fuel injector orifices. Free or unoccupied
volumes are held small so that, when the fuel diaphragm ruptures, fuel ine
jection starts with minimum delay and with Minimum "hammertt 86 the free
space fills with Propellant, This arran, Rot needed for the oxidizer

is
since oxidizer flow is accelerated Gradually, and oxidizer ie injected for an
interval before the fuel valve opens. .

4-75. ENGINE PROPELLANT-FLOW SYSTEM. (See figure 4-8.) The pro-
pellant flow system Provides a means of controlling propeliant flow from the
turbine pumps to the thrust-chamber Propellant injector and the gas-generator
combustion chamber, The system consists ing from the

ow control

8!
Paragraph 4-76) cau hy
to reepectively open the bipropellant valve and the fuel valve, thus admitting

Propellants to the G49 generator and fuel to the thrust-chamber Propellant
injector, .

NOTE
Prior to the Opening of these two valves,

oxidizer has already started to flow into
the propellant injector,
\

——Sterer— 4-23



SPACE COMMUNICA TIONS, See:

Ground-Space Communications and
Control

STRUCTURE, See: Samos Satellite
Tt

TABLE OF contents ,

Communication and Control, Samos

TELEMETRY SHIP, See: Operating,
Bases, Centers, and Stations

TRACKING AND ACQUISITION
STATIONS, See: Operating Bases,
Centers, and Stations

TRACKING STATIONS, See: Operat.
ing Bases, Centers, and Stations

v

VANDENBERG AIR FORCE BASE,
See: rating Bases, Centers
and seatnna "t Baers Centers,

VANDENBERG TRACKING STATION,
See: rating Bases Centers
and Steter , ,

VEHICLE, See: Samee Satellite, and
Samos Booster :

VEHICLE CHECKOUT, See: Samos
System Inte ‘ated ration, and
Samos Sareifice OFF

VISUAL RECONNAISSANCE PAYLO,
See: Samos Satellite an



SECTION Iv

LMSD-437237

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PROCESSING

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Figure 4-75. Somos Setellite Sequence of Events During Orbit

4-64

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a. A sequence timer signal immediately follows as backup to chut off
'w pneumatic controle, Also, the timer opens the gac-

-fire circuits and connects the inertial-reference-package

‘erometer to the velocity integrator,

@. The satellite rocket engine achieves steady-state thrust,

£. Shortly before satellite rocket engine shutdown, the equence timer
arms the pitch-and-yaw Pneumatic control circuit and activates the engine cut-
off safety switch which has been preventing premature engine shutdown com.
mand from the velocity integrator, The integrator determines when the
satellite has attained the required velocity to be gained and sends a shutdown
Signal to the engine relay dox, . .

&- The engine relay box cuts off electrical Power to cause the gas-
generator valve to close and effect satellite rocket engine shutdown, Simul.
taneously, the engine relay box actuates the Pneumatic control circuit to turn
on the pitch-and-yaw Pneumatic controls,

bh. The sequence timer dieconnects the accelerometer from the velocity
integrator to.complete Propulsion system Sequencing. .

4-49. POWER INPUT REQUIREMENTS, The Power input requirements uti-
Mxed during the Sequence.of system operation is supplied by the 22 to 29, Sevolt,,
unregulated, direct-current, tery power. Peak demand is 9.75 amperes

and minimam pow requirement for nominal 120-second operation is 500
ampere-seconds. .

4-50. PROPELLANT SUBSYSTEM.

4-51. GENERAL, The major propellant subsystem items and the order in
which they will be discussed are: Propellant-Tank Assembly and Mount,
Propellant Feed-and-Load and Pressurization Systems, Propellant-Tank
UNage Orientation System. The Propellant tank assembly, ite mount, and
the related Propellant systems comprise the propulsion system equipment

that {a utilized to Properly supply propellants to the tarbopumps of the
Satellite rocket,

section that are connected together by a truncated cone and a tankesy;

aingle installation unit, The forward tank has a volume of .
and is used to contain the oxidizer, The aft tank, with
volume of 37.80 cubic feet, is for the fuel. The o! Ingle unit forme an ¢,
el.aped tank The oxidizer sphere forme th
between the tanke, and the truncated cone completes
The oxidizer tank has a combined pressure inlet and v,
has a separate Pp
opening with a T.

of the complete assembly. Screen baffles are installed in the tanks to prevent
excessive liquid motion, The tank assembly is installed within the satellite

——SEGRET— 4-19

- ee



~~

SECTION vir

~~

SECTION VII

INTRODUCTION and SECTION ARRANGEMENT _7.2
COORDINATION AND CONTROL 7-2
INITIAL PROGRAMMING 7-4
SAMOS VEHICLE CHECKOUT AT VANDENBERG 7.6
PRELAUNCH OPERATIONS AND COUNTDOWN | 7-7

LAUNCH AND PREORBIT PHASES 7-31
ORBITAL AND RECONNAISSANCE PHASES 7-14
COMMAND AND DATA READOUT PHASES 7-18

DATA EVALUATION oO

——SECRET— 71


SECTION VI ~——SEG6RET__
LMSD-437237

makes specific recommendations regarding possible program redirection,
A monthly addenda supplements this report. .

7-136. TRACKING STATION REPORTS. All seed ptations report on
operation of their equipment, by teletype, to the Satellite Teat Center within

36 hours following launch or within 4 hours after a Premature termination af
the flight. . :

Ta24 __ SECRET —

co)


SECTION Iv LMSD-437237

4-65. ENGINE SUBSYSTEM.

4-66. GENERAL, The engine subsystem consists of the satellite rocket
engine, its related Propellant flow and control systems, and engine mounting
provisions, The Turbopump Propellant-Feed Assembly, Gas Generator As-
sembly, and the Thrust Chamber Assembly are the three major subassemblies
of the engine. Propellant-Flow Sys

Syste: Engine Mount and Gim!

ely related and interconnected
operation and control functions during engine operation,

4-67. TURBOPUMP PROPELLANT FEED ASSEMBLY, The turbopump pro.
pellant feed a. embly of the engine pumps fuel and oxidizer from the Propellant
tanks to the thrust chamber of the engine, The assembly consists of a single.
stage, impulse-type turbine, an oxidizer pump and a fuel Pump which are gear-
coupled to the turbine shaft, and a gear housing that serves as the assembly
frame. The fuel Pump shaft is sealed from the gear case with a primary seal
and a double-lip seal, To provide maximum Protection against oxidizer leake

ing into the gear case, this double-lip seal is Pressurized with low-pressure

helium gas which forces any oxidizer that leaks Past the primary seal to flow
into the overboard drain, :

4-69. The two Centrifugal-type pr. pe, for any given rotational

Speed, produce a flow rat Constant over large variations
in outlet pressure. This volumetric mixture
ratios of propellants de! ngin: chamber over the engine

Operating range. The

24,570 and 13,946 rpm, re nominal flow rates are 15. 1 pounds
Per second for fuel and 39, oxidizer, .

4-70: GAS GENERATOR ASSEMBLY. The G80 generator assembly provides
2 means of starting and mi intaining turbine operation of the satellite rocket
engine. The generator embly consists of a emall combustion er
aasembly; a solid-propellant, turbine starter assembly; a gas, erator, bi.

propellant valve and its associated solenoid valy ot and a pair cavitating
venturies,

4-71. The Solid-propellant container is attached to the combustion chamber,
Upon ignition of thie charge by a dual-equib igniter, the Fesuitant, hot combus.

yy
The bipropellant valve subsequently Opens (refer to paragraphee 4-75
allowing fuel and oxidizer to enter the combustion chamber where
they combine and ignite hypergolically, The venturies regulate the propellant
flow through the bipropellant valve to the enerator. After the solid-propellant
grain ie expended, the Gases from fuel and oxidizer ignition continue to drive
the turbine until engine shutdown, During operation, generator temperature

ia held to safe limits by film cooling and by regenerative cooling,

4-22 —srerer—_



SECTION VII —— SECRET —
ero LMSD- 437237

INTRODUCTION AND SECTION ARRANGEMENT

7-1. The information in this section is a summation of the various phases
of the Samos Satellite System that have been described in more detail in
Sections I through VI, and is intended to provide an integrated description of:
overall system concept and sequence of operation. The major phases of the
system are presented in descriptive sequence, beginning with system co-
ordination by the Air Force Ballistic Missile Division (AFBMD) and con-
cluding with the evaluating and utilizing of reconnaissance data obtained
during a Samos flight. The phases of operation include: initial programming
for the Samos System and the Samos Satellite, Samos Booster and Samos
Satellite checkout and preparation at the Guided Missile Assembly Building,
system prelaunch operations, countdown, launch and ascent phases of the
Samos Satellite, orbit and reconnaissance phases of the satellite, command
and readout of satellite data, plus evaluation and use of the data. The
principal agencies for coordinating and instrumenting system activities are
also presented.

SAMOS SYSTEM COORDINATION AND CONTROL |
7-2, AIR FORCE BALLISTIC MISSILE DIVISION.

7-3. The responsibility for overall coordination of the Samos Satellite
Syetem is delegated to Headquarters of the Air Force Ballistic Missile
Division (AFBMD) at Inglewood, California. The program goal is to pro-
duce and obtain reconnaissance data by means of photographic and electronic
detection. equipment installed in an orbiting ellite. In order to execute
program planning, the Air Force has delegated authority to two groups.
These are the Syatem Test Working Group )STWG) and the Flight Test
Ww roup (FTWG), (Additional information regarding these two groups
is presented in paragraph 7-18 under initial Programming.) Also, support-
ing facilities have been established to help implement efficient coordination
and function of system operations. The Principal coordinating and control
facilities are the Satellite Test Center in Sunnyvale, California and Vanden-
berg Air Force Base near Lompoc, California.

7-4. CONTROL OF SAMOS SYSTEM OPERATIONS.

T-5. GENERAL, Control of system operations is discussed in terms of
central control exercised by the Satellite Test Center, the launch base per-
sonnel and organization, Particularly prior to launch, and the coordination
of Gperation and control between these two facilities. Refer to Sections I and

il for additional intormation regarding bases, centers, and stations and their"
related personne},

3-6. CENTRAL CONTROL BY SATELLITE TEST CENTER. The Satellite
Test Center {STC) serves as the central facility for command and control

of all operational phases of a Samos flight. Specific command and control
prerogatives are further delegated to the launch base and tracking atations
under the control of the Test Center, Direction and control of ilighte are
effected jointly by a Lockheed System Test Director and an Air Force System
Test Controller at the Test Center. The Test Director has overall responsi-
bility for implementing the flight test program and for directing all elements
of the system in the prelaunch, launch, and orbit phases of operation. He
keeps the Test Controller informed of overall system atatus, and, as the

7-2 —— SEGRET—_


SECTION IV 437237

4-145. CONTROL DURING ORBIT PHASE.

4-146. GENERAL. Control during the orbit phase is provided by the gravity
gradient, the orbital angular momentum of the Samos Satellite, and the orbital
attitude damping syetem. The vertical, nose-down attitude of the satellite is
established and maintained as the equilibrium position by the gravity gradient
and the orbital angular momentum. The gravity gradient tends to pitch the
orbiting satellite so that its Z-Z axis continues to point toward the center of
the earth. The Z-Z axis is the axis of minimum moment inertia. The orbital
angular momentum which is given to the satellite by the pneumatic control
syatem during orbit reorientation also pitches the orbiting satellite to rein-
force the action of the gravity gradient. Disturba..ces about the equilibrium
position are controlled by the orbital attitude damping system. The orbital
attitude damping system operates in response to guidance signals from the
guidance system throughout the orbit phase except for the brief period when
the ferret payload is ejected. When the ferret payload is ejected, the pneu-
matic control system is reactivated to handle the large disturbances that re-
sult from the ejection. The operation of the pneumatic control system is the
same as described in paragraph 4-139, The operation of the orbital attitude
damping system is described in paragraph 4-147, :

4-147. ORBITAL ATTITUDE DAMPING SYSTEM. The orbital attitude damp-
ing system exerts torques on the Samos Satellite changing the angular ve-
locity, and consequently the momentum, of inertial reaction wheels. -These
torques damp irregular motions of t! tellite in pitch, roll, and yaw by
opposing the disturbing torques that jd the motions. The system con-
sists of two attitude damping packages and a constant-rate momentum wheel
for the initial Samos flights. These components are represented in

figure 4-22. A diagram is also included in figure 4-22 to illustrate the oper-
ation of the components and to show how they are“controlled by guidance sig-
nals from the pitch and yaw gyros. An-orbital attitude contro! system i
scheduled for the later Samos flights which consists cf one attitude damping
package, a constant-rate momentum wheel and a large gyro, and, for backup,
a low-amplitude gas reaction jet system. . .

4-148, The effectiveness of the orbital attitude damping system depends on
the momentum characteristics and reaction times of the orbital attitude con-
trol mechanisms. The optimum values of these parametere have been de-
termined by computer analysis at Massachusetts Inetitute of Technology. The
results of the analysis dictate the size of the inertial reaction wheels, the
minimum momentum that must be maintained about the reference pitch axis,
and the time delay that is required in the pitch attitude damping package. The
conetant-rate momentum wheel is used to ensure that the minimum pitch mo-
mentum is maintained, and a time-lag integrator is provided in the pitch atti-
tude damping package to delay the pitch reaction time for twenty mi:

time delay is not required in the roll/yaw attitude damping
momentum characteristics in roll and yaw differ from thos
pitch and they are satisfactory. The tachometer in each packa,
speed of the motor driving th
signal proportional to th
amplifiers in the packag:
the guidance signals.

senses the
inertial reaction wheel and provides a voltage
ed. This voltage signal is returned to the servo
feedback to improve the response of the motor to

4-149. The pitch attitude damping package ie controlled by an attitude error
signal from the pitch gyro which is electrically caged during the orbit phase

4-52 ——— SECRET —



- SECTION IV
THB

SPECIAL TEST EQUIPMENT FUNCTION
LOCKHEED AUTOMATIC CHECKOUT EQUIPMENT (LACE)
. (Continued)

tities. Controls other checkout con-

soles in addition to components being

tested. Provides time base, Pro-
grams test with punched cards, De-

. tects correct Sequencing and indexing
of each punched card, Monitors oper-
ation of card reader and facilitates
card search, Controls input power,
Selects mode of Operation. Provides
means of self-check against internal

Control and Monitor Console Provides visual indication of test quan. -

standards,
Vehicle Function Generator . Checks cabling from checkout equip-
Console ment to satellite, and simulates

various satellite functions, Simu-

. lates lite patch system, squib
actuator, battery and destruct

: clreuitry, hydraulic reservoir
cireuit, and guidance potentio-
meters and resolvers, Monitors
and checks satellite telemetry po-
sition, compartment Pressure and
temperature circuitry, gas valve
circuitry, pitch-yaw actuator cir-
cuitry, and other guidance and
control and payload circuits and
components,

launch operations building, and the performance of each subsystem -

graph 4-310, The. required signal interconnection tween the launch
Operations building and the Samos Satellite, and the required exte powe
distribution equipment are provided in the launch pad and service building.
Since the Samos Satellite is located near the launch pad an@bervice building
immediately Prior to launch, {t is convenient to station certain items of .
est equipment there. The signal interconnection, power distribution,

‘and special test equipment used in the launch pad and service building is
Nested in Paragraph 4-31),

4-308. The checkout operations performed at the launch pad complex includes
those performed before the Samos Sa with the Samos Booster
and those performed afterwards. checkout signal and power
connections are made between the quipment in the launch

+ Because of

——SECRET— 4-111



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SECTION IV  ABCRETS;

4-103. MAJOR ACCESS AND INSPECTION FEATURES, Four access panels
enclose the forward equipment area and form the outer surface of the satellite
between stations 304 and 326. The right-hand panel contains the acce:
for the air-conditioning coupling and the electrical umbilical connec! » Each
of the four Primary-battery installation areas of the aft equipment rack are en-
closed by an acceas panel,

4-104. INTERNAL ELECTRICAL POWER SYSTEM OPERATIONAL READI-
NESS CHECKOUT CONCEPT.

4-105. GENERAL. A Samos Satellite, upon arrival at the launch base area,
is received at the guided missile a mbly building in the squadron mainten-
ance area, where it is inspected an tested to verify the readiness of satellite
rior to transporting the satellite to the
eembly building, the internal electrical
visual inspection followed first by Preliminary check-
out at the m level and then by integrated systems level testing where.
in the overa: subsystems are sequentially checked. C

4-106. SUBSYSTEM CHECKOUT CONCEPT,
syetem initially undergoes subsystem checks
subsystem checks, Herein,
Console, the vehicle power monitor console, and the uni:
are used to check the output of
observed are the plu
phase ac; 2000-cycle, ingle-phase ac;
of 22 to 29.5 volts. In addition to volta,
cycle and 2000-cycle, a-c power frequency plus phase
three-phase inverter are measured, A programmed
are undergone in order to determine the Performance
regulator components. The load checkout se
accurate, and relatively quick checkout of the internal electrical power sys-
tem under actual Samos Satellite electrical load conditions.

4-107, ELECTRICAL CHECKS DURING SATELLITE SYSTEMS CHECKOUT.
The overall Samos Satellite systeme checkout is an eseentially complete,

ectrical
powcr system is one of performance monitoring in which the overall output of
the system is monitored to assure performance within the required limits.

4-108. In thie checkout Concept, the Samos Satellite is installed in the vehicle
tilt and roll stand and checkout is performed of all the subs im Components
while they are operated together ing & simulated flight. Dur-
ing this checkout,the output voltag ‘urrents of the internal electrical
power system are monitored to ch their response under actual load con-
ditions, Provisions are incorporated into the *quipment to

Permit the utilization should the
system monitor program irdicate a failure in the internal electrical power
system equipment, H i i the location of a fault in the

internal electrical power system at the satellite systems checkout level,

4-109. Upon success{ul completion of systems checkout, the satellite is then
taken from the guided missile assembly building to the launch pad complex.

4-34 —SECRET—



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SECTION Iv AB

(Continued) .
GROUND HANDLING EQUIPMENT FUNCTION

Vehicle Handling Yoke Fite around satellite for support
during transportation and hoisting.
Used with vehicle transporter,
handling dolly, and hoisting

. equipment,
Yoke Removal! Slings ‘Used to remove vehicle handling
(Horizontal and Vertical) yokes while satellite is either
. horizontal or vertical.
Vehicle Hoisting Slings Attach to vehicle handling yoke to
(Horizontal and Vertical) hoist satellite in either horizontal
Position or tilted, vertical Position,
Adapter Handling Dolly Used to carry adapter within build-
ings. Holds adapter in horizontal
position, .
Adapter Protective Cover Waterproofed cover to enclose

adapter while mounted on adapter

handling dolly for transportation
or storage. .

Tank Handling Dolly Used to carry « mopty fuel tanks
within buildings before mounting
in satellite.

Tank Sling Hoists empty fuel tanke duri:
removal from or installation into
satellite,

Engine Handling Dolly Used to carry satellite rocket
engine in horizontal position for
movement within buildings. Used
to install or remove engine from

satellite. . .
USAF Type B-1 Propellant Transports UDMH fuel from storage
Transport Trailer area to launch pad and service build-

ing. Has capacity of approximately
220 cubic feet, or 11,000 pounds of

UDMH fuel.
USAF Type B-2 Propellant Transports IRFNA oxidizer from
Transport Trailer storage area to launch pad and

service building. Has capacity
of approximately 350 cubic feet,
or 35, 000 pounds of IRFNA
oxidizer,

Payload Shipping Containers Enclose and protect payloads during
transportation.

4-118 ——__SEGRET


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SECTION IV

as

APPROXIMATE TIME
OF EVENTS IN MIN-
UTES AND SECONDS

EXACT TIME
OF EVENT
IN SECONDS

GUIDANCE AND CONTROL EVEN]

FROM SEPARATION TO START. OF SAMOS SATELLITE ROCKET ENGINE
(Continued) —- .

+8 minutes and 20
seconds

498

a. Hydraulic control system is
activated, : :

b. Ullage rockets are fired.’

FROM SAMOS SATELLITE ROCKET ENGINE START TO CUTOFF

+8 minutes and 30
seconds

+10 minutes and 30
seconds

$11

629.1

a Satellite rocket engine is
started by sequence timer. (The
Satellite rocket engine fires for
2 min.) :

b. Accelerometer is connected
to integrator,

¢. Pneumatic control system is ;
deactivated in pitch and yaw.

Satellite rocket engine is cut off
by integrator.

FROM SAMOS SATELLITE RO!
THROUGH ORBIT REORIENTATION

CKET ENGINE CUTOFF

+10 minutes and 40
seconds

+13 minutes and 10
seconds

638

788

a. Pneumatic control system is
reactivated.

b. Hydraulic control system is
deactivated.

¢. Horizon scanner is discon-
nected from inertial reference
package.

d. Samos Satellite is turned from
horizontal attitude to vertical,
nose-down attitude by sequencer
timer. (This orbit reorientation
Fequires 2 min, 20 sec.)

Orbit reorientation is completed
and Samos Satellite is pitched at
orbital rate by pneumatic control
system to maintain vertical, nose-!

down attitude.

Figure 4-24, Sequence of Guidance and Contre! Eveats (Shoot 2 of 4)
___SECRET—

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SECTION vi aay

Air conditioning control is checked. The launch complex communications
are checked, The Automatic Programmer Checkout Equipment {APCHE)

for the booster is checked. The launch sequencer is checked. The engines ;
and autopilot of the Samos Booster are checked,

7-50. X-1 DAY ACTIVITIES, The Satellite Test Center and tracking sta-
tions conduct detailed simulation and communication checks. The Vandenberg
reports telemetry ship On-station, the current status of the launch complex,
tracking systems, and weather are reported, Post-mating of pyrotechnic
installations a: ‘ompleted on the satellite. The destruct circuit functions
are checked. elemetry and command and Controls checks are made. Data
links are checked via r-f£ monitoring. Beacons and tracking station reaponse

are checked. Launch-complex cameras are tested and loaded with film.
ecor

Comple: Ommunications and the countdown clock are checked. Booster
APCHE system tests are performed. Guidance commands from the ground
station equipment and discrete command loop tests are performed, Batteries,
engine igniters, and charge accumulators are installed in the booster, and
range safety destruct command tests are performed,

7-5t. X-0 DAY ACTIVITIES, Range clearances are obtained for launch.
Vandenberg Communications Center control is assumed by the Launch Oper-
ations Control Center, The Satellite Test Center and tracking stations Per-
form final simulated tracking exercise and boresight checks. All stations

ther forecasts and reports
are transmitted at 6 and 2 hours before launch, Countdown activities com-

_ 1-53, Following receipt of Air Force approval to proceed with the launch,
in this Paragraph have been accom-

he Satellite Test Center verifies the
ready condition of the tracking Weather conditions must be 8
factory and above the minimum specified Conditions. Control of the Van-
denberg Communications Center has been assumed by the Launch Operations
Control Center. Verification that the telemetry ship is ready and is on
station, Range safety Clearance has verified that the land, sea, and air
conditions are satisfactory and that the range safety system is operational.
The booste r-guidance-system ground-station equipment has been Programmed
for a Samos launch at Point Arguello. The contractor launch conductors

@aunch Conductor/Agena and Launch Conductor/Booster) have verified
» systems readiness,

7-54, The Air Force Launch Controller initiates the countdown after ob-
taining the clearances described. Detailed countdown is published in the
Samos Preliminary Countdown Manual, LMSD- 445944, and consists of 14
tasks. The 14 tasks, their Tespective designations, and the inclusive times
for each task a indicated herein. Time of countdown is in terms of the
290 -minute period prior to launch. The minutes are indicated in consecutive
order, beginning with 290 minutes, designated as T-290 minutes, and ter-

minating at T-0 minutes, the time of Samos Vehicle release for liftoff from
the pad,

7-55. Task 1: Countdown Initiation, from T-290 to T-275 minutes. Task 2:
Electronic Warmup, from T-275 to T-225 minutes. Task 3; Samos Booster

7-10 —— SECRET-



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D TOK 1 ASECRER

SE
MSD -437237

DOWNRANGE TELEMETRY SHIP, See:
Operating Bases, Centers, and
Stations

ELECTRICAL POWER SYSTEM, See:
Samos Satellite

ENGINE, See: Samos Satellite
Fr

FERRET RECONNAISSANCE PAYLOAD,
See: Samos Satellite

FIGURES, See: List of Nlustrations
Page vi

FIRST STAGE BOOSTER, See: Samos
Booster

GENERAL DESCRIPTION OF SAMOS
SYSTEM, See: Samos System
General Description

GLOSSARY G-1

aa ONTROL 5-1

GROUND-SPACE SYSTEM 5-2

Alignment and Calibration Syetem 5-5

+ Description 5-5

+ + Boresighting 5-5

+ + Photogrammetric Trian-
Bulation 5.5

+ « Up-Converter 5.5

+ + + Transmitter 5.6

Angle Tracking System §-2

+ Function §-2

+ + Antenna Operation 5-2

Command System  §-3

+ Function 5-3

+ «+ Stored Program Command 5.3

+ + Verification 5.3

«+ + Command Signal 5.3

Communications Concept 5-2

+ Mission 5.2

+ + Ground Equipment 5-2

+. Satellite-Burne Equipment §-2

ata Receiving System 5-4

+ Description $-4

Digital Data Handling System 5-3

+ Description and Function 5-3...

+ + Antenna Slave Data Loop 5-4

+ « Central Timing and Data
Distributor 5.4

« » Computer 5-4
oe Repiote Timing and Data
4

Distributor
- » Time Comparator 5-4
Ground Communication System 5-4
+ Description 5-4
Range Measurement System 5-3
. mineare $-3 3
Tim: ‘stem = 5-.
. Funcine 5-3
INTRODUCTION AND SECTION

EMENT 5-2 :

+ + Scope of Section 5-2

GROUND suPPoRT EQUIPMENT, See:
Samos Satellite

GUIDANCE AND CONTROL SYSTEMs,
See: Samos Satellite |

t
ILLUSTRATIONS, See: List of
Dlustrations, Page vi

INITIAL PROGRAMMING, See: Samos
System Integrated Operation .

INTEGRATED OPERATION, See: :
Samos System Integrated Operation

INTERNAL ELECTRICAL POWER
SYSTEM, See: Samos Satellite

INTRODUCTION TO SAMOS MANUAL,
See: Samos System Manual Data
Presentation, Page ix

KAENA POINT TRACKING STATION,
See: Operating Bases, Center, and
Stations ;

KODIAK TRACKING STATION, See:
Operating Bases, Gentera, and
Stations .

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SECTION IV

SPECIAL TEST EQUIPMENT

SECRET __
LMSD.437237
(Continued)

FUNCTION

SYSTEM CHECKOUT COMPLEX 2A

Pulse Amplitude Modulation
Telemetry Console

Power Distribution Console
Command Beacon Console

VHF /FM/FM Telemetry Data
Processing Console Assembly

E-1 Visual Payload Interface
Console

F-1 Ferret Electronics
Payload Interface Console

Recording Consoles

Displays and records analog information
Content of 36 selected channels, Re-
cords reference and timing data,

Controls and distributes 208/120. volts,
three-phase, 60 Ps a-c power to vari.
ous consoles of system checkout com-
Plex 2A,

Provides for transmitting, receiving,
analyzing, and measuring input and
output signals of beacon system,

Consists of four consoles: telemetry
link and monitor console, dis-
criminator console, decommutator
console, and video recording console.
Measure, monitors, and records all
signals of VHF telemetry. system,

Receives wide band UHF signals from
satellite transmitter, Detects modu-
lation from carrier and applies to
ground reconstruction system for
checkout. Provides remote control
of E-1 visual payload, .

Receives narrow band UHF signals
from satellite transmitter, Detects
modulation and records for later
evaluation. Provides remote control
of F-1 ferret electronics payload,

Monitor and record guidance and con-
trol signals during system checkout,
Monitor and record dec power supply
voltages.

LOCKHEED AUTOMATIC CHECKOUT EQUIPMENT {LACE)

Measurements Console

Provides automatic awitching to each
sequential checkout function. Main-
tains proper signal levels for con-
verters, Measures voltages,

» frequencies, and time
Digitizes voltages, fre-
quencies, and time intervale. Oper-
ates ten-line code alphamerical printer

and bank of nixie tubes to enable visual
monitoring of voltages,

4-110 ~—SECRET—

oo



SECRET - SECTION Iv
LMSD-437237 .

ystem continues to control disturb-

pelt off. he orbital attitude damping
anes abou! the ‘sgulibrium position util the electrica! 1 power of the Samos

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SECTION vi ——-SECRET—
LMSD-437237

directly concerned. Additional data requirements, which may be necessary
because of flight events, are specified after a preliminary review of real-
time and operational telemetry records within 24 hours following launch.

7-119. FINAL DATA, A final Comprehensive compilation of all data re-
quired for detailed satellite and payload data are accomplis in the three-
to-five-day period following launch and subsequent payloads testa.

7-120. LAUNCH BASE EVALUATION,

7-121. Within four houre following a launch, available launch data are as-

7-123. Detailed analyses of the Samos Booster are conducted by the booster
contractor. To facilitate Planning of subsequent teats, problems revealed by

the various test reports produced (refer to para-
ry, design representatives for the booster con-

graph 7-131). If nec.

tractor attend the Sam, Subsystem Critique Meeting conducted at Lockheed
Missiles and Space Division at Sunnyvale, California.

7-124, PAYLOAD EVALUATION,

7-125. Samos payload evaluation may be considered to be in four areas of
evaluation: operations and real-time evaluation, on-line data evaluation
and analysis, off-line data evaluation and analysis, and data quality and |
payload-product analysis.

alysis of on-line data includes that data used in real time, .
or close to real time, in making analyses, Off-line data evaluation and
analysis are conducted Jointly by Lockheed Test Data and Reporte organi-
zation and the Samos Satellite subsystems personnel, using raw and pro-
Cessed data. Data quality and payload- product analysis is conducted by the
data processing contractor, which furnishes reports to cognizant Air Force
agencies and to Lockheed.

7-127, OVERALL SYSTEM EVALUATION.

7-128. The Operations Management organization of Lockheed Missiles and
Space Division Proceeds with investigation of Problems revealed by pre-
liminary evaluation and detailed analysis activity, and completes an inte-
grated evaluation of overall system operation, Ni sary remedial actions
affecting the planning and conduct of future tests are coordinated as early as
possible with all organizations concerned.

Taz ; —srcrer_



SECTION Vil ——SECRET—
LMSD-437237

system time checks, syatem runs, acquisition messages, tracking data, and
command and performance summaries. .

7-27. COMMUNICATIONS OPERATIONS PROCEDURES. The Communica-
tions Operations Procedures document defines the procedures to be used in

7-28. SAMOS SATELLITE PROGRAMMING.

7-29, Programming of Samos Satellite activities is effected through the com-
mand equipment of the satellite-borne communication system. The command
Program for the Samos Satellite is based on the mission schedule, in that
specific commands must be given to the satellite ona programmed basis in
order to accomplish the scheduled missions. The individual commands are
based, first, on the mission to be accomplished, second, on the vehicle
equipment, and third, on the vehicle orbit. The command equipment controls
the orbital operation of systems that comprise the satellite and payload equip-
ment based on prelaunch inset progr: . Th stored program com.
mands, and those introduced by tracking station during vehicle orbit, form
the basic cyclic Programming of events. This Programming may be modified
by subsequent introduction of ground- station-orginated, real-time commands
if additions, deletions, or revisions of the stored Program content is required.

7-30. The sequence rogrammer is the command Programming device in the
satellite. It is turned on during countdown by Samos ground operations com-
mand and stores Program commands fed into it during countdown and during
useful orbital lif The Programmer reads out these commands to the af-
fected satellite tem and payload equipment for execution of the required

ons. Refer to pertinent, subsequent discussions in this
section for sequential Operation of programmed events and to Section IV for
details relating to each pertinent system of the Samos Satellite,

SAMOS VEHICLE CHECKOUT AT VANDENBERG

7-31. A Samoe Vehicle is not received by Vandenberg Air Force Base in the
assemblied state, but rather, the Samos Booster, the Agena portion of the
Samos Satellite, the dual Samos payloads, and the vehicle adapter are inde-
pendently shipped from the respective contractor facilities, Upon arrival at
Vandenberg, a series of checkouts and inspections are conducted at the
Guided Missi. ly Building to ascertain operational readiness df all
systems that collect; vely comprise a Samos Vehicle. Initially, the vehicles
(booster and Agena) are separately checked at the subsystem level. Upon
Satisfactory completion of these checks, a final system checkout is then con-
ducted at this location. For this system check, both vehicles are completely
assemblied to assure complete systems operation, The Samos payloads are
installed in the Agena vehicle (thus constituting a Samos Satellite) and are
operated to also permit full evaluation of these systems. After successful
system checkout of the Samos Booster and Samos Satellite, the payloads are
removed from the Agena. The vehicles, payloads, and adapter are then
separately transported to the launch complex at Péint Arguello. Refer to

Section IV for detailed information regarding the checks and tests conducted
at the Guided Missile Assembly Building.

7-6 ——_SEGRET—

©




SECTION vz ——SEGRET—
LMSD-437237
7-105. DATA READOUT.

Y

Md

“Le

Y
i

Amma

Readout of the telemetered data is under
T and is eived on both the VHF and
activated prior to contact
ded data throughout the pass.
eption is attainea. The
asa matter of course.

lysis as coon as possible. This
analysis includes nominal tolerance checks to permit determination of eat.
ellite equipment status and evidence of equipment malfunctioning. Results
of postpase analysis are transmitted to the stored Program command for
coordination with Feports from all trackin stations.

DATA EVALUATION
7-112, GENERAL. .

7-113. Lockheed Missiles and Space Division has the overall responsibility
for Samos Syetem evaluation. The scope of thie evaluation encompasses all

7-20 ——stersr_

©

Cs



pee SECTION IV

flow to the fuel tank to supply a greater rate of 52s flow to the tank. This in-
creased flow which is needed during engine Operation is greater than can be
Supplied through the bypass regulator,

4-59. The combination of the bypass re; F and squib-operated valve are
incorporated into the system to help mi: ise the possibility of oxidizer and
fuel backflowing and mixing in a common ressurization line prior to engine
ignition, The two Components perform a kup function to the Pressurization.

‘ Mine check valves by blocking further backflow, should propellants leak throu...
the valves. After ignition, high gas flow in the nee prevent Propellant back.
flow and therefore any possi! of mixing in these lines, .

4-60. After satellite rocket engine shutdown, the Propellant tanks, the spheres

and the tank and aphere lines are vented when a signal from the i ance and
perated vent valves, causing them

sturbances to the attitude control system,

4-61. A line, which tape off of the low-pressure side of the main regulator,
routes helium through a lowepressure r. gulator to the oxidiser-pump double.
Up eeal, to force into the overboard drain any oxidizer that leaks past the
oxidizer-pump Primary seal, . :

4-63. ‘The System consists of two, Solid-propellant, allage-orientation Fock.
ets that are mounted on the after side of the aft equipment rack. The rockets
are fired by dance-and-control Sequence timer. (Refer to
Paragraph 4-4 Each rocket produces a fominal forward thrust of 128 #10
Pounds over rning time of 20% 2 seconds, After the Pockets have burned
for 16 raspadee en, This provides a thereuny burning.
ime overlap betw. engine thrust and thereby maintains
the forward thrust wity produced in the vehicle,

4-64. The Yocket-mounting brackets orient the rocket thrust vectors through
the Samoe Satellite Center-of-gravity, Each bracket incorporates a release
TMmechaniem that Jettisons its rocket casing when rocket thrust drops below a

Predetermined value, thus lightening the vehicle by approximately 17, § Bounds
total weight,

~——s8errT_ 4-21



Propellant Feed cand-Load and

ressurization Systems 4-20
+ + Backflow Prevention 4-21
+ + Gas Loading 4-20
+ + General 4-20
ary Preseurization System
Function 4-20
Oxidizer Draining 4-21

Propeilant Loading 4.20
Propellant Pressurization 4-20
System Shutdown Venting 4-21
Propeliant-Tank Ullage Orientation
System 4-21
« + Rocket Casing Jettison 4-23
os « Rocket Firing 4-21
Propellant System 4-19
+ General 4.19
. Tank Assembly and Mount 4-19
SATELLITE-BORNE COMMUNICATIONS
EQUIPMENT 4.80
Acquisition Signal Tranemitting
Systems 4-96
+ + Effectivity of UHF and VHF
Signal 4-96
+ « UMP Acquisition Signal 4.96
ve IF Acquisition Signal
+ SUHF and VHF Signal Trans-
mission 4-97
Sequence of Acquisition Events
UHF Syetem 4.98
+ UHF Acquisition Events 4-98
System 4-96
ommand Receiving and Processing
Systems
+ + Effectivity of UHF S-Band
Systems 4-92,
+ + Function of UNF and VHF
Systems 4-92
«© « Operation of S-Band
Equipment 4.93
+ + Operation of UHF Equipment 4-95
+ « Ranging Tones 4-93
+ « Real Time Commands
(RTC) 4.93
© + S-Band Tracking 4-94
+ « Stored Program Commands
ISPC) 4.93
+ S-Rand Syetem 4-96
+ SUF and S-Rand Command *
Reception. 4-95
« UHF System 4-94

eee

4-97

Ae see

-Purpose 4-80

+ + Antenna System 4-83
+ + « Received and Transmitted
Signals 4.86
+ «+ Effectivity of UHF and VHF
Equipment 4-80 :
+ Functions of UHF Equipment 4-80
. *Functional Schematic 4-82
Functions of VHF Equipment 4-83
+ Functional Schematic 4-84
Functions Performed by
jwipment 4-80
R + “Location of Components 4-81
e

VHF Transmission 4-88
+ + Effectivity of UHF and VHF
. Systeme 4-83
+ Ferret Reconnaissance
Data 4-83 .
+ + Functions of UNF and VHF
Systems 4.83
+ + Viewal Reconnaissance
4-85

jon Circuit 4-89
Sensing and .
Tranemitting Systems 4-89 .
« CUHF and VHF Transmission 4-91
+ + Commutation of Telemetry
Data 4-90
+ Effectivity of UHF and VHF
Syeteme 4-489
Recording of Telemetry
4-89

m 4-92

iming System 4-98

+ Ground Timing Complex 4-98

+ Ground Timing Signals 4.99 -

+ Satellite-Borne Timing
Generator 4-98

Satellite Time 4-99

« « Satellite Timing Signals 4-99

ee rs



SECTION VI
LMSD-437237

6-10, The Data Processing Section processes and stores large volumes of
data that have many different characteristics, Processing and storage is

Management Group by Providing large-screen, rapidly updated
dynamic information. The Central Reference Section supports all groups by
Ti

ion is partly automated and includes flexible organizational proce-
dures and equipment for indexing, storage, and retrieval,

6-11, MANAGEMENT GROUP,

6-12. The Management Group contains only one section, the Control Section,
This section plans and directs the activities of the Data Processing System,
It plans subsystem operations, supervis: echnical personnel, inspects
output products before their r, + and reviews the general subsystem per-
formance,

6-13, SYSTEM INPUTS,

6-14. System inputs consist of reconnaissance data, simulated audlary
satellite data, and reference data. .

6-15. The reconnaissance data includes the 35mm, positive film strip,

primary records recorded by visual equipment, and the digital magnetic
tape records of e!

6-16. Input reference data for pHfoto data reduction include the best map
Coverage available, geodetic control data, selected Photography, and Photo
interpretation keys. For electronic intelligence data reduction, reference
data consist of external reports and documents that are useful to an Elec-
tronic Intelligence Analyst. Reference data for photo data reduction and
electronic intelligence data reduction will, in many cases, be useful in
Correlating these two functions. . so

6-17, SYSTEM OUTPUTS,

6-18, System outputs include data process products, duplicate repro-
ductions of satellite reconnaissance data, and ease information (refer
to paragraph 6-3),

6-19. Data Processing products include photographs, initial photo ivter-
pretation reports, target position analysis reports, and electronic intelli-
gence initial analysis reports,

——SEGRET— . 63



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SECTION V —— SECRET _
LMSD-437237

versa, programs and controls the flow of digital data through Subsystem H

equipment at the tracking station, and provides intrastation digital data trans- C .

mission facilities. The Digital Handling Subsystem includes the station come
puter (a large capacity, high epeed digital machine) and its input-o:
equipment, the time comparator equipment, the central timing and data dis-
tributor equipment, three remote timing and data distributors, and three
antenna slave data loops. .

5-20. The computer converte the predicted satellite orbit parameters to
antenna positioning data and to visual display signals relative to local station
coordinates and estimated time of acquisition, time indexes the stored Pro-
gram commands with reference to the vehicle time cycle, generates certain
real-time commands, and converts the corrected vehicle path to a corrected
vehicle orbit.

5-21, The time comparator equipment compares the vehicle time code word
with “he ground time code word and supplies the computer with corrected
time information needed to time index the stored program commands.

5-22, The central timing and data distributor equipment controls the flow of
signals and of antenna slave data, to and from the remote timing and data dis-
tributors, one of which is located at each antenna site,

5-23. Each of the remote timing and data distributors separates and routes,

to the using equipment, signals received from. the central timing and data

distributor equipment, The remote timing and data distributor equipment

associated with the Angle Tracking Antenna also supplies the master timing C
and data distributor equipment with antenna Position data,

5-24, Each of the antenna slave data loops compares the position of its asso-
ciated antenna with the slaving signals and generates error signals that are
used by the antenna. servos to Position the antenna,

5-25. GROUND COMMUNICATION SUBSYSTEM.

5-26. The Ground Communication Subsystem provides facilities for inter-
station and intrastation ground communications. Interstation facilities include
leased toll telephone lines and leased teletype lines between the Satellite Test
Center and other stations for ministrative and operational control. This
network will eventually includ pecial wideband facilities for the transmission
of reconnaissance data from certain tracking stations to the Satellite Test
Center, The intrastation facilities include an administrative telephone net,
various maintenance and operational telephone nets, and a voice paging
system, .

5-27, DATA RECEIVING SUBSYSTEM,

5-28. The Data Receiving Subsystem receives signals transmitted from the >
vehicle to the ground. The subsystem includes the Telemetry and Data Re-

ceiving Antenna and ite os, the wideband data receiving channel, the

narrow-band data receiving channel, and the receiver and antenna control

panels.

5-29, If the vehicle being tracked is equipped for visual Feconnaissance,
wideband and narrow-band data signals are transmitted simultaneously by

5-4 SECRET —



SECTION IV

Figure 419, Pneumatic Control System, Principles of Operoti

4-48 SECRET



+ SECRET-
LMSD~437237

SECTION IV

|
i
oa . {
|
I

Figure 4-17, Essentiol Features of Gyre ead Orientation of Gyres in Semos Sotellite
-
—_—SECRET_

4-43



SECTION v1 —7TaBCRET

PAYLOAD DATA PROCESSING SYSTEM
6-2, GENERAL.

6-3. Purpose of the Payload Data Processing System is to provide proces-

" sing and initial analysis of satellite-derived data, and also to convert these
data to a format which is compatible with standard Air Force usage. An
important additional function of the system ie to ald in Proper evaluation and
control of the satellite collection s eedback information
to the Samos System. Feedback consist
parts of the Samos System, These reports relate mission accomplishment,
coneite collection, orbital correction data, and ferret calibration and
control data,

6-4. The Payload Data Processing System interacts principally with the
Visual Reconnaissance System and the Fer: Reconnaissance System by
Processing the data received from both of these systema.

6-5. The Payload Data Processing System functions are accomplished by
three major Operating groups: the Operations Group, Support Group, and
Management Group. These Sroups and the Payload Data Processing system
outputs and inputs are described in the following Paragraphs,

6-6. OPERATIONS GROUP,

6-7, The Operations Group consists of sections responsible for Preparing
output information. Sections within the 0up are the Photo Reduction
Section and the Electronic Intelligence Data Reduction Section. The Photo
Data Reduction Section converts and Processes photographic and Printed

6-8. SUPPORT GROUP,

6-9 The Support Group has three sections which perform Processing, dis-
Play and storage functions. These sections are identified as the Data Pro-
cessing Central Section, Display Section, and the Reference Section,



.

“

— Da

SECTION IV

APPROXIMATE TIME
OF EVENTS IN MIN-

L_IN SECONDS _

BEFORE LAUNCH |

EXACT TIME
OF EVENT

GUIDANCE AND CONTROL EVENT

-2 minutes; before
launch .

Power is applied to Samos Satellite

FROM LAUNCH TO SEPARATION

guidance and control equipment.

Time zero; at launch

+ 4 minutes and 10 ©
seconds; after launch

+4 minutes and 20
seconds

“0

“248

259.6

a. Samos Booster engine:
started. (The Samos B.
Operates for 4 min, 20 sec.)

lb. Power remains applied to Samos
Satellite guidance and control. equip-
ment, but the equi; nt performs
‘no function until aher. separation.

4

Sequence timer brake is released.

Separation of Samos Satellite from
Samos Booster is commanded,

FROM SEPARATION TO START OF SAMOS SATELLITE ROCKET ENGINE

_ seconds

+ 4 minutes and 20

+ 4 minutes and 35
seconds

+ 5 minutes and 15
seconds

260.6

274

315

_| timer,

a. Samos Satellite vehicle is separ-
ated from Samos Booster. (The .
satellite coasts for 4 min, )

b. Guidance system and pneumatic
control system are activated.

a. Samos Satellite is turned from
attitude 25.6 degrees above horizon-
tal to horizontal attitude by sequence

b. Nose cone and horizon scanner
fairings are ejected.

Morison scanner and integrator are
connected to inertial reference

package. .

Figure 4-24, Sequence of Guidence and Control Events (Sheet ¥ of 4)

4-57



SECT
ION VIL REE .

" guidance discrete command tests; instrumentation and telemetry cross-
checks; and backup separation signal tests. In addition to these interface
tests, numerous launch control interconnection test are required, such as
liftoff signal and tone tests; start booster countdown enable switch; proceed .
with commit signal and control; launcher power on signal; commit start
signal; countdown clock interconnection and control; and tiftoff, loop test
start, and countdown clock interconnections to ground station equipment.

7-39. The range safety tests establish that the spaceborne components
meet requirements for flight safety. Destruct simulators are used in both
vehicles and the receipt of destruct command is monitored at both vehicles.
Launch Operations Building remote control of the vehicle safe-arm function
ia checked,

7-40. Before proceeding with the initial launch sequence Preparations, a
complete dress rehearsal is conducted, This simulates a launch (without
engine firing) with all contractor and range personnel participating. The
success of this rehearsal is a major factor in determining flight readiness
and the flight date.

7-41, INTEGRATED FACILITY PRELAUNCH PREPARATIONS,

7-42. During the prelaunch period, the Vandenberg Communications Center
is responsible for monitoring and coordinating those operations at Vanden-
berg Air Force Base which relate to launch. Prelaunch activities are con-
tinuously monitored by the Air Force field office at Vandenberg, Tracking
equipment at all tracking stations is independently and repeatedly calibrated
and adjusted during prelaunch. Independent calibration is required to
that accuracy of tracking data acquired during orbital tracking is within
tablished limits of accuracy for measuring orbital parameters. In general,
primary tracking system calibration is accomplished by the use of fly-by air-
craft which carry equipment to permit tracking by tne tracking station. Dur-
ing prelaunch preparations, the Vandenberg Commuinications

status of the launch complex and the Vandenbe rg Tracking Station. During
the precountdown period of 6 days. prior to launch, the Communications
Center maintains the direct communications link to the telemetry ship and to
the Satellite Test Center, and also reports on weather forecasts and con-
ditions at Vandenberg. Before commencement of the countdown, at approxi-
mately 8 hours before launch, all local control and coordinating reaponsi-
bility at Vandenberg is vested in the Launch Operations Control Center with
the Vandenberg Communications Center performing necessary communi-
cations functions under jurisdiction of the Launch Operations Control Center.
Starting at 6 days prior to launch the Satellite Test Center directs a program
of checkout procedures to assure station readiness for tracking. Procedures
include direction of simulation tests and exerci it each station, limited
interstation communications checks, and d rehearsal c nications
routines simulating the mission requirements which exercises all station
functions under the direct control of the Satellite Test Center. The intra-
station communications checks consist of both the voice-teletype and digital
data communications link: Each tracking station conducts independent,
simulated communicatio: xercises within its respective areas under the
direction of the System Test Director at the Satellite Test Center, Typical
mission operating procedures are emphasized. The interstation communi-
cations check includes both voice-teletype and digital data communications
links to each station under direction of the System Test Director.

7-8 ___SECRET-



ian

irene
LMSD-437237

GLOSSARY (Continwed)
UMBILICAL—Connections that allow fluids and lectrical signals to fow
between a vehicle and the ground Support equipment at the launch complex,

UHF-Ultra high frequency, Electromagnetic radiation in the radio-frequency
range from 300 me to 3000 me, : an .

UDMH-Unsymmetrical dimethylhydrazine; a fuel Propellant.

VHF-Very high frequency. Electromagnetic radiation in the radio-frequency
Fange from 30 mc to 300 mc.

VERNIER ENGINES-Two rocket engines of the Samos Booster that perform
two functions: control of Samos Vehicle attitude, and adjustment o: Samos
Vehicle: velocity before separation,

YAW-Motion of the @atellite about the reference yaw axis.

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~—SECRER?— SECTION VII
LMSD-437237

7-18. SAMOS SATELLITE SYSTEM PROGRAMMING,

7-19, Initial programming for the Samos Satellite System is performed by
such functioning groups as the System Te. Working Group and the Flight
Test Working Group. Various documents are Prepared to effect proper plan-
ning, coordinating, and implementing of system flight objectives. Such doc-
uments include the Detailed Test Objective, the System Test Directive, the
Countdown Manual, the Readout Plan, the Satellite Program System Pro-
cedures, and the Communications Operations Procedures. .

7-20. SYSTEM TEST WORKING GROUP. The System Test Working Group
at Lockheed in Sunnyvale is delegated the authority for preliminary ing
and coordination of the overall system activity is composed of Lockhee:
(LMSD) and Air Force (AFBMD) representatives, with an AFBMD represer.-
ative serving as chairman, The group is responsible for preparing the
System Test Directive document. The.document is based upon the Detailed
Test Objective document which is previously coordinated at AFBMD Head-
Quarters with all contractors involved in the program. Major emphasis of
the document is on the orbit phase of the program.

7-21, FLIGHT TEST WORKING GROUP, The Flight Test Working Group at
Vandenberg Air Force Base is delegated the authority for launch and prelaunch
planning and coordination. The group is composed of Air Force, Loc! d.,
Pacific Missile Range, and associated contractor personnel. A
tive of the AFBMD field office at Vandenberg Air Force Bas ve:
chairman. The group is responsible for preparing the Flight Test Directive,
@ document based on Pacific Missile Range requirements and'on information *
of the Detailed Test Objective to cover detailed launch requirements.

7-22. DETAILED TEST OBJECTIVE. The Detailed Test Objective is the
formal planning source for use by the working groups, Satellite Test Center,
launch base, and tracking stations in planning flight operation Procedure

The document Provides such information as test objectives, vehicle conf: gu-

ration, flight plan for launch and orbit, telemetry instrumentation schedules,
and nominal orbit tracks. .

7-23, SYSTEM TEST DIRECTIVE. The System Test Directive is used in
the conduct of the flight define in the corresponding Detailed Test Objective.
It provides a general description of the overall flight program and specific
details regarding tracking, data recording and instrumentation, and vehicle
command operations under the direction of the Satellite’ Test Center.

7-24, COUNTDOWN MANUAL. The Countdown Manual outlines the tasks
to be completed during the countdown for launch of a specific vehicle. The

sequence of tasks and the nominal time spans for completion of each function
are shown.

7-25. READOUT PLAN. The Readout Plan provides planning information to
establish acquisition and data reception and readout procedures for the system.

7-26. SATELLITE PROGRAM SYSTEM PROCEDURES, This document de-
fines in detail the communications required to conduct system flight teats and
include system countdown, weather reporting, discrepancy summaries,

. ——SEGRET__ 7-5

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SECTION Iv

POWER PACKAGE
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INTEGRATED
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LMSD-437237
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VALVE. ACTUATOR

> ~ MECHANICAL Lt

—— ELECTRICAL CONNECTION

INKAGE

HORIZON
SCANNER

PITCH DEMOOULATOR |
AND RATE NETWORK

PITCH A-C
AMPLIFIER

COMPUTER

bene eee

Figure 4-20. Hydraulic Control System, Functional! Schematic
——srerer—



SECTION Iv

“SECRET _
LMSD-437237

Components in the forward equipment rack, The fluid fittings include four .
for propellant filling and venting, and one each for helium filling, nitrogen/ C :
freon filling, and air conditioning. The fittings are made of stainless steel]

and aluminum, and they are provided with internal sealing materials that are

compatible with their individual applications. The Couplings are disconnected

at launch by a lanyard pull system. All Sittings except the air Conditioning

duct, which is covéred by the cooling access loor, are self-sealing when
disconnected. .

4-321. SATELLITE SERVICE REQUIREMENTS, The satellite service
quirements include loading the Propellants into the satellite tanks, fli:
helium and nitrogen/freon spheres, conditioning the air and maintaining the

temperature within the satellite, and connecting the satellite Pyrotechnic de-

vices. Provisions are aleo required for removing satellite fluids in the event
of an aborted flight.

4-322. The only satellite service requirement at the guided missile assembly
building is for air conditioning during system checkout when the electronic
Components are being operated in place within the Satellite spaceframe, Dur-
ing this time, conditioned air is provided by a type 14, mobile, air condition-
ing trailer. The conditioned air maintains the temperature within the space-
t .

4-323, All satellite Service requirements mentioned in Paragraph 4-32) are
met at the launch pad complex by the special ground servicing equipment dis-
cussed below in Paragraphs 4.324 through 4-327, This equipment is stationed
at the launch pad and service building,

4-324, Loading the_propellants into the satellite tanks is accomplished with
the use of two Propellant transfer sets, One is used for the UDMH fuel, and
the other is-used for the IRFNA oxidizer, Each one transfers its propellant
from the propellant transport trailer into a Propellant weighing tank and, upon

Propellant into the satel! lant tank. The approximate maximum weight
is 1850 pounds of fuel and 4800 pounds of oxidizer. In addition to weighing the
ions for maintaining the Propellant

T sets are also used for removing the pro-
vent of an aborted flight, and for flushing ©

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4-325. Filling the helium and nitrogen/freon spheres of the satellite is
accomplished with the use of gas pressurization equipment located at the
launch pad and service building, The equipme: apable of delivering

it pr up to
ings that are used for filling
are specially designed to Prevent high shock loading when the couplings are
disconnected at these high pressures,

cooling access door,
the outer surface of



, .
LMSD-437237 SECTION Iv

4-91, PRIMARY BATTERIES, TYPE |, (See figure 4.10.)
silver-peroxide/zinc batterie ve as the ime electrical
for Night configuration I Samos
aft equipment rack (figure 4.4)
Each battery weighs 106 pou:
closed within a magnesium ca,
hours at a nominal pot.
energy-to-weight fact
Parallel-connected diodes in series
vide circulating current Protection,
unit in the event of malfunction, to thu
battery bank. Each battery is eq

to maintain internal Pressures within the range of 5 to 15 pounds Per square
inch, :

4-92. 2000-CYCLE, STATIC INVERTERS, (See figure 4-10.) Two 2000-
cycle, single-phase, static inverters, Controlled by an inductance- capacitance
circuit are used in the iaternal power syatem to coavert d-c batte. power to
alternating current power. The two inverters are transistor-oecillator Power
amplifiers that furnish the main 115 volt, 2000 cycle, Single-phase a-¢ Power
for the vehicle telemetering equipment and for the Payload. Dual inverters
@re used to meet the heavy electrical loads of such equipment,

ater-series combination. The booster operates from the 201 eye!
verter power source, thu: liminating the requirement for an associated in-
verter. The regulator instructed with efficient heat transfer detween
diodes and the metallic case Structure. All Components are mounted mechan-
feally rigid, and the entire unit is potted to Provide a sealed construction,

4-94. NEGATIVE 28. VOLT POWER SUPPLY. (See figure 4-10.) The = 28-volt
power supply is a conventional full-wave transformer- rectifier arranged for |
Operation from the 200-cycle inverters. Regulation is accomplished through

the Magnetic-amplifier circuit of the No. 1 +28-volt regulator.

4-95. 400-CrcLe, 3- PHASE INVERTER, (See figure 4-10.) The 400-cycle,
Phase inverter is a transistorized Power inverter. its frequency is control-
19. 2-kilecycle, quarts-cryetal- controlled Cecillator, The unit incor.

power input filters to Featrict noise reflection into the battery circuit,

ore eliminate noise feedback that would pass from the inverter tran.

oF ewitching circuits through the battery to the low-signal-level electronic

‘ment. :

400-CYCLE, SINGLE-PHASE, A-C REGULATOR, (See figure 4-10.)
400-cyecle, single-phase, a-c regulator accepts and maintains a root
Square output of 115 vo! @. The voltage output is controlled by means
of an input series magnetic amplifier, Integral: circuits Prevent phase
shift between the input and output voltage from exceeding 3.5 pes-ent.
All Power-handiing elements in the unit are located on the base of the

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4-27



SECTION vii
Ma

operations. He conducts the booster countdown, monitors activities of
Booster Senior Launch Engineers, makes Operational decisions relative to
the booster, and reports progress Periodically to the Launch Controller,

7-12. Pad safety is under Cognizance of the Pad Safety Officer, an Air

responsible for evaluating pad operations during countdown to assure maxi-
mum safety of personnel, the launch complex, and vehicle. He advises the
Launch Controller on pad safety matters luring the countdown,

7-13. Overall flight safety is under cognizance of the Range Safety Officer
for the Pacific Missile Range. He is responsible for evaluating range status
to determine whether or not the launch meets all safety requirements.
Through the media of the Missile Flight Safety System, he follows the tra-
Jectory from launch and may command destruct if it deviates from the pre-
scribed safe path. He advises the Launch Controller and is the final author.
ity on launch-countdown holds necessitated by unsafe range conditions. He
carries full responsibility for commanded destruct,

7-14, SATELLITE TEST CENTER AND LAUNCH BASE COORDINATION,

The Satellite Test Center (STC) and the Launch Operations Control Center

(LOCC) exercise command and control of Prelaunch and launch coordinated
Operations. Vandenbe rg launch control is executed in accordance with the

applicable Flight Test Directive. The tracking stations are involved in the

Prelaunch and launch phase to the extent that they relay information and

Participate in time checke and checkouts as directed by the Satellite Test (
Center, During these times the Test Center maintains overall command “ot
and cyntrol of system operations; however, Specific command Prerogatives

are delegated to the Launch Ope rations Control Center. During launch, sole
responsibility for range safety and destruct of the Samos Vehicle rests with

the Range Safety Officer of the Pacific Missile Range, who may command

deatruct at any time between liftoff and sustainer engine cutoff of the Samos

Booster,

7-15. During the orbit Phases of operation, command and control over a
Samos flight executed by the Satellite Test Center in accordance with
the System T. Directive. Control over the tracking stations is accom.

plished through a communications network of which the Te.
hub. A ground- Space communications system is used whe
stations relay commands and readout satellite and reconna: ance data. in
thie capacity, the stations serve as instruments of the Test Center, .

INITIAL PROGRAMMING

Center is the

7-16, GENERAL,

7-17, For convenience of pre tation, initial Programming is discussed -
‘under two; main categori: first, the Planning and Preparation that is re-

quired on an overall Samos Satellite System basis, and condly, the pro-

gramming which is directly applicable to a Samos Satellite. Of necessity, a

close relationship eiists between system programming and vehicle Program-

ming. :

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SECTION VI
LMSD-437237.

6-20. The photographic Products include copies of the Data Processing
Section records, reproductions of reassembled primary records, lected
frames from these rolla, and annotated photography. The Photographic
products are appropriately titled and produced in Proper formats,

6-21. The photo interpretation reports produced include the first-phase

intelligence reports in formats that can be readily utilized by using agencies.
Indexed overlays and machine-readable tapes containing information derived
from photo data reduction are also produced. Electronic intelligence analy-
sis reports include flash reports, immediate reports, and detailed interpre-

tation reports, similary conforming with the anticipated operation require-
- ments of the using intelligence agencies,


SECTION IV yARRETS

4-314. SAMOS SATELLITE MAINTENANCE, Samos Satellite maintenance
includes repair of a specific Category of defects, and installation nf Tequired
modifications. All repair work is performed within the guided missile as-

POWER AND SERVICE REQUIREMENTS
4-315. POWER REQUIREMENTS, ,

4-316. SATELLITE ELECTRICAL POWER REQUIREMENTS AT GUIDED
MISSILE ASSEMBLY BUILDING. The satellite electrical power requirements
at the guided missile assembly building are for the voltages that are nec:

to operate the aystem and subsystem componente du:

ages: +28 and -28 volts, d.c.; 115 volts, 400 cps, sin le-phase; 115 volts,
400 cps, three-phase; and 115 volts, 2000 cps, s phase. These voltages
are provided by power supplies associated with the special test equipment
used for the checkout procedures at the guided missile assembly building.

4-317. SATELLITE ELECTRICAL POWER REQUIREMENT AT LAUNCH PAD
COMPLEX, The satellite electrical power requirement at the launch pad com.
plex is exclusively for 28 volts, d.c. This voltage is required to substitute
for the satellite internal battery power during the days before launch. It is
used to operate the system and subsystem components while the final Prepara-
tions are being completed and the performance of the components is being
monitored. The required 28-volt, dec power is provided at 100 amperes by
power supplies in the pad Power supply and signal Processing set at the launch
pad and service building, .

4-318. SERVICE REQ UIREMENTS,

4-319. ACCESS AND INSPECTION FEATURES FOR SAMOS SATELLITE
SERVICING. The access and inspection features for Samos Satellite servic-
ing consist of removal ‘8, and spring-loaded doors
i stings. Four removable
rward equipment rack to

4-320. All ground servicing lines connect to the Samos Satellite through the
quick disconnect fittings that are covered by spring-loaded doors, {See
figure 4-37.) The electrical fitting is located directly over the electronic

4-114 —-SEGRET—



“
:
:

|



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Y

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ECTION IV
Sze CMB AEs

period during the orbit phase when the ferret pay-
jected. The hydraulic control system ie located at the rear of the
lite and the two hydraulic actuators of the system are mounted on the
tellite rocket engine. The hydraulic control system is activated only when
the satellite rocket engine is firing. The orbital attitude damping system is

tellite and the two inertial reaction whe
mounted in the forward equipment rack. The orbital attitude damping
tem is activated only during the orbit phase. The equipment that forme the
control system and the functions they Perform during the ascent and orbit
phases are described in Paragraphs 4-137 and 4-145,

4-119,. OPERATION OF GUIDANCE AND CONTROL SYSTEMS, The Operation
of the guidance and control systems is illustrated in figure 4-16. The pitch,
roll, and yaw gyros in the inertial reference package provide attitude error

attitude error signals are amplified by the flight control electronics unit and
distributed to the Pneumatic and hydraulic control systems which

the Samos Satellite at the prescribed attitude. The sequence timer is the
computer provides the Starting signal for the satellite rocket engine, and the
accelerometer in the inertial reference package ie connected to the integrator

in the computer to meas: ity gained by the satellite. When the
velocity of the

engine. The «.

4-121. GENERAL, Guidance during the ascent, Phase is provided by the iner-
tial reference Package, the flight control electronics unit, the computer, and
the horizon scanner. The operation of these major equipment items is des-
cribed in Paragraphe 4-124 through 4-132, .

4-123, During the Operation of the satellite rocket en; + the gyros, the
flight control electronics unit, and the hurizon scanner operate as they did
during the 6t period, However, the pitch attitude Ptogram command sig-
hal is provided by the integrator in the computer rather than the sequence
integrator also detects the velocity gained as the Samos Satellite
‘ates by measuring the input from the ac erometer in the inertial
reference Package. The integrator cuts off the satellite rocket engine when

4-40 —SECRET_



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P Tos

POINT MUGU TRACKING STATION,
See: Operating Bases, Centers,
and Stations

POWER REQUIREMENTS, See:
Satellite

POWER SYSTEM,
Samos Satellite

PRELAUNCH OPERATIONS AND
COUNTDOWN, See: Samos System
Integrated Operations

PREORBIT PHASE, See: Samos
System Integrated Operation

PROGRAMMING, See:
Integrated Operation

Samos

ELECTRICAL, See:

Samos System
PROPELLANT SYSTEM, See:
Satellite

PROPULSION SYSTEM, See:
Satellite,

Samos

Samos
and Samos Booster

RECONNAISSANCE PHASES, See:
Samos System Integrated Operation,
Ground-Space Communications and
Control, and Samos Satellite

RECONNAISSANCE SYSTEMS, See:
Samos Sateilite

ROCKET ENGINE, See: Samoe
Satellite

$s
SAMOS OSTE 3e1

ATLAS MISSILE 3-2
Airframe 3.2

+ Description 3.2
Electrical Power

. Description 3-3
Flight Control System 3.3
+ Description 3-3

3-3

“SECRET _
LMSD-437237

Guidance 3.3
- Description 3-3 : C ‘
Hydraulics 3-3
+ Description 3.3
Instrumentation and Range Safety 3-3
+ Description 3-3
Propulsion System 3-2 Sos
- Descri 3-2
INTRODUCTION AND SECTION
EMENT 3-2
ee Scope of Section 3-2

YLOAD 4-65 :
General Characteristics 4-65
+ Calibration 4.63
+ Determination of Transmitter

Location 4-66
ory sauyels Accuracy 4.68
ee ception and

auipae 4-66
+ *Location of Com, ments 4-67
+ Measurement of Signal

Facteristics 4-66
+ Orbital

jor Equipment 4-69
Major Eaulpment, . .
oe Configuration I Modification 4-70
+ Data Handling Unit 4-70
+ + Computer Measurements 4-71
+s Signal Verification 4-71
o Ferret Equipment Pro, ‘ammer 4-7]
+ + Calibration Phase 4-72
++ Emergency Calibration
Phase 4.72

Magnetic Tape Recorder 4.71
Receivers 4-79
Purpose 4-65 _
+ Frequency Bands 4-65
« General 4.65
GROUND SUPPORT EQUIPMENT 4-117
+ General 4-117
+ Ground Handling Equipment 4-117

: General 4.69

____ SECRET—


LMSD-437237

GLOSSARY (Continved)

RETROGRADE ROCKET-A device, that when fired, retards forward move-
ment of the Samos Booster,

SQUIB-An electrically detonated explosive device,

SUSTAINER ENGINE-The main rocket engine Permanently attached to the
Samos Booster.

SAMOS BOOSTER-A Convair Atlas missile (SM-65D) modified to mate with a
Samoa Satellite and provide first-stage thruet for Placing the satellite in orbit.

SAMOS SATELLITE-A Lockheed Agena orbital vehicle (Agena) contai,
Samos Payload and using a Samos Booster for first-stag th.

rocket engine that is Part of the Samos Satellite provid
for placing the Satellite in orbit, .

SAMOS SATELLITE SYSTEM-An array of equipment, Personnel, and facilities
designed to obtain and process electronic and Photographic reconnaissance d
The array includes the Samos Vehicle, the launch complex, the operating

and tracking Stations, and the world-wide, Sround -space communications
network,

SUBSYSTEM~A Combination of components that performs a single function or

a number of related functions, A subsystem is interrelated with other eub-
syetems to form a stem, :

S-BAND~-Electromagnetic radiation in the radio-frequency range from 1550 mc
to $200 me,

STORED PROGRAM COMMAND~A command stored within the Samos Vehicle

and executed at some later time. Used to initiate an equipment function
automatically,

TELEMETER-The function of measuring a quantity and tranemitting the

measurement toa station some distance away. Such measu: iments include
those of temperature, pre ure, velocity, and voltage. A Dn electronic
instrument designed to Perform thie function,

THRUST CHA MBER-The Combustion and expansion chamber of a rocket engine,
in which Propellants are burned to produce thrust.

ULLAGE~th a propellant tank, that part of the volume not occupied by
Propellant,

G-4 ——~srenee

(>


—rMBSis7 SECTION IV

4-110. LAUNCH PAD AND COUNSDOWN PHASE CONCEPT. The rincipal
activities relating to the internal electrical power system while the os
Satellite is at the launch pad include primary battery installation Prior to
hoisting and mating of the Samos Satellite to the Samos Booster. Also, final
checkout of Samos Satellite functions is made through auto-sequenced program.
mer and monitor consoles and continued as required until all equipment checks
out properly. During final Countdown, internal and external electrical power
ia monitored during the final testing, servicing, and monitoring of the Samos
Satellite and Samos Booster equipment and supporting facilities.

4-111. SUPPORT PERSONNEL. The Samos Satellite during its time at the
launch base is handled by a satellite cre: tellite bench maintenance crew,

and a launch Pad complex maintenance The satellite crew has complete
responsibilit: rom arrival at the base until
launch, This crew is su; Of the guided missile as-

lite
for launch, is pe unch pad complex maintenance
crew. The Samos Satellite and its adapter are hoisted mated to the Samos
Booster by t ellite crew, whick then completes all connections and instal-
lations. The launch pad complex crew Provides satellite testing during prep-
aration for launch and, therein, is augmented as needed by the satellite crew
in preparing the Samos Satellite for launch,

4-112, SUPPORT EQUIPMENT. The various support equipment that is used
to check out and prepare the internal electrical power system for launch in.
cludes the vehicle Power monitor console, the auxillary power checkout unit,
the universal power Support, the discharge load tester for primary batteries,
the dual battery charger, battery activation and handling equipment, and the
overall systems checkout complex.

SAMOS SATELLITE-SORNE GUIDANCE AND CONTROL sysTexs .
4-113. PURPOSE,

to perform the vital function The guidance sys-
signals for the control system, and the
control system chan, of the satellite in reaponse to the guidance

signals. The location of the major equipment items that form the two systems
is shown in figure 4-13,

4-115. Since the operations Performed by the
during Samos Satellite ascent are different from those performed whilethe
eliite isinorbit, and since different guidance signals

rately. The ascent phase includes the separation of the Samoe
the Samos Booster, the Operation of the eat ket. engine, and. the
orbit reorientation of the satellite from ite horizontal attitude to a vertical,

——__SECRET— 4-35

th



oO

LMSD-437237

TR cen ene ener nega

ALPHABETICAL INDEX
Systems and major headings concerned with the Samos System are listed

alphabetically in the following

Items in the index are listed in

alphabetical Groupings under each system or major heading and corre-

"pond to titles and subjects covered

Asan example: To

find i;
look up POINT ARGUELLO LAUNCH

in the manual,

information on the Missile Service Tower, first
BASE under the major heading of

3 under the Launch Base

Meting, find the Launch Pad and Service Building group, and the Missile

Service Tower is lsted in alphabetical order in that

group.

Cross references are Usted throughout the index and refer to major
headings and groups. An asterisk (#) Preceding a listed item indicates

the location of an illustration.
A-

ACQUISITION STATIONS, See: TRACK-
ING STATIONS in Operating Bases,
Centers, and Stations . :

ADMINISTRATION OF SAMOS SYSTEM,
See: Samos System General
Description .

AGENA VEHICLE, See: Samos Satellite

AIRBORNE COMMUNICATIONS, See:
Samos Satellite .

AIRFRAME, See: Spaceframe in Samos
Satellite

ATLAS MISSILE, See: Samos Booster
8
BELL ENGINE, See; ‘Samos Satellite
BOOSTER, See: Samoe Booster
c

CHECKOUT, See: Samos Satellite, and
Samos Satcilite Integrated Operation

COMMAND AND CONTROL, s.
Samos System Integrated Operation,
Ground -Space Communications and
Control, and Satellite Test Center in
Operating Bases, Centers, and
Stations .

COMMUNICATIONS AND CONTROL,

See: Ground -Space Communications
and Control

COMMUNICATIONS EQUIPMENT, See:
Samos Satellite, and Ground-Space
Communications and Control

CONTROL AND COORDINATION, See:
Samos System Integrated Operation

COORDINATION AND CONTROL, See:
Samoe System Integrated Operation

COUNTDOWN, See: Samos System
Integrated Operations

DATA EVALUATION, See: Samos
System Integrated Ope:
Payload Data Proces

DATA PROCESSING, See: Payload
- Data Processing System, Ground-
Space Communications and Control, ;
and Samos System Integrated ;

Operation H
PR UNITION OF TERMS, See: Glossar

Page G-

DESCRIPTION OF SAMOS SYSTEM,

See: Samos System General
Description

"SECRET x1


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SECTION Iv ——sBenrr—
LMSD-437237
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as AND VALVES

Figure 4.9, Setellite Rocker Engine, Electric! Control

———SEGRET—

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LMSBS 37 SECTION Iv

e .
The pressure regulator in the pressure
shown in the diagram of figure 4-18, to

r) P re between 150 to 190 psig. Of the six
gs control valves and gas reaction jets, two are used exclusively for yaw con-
trol, and the other four for beth pitch and roll control. Only two of
the roll gas reaction Je hown in the diagram. When a roll torque is
commanded by the d-c ra: ignal, the correct pair of roll gas control valves
are opened. Gas is expelled through the roll gas reaction jets to correct the
attitude of the satellite in roll, The operation is the same for the pitch and
yaw reaction jets. The principles of operation of the pneumatic contro}
system are illustrated in figure 4-19,

4-141. The high Pressure relief valve opens {f the supply pressure to the
pressure regulator exceeds 3500 psig. The low pressure relief valve opens

if the regulated Pressure to the gas control valves exceeds 200 psig. The two
relief valves are u: to prevent damage to the system due to thermal expan-
sion of the gas or irges of p: lure caused by rapid cycling of ng gee con.
trol valves. The filter removes particles. from the ges that exceed 25 microns
in size.

4-142. HYDRAULIC CONTROL SYSTEM, The hydraulic control system po-
sitions the satellite rocket engine while the engine is Operating to control the
attitude of the Samos Satellite in pitch and yaw. The system consists of the
hydraulic power unit and integrated hydraulic package; two servo valves,
hydraulic actuators, and position sensors; and the nec sary hydraulic con-
nections, These components are sented in figure 4-20. A diagram is
also included in figure 4-20 to illus atlon of the hydraulic power
unit and integrated hydraulic package with the fitch actuator, and to show how
the actuator is controlled by guidance signals from the flight control electronics
unit.

4-143. The hydraulic control system is pressurized by the. 28-volt d-c motor
which drives a hydraulic pump. The MIL-H-5606 hydraulic fluid is applied, |
as shown in the diagram of figure 4-20, to the pitch servo valve ata pressure
of approximately 2400 psi. When a change in the pitch Position of the satellite
rocket engine is commanded by the d-c rate si » the pitch servo valve is
Opened and the pressurized hydraulic fluid drives the Pitch actuator. Move-
ment of the pitch actuator turns the satellite rocket engine on its gimbal to
Correct the attitude of the satellite in pitch. The operation is the same for
the yaw actuator. The principles of operation of the engine gimballing sys-
tem are illustrated in figure 4-21. .

4-144. The Position sensor in each actuator detects the position of the actu-
ator and provides a voltage signal Proportional to the
signal is returned to the d-c amplifiers in the flight control electronics unit
as feedback to improve the Fesponse of the actuator to the gui

from the hydraulic fi
ransducer provides a Pressure telemetry

——sBeRET_ , 4-49



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LMSD-437237
SPACEFRAME 4.3 VISUAL RECONNAISSANCE (
Description 4.3 PAYLOAD 4-72
+ + General 4.3 Combined Ferret and Visual
+ + Purpose 4.3 Payloads 4.74
+ “Inboard Profile 4.4 Functional Description 4-79
- *Major Assemblies 4-2 « *E-1 Payload Operation Phases 4-78
Flight Destruct System 4-9 « . Film Operation During
: + Description 4.9 Readout 4-80
Midbody Equipment 4-7 + « Film Path Camera to
+ Aft Midbody Equipment 4-9 Processor 4-79
+ Aft Rack Equipment 4-9 ee Film Path Supply to Camera 4-79
+ Aft Equipment Rack and Installed « « Film Synchronization Payload to
Components 4-8 curound ‘80 8
« Forward Midbody Equipment 4-7 Major omponente 4-
+ General 4.7 « Film 4-75
Miscellaneous Equipment 4-13 + + Test Film 4-75
+ Antennas 4-13 + Film Transport Mechanism, Reels,
+ General 4-13 . and re 4-75
+ Intercabling and Wiring « General 4.75 .
Harness 4.13 + Horizon Optice 4-75 .
+ + Construction 4-13 « Processor 4-76
+ Plumbing Systeme 4-13 « Readout Mechanism 4-76
+ + Attitude Gas Reaction Jet + Reconnaissance Camera 4-75
System 4-13 Principles of Operation 4-72
+ + Propellant Feed-and-Load Purpose 4-72
System 4-13 » *Location of Components 4-73 -
+ + Propellant Preseurization Special Provisions 4-77 . Ca
Syetem 4-13 + Attitude Indentification 4-77
. Pressurization Spheres 4-13 ee Horizon Optice 4079
Nose Section Equipment 4-7 + « Horizon Images 4-79
«+ Description 4-7 + Exposure Control 4-77
Pyrotechnic Devices 4-01 « Focus Control 4.77
Pyrotechnic Devices 4-12 + General 4.77 .
: + Description 4.11 + Image Motion Compensation 4.
Satellite and Booster Adapter + Time Indentification 4-79
Equipment 4.9 2 + Time Channele 4-79
+ General 4.9 Structural Description 4-74:
+ Separation System 4-9 + General 4-76
+ + Separation Sequence 4-10 + Lower Pan and Supporting
Samos Satellite Structure 4-3 Structure 4.74
+ General 4.5 + Mounting Grid 4.75
Nose Section Structure 4-3 5 + Pressure Shell 4.74
+ Midbody Section Structure 4.
+ + Aft Equipment Rach 6-8 SAMOS SYSTEM GENERAL
+ + Aft Midbudy Structure 4-5 DESCRIPTION 1-1
- + Forward Midbody Structure 4-5 INTRODUCTION AND SECTION -
+ + *Forward Equipment Rack and ARRANGEMENT 1.3
Installed Componente 4-6 + « Scope of Section 3-3

- Propellant Tank 4-5
Satellite and Booster Adapter
Structure 4.7

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SECTION IV BDI

the satellite reaches the correct velocity.. After the satellite rocket engine is

cut off, the satellite is changed from its horizontal attitude to @ vertical, nose- Cc

down attitude for the orbit phase, During this orbit reorientation the horizon

scanner is deactivated and the gyros in the inertial reference package are con-

trolled by the pitch attitude Program command signal from the sequence timer,
fPamplsfied by the flight con.

trol electronics unit and distributed to the control system. Guidance after

orbit reorientation, during the orbit Phase, is described in Paragraph 4-133,

Syros from the computer and the horizon Scanner. Thay also amplify.
i that are sent from the gyros to the control system and the signal
that is sent from the accelerometer to the integrator in the computer,

4-125, The three gyros in the inertial reference Package are Operated ag.
integrating, displacement Syros during the ascent phase and they are elec
trically caged during the orbit phase and operated as rate Gyros. See

figure 4-17 for an illustration of the gyre Parts and for the general orientation
of the gyros in the Samos Satellite. Each id
dicular to the *pin axis of

gyro. The input axis of the gyro is Perpendicular to bath the spin axis and the

4-126. The output axis rotor of each gyro pas
on one side of the gyro,

rotor directly, and an attitud,
erator to correct the consta:
pitch attitude Program com:
in this way.

4-127. The accelerometer is a Pendulum that is electricaliy restrained by a
restraint motor, The unbalanced weight of the Pendulum is free to move

around a rotor that also pa.
Generator, Aa the Samoa

‘8 through the restraint motor and an a-c signal
ellite is accelerated by the satellite rocket
~

4-42 ——_SECRET—



——SEERET— SECTION .V
LMSD-437237
C APPROXIMATE TIME | EXACT TIME .
, OF EVENT IN MIN- OF EVENT | GUIDANCE AND CONTROL EVENT

UTES AND SECONDS

IN SECONDS

ORBIT PHASE

+13 minutes and 40
seconds

+13 minutes and 50
seconds

818

a. Pneumatic control system is
deactivated,

b. Gyros in inertial reference
package are electrically caged.

c. Samos Satellite is pitched at
orbital rate by gravity gradient
and orbital angular momentum to
establish and maintain vertical,
nose-down attitude as equilibrium
position.

d. Signal from computer is appli-
ed to pitch gyro to compensate for
orbital pitching of Samos Satellite.

a, Orbital attitude damping system
is activated to control disturbances
about equilibrium position,

b. Sequence timer turns itself off.

c Orbital attitude damping system]
continues to control disturbances
about equilibrium position until
ejection of ferret payload.

EJECTION OF FERRET PAYLOAD

Ejection time sero ETZ +0 a Sequence timer is re-started
(ETZ); chosen to be by sequence programmer.
approximately one
and one-half days bd. Pneumatic control system is
after launch, (See reactivated,
par. 4-171.)
c. Orbital attitude damping system!
is deactivated,
d. Cable wires to ferret payload
are cut.
Cc Figure 4-24, Sequence of Guidance ond Contro! Events (Sheet 3 of 4)
———SECRET— 4-59



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LMSD-437237 SECTION Iv

The synchronous motor is supplied with a 28-volt, dec clutch/brake. The
brake can stop the running motor within 0.05 second, and when it is Feleased,
the motor gains full speed within 0.1 second. .

4-131. The integrator is a Motor-tachometer unit that is controlled by the
signal from the accelerometer in the inertial reference package. The signal
from the accelerometer is amplified by transistor amplifiers in the integrator
package and applied to the integrator motor. The torque on the motor is Pro-
Portional to the amplitude of the input signal, and the number of revolutions

it turns is proportional to the velocity gained by the Samos Satellite.’ After
the motor has completed a pre-determine

is closed through a gear train to cut off the satellite rocket engine. The motor
also adjusts, through another gear train, a pre-determined, variable voltage
that ie sent to the inertial reference package. This voltage is the pitch atti-
tude program command signal from the integrator, The tachometer senses
the speed of the motor and Provides a voltage signal ‘oportional to the speed.
This voltage signal is returned to the transistor am; ers in the inte, rator
Package as feedback to improve the response of the motor to the signal front
the accelerometer. A velocity meter is scheduled to.replace the acceler-
ometer on later Samos flights. The velocity meter will also be controlled by
an accelerometer signal, but it will be capable of handling more than one
velocity setting, .

4-132. HORIZON SCANNER. The horizon scanner is an optical, infrared de-
vice used to detect the Position of the earth in relation to the attitude of the
Samos Satellite. It consists of a scanning mirror, an infrared radiation de-
tector, and an electronic circuit to generate signals proportional to any infra;
red radiation that is detected. The amount of infrared radiation detected from

the earth, horizon to horizon, is greater than that detected from Space. When
the Samos

in front of the reference pitch axis ie equal to that behind, Using these infra-
red radiation fields, the horizon scanner detects any constant deviation about
the reference pitch and roll axes from the required horizontal attitude and it
generates voltage signals Proportional to the deviation, These voltage sig-
nals are the pitch roll attitude reference signals that are sent to the gyros
in the inertial reference package to correct for constant Pitch and roll attitude
errors. The horizon scanner is used on the initial Samos flights, but it is
scheduled to be replaced by the Model 2 horizon sensor on later flights.

4-133. GUIDANCE DURING ORBIT PHASE,

4-134, GENERAL, Guidance during the orbit phase is provided entirely by
the gyros in the inertial reference package. Throughout the orbit pha, ex-
cept for the brief period when the ferret payload is ejected, the sytos are
electrically caged and operated as rate ros, and the attitude error signals
from the pitch and yaw gyros are a; directly to the orbital attitude damp. -
ing system. When the ferret payload is ejected the Sequence timer and the
flight control electronics unit are reactivated to distribute attitude error sig-
nals from the gyros to the Pneumatic control system. The Operations of the
inertial reference Package, the flight control tronics unit, and the se-
quence timer in the computer are the same as deacribed in Paragraphs 4-124
through 4-130. The horizon sensor, which Performs no guidance functions, *

is operated during the orbit Phase. Its operation is described in Paragraphs
4-135 and 4-136,

——SEGRET_ . 4-45

el seemensaie



Oper-tional Readiness Concept 4-34
+ Electricat Checks During Systems
4

4-34

+ Subsystem Checkout Concept 4-34
ower and Service Requirements “4-32

+ Electrica} Power Requirements 4-32

+ + *Power Requirements 4-33

Inspection
Features 4-34

+ + Satellite ang GSE Air

‘Onditioning

s+ Scope Of Section 4.3 :
PERATIONAL CHECKOUT AND
MAINTENANCE CONCEPT 4-99
rational Checkout Concept 4-99
+ Checkout at Launch Pad 4-11h
. Mating 4-11)
+ + Betore Launch 4-112
+ + Before Mating 4.111
+ + Control and Monitoring Equipment
atLOB 4.112
Equipment at Pad and Service
Building 4-113
+ Checkout Sequence 4-100 |
+ + Payload Checkout 4-100
+ Yeneral 4.99
+ System Checkout at GMAB 4-108
. Quipment 4-109 .
* sup Totedure 4-109
. Subsystem Checkour at GMAB 4-100
oe Communications Equipment 4-106
soe). Checkout Equipment 4-107
s+ Electrical Power 4-102
+ + + Checkout Equipment 4-102
+ + Geophysical Environmental
Inatrumentation 4-108
widance and Control
+ + Chechout Equipment
+ Payloade 4-105 .
++ Checkout Equipment
+“ Populeion 4-101
+ + Checkout Equipment
+ Spaceframe 4-108
aFatiunal Maintenance
oncept 4-113
General 4-013
+ Satetite Maintenance 4-1he

MENTS 4-114

POWER AND SERVICE
REQUIR

.

.
.
.
.
.
.
.
.

. Power Requirements 4-114

Requirements atGMAB 4.114
wirements at Launch Pad 4-114

Access and Inspection Features 4-114

rings and Access

« Quick Disconnect Fittings
Satellite -

Pad 4-116
+ Equipment at Launch Pag 4-116
« Gas illing 4-116
+ - Propellang Loading 4-116
5 Frotechnic res 4-117
sy: 4-

Electrical Controt System 4-25

+ Power Relay 4.25

+ System Schematic 4-24

+ Thrust-Chamber and Shutdown

. ays 64.25

Gas Generator 4-22 :

« Generator Operation 4022
val 4-22 .

Turbopum Propetiant Feed 4.22
+ Propeltan Pumps 4-22
+ Turbine 4-22

Familiarization 4016

.

Functional De

neral 4.14
« Scope of Text 414
Leading Particulars 4-14
scription 4.14
*Fliow Diagram 4-16
General 4.14 .
+ Ignition 4-18
. ‘-Minimum-Pressure Switch

Actuation

° Oxidizer Fiow 4-18
‘s Shutdown Sequence 4-18
Power Input Requirements 4o19
Sequence of System Ope ration dole



SECTION IV

used previously, The
and many additional ones

4-309. Approximately two hours before launch;

and monitoring functions are Perfo:
umbilical cable connections enable
until launch. The groundspace co:
control of the launch operations bi

rmed in the launch operations
continued monitoring
vidoe cation Links are
ding, first equipment in the ded
missile assembly building, and, finally, by the Vandenber, trac’ stein
Minutes before launch, a command from the launch operationa bui: re-
moves the external power from the Samos Satellite and applies the internal
battery power, Except for the monitoring, which continues until launch, the

LMSD- 437237

the inspection and access features of the satellite,
ne component replacement can be accomplished at this time. Special
subsystem test equipment from the guided missile assembly building can be
used to assist in isolating any malfunctions. After mating, many signal and
power connections are still provided, but they are fewer in number than those

Samos Satellite checkout is Completed at this time,

4-310. The launch operations building special contro! and monitoring equip-

ment is tabulated below:

CONTROL AND MONITORING
EQUIPMENT

Electrical and Guidance
Console

Facilities and Propulsion
Console

Communications and Payload
Console

Launch Conductor, LMSD,
Console

Operations Inte reconnection
Box

FUNCTION

power equipment in sai
Operations building, and launch pad
and service building. Controls and
monitors guidance and control com-
ponents and functions,

Controls facilities water deluge sys-
tem, missile service tower, and um-
bilical mast. Controle and monitors
Propulsion subsystem fueling and pro-
Pulsion components and fanctions,

Controls and monitors satellite-borne
communications equipment and E- 1
visual payload.

Provides Samos Satellite countdown
status and launch status control, Pro-
vides destruct control and monitoring
devices.

Provides circuit distribution center at
launch operations building,

4-112 ~—SECRET—

detailed checkout and, if

when the launch pad and ser-
vice building is evacuated in Preparation for launch, all test connections other
than umbilical cable connections are removed from the satellite and the Samos
Satellite access doors are installed. From this time until launch, all control

building. The
of all critical voltages
verified, under the



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SECTION IV

4-302. The satellite-borne communications equipment special test equipment
used for the subsystem checkout is tabulated below: :

SPECIAL TEST EQUIPMENT
Rolling Ground Station

Rolling Power Console

Rolling Pressure Console

" UHF Command Receiver, and ©

Telemetry and Data Trans-
mitters Checkout Console

PAM/FM Data Link Checkout
Console. .

PAM Demultiplexer
Equipment

Sets of Antenna Couplers

Command Decoder-Sequence
Programmer and Clock Check-
out Console

—SEGRET—

FUNCTION’

Calibrates and checks VHF FM/FM
unitized telemeter circuits. Meas-
ures input and output signals of sub-
carrier channels and VHF transmitters.

Supplies all power required by VHF
FM/FM unitized telemeter circuits
during checkout. .
Applies variable pressures to pressure-
sensing transducers of telemetry instru-
mentation to provide known inputs to
subcarrier channels.

Provides method of radio-frequency
communication with satellite UHF
equipment for ground checkout,
Checks operation and performance
of UHF command receiver, UHF
Barrow band tranemitter, and UHF
wide band transmitter, Measures

. frequency, power, bandwidth, and

noise figure.

Tests sation of PAM multiplexer
and UHF telemetry FM circuits
ceives and demodulates radio-f; acy
telemetry signals, and records results.

Used with PAM/FM data link checkout
console to demultiplex telemetry sig-
nals received by console.

Used to couple radio-frequency énergy
between radio-frequency checkout con-
soles and Samos satellite. Prevents
scattered and spurious radiation through-
out checkout area, :

Tests operation of command decoder
and sequence programmer, and checks
frequency of satellite timing generator.
Injecte messages into command decoder-
with punched paper tape and monitors
Fesponse of command decoder and
quence programmer logic circuits,
Furnishes d-c and a-c power required
by components.

4-107





SECTION Iv Da:

(Continued)
SPECIAL TEST EQUIPMENT . FUNCTION
Beacon Transponder/ Interrogates transponder, measures
Decoder Checkout Console Parameters of transponder responses,

and checks operation of decoder. Moni-
tors voltages and currents of trans-

tT, decoder, and acquisition trans.
mitter. Furnishes d-c and a-c power
required by Components,

Secondary /Interim Pro- Ensures Fairchild int
grammer Checkout Console is prepared to recei
ks

mer operation, Fu;
Power required by Programmer,

Command Programmer Checks stored Program command fune-
Checkout Console tion of sequence Programmer and inter.
mediate storage unit. .
Intermediate Storage Unit Simulates commands to position relays
Checkout Console within intermediate stora, unit and

monitor reaponse of the re lays. Also
commands relay closures in power

control unit,
Tape Recorder Checkout Provides simulat. id, commutated tele-
Console metry signals to each recorder and

monitors output signals and tform-
ance of each recorder. Furniches d-e
and a-c power required by Fecorders,

4-303, Subsystem checkout of the satellite geophysical environmental
inetrumentation comporents { accomplished at the guided missile
assembly building, The ig devices that constitute the instrumenta-
tion are assigned by the Geophysical Research Directorate, and checkout
is performed in accordance with its specifications. Tes are performed
for the following geophysical environmental instrumentation components:
radiometer, densitometer, cosmic ray monitor, electric field meter, ion -

Re, micrometeorite detector-grid type, and the micrometeorite detector -
acouatic type, .

4-304. SYSTEM CHECKOUT aT GUIDED MISSILE ASSEMBLY BUILDING,
System checkout at the guided missii. embly building is performed with
Special syeteme test equipment that en; ‘St Operations to be com-
pleted semi-~automatically ina Single, integrated Procedure. The major
elemente of the special systems test equipment are the system checkout
complex 2A, the Lockheed automatic checkout equipment (LACE), and the
vehicle function generator console. These major elements, the checkout

consoles which compose them, and the Specific functions they perform
are listed in Paragraph 4-306, .

4-108 —SECrer—



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SECTION vi ——SECRET—
LMSD-437237

SAKOS SATELLITE ORBITAL AND RECONNAISSANCE PHASES
7-67. GENERAL,

7-68. At the time of this writing, the planned sequence of events during the
orbit phase includes a nosedown reorientation of the vehicle from the hori-

zontal attitude, as
state orbiting.

and fuel venting. Then, Feorientation would be established. The pertinent
information that follows is Presented on the basis of the modified events
which are discussed as Pre-nosedown and then post-nosedown orbiting.

7-69, PRE-NOSEDOWN ORBITING PHASE.

7-70. Following engine shutdown, the satellite enters pre-nosedown orbiting.
During this phase the helium, oxygen, and fuel vent equibs are fired, teleme-
try calibration starts and stops, the antenna boom squibs are fired, the Geo-
Physica Research Directorate (GRD) equi; at is armed, 28-volt power is
removed { the engine cutoff relay, Gnd the tape recorder is armed. The
satellite retains a horizontal attitude until receiving a real-time command

for nosedown reorientation during the first-orbital pase over Hawaii.

7-71. POST-NOSEDOWN OR BITING PHASE.

7-72. During this phase of Operation, the satellite maintains a nosedown
attitude and a steady-state condition, The vehicle is maintained in the nose-
down position by the attitude damping system which becomes activated during
orbit. The flight control and Pneumati. e

and yaw gyros of the inertial reference package are switched to the rate mode
for operation in conjunction with the attitude damping system. The damping
system, by controlling the angular momentum of the inertia wheels, mini-
mizes satellite oscillations within six hours to predetermined deflections.
The horizon 8 and periodically provides indications by means
of telemetry, his orbital phase, the nitrogen valve is closed,
2000- and 400. cps inv control is provided, the gyros are caged, th
pneumatic system ented, switch is made to orbit functions, the attitude
damping aystem is activated, switch to orbit antennas is-made, and initial
shutdown of guidance timer is effected.

7-73. RECONNAISSANCE PHASE, (See figure To2.)

7-74, The payloads are activated after nosedown of the satellite over areas
of predetermined inter it. the pre-programmed commands stored in the
satellite-borne progr, Visual and ferret reconnaissance data, geo-
physical environment da and equipment performance parametere are re-.
corded on satellite-bo: torage media, When a satellite pasees within a
range of tracking stations at Vandenberg and Kaena Point, the area pro-
grammer activates the command receiver, The tracking station then com-
manda readout of all Payload and telemetry data and dispatches real-time
commands and new stored commands to the satellite programmer.

7-75, Under nominal conditions, complete functional checking of the pay-
loads and the satellite command and control equipment occure within the :

Ta14 ——seerer-


SECTION vn

UU

Wt ayyyyyypyypy OO Wdddtdenrparrrpatararerevecaren MV). i;My YL

NLL TOD he be “Aeerttr itil WULLELLLU LLL LLL LLM tty WM ttt eeevareecerarteenrtraneennrp WU

Sette teen,

ae WM tpt bt eee sere,

CM «AE el Aenean erneeeee 0

g

AMD Uday
Mitittt lf, “yy

MILLA Lh tpnpy



—SECGRET——_ ‘SECTION IV
LMSD-437237
| "CONST ANT-RATE
ROLL/YAW ATTITUDE MOMENTUM WHEEL

DAMPING PACKAGE ———“

_ PITCH ATTITUDE

DAMPING PACKAGE
Po ~~ “Piven ATTiTUOe DauPmo PACKAGE 7
!T yuetao : = H
H INTEGRATOR H
PITCH GYRO
! meRTuL |
] REACTION !
! HEEL i
I [ tacoueren | |
bee TT.
por o eee -4
H ROLU/VAW ATTITUDE DAMPING ei i
'
1 SERVO H
H in| fre H
YAW GYRO (DURING
ORBIT THE INPUT | el ]
AXIS OF THIS GYRO 1 enacrion 1
1S ALONG THE REF- { t
ERENCE ROLL Axrs) { Tacoma ea
Loi ——— pgm — Sf

om ae

Figure 4-22. Orbitel Attitude Demping System, Functional Schematic .
—— SECRET . 4-53



~~ SECRET

SECTION Iv
LMSD.437237
Mv ac
. WO~, 16
—= =>] ¢ A > TO FLIGHT
a CONTROLS,
Guipance,
ovate
“~~
mventen "LOAD Luster :
MSV ac
anon, 16
ay
TRANSPER RELAY
a >
Rw TV ac uence
SYSTEM
~, 1
POWER AMPLIFIER WO. 1)
—
100,16
00~, 16 REGULATOR
POWER AMPLiFiCR ino, 2
———S=_—=—=—==>. TO PAYLOAD
. Vsv ac Svstems
On, 1h
Figure 4-11. Somos Satellite Interne! Electricel Power System, Functional Schematic (Shoot 2 of 2) _
—SEerer- 4-29

—


C

s L tao
LMSD-4372,
L Operational Checkout Concept 2-39
+ « Following Orbit Pass 2-39
LAUNCH AND PREORBIT PHASES, + + Prior to Acquisition 2-39
See: Samos System Integrated + + Prior to Launch 2-39

Operation

LAUNCH BASE AT POINT ARGUELLO,
See: Operating Bases, Centers,
and Stations : . .
LAUNCH BOOSTER, See: Samos .

‘Booster

LIST OF ILLUSTRATIONS vi

: *Launch Complex No.

Receiving Area 2-39
Transmitting Area 2-37

OINT ARGUELLO LAUNCHBASE 2
Complex Service Building 2.8

+ Description 2.8

General Description of Base 2-3
jor

je 2-5
Purpose and Location 2-3
1 Point
Arguello 2-4 :

Launch Operations Building (LOB) 2.7

a « “Launch Operations Building and

MAINTENANCE, See:
t]

OPERATING BASES, CENTERS, AND
2-1

DOWNRANGE TELEMETRY SHIP
40

Description 2-40

+ Function 2-40

+ General 2-40

> *Telemetry Ships 2-4)

INTRODUCTION AND SECTION .
ARRANGEMENT 2.3 .

Introduction 2-3 .

+ + Facility Functions 2-3 .

* 2 Major Facilities 2.3

- Samoa Facility Location 2-2 °

Section Arrangement 2-3 .

+ Scope of Section 2-3 .

KAENA POINT TRACKING .

STATION 2-35 .

Samos Satellite .

Consoles zo
Major Equipment <7
. li sv and Payload
Console 2-8
+ Electrica! nd Guidance
and Propulsion
Console 2-8
+ Launch Conductor Console 2-3
+ Television System 2-8
Pu: and Construction 2.7
and Service Building 2-5
and Service Building 2-6
Major Equipment 2-5 .
Purpose and Construction 2.5
« Air Conditioning Equipment
+ Gas Preseurization
Equipment 2-7
Launcher Aseembly 2-5
Missile Service Tower 255
“7

2-7

: Propellant Transfer Set
+ Television Cameras 2-7

Umbilical Coupling Set 2-7
*Kaena Point Tracking Station 2-36 +s + Umbilical Mast 2-7
Description 2-35 Maintenance Concept 2-8.
++ Location 2.35 + + General 2.3
+ + Purpose 2.35 . Normal Operation Phases 2-9
Operational Checkout Concept 2.35 + Communications 2-9
+ « Following Orbital Pass 2-27 + Launch 2.9
+ + Prior to Acquisition 2-37 + Post Launch 2.9
+ « Prior to-Launch 2-35 + Samos Vehicle Preparation 2-9
Receiving Area 2.35 Operational Checkout Concept 2.8 >
Transmitting Area 2-35 + General 2-9 ~
KODIAK TRACKING STATION 2-37 + Post-Launch 2.8
*Kodiak Tracking Station 2-38 + Pre-Launch 2.8
Description 2-37 + Required Personnel 2-9
. +. General 2.9

“SEcrer_ X23



E SYSTEM

SATELLIT
DESCRIPTION 1.3

Data Handli 4 by Satellite Test
32.

ata mens ting Analysis, and
Analysie and Evaluation
A 1.23

+ Automatic Controle Area 1

rocesel:
21 ng

rea
General 1.
Receipt of Data 1-2)

STC as Ground-s

c Pace Control
s+ Operation

Pattern 1.19
Centers and

Centers, and
lemetry Ship I-19

8 Station | 1216
ol by STC 1-17

lions. bo15
ownrange TT

ena Point T:
¢ Orbit Phas,
ching Stat:
ello 3616

u Tracking Station 1-19
Satellite Test ¢

adout Program
Onnsissance 1.9
Configuration f 1-9
Flight Confi
pal Objectives
aht Vehicles

+ Samos Vehicle

Vieual Reconnals
Samoe Booster
+ Engines and

Ope rational
feu
Main Features

Samos Satellite 1-9

Agena Vehicle Major

ubeysteme 1-11 . .
Electrical Power System 1-13

dance and Control System 1-13

nae Satellite 1-10

ee Leading Particulars aeu
« + Satellite-Borne Communication

: al
_ 2 Major tells Ba22

Samos System Concept 1-5

8
oe

os System Concepe 1-6

ral 1.3

Launch Vehicte 1-3
oe Satellites 1.5

Surveillance 1-5
+ Polar Orbits 1.3

« %yetem Organization for

uration! 1.4

Co
Satellite System 1-2
SYSTEM ADMINISTRATION 1.23

Administrative Structure 1.23

.
.

iniatration for Command and
Control of System 1.24

General 1.23 .
Command and Controt 1.27
+. Technical Direction 1.23
Working Groups 1.27
unch Base Organization 1.27
+ “For Pre

Sateltite
Organization 1-27

Tracking Station Organization 1-27

+ °Typical Station Organization 1-26

patnch and Launch
¢ Center

BAos SISTEM INTEGRATED

COMMAND AND DATA READOUT
” PHASES :

7-18

Acquisition 7-18

Deecription 7. ls

Data Readout 7-20
+ + Function 7.20
+ Ferret 7.20

Telemetry 7-20

+ Vieual 7.20



.

~~

|
|
|

N
N
N
N
N
N
N
N
\
\
N
\
N
N
N
N
N
N
N
N
N
N
N



SECTION Iv

LMSD-437237

PRIMARY GATTERY, Type;
{1 OF 12 IN PARALLEL)

EXTERNAL
GROUND Power

== - NN

nn

$28.39 pe 28,3v De a 20.3v oc
REGULATORS (NO. 1 AND 7 Power
SUPPLY 7
+78.3v DC i To FucHT

suroane
rT
2m 0c ano PAYLOZD
S¥sTems
a a
Figure 4-11, Somes Seteilite Internal Electrical Power System, Functional Schewetic {Sheet 1 of 2
4-28

—SECRET—


SECTION IV
LMSD-437237

and yaw attitude control function from the gas reaction jets to the hydraulic
control system, The pitch and yaw gyros in the inertial reference package
continue to provide error signals for guidance in Pitch and yaw. The hydrau-
lic control system maintains the pitch and yaw attitude within 0.4 degree of
the reference pitch and yaw axes, and the gas reaction jets maintain the roll
attitude within 1.5 degrees of the reference roll axis, Throughout the time
that the satellite rocket engine is firing, the pitch attitude Program command
signal from the integrator and the pitch and roll attitude reference signals
from the horizon scanner Continue to be applied to the inertial reference
Package,

4-164. When the satellite rocket engine is started the Sequence timer con-
nects the accelerometer in the inertial reference package to the integrator

in the computer to me he velocity gained by the Samos Satellite. When
the velocity gained byt atellite matches the value Preset in the integrator,
the integrator provides a cutoff signal to the satellite rocket engine. The per-
formance. of the Samos Booster is so predictable that it is not necessary to
Provide a way to adjust the velocity setting in the integrator after launch. The
Satellite rocket engine Operates for slightly less than two minutes when the .
integmtor cutoff signal is received. Ten seconds after this nominal time a
back-up cutoff signal is provided by the sequence timer,

4-165. FROM SATELLITE ROCKET ENGINE CUTOFF THROUGH ORBIT
REORIENTATION. From Satellite rocket engine cutoff through orbit re-
orientation, the attitude of the Satellite is controlled about the Pitch, roll,
and yaw axes by the gas reaction jets of the pneumatic control system. Ten
seconds after satellite rocket engine cutoff, the sequence timer changes the
pitch and yaw attitude control function from the hydraulic control system to
the gas reaction jets. At this time the pitch and roll attitude reference sig-
nals from the horizon scanner are removed from the inertial reference
Package.

4-166. Orbit reorientation is performed by removing from the inertial refer.
ence package the pitch attitude Program command signal of 3.82 degrees per
minute which was applied from the integrator and re-applying the pitch atti-
tude program command signal of 40 degrees per minute from the sequence
timer. The Samos Satellite is turned from its horizontal attitude to a vertical,
nose-down attitude within two minutes and twenty seconds. Orbit reorientation
is ended by removing {rom the inertial reference package the pitch attitude
Program command signal of 40 degrees per minute and re-applying the pitch
attitude program command signal of 3.82 degrees per minute from the
integrator.

4-167. For thirty seconds after orbit reorientation, the gas reaction jets
respond to the pitch attitude rogram command signal and pitch the orbiting
Samos Satellite at the rate of 3. 82 degrees per minute so that its nose con-
tinues to point toward the center of the earth. After thirty seconds, the pneu-
matic contro? system is depressurized. The vertical, nose-down attitude of
the satellite is established and maintained as the equilibrium position by the
gravity gradient, the orbital angular momentum of the satellite, and the
orbital attitude damping system,

4-168. During the orbit reorientation, which is completed approximately
thirteen minutes and ten seconds after launch, the UHF omni-directional
antenna boom is swung out and any remaining oxidizer or fuel is released by *
Opening vents in the oxidizer and fuel tanks, Thirty seconde after orbit

4-62 —_SECRET—


ay

o

SECTION vir
LMSD-437237

Figuwe 7-1. Somos Vehicle Flight Prolite
————SECRET-


8)
LMSD-437237

SECTION Iv

PITCH, ROLL, AND
\ ‘Yaw GYROS

PITCH ATTITEN
PROGRAM

an i

ELECTRONICS
UMT

comeing  Pesacc

[= ATTITUDE ERROR
Soma

ORBITAL ATTITUDE

DAMPING SYSTEM

ea
Peescocooscsesosms .

SCOO CONTROL Lease
= ELECTRICAL CONNECTION

womzon ATTITUDE Ter.eueTay
Semon Li. HONALS

Frowre 416. Somes-Setellie-Bamne Cuidence ond Cantvel Systems, Functional Schanetic

—SEcRET-

a



SECTION IV ~~ SECRET—
e LMSD-437237

‘GAS REACTION JETS

WTORAULIC ACTUATOR °
WYORAULIC POVER LNT AND INTEGRATED HYDRAULIC PACKAGE
HORIZON SENSOR .

CONSTANT -RATE MOMENTUM WHEEL .
WHERTIAL REFERENCE PACKAGE (BEHIND (vem 8)
ATTITUDE DAMPING PACKAGE

WORIZON SCANNER

COMPUTER

PLIGHT CONTROL ELECTRONICS LeaT

SEQUENCE TIMER

GAS SUPPLY SPHERE

Seevevsuns

Figure 413. Sames-Setellite-Berne Guidence ond Control Systeme, Location of Compencats
4-36 —— SECRET —



as, SECTION Iv

Conversely, opening of the solenoid valves (refer to Paragraph 4-76) causes

C “hydraulic-balance pressure in the Secondary lines to close the respective flow-
a control valves. Thus, combustion stops in both the

thrust chamber, and the turbine rapidly decelerates

sure. The oxidizer vaive closes last under spring pressure to provide after.

flow of oxidizer during shutdown. The deriving and a

actuating power from the Propellant system in this manner eliminates the

added weight and complexity of an auxiliary hydraulic syetem.

4-76. ENGINE ELECTRICAL-CONTROL SYSTEM. (See figure 4-9.) The
electrical-control system provides the electrical power for satellite rocket

assembly. Connected to the circuit are the dual- squib igniter assembly for
engine starting, three pressure switches (fuel case, oxidise manifold, and
thrust chamber minimum-pressure) for flow-control-vaive ration, and
the two solenoid valves described in paragraph 4-75. in on, output cire.
cuits, for monitoring the sequence and function of the control syetem, and
signal inputs are associated with the circuit. Provisions are made whereby
& ground safety assembly can be plugged into the system for frosne testing
for the prevention of damage to the engine in the event of malfunction.

4-77, The power relay serves as an arming device for satellite rocket engine
starting. While the power relay is de-energised, the squib firing circuits in-
C side the engine relay box receive no exte power, but when the relay is
energized, the firing circuits can be fired by the sequence timer to ignite the
‘solid- propellant turbine starter charge. Gases from this ignition start the
turbine, causing pressure buildup that actuates the three Pressure switches.
(See paragraph 4- 76.) The thrust-chamber pressure switch is the last to be
actuated, It is actuated, only after it senses a sure start, to complete the
circuit that energizes the thrust chamber relay. Prior to energising this re-
lay, power is connected through the fuel-case Pressure switch and
oxidizer-manifold pressure switch to ane rgize and close the gae-generator
solenoid valve and the fuel-control solenoi
turn, causes the bipropellant valve and the fuel valve to be opened (refer to
paragraph 4-75). After the thrust chamber relay is energized, power passes
directly through the relay to the solenoid valves, thereby bypassing the pres-
sure switches, To meet the fast actuation requirements of the fuel valve, its
Sesociated solenoid valve is pilot-operated rather than direct-acting to meet
substantial increased fuel flow with only a 70-percent increase in ele
current. : .

4-78. The energized thrust-chamber relay, in connecting electrical power

to the two solenoid valves, also establishes conditions so that now only the

shutdown relay or the power relay can de-energize the solenoid valves and

thus cause engine shutdown. The engine will then shut down if the power in-

put to the contro! box is interrupted; the voltage drops sufficiently to allow

the power relay to drop out (be de-energized); the shutdown relay ie energised

by an external 28- volt command signal; or the chamber Pressure decreases

because of pump cavitation or any other cause that returne the thrust-chamber
Cc minimum- pressure switch to its original, de-actuated position, In the de-

actuated position, the ewitch complete

circuit that energizes the shutdown

——SEcRET—_ ; 4-25



wy

LMSD-437237

GLOSSARY

ACQUISITION AND TRACKING-The Searching for, locating, and subsequent
Continuous observance of satellite flight position during a limited span of time.

ARM-To Prepare a fuze or detonator for

ignition; to Corsplete an electrical

Circuit toa fuze or detonator so that an electrical impulse can cause ignition.

ATTITUDE-The Position of a vehicle or
Pitch, roll, and yaw,

satellite, expressed in terms of

BOOSTER SECTION-A jettisonable aft Portion of the Samos Booster; contains
the booster engine of the Samos Booster and Associated equipment.

BORESIGHT TOWER-A fixed target used for visual alignment of a radar
tracking antenna used in Conjunction with a sighting telescope to adjust the

orientation of the antenna to correspond
t

with the direction of the radiation

BLACK BOXx-a Combination of Parts that performe @ particular function,

(See Component, ) Refers to an enclosed
that can be readily inserted or removed
components,

component, Ueually electronic,
froma Stroup of interconnected

BOOSTER ENGINE-The dual-chamber rocket engine of the Samos Booster
that assists in accelerating the Samos Vehicle from launch. It is a part of

the jettisonable booster section,

CHECKOUT-The Performance of a test,
verity and validate that the Operational ¢!
and structures are within specified limit

COTAR—Correlation tracking and Fang
measuring system for determining vehic

Or series of tests, designed to"
haracteristics of vehicle equipment
8.

ing, A Phase-comparison Padio
le location,

COUNTDOWN-A Controlled sequence of events ending in the launching of a

vehicle,

DIPLEXER-A mechanical or electronic component which enables two trans-
milters to use a single antenna; Prevents either transmitter from dissipating

a@ large pertion of ite transmitted energy

__ SECRET | G-1


C

——-sEcREr_ SECTION VII
LMSD-437237

countdown approaches the terminal point, advises him that the orbital portion
of the test complex is ready to support operation. The Director informs the
Controller of conditions which could deter launch, The Controller ascer-

i le directive documents. He acts .
rial pickup of data from the telem-
ight operations released by Lockheed
is presented to the Controller for transmittal to AFBMD. Headquarters.

7-7. LAUNCH BASE OPERATION AND CONTROL. Launch base operation
and control, in preparation for a Samos flight, is conducted by Air Force,
Lockheed, and associated contractor personnel who cooperate under control
of the Satellite Test Center, Key personnel inciude a Launch Operations
Coordinator and an Operations Director located in the Launch Operations
Control Center, plus a Launch Controller, two Laurich Conductors (one each
for the satellite and the booster), and a Pad Safety Officer, all located in the
Launch Operations Building. A Tracking Station Commander (AFBMD) and

a Tracking Station Manager (LMSD) ‘are located at the Vandenberg Tracking
Station, Provision is made to locate the Range Safety Officer in this building.

7-8. The Launch Operations Coordinator (AFBMD) is responsible for con-
trolling prelaunch and launch operations. He receives recommendations
from the key representatives in the Launch Operations Building and the
Vandenberg Tracking Station upon which he makes decisions; these are re-
layed, when applicable, to the System Test Controlier at the Satellite Test
Center, The Coordinator and the Launch Director (LMSD} serve bilaterally
in directing prelaunch and launch operations, .

7-9. The Launch Controller hae overall supervision of the launch phases at
the Launch Operations Building. He ie the direct contact between all
members of the AFBMD/contractor launching team, launch base, and Sat-
ellite Test Center, During a launch operation, he receives recommendations
from Launch Conductor/Booster regarding the Samos Booster and from
the Larach Conductor/Agena regarding the Samos Satellite. He weighs these
recor nendations against the status of the local range and makes decisions
accordingly, For coordination of the launch phase with other pha of
Samos operations, he is subject to direction by higher Air Force authority,

7-10, The Launch Conductor for the Samos Satellite is located adjacent to

the Launch Controller, and has both technical and operational supervision of

the Satellite portion of the launch operation. He is Freeponsibie to the Launch
Controller for technical adiness of the Samos Satellite System and accom-
plishment of Satellite Test Objectives. He makes technical recommen-

dations relating to satellite readiness and teat Objectives. He is also Fespon-

sible for preparation and checkout of the satellite during test and countdown
Operations, He conducts the satellite countdown, monitors activities of the

Agena Senior Launch Engineers, makes operational decisions relative to the -
satellite and reports Progress periodically to the Launch Controller.

7-11, The Launch Conductor/Booster ie located adjacent to the Launch
Controller and has both technical a Operational supervision of the booster
Portion of the launch operation. He is responsible to the Launch Cont roller
for technical readiness of the booster system and accomplishment of booster
test objectives, He makes technical recommendations relating to booster
readiness and teat objectives. He is also responsible for preparation and
Checkout of the booster and launch complex during test and countdown

—_SECRET 7-3



\ W —
N N

\

H Ny
| N \

s

WV
N
. Xd
N

N
\

—_

\
we
\

\


4-135. HORIZON SENSOR. The horizon sensor used on the initial Samos
flights is the Model 1 which consists of three optical, infrared devices. It is
used after orbit reorientation to Provide attitude telemetry data but performs
The horizon aensor used on later Samos flights is the

of two optical, infrared devices. It replaces the ho-
rizon ecanner on the later flights and provides the pitch and roll attitude ref-
erence signals that were provided by the horizon scanner. .

4-136. The operation of the horizon sensor is similar to that of the horizon
scanner described in paragraph 4-132. Each optical, infrared device consists
of a scanning germanium prism, an infrared radiation detector, and an elec-
tronic circuit to generate signals proportional to any infrared radiation that

is detected. The three optical, infrared devices of the Model 1 horizon sen-
sor measure the total infrared radiation field around the reference pitch and
roll axes and generate pitch and roll attitude telemetry signals that corre-
spond to the measured radiation. Consequently the pitch and roll attitude
telemetry signals correspond to the attitude of the Samos Satellite in pitch

and roll. One of the two optical, infrared devices of the Model 2 horison
sensor measures the total infrared radiation field on the left Of the reference
roll axis, and the other one measures it on the right. The voltage signal
generated by the combined operation of the two devices is the roll attitude
reference signal. Each device measures the total infrared radiation field
around the reference pitch axis independently, and the two measurements

are averaged to generate the pitch attitude reference signal. s

4-137, CONTROL DURING ASCENT PHASE,

4-138. GENERAL. Control during the ascent phase is provided by the pneu-
matic control system and the hydraulic control system in response to guid-
ance signals from the guidance system. During the coast period following
separation, the gas reaction jets of the Pneumatic control system maintain
the attitude of the Samos Satellite in pitch, roll, and yaw. During the oper-
ation of the satellite rocket engine, the hydraulic control system maintains
the attitude of the satellite in pitch and yaw, and the gas reaction jets control
the attitude in roll. This is the only time that the hydraulic control system
is operated. After the satellite rocket engine is cut off, the gas reaction jets
change the satellite from its horizontal attitude to a vertical, nose-down atti-
tude for the orbit phase. Control after orbit reorientation is described in
paragraph 4-145. The operation of the pneumatic and hydraulic control sys-
tems are described in paragraphs 4-139 and 4-142,

4-139.. PNEUMATIC CONTROL SYSTEM. The pneumatic control system
exerts torques on the vehicle by forcing yas through gas reaction Jets to con-
trol the attitude of the lite in pitch, roll, and yaw. The system cons
of the supply sphere and the pressure regulating package; six gas val
and reaction jets that are mounted on the aft equipment rack; and the
necessary pneumatic connections. The components are represented in
figure 4-18. A diagram is also included in figure 4-18 to illustrate the
operation of the pressure regulating package with two of the roll gas reaction
jets, and to show how the jets are controlled by guidance signals from the
flight control electronics unit, . .

4-140. The gas used in the Pneumatic control system is a gaseous combina-
tion of dry nitrogen (N,) and tetrafluoromethane (CF). Tetrafluoromethane
{s commonly known asTreon. It is stored in the gas supply aphere at a

4-46 —_SEGRET—



LMSD-43

GUIDANCE AND CONTROL
SYSTEMS 4-36
Control During Ascent Phase 4-46
+ General 4-36
+ Pneumatic Control System 4-46
+ + “Functional Schematic 4-47
+ «© Gas Action 4.46
+ + "Principles of Operation 4-48
+ + Relief Valve and Filter
Action 4-49
+ Hydraulic Control System 4-49
+ + “Functional Schematic 4-50
+ « Principles of Operation 4-49
+ + *Principles of Operation 4-5]
Control During Orbit Phase 4-52
+ General 4-52
+ Orbital Attitude Damping
System 4-52
+ + *Functional Schematic 4-53
+ + Pitch Attitude Damping
4-52

+ Computer 4-44

+ « Integrator 4-45

+ «+ Sequence Timer 4-44

+ Control System 4-37

+ Flight Control Electronics
Init 4-44

+ General 4.40

+ + Coact Period 4-40 :

+ + Orbital Boost Period 4-40

+ Horizon Scanner 4-45

+ Inertial Reference Package 4-42

+ + Accelerometer 4.42

+ « Control Signals 4.42

+ + *Gyro Features and Orientation
in Satellite 4.43

+ + Gyros 4.42

Guidance During Orbit Phase 4-45

+ General 4.45

+ Horizon Sensor 4.46

+ + Operation 4-46

Purpose 4-35

+ General 4.35

+ « Location of Components 4-36

+ + Scope of Text Coverage 4-35

+ Guidance System 4-37
+ + Reference Pitch, Roll, and
Yaw Axes (4-37
+ + Definition of Pitch, Roll and
‘aw ments 4-39
+ « Definition of Reference Pitch,
Roll, and Yaw Axes 4-38

+ Launch Pad and C Phase
Concept 4.35
+ Operation o, Guidance and Control

+ . *Functions Schematic 4-41
rt Equipment 4-35
« Support Personnel 4-35
Sequence of Guidance and Control
Events 4-54

Events 4-87 —
+ Before Launch 4-56
+ Ejection of Ferret Payload 4-63
+ Engine Cutoff Throu th
1p, neotlentation 4-62 5
ingine Start to Cutoff 4- 1
General 4.54
eageah to Separation 4-56 6
. uence Timer Operation 4-5
Orbit Phase 4-63 . .
+ “Sequence of Events 4-64
Separation to Engine Start 4-61
INTERNAL ELECTRICAL POWER
SYSTEM 4-26
‘amiliarization 4-26
« Design Concepts 4-26
« General 4-2
+ « *Lecation of Components 4-27
Functional Description 4-26
+ *Functional Schematic 4-28
+ + Batteries 4-30
+ « General 4-26

eee wees

+ Ground Connections 4-30
basi Power Junction Box ort, 30
jor Equipment Description 4-

+ General 4.39

+ Negative 28-Voit Power Supply 4-3) -

+ Plus 28-Voit Regulators 4-3}

+ Primary Batteries 4-31

+ 400-Cycle Sin, Phase A-C
Regulator 2.

+ 400-Cycle Single-Phase
Synchronized Power Supply 4-32

+ 400-Cycle 3-Phase Inverter 4-31

+ 2000-Cycle Static Inverters 4-31



—~-SEGRET___
SECTION IV LMSD-437237

assembly for maximum heat transfer. All components are rigidly mounted, .
and the entire assembly is potted to insure vibration-proof Operation. C .

4-97. 400-CYCLE, SINGLE-PHASE, SYNCHRONIZED, POWER AMPLIFIER,
(See figure 4-10.) The Power amplifier is a transistorized, tatic-electronic
unit which receives a Precision 400-cycle frequency signal from one phase of
the three-phase inverter in order to synchronize and maintain its frequency to
that of the three-phase Power distribution system, One watt of Power is drawn
from the three-phas. supply. The main power output is supplied by the un-
regulated d-c voltage through a Push-pull, transistor switching-circult, The
output voltage is maintained at 115 (41.15) volts root mean equare, and the
output current is traneformer-sensed So that if the current exceeds the rating
of the power amplifier, a current-limiting action occurs to Prevent unit dam-
age. All components are mechanically secured, and the entire unit is potted
to provide maximum protection against vibration. All heat-producing com-
Ponente are in good thermal contact with the baseplate Sesemb', for efficient

4-98. POWER AND SERVICE REQUIREMENTS,

4-99. ELECTRICAL POWER REQUIREMENTS, During countdowa and until

just prior to launch, satellite electrical power requirements are met by ex-

ternal auxiliary power supplied by the pad power supply and signa! Processing

set. From just prior to launch, when the external-internal power transfer

switch ie actuated, until the end of Samos Satellite reconnaissance lifetime,

the power requirements of the Satellite subsystems and Payload subsystems :
are met by the satellite internal electrical power system. Figure 4-12 liste ran)
the estimated electrical loads and energy requirements during the time the

internal electrical Power system supplies power. Three general groupings

are shown: the satellite subsystems, the payload subsystems, & summa.

tion of values indicated in the first two groups, .

100. SYSTEM SERVICE REQUIREMENTS, Upon verification of internal
system operability, no service requirement Beeded inasmuch as the
satellites are used for single mission Purpos

4-101. AIR CONDITIONING OF SAMOS SATELLITE AND GROUND SUPPORT
EQUIPMENT, Air conditioning provisions are utilised when testing and check-
ing out the Samos Satellite electronic equipment at the guided missile. aseembly
building and when the Satellite is located at the launch pad Somplex. During

ed to the
Satellite connector located on the right-hand access panel of the forward
equipment rack, At the launch site, when the ground air-conditioning unit is
Operating and prior to the lectronic equipment being turned on, the conditioned
alr delivered to the satellite mainly performe a drying function. Whea the
equipment is turned on, in addition to drying, cooling of the equipment begins
and continues until separation of the coupling at launch.

in the launch pad and service building, It provides conditioned air for the in-
ternal electrical power system of the satellite and for the checkout and service
equipment located within the launch pad complex.

6-32 SECRET


LMSD-437237

Command Transmitter Building 2-31

+ *Command Transmitter and Angle
Tracker. Buildings 2-32

Equipment . 2-3)

+ Antenna Positioning System 2-33

+ Command Encoder 2-33

+ Command Monitor Receiver 2-33

+ Command Multiplexer 2-33

+ Command Transmitter .
Antenna 2-32 co

+ » UHF Command Transmitter 2-33
- Purpose and Construction 2-31
Data Acquisition and Processing
Building 2-27

“Data Acquisition and Processing
Building 2-28

+ Equipment 2-27

+ + Data Sequencer 2-27

+ + Master Control Console 2-29

.

+ Slave Data Control System 2-27
Purpose and Construction 2-27
General D.

cription of Station 2-25
s and Buildings 2-25
r Area 2-27
+ Transmitter Area 2-27
Purpose and Location 2-25
Maintenance Concept 2-34
Norma} Operation Phases 2-34
+ + Control 2-35
+.« Ground-Link

Communications 2-35
+ « Ground Satellite
: Communication 2-35
+ « Readout 2-34
Operational Checkout Conce, 2-33
« .« Following Orbital Pass 2-34
+c» Prior to Launch 2-33
+ « Prior to Satellite

Acquisition 2-33
Required Personnel 2-34

+ + Operator Maintenance Crews 2-34.

+. Organizational Groups 2-34
+ + Payload Crews 2-34
UHF Telemetry Receiver
Building 2-29
+ Equipment. 2-29
+ « Antenna Positioning .
System 2-29
+ + Telemetry and Data
Antenna 2-29
+ « Telemetry Receiver 2-29
+ + Wideband Data Receiver 2-29
- Purpose and Construction 2-29
+ “UHF and VHF Telemetry and
Receiver Building 2-20

0 To P

VHF Telemetry Receiving
Building 2-29

+ Equipment 2-29

. Purpose and Construction 2-29
+ + TLM-18 Telemetry Antenna 2-31

+ Tri-Helix Antenna System 2-31
and UHF Telemetry and
Receiver Building 2-30

OPERATION, See: Samos System
Integrated Operation

OPERATIONAL CHECKOUT, See:
Samos Satellite, and Samos System
. Integrated Operation .

ORBIT AND RECONNAISSANCE
. PHASES; See: Samos System
Integrated Operation, and Samos
Satellite

ORGANIZATION FOR SAMOS SYSTEM,
See: Samos System General :
cription .

INTRODUCTION AND SECTION
ARRANGEMENT 6.2

o + Scope of Section 6-2

PROCESSING SYSTEM 6-2

General 6-2

« Major Operating Groups 6-2

. Pu: je and Function 6-2

+ System Relationships 6-2

Management Group 6-3

« Description 6-3

Operations Group 6-2

+ Description 6-2

System Inputs 6-3

+ Input Components 6-3

+ + Description 6-3

System Outputs 6-3

+ Output Components 6-3

+ + Description 6-3

Support Group 6-2

+ Major Sections 6-2

+ + Description §-2

POINT ARGUELLO LAUNCH BASE,
See: Operating Bases, Center, and
Stations

__SBCRET X-5



LMSD-437237

SECTION Iv

ENGINE MOUNTING RING

HYDRAULIC ACTUATOR

ENGINE THRUST CHAMBER
ASSEMBLY

ENGINE MOUNT-AND-GIMBAL bad vew
RING

us MAXIMUM
DEFLECTION PATTERN
OP GMB ALLED ENCINE
(SINGICATED ey
GROKEN Line .

Figure 421. Hydroulic Control System, Principles of Operetion

—seeretT—

4-51



SECTION V

LMSD-437237
c
¢
SECTION V
INTRODUCTION and SECTION ARRANGEMENT 5.2
GROUND-SPACE COMMUNICATIONS SYSTEM 5.2
GROUND-SPACE COMMUNICATIONS CONCEPT $2
ANGLE TRACKING SUBSYSTEM $2
RANGE MEASUREMENT SUBSYSTEM. ti a)
TIMING SUBSYSTEM 33
COMMAND SUBSYSTEM $37
DIGITAL DATA HANDLING SUBSYSTEM ; $3
GROUND COMMUNICATION SUBSYSTEM 54
C DATA RECEIVING SUBSYSTEM $4
ALIGNMENT AND CALIBRATION SUBSYSTEM 5

5
————_ SECRET — 5-1


C) SEC.
LMSD-43723

Samos Technical Support ‘Building 2-8 Operational Ch, 2-22

s+ Description 2-8 oa
POINT. MuGU TRACKING STATION
2-39

. Telemetry Data Processing 2-21

Description 2-39 Payload Data Processing Element 2.22
Receiver Area 2-40 + Equipment 2-22
Transmitter Area 2-40 + General 2.22
SATELLITE TEST CENTER 2-17 VANDENBERG AIR FORCE BASE 2-12
Analysis and Evaluation Element 2.22 General Description 2-12
+ Equipment 2-22 + Major Areas 2-12,
+ + Comparison Viewer 2.22 . se and Location 2-12
. Primary Record Viewer 2.22 + Samos Facilities at VAFS 2-14
+s Quality Evaluation Viewer 2.22 Guided Missile Assemb)
+ Genera? 2-22 Building (GMA) 2. 2
Automatic Controls Element 2-21 . ‘pment 2-12
+ Equipment 2-21 * Purpose and Conetruction 2-12
o pigital Computer 2-21 Launch Operations Controt
. ICE 2.2)
+, General 2.2) Maintenance Concept 2-16
Ommand and Controi Element 2-17 s+ General 2.16.
+ Equipment 2-17 Normal Operation Phases 2-16
+ + Assistant Test Director o + Recei: atGMAB 2.16
Onsoles 2.20 + Subsystem ckout 2.17

+s Audio System, 2-20 + System Tests 2-17.
+ + Closed Circuit Ty System 2.2) oe ckout and Shi ent to Pad 2.17
+ + Data Boarg 2-20 Operational Checkout Oncept 2-16
+ + Dispatcher Console 2-20 5° General 2.16

+ Projection System 2.20 Propellant Storage Area 2-16

+ Test Controller Console 2-29 © ¢ General 2-16

+ Test Director Console 2.29 tired Personne} 2-16

Description Spe 2-17

Flight Test Operation

Fiod 2.23

° Preflight Checkout Period 2-23

Center G@0Cc) 2.23

Purpose 2.13

+ Timing System 2-20 2 + General 2-16
General 2-17 Samos ding cal, Support
juild: °
+ General 2.17 oa General 2-16
+ Major Eleme, te 2017 Vandenberg C, nications

+ Satellite Tes Center ang Typical
Console 2-18 .

omm
Center (VCC) 2-15
Equipment 2.13

Maintenance Concept 2-23 + Purpose 2.13
+ + Field Maintenance 2-23 VANDENBERG TRACKING
+ + Maintenance Organization STATION 2.25
Areas 2.24 “Vandenberg Tracking Station 2.26
+ Personnel Organization 2-23 Angle Tracker Buildin, 2-31
Preventive Maintenance 2-23 + Angle Tracker and 03

Outine Maintenance 2-23
Maintenance Organization
3

.

Norma! Operation Phases 2.24 :
oe Coordination Areas 2.24 .
+ Modes of Operation 2-24 .
24 .

: Orbit 2-25
+ Post-Launch 2-25

X-4 : : __SEGREr—

Equipment

T: it B et 32
Transmitter uildings 2.
Purpose 24 Construction 2-31

Angle Tracker Antenna 2.3)
Angle Tracker Console 2-31
ngle Tracker Receiver 2-32
‘enna Positio, .
Equipment 2.3]

ie

4

ae Be



DER, SECTION Iv

(Continued)
C CONTROL AND MONITORING FUNCTION
EQUIPMENT
Operations Power Supply Supplies d-c and a-c power to equip-
Set pipiy in launch operations building,
Signal Data Recording Set Records data modulated signals of
control and monitoring functions.
Countdown Clock System Provides countdown time in real
time, :

4-311, The launch pad and ervice building signal interconnection, power
distribution, and special test equipment are tabulated below: .

EQUIPMENT FUNCTION’

Pad Interconnection Box Provides circuit distribution center
in launch pad and service building.
les making connections be-
tween Samos Satellite and launch
operations building for any satellite

configuration,
. Pad Power Supply and Signal Supplies d-¢ and a-c Power to equip-
( Proceasing Set ment in launch pad and service build-
. ing, and to Samos Satellite.
Battery Power Supply Provides backup for d-c power
supplied by pad power supply and
' signal processing set.
Destruct Squib and Simu- Checks squib elements of all Pyro-
lator Test Set technic devices. Measures resis-

tance of squib elements on "GO, NO
GO" basis. Monitors destruct primer
circuit,

R-f Data Link Establishes r-f communications links
between Samos Satellite at launch pad
complex and guided missile assembly
building and Vandenburg tracking
station, :

4-312, OPERATIONAL MAINTENANCE CONCEPT,

4-313. GENERAL, The Operational maintenance concept for the Samos sys-
tem specifies that no unreliable or 5
included in any Samos Satellite. All ess
tions shall be performed within the equa

——SEGRET— : 4-113



LMSD-437237
ESTIMATED VALUES FOR SATELLITE SUBSYSTEMs;

SECTION IV

Ascent Phase, Satellite Electrical Load

SUMMATION OF Pow

Component Duty Cycle Maximum Demand Energy
(minutes) (kilowatts) (kw-hr)
Ascent Guidance 17.0 ~ 0.020 0.006
Engine Ignition 1.7 0.526 0.015
Flight Control 20.0 0.013 0.004
Hydraulic System 2.0 1.960 0.065
Horizon Scanner 10.0 0.012 0.002
Subtotal 2.531 0. 092
Orbit Phase, Satellite Electrical Load
Component Duty Cycle [Maximum Demand Energy
(minutes per day) (kilowatts) (kw-hr/day)
Attitude Control Continuous 0.440 1.058
System . .
Data-Link 100 0.100 0. 167
Transmitter
Command Trans- 260 0.100 0.433
_ mitter
Programmer and Continuous 0.009 0.216
Timer
Decoder 262 0.025 0.066
Telemeter 130 0. 230. 0.500.
Tape Recorder 300 0. 078 0. 390
Subtotal 0.972 2.830
ESTIMATED VALUES FOR PAYLOADS SUBSYSTEMS
Payload Subsystem Maximum Power Demand Energy
(kilowatt) (kwehr/day)
Ascent Orbit
E-1 Payload Readout ) 0.076 0.320
F-1 Payload Readout ° - 603 32

ER REQUIRED FOR SATELLITE

AND PAYLOADS SUBSYSTEMS .

Satellite Subsystems Accent Ener, Maximum Orbit Power Total Energy
Plus Payloade (kwehr) (kilowatt) (kw-hr/day)
E-1 0.092 1.048 3.150
Fel 0.092 1.573 3. 662
E-l and F.} 0.092 1. 649 3.982

Figure 4-12,

——

Somes Serellite Interne! Electricel Power Requirements

—SECRET_

4-33



— BS: | SECTION Iv

. (Continued)
C GROUND HANDLING EQUIPMENT FUNCTION

Payload Hoisting Slings Lift payloads from shipping con-
tainers. Mount payloads in satel-
lite in either horizontal or vertical
Position.

Payload Dollies Used to carry payloads within build-
ings, or to support them for storage.

Nose Cone Handling Dolly Used to carry nose cone portion of

satellite spaceframe for movement
within buildings,

Vertical Work Stand Used in launch area for handling,
assembling, and disaszembling
satellite. Supports llite in
vertical position. Enables
measurements of center of gravity.

Vehicle Tilt and Roll Stand Supports satellite during system
checkout of guidance and control
subsystem. Enables arbitrary
motions of the satellite to be made,

C

_—SEcRET 4-119


SECTION V -

LMSD-437237

the Subsystem H ‘equipment in the vehicle and are picked up by the Telemetry
and Data Receiving Antenna, a : .

5-30, Filters in the receiver input circuits route each eignal to the appro-
priate receiving channel, The wide-band channel output (visual reconnais-
sance data) is sent directly to the visual reconnaissance user equipment.

5-31, Since the narrow-band data signal is a composite of several signals,
the narrow-band data channel includes equipment for Separating the composite
signal into its various components and routing them to the proper destinations.
If the vehicle being tracked is equipped for electromagnetic reconnaissance
only, the reconnaissance data is included in the narrow-band data signal, and
the vehicle equipment does not transmit a wideband data signal,

5-32, Although the Telemetry and Data Receiving Antenna can be positioned
manually, it is slaved during a pass to the Angle Tracking Antenna, if the
latter is tracking. Otherwise, it ie Positic in accordance with computer-
Predicted data by the antenna slave data network, . .

5-33. ALIGNMENT AND CALIBRATION SUBSYSTEM,

5-34, The Alignment and Calibration Subsystem provides a means of check-
ing and adjusting the accuracy of Subsystem H antennae and range measure- |
ment equipment. It includes the following: photogrammetric triangulation
equipment, boresighting equipment, and Up-Converter, .

5-35. The photogrammetric triangulation system provides a means of check-
ing the dynamic tracking accuracy of the Angle Tracking Subsystem and the .
accuracy of widely separated cameras, The cameras take simultaneous night
Photographs of a flashing lamp mounted on an airplane, and the airplane is
tracked by the tracking station being calibrated. Since the Photographs have
a background of stars, and the star positions are Precisely known, the Photo-
graphs can be used as the basis for an equally precise computation of target
Position. The target position thus calculated is compared with the target
Position as determined by the tracking station equipment. In this manner,
tracking etation equipment operating errors can be ascertained,

5-37. The Up-Converter Provides a. means of operating the Range Measure-
ment Subsystem in a closed-loop configuration in order te-compen: flor
phase shifts introduced by components in the ground path of the ranging tones.
The Up-Converter is a combination receiver and low-power transmitter, The
receiver, which ie tuned to the frequency of the Command Tra: emitter, re-

Seives its input from a small antenna on the base of the Data Receiving
Antenna,

. __- SECRET , 5-5



~

SECTION VI.
LMSD-437237

Telemetry Checks, from T-255 to T-225 minutes, Tasks 4 and 5; Samos
Satellite Guidance and Flight Control Checkout, and Samos Satellite R-F
Checkout, both from T=225 to T=215 minutes. Task 6; Range Safety Com-"
mand Tests, from T-215 to T-185 minutes. Task 7: Samos Satellite Arming,
from T-185 to 155 minutes. Task 8: Samos Vehicle Flight Safety System
Connect and Arm, from T~185 to T-165 minutes. Task 9: Countdown Evalu-
ation, from T-155 to T-125 minutes. Task 10: Plug Disconnect and Tower
Removal, from T-125 to T-90 minutes, Tack 11: Samos Satellite T: .
from T-90 to T-60 minutes. Task 12: Samos Satellite Pressurization, from
T-60 to T-40 minutes. Task 13: Countdown Evaluation, from T-40 to T-25
minutes. Task 14: Terminal Count (which includes Samos Booster fueling
and pressurization), from T-25 to T-0 minutes. .

SAMOS SATELLITE SYSTEM LAUNCH AND PREORBIT PHASES
7-56, GENERAL. ,

7-57. Ascent operations begin at liftoff, which is the first, two-inch, ver-
tical motion of the Samos Vehicle. This motion actuates the launcher release
mechanism, releases the umbilicals, starts retraction of the umbilical tower,
and supplies a tone or signal to all support faclities to indicate sero time.

At this instant all control is divorced from the launch complex. Then, only
the ground station equipment of the booster guidance system and the range
safety have control over the vehicle. The principal activities that relate to
launch and preorbit phases of operation is presented under two areas of

coverage: those activit that relate to the ground facilities and those that
relate to the Samos Ve , .

7-58. BASES, CENTERS, AND STATION ACTIVITIES,

7-59. Exit tracking of the Samos Vehicle commences at Vandenberg Track-
ing Station. Three basic tracking systems are involved: VERLORT track-
ing, VHF tracking using tri-helix and TLM-18 antennas, and optical track-
ing. Starting with liftoff, the optical tracker follows the vehicle, and through
the slave data bus, positions the VERLORT and TLM-18 antennas until the
vehicle is high enough to permit lock-on by radar autotrack equipment. The
three systems track during ascent to the limit of their respective ran;

The VERLORT radar tracks through separation and the initial coast phase.
The TLM-18 antenna tracks to the limit of line-of-sight. Both position

data and telemetry readout are monitored by this antenna, The tri-helix

7-60. The Pacific Missile Range (PMR) facilities at Point Arguello and
Point Mugu, plus the telemetry ship positioned some 1500 miles downrange
by PMR, are also used for tracking during launch and ascent. The ending
vehicle is tracked by skin-tracking radars and metric Optical equipment at
Point Mugu and Point Arguello, with the optic equipment tracking to the limit
of visibility. The ship collects and records FM/FM telemetry and provides
Doppler and angular tracking of the Doppler beacon in the Samos Satellite.

7-61. The Range Safety Officer monitors tracking and, if necessary, com-
mands destruct of the vehicle during the ascent phase. He uses facilities of
both the Vandenberg Tracking Station and the Pacific Missile Range.

—— SECRET Tell



Nea \ at
TL 2g eee

\N

xan - - . § —


\ \\
\ |

*



SECTION IV ——-SEeRneT—
LMSD-437237
or NOTION
/ WAL AS
/
REFERENCE
Yaw AXIS \_
aN
*
bd
/
/
f
THE REFERENCE PITCH AXIS
ANO THE VEHICLE Y-Y Axis
ARE ALONG A LINE PERPEN.
DICULAR TO THE PLANE OF
THE PAPER.
REFERENCE VmOCLE MOTION

YAW AXIS:
N

THE REFERENCE PITCH AXIS
AND THE VEHICLE Y.Y axis
ARE ALONG A LINE PERPEN.
DICULAR TO THE PLANE OF

Onett Prase

~N

CENTER OF
we

NN

warn
4

Figure 4-14. Definition of Reference Pitch, Rell, end Yow Anes

—sEenet—



SECTION V -
LMSD-437237

INTRODUCTION AND SECTION ARRANGEMENT

5-1, Section V presents a brief description of the Ground-Space Communi-

cations and Control Network, Subsystem H, of the Samos Flight Test Pro-
gram. The information in this section is general and presents Subsystem H
in terms of its eight subsystems: Angle Tracking, Range Measurement,
Timing, Command, Digital Data Handling, Ground Communication, Data
Receiving, Alignment and Calibration. A more comprehensive discussion
of Subsystem H can be found in document LMSD-446936.

GROUND-SPACE COMMUNICATIONS AND CONTROL. SYSTEM
5-2, GROUND-SPACE COMMUNICATIONS CONCEPT,
5-3. The overall mission of Subsystem H during the test phase of the Samos
Program is two-fold: to control the operation of reconnai equipment
s

carried by Samos Satellites, and to obtain and record data gathe~ed by thi:
equipment.

5-4. Satellite-borne Subsystem H equipment receives from the ground the
commiands that control operations of satellite equipment, directs these com-
mands to the proper equipment at the Proper time, and transmits reconnaie-
sance and operational /environmental telemetry data to the ground,

5-5. Ground-based Subsystem H equipment acquires and tracke the satellite,
up-dates predicted orbital parameters on the basis of tracking data, and
generate id transmits up-dated data, in the form of commands, to the
The ground equipment also receives, records, and distributes

ce and telemetered operational/environmental data, and pro-
vides interstation and intrastation communication facilities. :

5-6. ANGLE TRACKING SUBSYSTEM.

5-7, The Angle Tracking Subsystem acquires and automatically tracks the
satellite, and generates precision antenna position data. The position data

are used by a computer, a part of the Digital Data Handling Subsystem, to

correct the predicted satellite orbit. Tha data are also used by the Antenna
Slave Data Equipment to slave the Data Receiving and the Command Trai
mitter Antennas to the Angle Tracking Antenna. The Angle Tracking Sub.
system includes an Angle Tracking Antenna, an Angle Tracking Receiver,
Angle Tracking Antenna Servos, and an Angle Tracking Console. on

5-8. In the automatic tracking mode of operation, the Angle Trackin;
Antenna servos position the antenna by means of satellite -t ransmit error
signals, which the Angle Tracking Receiver derives from the narrow-band
data signal, The antenna can also be slaved to the tracking station computer
to follow the predicted path of the satellite, or the antenna can be positioned
manually, or be made to search in a preset pattern about the predicted
point-of-arrival of the satellite. Angle Tracking Antenna position data are
supplied to the Digital Data Handling Subsystem by means of a mechanical
connection. :

52 —— SECRET



engine, the pendulum is swung back from its static position to cause an output
from the a-c signal generator. The restraint motor ensures that the move-
ment of the pendulum is proportional to the acceleration, since with no re-
straint the pendulum would swing the maximum amount with only a slight
acceleration, The output from the a-c signal Senerator is amplified by the
transistor amplifiers in the inertial reference package and sent to the inte-
grator in the computer. When the integrator determines that the satellite
has reached the correct velocity, it cuts off the satellite rocket engine.

4-128. FLIGHT CONTROL ELECTRONICS UNIT. The flight contro! elec-
tronics unit processes the attitude error signals from the inertial reference
package and distributes them to the control system, Processing consists of
amplifying the a-c attitude error signals from the pitch, roll, and’yaw gyros
in the inertial reference package, converting the amplified a-c signals to d-¢
and amplifying the d-c rate signals. The amplified d-c rate
pplied to the control mechanisms of the pneumatic and hydraulic
control systems to adjust the attitude of the vehicle in accordance with the
guidance attitude error signals from the inertial reference package. Separate
channels are provided for the pitch, roll, and yaw mechanisms the pneu-
matic control system and the itch and yaw mechanisms of the hydraulic con-
trol system. The roll channel for the pneumatic contro! system and the pitch
channel for the hydraulic control system are shown in the diagrams of

figures 4-18 and 4-20. Feedback signals from position sensors in the hydrau-
lic control system are. returned to the d-c rate amplifiers in the flight control
electronics unit to improve the response of the hydraulic actuators to the sig-
nals from the amplifiers. The a-c amplifiers, the demodulators and rate
networks that convert the a-c attitude error signals to d-c rate signals, and
the d-c amplifiers that constitute the flight control electronics unit are com-
pletely transistorized,

4-129. COMPUTER. The computer includes the sequence timer and the :
integrator. The sequence timer provides switching signals during the ascent
and orbit phases to perform such functions as firing the ullage rockets and
connecting the accelerometer in the inertial reference package to the inte-
grator. It also provides a pitch attitude Program command signal to the iner-
tial reference package. The integrator provides a pitch attitude program
command signal to the inertial reference package and detects the velocity
gained by the Samos Satellite during the operation of the satellite rocket
engine.

4-130. The sequence timer is a set of mechanical counters and ewitches that
is driven by a synchronous motor, There are 24 mechanical counters. They
are driven by the synchronous motor and each can be set to actuate from one
to five switches at any time throughout an Operational range of from 0.1 sec-
ond to 1600 seconds after the start of th quence timer. The initial Samos
flights require a maximum time s: than 700 seconds. There are
twelve banks of switches with s each bank. Two of the 24 me-
chanical counters are associ: bank. One counter is set to actuate
from one to five of thi itches in the bank at a specific time, and the other
counter is set to actuate the remaining switches, usually at a different time.
One of the functions of the ence timer is to connect a pre-determined
fixed voltage to the inertial reference package at the Proper time. This volt.
age is the pitch attitude program command signal from the sequence timer.

4-44 ——SEGRET—



lll —_ downy

CLE EKJ@. —

yy

"ddd CECH tw

a ccM::E;C-;/EZEEEL

WLSILLLELLSLLESL TD

ron

et



on

, ullage rockets plug, and installing and connecting the d.

——SEGRET— SECTION IV
LMSD-437237

conditioned air performs a drying function when the electronic components
are not operating, and when they are turned on, the air. performs a cooling
function also. The temperature within the satellite is maintained between
30°F and 65°F, The temperature of the payload
Critical, and it is maintained between 450F and 8: of a tem-
perature control blanket. The temperature control blanket, which consists
of a network of passageways for conditioned air, is fitted over the nose cone
of the assembled satellite, Conditioned air from the type 60M air condition.
ing system is forced through the passageways of the blanket to control the
temperature within the payload compartment,

4-327. Connecting the satellite Pyrotechnic devices is accomplished after the

squib circuits have b checked for electrical continuity and when the ab-

‘nt and spurious voltages is ensured. Pyrotechnic servicing

consists of connecting the Samos Satellite rocket engine arming plug, the
truct unit. The

destruct unit arming plug and safety wire are connected and the fairing is

installed over the destruct unit to complete the pyrotechnic servicing
Procedure,

GROUND SUPPORT EQUIPMENT

4-328. GENERAL. Ground support equipment for the operational Samos
system is stationed at the guided missile assembly building and the launch
pad complex and includes all ground equipment involved in checking out,
servicing or handling the Samos Satellite. The ground checkout equipment
ed under Operational Checkout and Maintenance Concept, para-
-291 through 4-311, The ground servicing equipment is discussed
under Power and Service Requirements, paragraphs 4-321 through 4-327.

The ground handling equipment is discussed below, Paragraphs 4-329 and
330. .

4-329. GROUND HANDLING EQUIPMENT, The ground handling equipment
is used to transport, hoist, and position the Samos Satellite. It is used to
facilitate the checkout and Preparation operations that are performed on the
Samos Satellite at the guided missile assembly building and the launch pad.
complex. Certain items of ground handling equipment are necessary in order
to check out the satellite thoroughly, All items of ground handling equipment
and the specific functions they perform are listed in paragraph 4-330,

4-330. The ground handling equipment used at the guided missile assembly
building and the launch pad complex for the Samos Satellite is tabulated below:

GROUND HANDLING EQUIPMENT FUNCTION

Vehicle Transporter Used to carry satellite over Primary
and secondary roads.

Vehicle Protective Cover Waterproofed cover to enclose satel-
lite and handling yoke during trane-
Portation and storage. .

Vehicle Handling Dolly Positions satellite for horizontal
hoisting. Used to carry satellite
within Duildings for checkout and
mating with adapter,

——SEGRET— 4-117

es te


SECTION IV

LMSD-437237
otstuRBING
VeHICLe Torau
MOTION
RESTORING
Prretans TORQUE
Pitch PITCH INERTIAL .
Gyro /: ~ REACTION WHEEL
input ACCELERATION OF
AXIS PITCH INERTIAL
_ REACTION WHEEL
SAMPING IN PITCH
REFERENCE
SAW AXIS DISTURBING
TORQUE
venice
MOTION
tweut
; ARIS
ASCELERATION or
LL/YAW INERTIAL
raw CONSTANT-RATE
Gyro MOMENTUM WHEEL REACTION WHEEL,

ROLL/YAW INERTIAL

REACTION WHEEL aestorine
TORQUE
CAMPING IN YAW
RESTORING TORQUE
ORBITAL PITCHING RESULTING FROM OnpiTat. RESTORING
OF VEHICLE GYROSCOPIC ACTION §—_PITCHING. TORQUE
VEHICLE MOTION
(REFERENCE ROLL
axis)
. ACCELERATION ;
a
O1STURBING ACCELERATION OF OISTURBING
Re ROLL/YAW INERTIAL ORBITAL TORQUE L.
Tonaue REACTION WHEEL PITCHING
ACCELERATION

CAMPING In ROLL

Figure 423. Orbital Attitude Damping System, Principles of Operation

4-55


—EMsD 437837
GLOSSARY (Continved)

DISPLACEMENT GYRO-A gyro that Provides an electrical signal proportional
‘to the angular displacement of the gyro from a reference Point.

mitter and a receiver to use @ Common antenna; prevents a large portion of
the transmitted energy from being dissipated in the receiver, ard prevents
the received energy from being dissipated in the tranemitter,

EPHEMERIDES-Tables that show the geographical location of the Samos
Satellite as a function of time. The location of the ite can be deter-
mined for any past, Present, or future time,

GANTRY-a crane-type structure, with Platforms on different levels, used
to erect, assemble, and service large rockets or missile It may be placed
directly over the launching site and rolled away before firing.

GIMBAL-A mounting that uses universal Joint suspension to Permit inclination
of the mounted item to a desired position.

GIMBALLING-The action of moving, or inclining, an item mounted in a
gimbal,

of the remaining two orthogonal axes; a Syroscope. A significant character-
istic ie that the spin axis remains fixed in space when the Fotation is main-

HYPERGOLIC PROPELLANT-A rocket fuel that ignites Spontaneously upon
Contact with an oxidizer,

IGNITER-A Pyrotechnic device that ie electrically ignited to start the burn-
ing of an explosive charge ora quantity of propellant.

INVERTER-An electrical device for Converting direct current to alternating
current,

INF RARED-Electromagnetic radiation ina frequency Fange slightly below
the frequencies of visible Nght. gThe frequency range is approximately one
thopgand million megacycles (10 mc) to ten thousand million megacycies
(QO mc}. The wavelength range, from shorter to longer wavelengths, is
@pproximately 3 to 30 microns, .

IRFNA-~tnhibited red fuming nitric ‘acid; an oxidizing propellant.
LAUNCH COMPLEX-The area immediately surrounding the vehicle launch

Pad: includes the buildings and Support equipment directly associated with
countdown and launch Operations.

Gee a ae


LMSD-437237

REFERENCE

ROLL axis
q . ores
ToS oe _
~~ . Yaw
NY, A EB
onarr pase mou J ! [
” Ny
Zz SA ae
mn : ‘s
: 2

Figure 4-15, Doliaition of Pitch, Rell,
——SEGREer—

ond Yow Movements

4-39



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LMSD-437237
7-129. PROGRAM EVALUATION AND REDIRECTION.

7-130. The System Operations Integration or
and Space Division maintains accurate and co;
activity and results, and conducts a continuing evalua
tions on a flight-to-flight basis, Pertinent records

7-131, Test REPORTS.

7-132. Test Feports that are required for official distribution are divided
into four groups--launch reports, booster contractor reports, Samos reports,
and tracking station reports, .

7-133. LAUNCH REPORTS, The types of launch Teports that are prepared
by the Flight Test Working Group and transmitted to the Air Force Ballistic
Missile Division include a flash report, a follow-on report, and a final report.
The flash report briefly describes launch operations, and the results are
transmitted by teletype. The follow-on report gives a more complete des-

type after a Preliminary review of raw launch data. The unch report
Provides formal documentation and Flight Test Working evaluation of
launch operations and Fesults, in pertinent launch data,

7-134. BOOSTER CONTRACTOR REPORTS. The booster contractor pre-
Pares a quick-look Feport and a final flight test report. The quick-look re-
Port contains results of a Preliminary review of Samos Booster launch data,
with particular emphasis on indicated problem areas. The report is trans-
mitted by teletype. The final flight teat report provides a final analysis of
booster equipment Operation and performance during countdown, launch, and
flight, and encompasses all factors involving possible modification of equip.
ment, plans, or Procedures on future flights. It is due 20 daye after launch,

7-135. SAMOS REPORTS. The types of Samos reports include a prelimi-
nary flight information report, a preliminary system test.report, and a pers
formance analysia and system test report. The preliminary flight informa.

brief summary of system test r,
anda Preliminary Operational
of teat objectives, Problems

It
also contain a final Operations evaluation of overall system performance and

7-23

ee ew ante
tee ee



SECTION vit
“7

SAMOS SYSTEM PRELAUNCH OPERATIONS AND COUNTDOWN
7-32. GENERAL,

7-33. Samos System prelaunch operations are presented under the following
Phases of flight Preparation: the initial vehicle Preparation conducted at the
TJaunch complex which provides final verification of Samos Booster and Samos
Satellite readiness for launch; the integrated facility prelaunch operations
that are conducted to Prepare Samos System equipment and facilities for
applicable system function; the Precountdown operations that are conducted
during the six days prior to launch; and finally, countdown operations, which

are considered to commence approximately 290 minutes before launch of the
Samos Vehicle. . .

7-34. INITIAL VEHICLE PREPARATION AT LAUNCH COMPLEX,

7-35. Although both the Samos Booster and Samos Satellite have completed
@ system checkout in the Guided Missile Assembly Building, extensive re-
testing is required at the launch complex using the launch facilities and
ground- support equipment. Interface and mating tests between t!

and between the vehicles and range tracking facilities are performed for the
first time. Between the time the vehicles arrive at the complex and the

Properly, The Satellite-borne r-f communications, gui
data transmission subsystems, and the Payloads are te

spaceborne and ground equipment, R-F interference between various con-
tractor and range-controlled equipment is also checked.

7-36. The Samos Booster is the first vehicle to be transported to the launch
complex, Here, after initial facilities and Froud support equipment Prepas
ration, the booster is erected within the pad service tower and tests are
conducted over a period of several days. The special booster transport
adapter is removed and replace with the satellite adapter, When booster
tests have been completed, the Samos Satellite (less payloads) ie then trans-
perted to the complex where teets are conducted on that vehi:
cessful completion of vehicle testing, Preparations are then essentiall: com-
plete for erecting and m sting th tellite. Booster and satellite interface
umbilicals are connected,

3
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mt

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5

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2
cy
2
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=
2
&
3
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4
z
2
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ra
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Fy
a

Samos Satellite, but does not cor its and -
electrical equipment. The prin
test is to check the booster propulsi
static, hot firing of the booster.
down and normal launch Sequence operations within limits of the captive

conditions, and are programmed through the automatic launch sequencer or
the Samos Booster. :

7-38. Interface teste between the booster and satellite consist of adapter
mechanical and electrical mating tests; range safety command destruct tests;

——— SEGRET. 7-7



“

SECTION VII
BEF

7-43. PRECOUNTDOWN OPERATIONS,

7-44. Precountdown activities during the 6-day period prior to launch are
briefly presented. The days are indicated in consecutive order, beginning

with the sixth day,indicated as X-6 days,and continuing in sequence to X-0

day, the day of launch. . .

7-45. X-6 DAY ACTIVITIES, The activities during this day include Satellite
Test Center and ‘acking station simulation tests performed for tracking and
communications systems; Vandenberg Control Center voice, teletype, and
timing checks are performed with the Satellite Test. Center and Vandenberg
Tracking Station; Vandenberg Control Center reports on Samos Booster and
Samos Sate lite launch complex status; the satellite is demated and lowered
to the pad for final flight preparations; the umbilical mast is lowered; mast
and connections are checked; final launch preparations are begun for the
Samos Booster. Any descrepancies in booster, launch controls, tanking,
and launch- related equipment are corrected and modifications are completed,

7-46. X-5 DAY ACTIVITIES, Tracking station aircraft fly-by tests are
performed, Telemetry ship departs for downrange position, The Vanden-
berg Communications Center performs communications checks with the
telemetry ship for position, weather, and equipment status reports. The
checks of the previous day are repeated. Internal inspection and final pre-
paration of the Samos Satellite are started, the engine is functionally check-

ed, final check of pyrotechnic circuits is begun, and tanking and pneumatic
systems are serviced,

7-47, X-4 DAY ACTIVITIES. The Satellite Test Center and tracking station
tests are repeated. The Vandenberg Communications Center repeats the
X-6- and X-5-day tests. The Vandenberg Communications Center eives
status from the telemetry ship and relays the report. The Samos Satellite
engine checks are completed. The Pyrotechnic circuits are tested with guid.
ance programmer and squib simulator, The launch control functions are
checked. The umbilical release functions on the launch pad are checked.
Propellant and pneumatic loading systems are functionally checked.

7-48. X-3 DAY ACTIVITIES, The Satellite Test Center and tracking
station tests are repeated, The Vandenberg Communications Genter repeats
the functions of the Previous-day, In respect to the Samos Satellite, complete
guidance and control function tests are run; propellant loading tanks are
filled; final leak tests are performed; the hydraulic: system is recharged and
serviced; blockhouse control functions are checked; and air conditioning
every are performed. The Samos Booster final premating tests are com-
leted, : : :

7-49, X-2 DAY ACTIVITIES. The Satelite Test Center transmits pre-
dicted command and acquisition messages, and exercises tracking station
equipment. The Vandenberg Communications Center repe: communication
checks, and the status of the telemetry ship, tracking stations, launch Prepa-
rations, weather, and other pertinent data revelant to launch are maintained
current for general distribution. Satellite Squibs, ullage rockets, and turbine
starter are installed. Final Premating inspections and safety wiring are com-
pleted. Satellite batteries are installed Payload components are installed
in the satellite and the vehicle is then hoisted and mated to the Samos Boos
The umbilical tower is raised and umbilicals and test plugs are installed.



N ——seeanr_
SECTIO Vv LMSD-437237

relay. Any of these four actions remove Power from the solenoid valves,

causing them to open, Thus, the §as-generator valve and the fuel valve will
close to effect satellite rocket engine shutdown,

4-79. ENGINE MOUNT AND GIMBAL ASSEMBLY, (See figure 4-21.) The
engine mount consists of a tubular frame integral with the Gatellite rocket en-
gine to attach th agine to the satellite. The engine thrust chamber assembly
is attached to the mount and gimbal assembly by the two raulic actuators
of the gimbal system to thus provide for satellite pitch -yaw control,
(Refer to. the guidance and control system, Paragraph 4-143, for a discussion
of the actuator control, ) .

4-80. The gimbal system nominally provides & 5 degrees of defection ina

square pattern. A ting, containing four equally spaced dearinge, surrounds
the thrust chamber forward of the injector, The Pea: 4

rings orgeee four pins,
two located on the forward face of the injector, and two, on engine mount
structure. Rotation of two Opposite pins provides deflection in one plane. In
a similar manner, rotation of the other two pins permit deflection in the per-
pendicular plane, Thrust is thus transmitted through each pair of bearings
to the satellite via the engine mount structure, ‘ :

SAMOS SATELLITE INTERNAL ELECTRICAL POWER Syste
4-81. INTERNAL ELECTRICAL POWER SYSTEM FAMILIARIZATION,

4-82. GENERAL, (See figure 4-10.) The Samos Satellite internal electrical
power system includes the satellite-borne equipment that furnishes electrical
Power for ope Fating the lite and payload reconnaissance equipment
from immediately launch and th: fhout the satellite reconnaissance
lifetime. Prior to of the satellite-borne equipment, ground power

gro
vides the means of switching from ground to the satellite-borne equipment at
the appropriate time Prior to launch, .

4-83. DESIGN CONCEPTS OF INTERNAL ELECTRICAL POWER SYSTEM,
The major objective of the internal power System is to furnish electrical
Power in the proper form to the associated Samos Satellite and Payload eub-

a

which incorporated and universally satisfied the requirements of all Satellite
Subsystems insofar as Practicable, while Preserving a high
ciency and compatibility between the associated subsystems employed,

since the various Satellite systeme have electrical load requirements of the
eame relative magnitude,

4-84. FUNCTIONAL DESCRIPTION OF INTERNAL ELECTRICAL POWER
SYSTEM. (See figure 4-11.)

time energy equipment consists
of 12 silver-peroxide/zinc. Primary batteries, The power conversion equip-

iP:
ment includes two 2000-cycle and one 400-cycle inverter, two 2000-cycle and
one 400-cycle load limiter, two +28-volt and one 400-cycle regulator,

4-26 "—SECRET—

C:



o~

ae

LMSD-437237

SECTION IV |

BLACK

STRIPABLE PAINT
REMOVED PRIOR
TO LAUNCH -

!

COOLING ACCESS
DOOR

"ANTENNA SURFACES
(UNPAINTED)

WHITE

ELECTRICAL
UMBILICAL
ACCESS DOOR

my

FUEL VENT
ACCESS DOOR

\\ OXIDIZER VENT

ACCESS DOOR
t FUEL FILL
DATA LINK ACCESS DOOR
UNPAINTED | OxIOIZER Pius
t ROLLER AOJUST
ACCESS DOOR
1 NCAL 6 PLACES)
HELIUM FILL
P-ACCESS DOOR
mS
Po FILG at DOOR
WHITE PAINT
WITH BLACK |
MARKING

Fe :
QUADRANT MUMBER QUADRANT NUMBER QUADRANT NUMBER QUADRANT NUMBER
L \ ! i] \

Figure 4-37, Somos Sotellite Morkings ond Access Features

—_——__SECGREF—

42115



SECTION IV

—88eRer—
LMSD-437237

OF EVENT IN MIN-

APPROXIMATE TIME EXACT TIME
UTES AND SECONDS | IN SECONDS

OF EVENT

GUIDANCE AND CONTROL EVENT

EJECTION OF FERRET PAYLOAD

(continued)

Plus 40 seconds

plus 50 seconds

plus 60 seconds

Ejection time zero ETZ + 40
Ejection time zero ETZ +450

Ejection time zero ETZ + 60

Ferret payload is ‘ejected,

acth

a Orbital attitude dampi; oye
is. reactivated, TM

b. Sequence timer turns itself off,
©. Orbital attitude damping ee]

Pneumatic control system is de.
vated. :

Continues to control disturbances
about equilibrium Position until
electrical power of Samos Satellite
is exhausted. .

Figure 4.24, Sequence of Guidence ond Control Events (Sheet 4 of 4)

4-60 ——"sreRE?r—__



SECTION vz

WI LMSD-437237
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A Wb )!

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——seener—



——-SECRET— SECTION v
LMSD-437237

5-9. RANGE MEASUREMENT SUBSYSTEM,

5-10, The Range Measurement Subsystem measures the slant range from
the tracking station to the satellite. The subsystem continuously generates
four ranging tones that are transmitted to the satellite by the Command
Transmitter and transmitted back to the tracking station as part of the
narrow-band baseband signal. The received tones are continuously com-
pared with the tones being generated, Because of the time required for the
round trip, there is an apparent phase difference between the received tones
and the reference tones. The difference is measured at regular in-
tervals and converted to slant ran The measured range is supplied to the
station computer, which uses slant range as part of the data necessary to
Correct predicted orbital parameters, Range measurement is also displayed,
in nautical miles, on the Master Control Console,

5-11. TIMING SUBSYSTEM,

erates time signals for indexing data and
The subsystem also synchronizes signals
used to synchronize Subsystem H operations to a common reference,
The operation of the Timing Subsystem is in turn synchronized with the tim.
ing signals broadcast by the National Bureau of Standards radio station,
wwy.

5-13. COMMAND SUBSYSTEM, ..

5-14. The Command Subsystem transmits to the Satellite the commands
that control satellit equipment operation. A synchronized signal and four

station,

buf.
fer storage, and supplied to the Command Subsystem for transmission to the
Satellite during a pase, :

5-16. Command verification signals, denoting receipt of the narrow-band
data signal, are routed to the Command Subsystem by the Data Receiving
The Command subsystem includes the Command Encoder, the
Command Multiplexer, the Command Tranemitter, the Command Trans-
mitted Antenna and ite servos, and two bays of the Master Control Console.

5-17. The Command Encoder sequences the flow of commands through the
Command Multiplexer (which ad the Synchronizing signal and ranging
tones) to the Command Transmitter, The Command Transmitter Antenna
can be positioned manually, but it is formally slaved to either the station
computer or the Angle Tracking ‘Antenna,

5-18. DIGITAL DATA HANDLING SUBSYSTEM,

5-19, The Digital Data Handling Subsystem performs arithmetic Operations
upon data in the digital form, converts digital data to analog data and vice

—SECRET— 5-3



* to operate as rate gyros with the orbital attitude damping

= 37 SECTION IV
reorientation when the sequence timer de-pressurizes the pneumatic control
syatem it also electrically cages the gyros in the inertial reference package
‘stem. Forty sec-
onds after orbit reorientation the sequence timer switches from ascent guid-
ance and control equipment to the orbital attitude damping system, and then
it turns itself off. The Samos Satellite is prepared to perform its orbit
functions.

4-169. ORBIT PHASE. (See figure 4-25.) During the orbit phase the vertical,
nose-down attitude of the Samos Satellite is established and maintained as the
equilibrium position by the gravity gradient and the orbital angular momentum
of the Samos Satellite. Disturbances about the equilibrium position are con-
trolled by the orbital attitude damping system in response to guidance signals
except for the brief period when the ferret payload is ejected. The pneumatic
and hydraulic control systems are deactivated, and power is removed from

the horizon scanner and the flight control electronics unit. The horizon sen-
sor is activated to provide attitude telemetry data, but it does not perform

any guidance function on the initial Samos Nights. .

4-170. The pitch and yaw gyros in the inertial reference package are oper-
ated as rate gyros to provide guidance signals for the orbital attitude damp-
ing system, A signal from the computer is applied to the torque generator

of the pitch gyro to turn the gyro at the orbital pitch rate of the Samos Satel-
lite. This compensates for the constant orbital pitching of the satellite so
that the gyro does not detect the orbital pitching as a disturbance. The orbital
attitude damping system is capable of damping the anticipated peak disturbances
about the equilibrium position to less than two degrees within six hours. Such
disturbances are held to a rate of 0.4 degree per minute. The orbital attitude
damping system operates until the ferret payload is ejected. it does not oper-
ate during the ejection sequence but it is reactivated afterwards and continues
to operate until the electrical power of the satellite is exhausted.

4-171, EJECTION OF FERRET PAYLOAD. (See figure 4-25.) Ejection of
the ferret payload from the dual payload configuration is initiated by a stored
Program command from the satellite-borne communications equipment. The
time for ejection is determined at the Satellite Test Center. ff telemetry data
verifies that the ferret and visual payloads are operating satisfactorily, the
ferret payload will be scheduled for ejection after the Samos Satellite has
completed 24 orbits following launch. This will allow the ferret payload
approximately one and one-half days to successfully perform the functions.
The chosen time is transmitted to the satellite and stored in the sequence
programmer of the satellite-borne communication quipment. At the time
chosen for ejection, the sequence programmer re-starts uence timer
(which had turned itself off after orbit reorientation) and th ction sequence
is started,

4-172." When the sequence timer is turned on, power is re-applied to the
flight control electronics unit, the Pneumatic control system is again pres-

 surized, and cable wires to the ferret payload are cut to prepare it for ejec-

tion. Forty seconds later, the sequence timer switches the signals being
transmitted from the ferret payload antenna to the visual payload antenna,
and the ferret payload is ejected. Ten seconds after ejection, the sequence
timer de-pressurizes the pneumatic control system. Twenty seconds after
ejection the sequence timer removes the power from the flight control elec-
tronica unit, reactivates the orbital attitude damping em, and then turns

—— SECRET— 4-63