SURVEYOR MODEL
SPACECRAFT A-21A DESCRIPTION
, .o. _67- 392 9 0
i _ :E (ACCESSIOlt NUMBER] (THRU) (PAGES) (CODE)
C,_(NASA CR OR
<_VI96
TMX OR AD NUMBER)
3 I
(CATEGORY]
.............................
"1 i
HUGHES
.-.. ......................... /
:
i
NIIliNES SPACE
AI|CIAFT SYSTEMS
COMPANY 9WISIOII
224848
COPY
1 SEPTEMBER 1964
SURVEYOR MODEL
SPACECRAFT A-21A DESCRIPTION
HUGHES
L ............................ HIIGHES SPACE AIICIAFT SYSTEMS COMPANY DIVISION J
JPL950056
This work was performed for the Jet Propulsion Laboratory, California Institute of Technology, sponsored by the National Aeronautics and Space Administration under Contract NAS7-100.
CONTENTS
Page
I,
INTRODUCTION Scope Mission General Mi s s ion
...............................
1 I 1 2 4 .............. 9 9 15 15 ........... 17 17 17 19 19 19 20 20 21 33 35 35 38 45 45 45
...................................... Objectives De scription ............................. .............................
..................................... DESCRIPTIVE Arrangement SUMMARY
II.
SPACECRAFT General Basic Payload
........................... ........................... ............................. VEHICULAR SUBSYSTEM
Design
Concepts
Integration AND
III.
STRUCTURAL Spaceframe Landing Crushable Leg
................................... Mechanism ..........................
Blocks
.............................. Antennas Mechanisms (A/SPP) ................ ...............
Omnidirectional Antenna/Solar Pyrotechnic Electrical Thermal ENGINEERING PROPULSION Vernier Main VI. Engines Retro-Rocket
Panel Devices Cabling
Positioner
............................ .............................. .......................... ................. ........................
Compartments
INSTRUMENTATION SUBSYSTEM
............................... ............................. SUBSYSTEM .................
ELECTRICAL Solar Battery Panel
POWER
..................................
.....................................
Page Battery Boost VII. Charge Regulator Regulator Unit UNICATIONS Group Decoding Processing CONTROL Control Sun s Sensor Sensor Unit ....................... 47 48 .................. 49 49 54 59 ..................... ........................ 67 67 71 72 77 78 SUBSYSTEM ........................... ................. 79 85 85 Subsystem Experiment Subsystem Experiment Subsystem PROPERTIES ............... Subsystem ................ Subsystem ................... ................... ........... ..... 87 98 104 114 121 129 129 Inertias ....................... ..................... ....................... CONTROL ........................ .................. 129 135 136 139 139
............................. SUBSYSTEM
TELECOMM Data Link
................................ Group Group .......................... ...........................
Command Signal VIII. FLIGHT Flight Secondary Radar Roll Attitude APPROACH SCIENTIFIC Introduction Survey Soil Alpha
SUBSYSTEM Group
.............................
...................................... .................................. .................................. TELEVISION PAYLOAD ................................... Television Experiment Sampling Experiment Detector Experiment MASS ...................................... of of Gravity Payload Coordinate THERMAL of Thermal Control and
Actuator Jets
Mechanics-Surface Scattering
Micrometeorite Seismometer XI. SPACECRAFT Weight Center Effects Spacecraft XII. SPACECRAFT Methods
Combinations System
ii
Page XIII. OPERATIONAL Prelaunch Transit Lunar APPENDIX APPENDIX APPENDIX A. B. C. Phase Phase Phase Document SEQUENCE AND FUNCTIONAL PEKFORMANCE 143 143 145 18Z 199 205 ....... 215
................................. ..................................
.................................... and Specification Data Between List ..................
Characteristics Differences
......................... A-21 and A-21A Spacecraft
iii
ILLUSTRATIONS
Page
Frontispiece i-I l-g 1-3 2-I 2-2 3-I 3-2 3-3 3-4 3-5 3-6 3-7 3-8 3-9 5-1 5-2 5-3 5-4 5-5 5-6 5-7 6-1 7-I 7-2 7-3 Atlas/Centaur Earth/Moon Spacecraft Spacecraft Spacecraft Basic Landing Electrical Typical Harness Thermal Typical Thermal Compartment Elements Vernier Vernier Vernier Vernier Main
- Surveyor Launch Trajectory Terminal System General
Spacecraft Vehicle
A-ZIA
..............
viii 3 4 7
.......................
........................... Descent ....................... .................... .....................
Block
Diagram
l0 II 18 18 ZZ Z3 Z5 Z6 Z7 Z9 30 ....... 36 . . . 37 39 39 4O 41 43 46 50 51 55
Arrangement
Spaceframe Leg and
............................... Foot Pad ......................... Device ..................
Harness Harness Removal Tray Thermal Switch
Disconnect Interconnection Concept
.....................
.......................... ........................... Design .................
Assembly
Compartment
................................. Dis sipating Capability and Flight ................... Control Subsystem
of Propulsion Propulsion Engine Thrust
System,
Functional
Schematic
Diagram
Assembly Chamber Tanks Engine
.......................... .......................... and Spaceframe ..............
Propulsion Retro-Rocket
......................... ........................... Block Diagram ..............
Retro-Rocket Electrical Command
As sembly Power Receiver Subsystem and
Transmitter Block
.................. Diagram ...........
Telecommunications Command Decoding
Subsystem Block Diagram
....................
iv
Page 7-4 7-5 8-I 8-2 8-3 8-4 8-5 9-1 9-2 9-3 i0-I 10-2 10-3 Central Central Flight Altitude RADVS _ADVS RADVS Surveyor Approach Approach Survey Survey Survey Block Survey Survey Command Signal Control Marking Assembly Beam Block Decoder Processor Block Diagram Block ......................... .......................... ....................... Diagram ................. Cover Removed ........ 57 64
68
73 74 75 76 80 81 83 ..... 88
Radar with
Preamplifier
Orientation Diagram
.......................... ............................
Television Television Television Television Television
Approach Subsystem
Camera Block
................. Diagram Sequence Block ........... ............ Diagram
Picture-Taking Subsystem, Camera
Functional
......................... Functional
89
TV Camera/Ground Equipment Interface, Diagram ................................. Television Television Composite Video Output Sequence
91 94 97
10-4 10-5 10-6
.............. ..............
Picture-Taking
Soil Mechanics Surface Sampler Experiment Subsystem, Functional Block Diagram .......................... Soil Mechanics Surface Sampler Instrument Partially Functional Extended Block
99
i01
10-7 i0-8
Alpha Scattering Experiment Subsystem, Diagram ..................................... Alpha Scattering Sensor Deployment
iii ........... 113
10-9 i0-i0
Mechanism
Potential Scattering
Azimuth Mechanical Interference and Surface Sampling Experiments
between Alpha ............ Functional
115
I0-ii
Micrometeorite Block Diagram View Factor
Detector Experiment Subsystem, ................................. of Micrometeorite Instrument Experiment Instrument Payload Detector Instrument
117 Assembly . 119 122 Block Diagram 123 127 134 140 152 154
10-12 10-13 10-14 10-15 II-I 12-I 13-i 13 -2
Seismometer Seismometer Seismometer Atlas/Centaur Lunar Sun Surface Sensors, Sun
.......................... Subsystem, Auxiliary Envelope Functional Unit
.................
.....................
Temperature Locations Sensor and
........................ Orientation Logic ................ ................
Secondary
Orientation
v
Page 13 -3 13 -4 Typical Typical Transit Standard Transit Operations and Power Profile ....... 157
Compartment Thermal Dissipation During Phase .................................. A and B Temperature Profiles
Standard 158 159 165 Magnitude for 166 168 174 176 178 180 of Touchdown 183
13 -5 13 -6 13 -7
Compartments Po s sible Typical Normal Typical Approach Typical Landing
.............
Site s
.............................
Miss as Function of Midcourse Maneuver Impact .................................. Midcourse Television Retro Sequence of Main Phase Maneuver Geometry Capability
13 -8 13 -9 13-10 13-11 13-12 13-13
..................
........................ ..................... Velocity .............
of Events Retro Burnout
Determination Vernier Maximum Incidence Maximum Landed Lunar Lunar Descent
.............................
Allowable Lateral Velocity as Function Angle .................................. Allowable Vertical Center of Gravity Weight .................................. as
13-14
Function
of 184
13-15 13-16 13-17
Day Night
Operational Survival
Capability Capability
for
Equatorial
Landing
......
187 192
....................... Landing
Hypothetical Lunar Phase Sequence - Equatorial Assumed ...................................... Surveyor Compartments Differences Compartment Vernier Descent Physical Differences
193 Z19 Profile 221 Lunar Day . . . 222 223
........................
A and B Thermal Tray Temperature .................................... Dissipating Phase Capability Differences Differences,
....................
vi
TABLES
'Page 3-1 3-Z 7-1. 7-2 I0-i II II -i -2 Pyrotechnic Thermal Portion Time Soil Devices Compartment of ESP Required .............................. Component Data Frame Sampler Summary Weight Installation Frame ............. Zl 31 60 6Z 105 130 130
Commutator for One
................. Data .........
of Commutated Commands
Mechanics-Surface A-21A A-ZIA Weight Detailed
...............
Surveyor Surveyor Itemized Experiment Payload Summary Attitude Coast Typical Power Expected Vernier Maximum A-21A
..................... Status .................. Payload
11-3
Weight Summary for Each Scientific .................................... Combinations of Operational Control System Summary
137 138
1 1 -4 13 -1
...................... Phase
Modes Provided by Coast ............................ Control Capability
147 156 ....... 160 163 164 170 175 191 ZI6 217
13-Z 13-3 13 -4 13-5 13-6 13 -7 13 -8 Cfl C-2
Phase
Attitude
................. During Transit
Expected Into Antennas Quality Engine Shade Compartment
Thermal
Performance
.............................. ....................... Modes .................. Equipment Loss ............ .......
of Telemetry Thrust Time Control for Lunar
Spacecraft Night Heat
Summary
Component Weight
Differences, Differences
A-21 and
to A-ZiA A-ZIA
.................. ..................
of A-Z1
vii
&IN A
NlDlRECTlONAL rENNA ( 2 )
.
-
VERhIlER FUEL TANK
DESCENT RADAR ANTENNA
(LI
innu31
I
CHAMBER ( 3 )
ATITUDE JET
L
SURVEYOR S P A C E C R A F T A-21A viii
I NTRODUCTI
ON
S C OPE
This A-ZIA model
document
describes
the
general The
design material
and
overall
performance in the
of the
of Surveyor spacecraft expected
spacecraft. design of the
presented
document predicts condi-
describes
the
in its present spacecraft Although by
state
of development under
and
the performance tions in a
when the
operated spacecraft
realistic
specified
environment. or design influenced and
performance other
is conthan those
trolled, associated defines
constrained, with spacecraft The current of these context the
a large
number
of factors
construction in terms
of the of the
spacecraft
itself, design provided
this document mechanizaJPL and repre-
performance
spacecraft
and by
tion only. sents the
description Hughes
of those Aircraft A
scientific Company complete functional,
instruments understanding definition
of the design of spacecraft and
operation in total interfaces
instruments. the numerous
performance
with
operational,
administrative
that exist
is beyond
the
scope
of this document.
MISSION
OBJECTIVES
The Company sion
Surveyor under the
spacecraft direction for
is being of the
designed
and
built by
Hughes
Aircraft Jet Propul-
California
Institute and
of Technology Space
Laboratory The a transit
(JPL) Surveyor from
the National
Aeronautics has been
Administration and designed and to trans-
(NASA). effect mit
spacecraft earth to the
vehicle moon, and
conceived
perform engineering
a soft lunar data relative
landing,
back
to earth and
basic
scientific
to the moon's
environment
characteristics.
To date
obtain
maximum
utility, payloads.
the
spacecraft basic
has
been
designed elements
to accommoof structure, provide the
various
alternative
The
spacecraft and make
telecommunications, capability maintaining to perform two-way
power the
generation, earth-moon
propulsion, transit This and basic
flight control a soft lunar
landing
while
communication.
grouping
of spacecraft
elements,
designated as the "basic bus," can thus provide transportation, power, and communication services to the designated variety of payloads. The A-Z1 series of spacecraft, which constitutes the first four Surveyor launches, carries an engineering payload. The purpose of the A-21 series is to demonstrate successful transit and soft lunar landing and to gather basic engineering data relative to the performance of the spacecraft in the environments encountered in transit. The collection and transmission of scientific data is a secondary objective for this series of spacecraft. The A-21A series of spacecraft utilizes essentially the same basic bus but carries a different payload, consisting of various scientific instruments. The primary purpose of the A-21A series is the collection and transmission of scientific data relative to the lunar environment. The design and performance of the A-21 spacecraft is covered in HAC document 224847, "Surveyor Spacecraft A-21, Model Description."
GENERAL DESCRIPTION
Spacec
raft The general are and shown arrangement in the of the spacecraft The mounted members. of a tripod of gravity and identification is composed of its various of several of
elements electronic thin-walled frame for use
frontispiece. assemblies tubular
spacecraft on The landing
mechanical alloy the
a spaceframe configuration gear with
constructed of the three space-
aluminum by
is dictated in the over
selection Center
foldable low
legs
soft landing. a wide range
of the vehicle Thermal lunar
is kept control
to obtain
stability ment over
of landing
conditions. range of the
of the equip-
the extreme
temperature
surface and
(+Z60 ° to -Z60 ° F) active methods.
is accomplished This and design thermal
by a combination represents design the latest
of passive,
semi-passive
state of the art
in the application structures.
of structural
principles
to light-weight
space
Launch
Vehicle The spacecraft boost will be vehicle launched (figure on l-l). vehicle to meet its 66-hour Under transit to the moon by the Lewis by
Atlas/Centaur Research General mission.
the direction
of the NASA designed
Center,
the Atlas/Centaur
is being
specifically requirements
Dynamics/Astronautics The folded spacecraft
the launch within
of the Surveyor shroud on
is housed
a conical
breakaway
FIGURE 1 1.
-
ATLAS/ CENTAUR LAUNCH V E H I C L E
3
top
of the
second and
stage liquid
Centaur. hydrogen
The
Centaur,
complete
with
its guidance atop the
system,
fuel tankage, Atlas rocket.
engines,
is located
directly
first stage
MISSION
Launching
and
Tracking will
Operation originate from boosted nose from JPE by Cape Kennedy, Flight Florida, with computation (SFOF) 60 then inject of will antennas. initial South protranspanel in
Launchings and space
flight control Upon being
Space
Operations altitude The l-Z). Centaur gear on
Facilities
Pasadena. miles, the the
the Atlas will be
to an
of approximately Centaur At the will
the breakaway into an thrust
shroud
jettisoned. (figure the landing
spacecraft Centaur
earth/moon but extend mode
trajectory
conclusion
phase, that
before the
separation, spacecraft
programmer and the omni
generate The
commands
high-power
transmitter
is also
commanded will occur (DSIF) and are
at this time
to aid
spacecraft Africa vided mission Deep by
acquisition. Space
Initial acquisition Facility Australia, this phase
at the
Johannesburg, later
Instrumentation at Canberra, during
station
with
tracking Both
stations and
Goldstone, the
California. The Centaur.
reception automatically
from
spacecraft. from
solar
is deployed
at separation
of the
spacecraft
POSITION
OF
MOON
AT
IMPACT I
TOUCHOOWN
TOSUN PHASE RETR0 ,,N,T,ATE0\ \t
_ / _ / GOLDsTONE _ / _ _ AT 60 NOMINAL MILES RANGE FROM OF _ _ l t MOON) _L _..'a',
............
mlJlr.t.vluN SEPARATION
_
Anli,I/ / /
_jl-..("._1
• ._1_. ":'/ / "_
..."- _ET.OMA__U_VEpS
_" ; / "_-_ \ • _X X OTHER MINUTE REACOUISITION OPERATIONS S BEFORE SUN T
A_
(NOMIN&LL¥ TOUCMOOWN, AN{)
_i
30 , 4#
///I
OF
ACQUISITION / SUN ACQUISITION (NOMINALLY <1 AFTER LAUNCH / HOUR / STAR ACQUISITION AND VERIFICATIO_d (NOMINALLY 6 HOURS AFTER LAUNCH) M,DCOURSE CORREC
,ON) ,o _ _ _
_,_ \'POSITION MOON AT OF LAUNCH
MIOCOURSE CORRECTION (NOMINALLY I_) HOURS AFTER LAUNCH)
FIGURE
I-2.
EARTH/MOON
TRAJECTORY
Midcour
se Correction Three reaction and gas track jets, the located sun and on the the star a landing Canopus. gear legs, When position the space-
craft
to acquire lock which The on
the appropriate system is established
sensors in space maneuver. begins Tracking and from used
to these will
celestial be
points,
space
coordinate maintained direct sun
thenceforth panel
automatically to achieve
until the midcourse illumination and battery are and charging. processed
solar
is oriented energy for from
to generate data by
electrical
spacecraft the DSIF
operation tracking
received
in sequence
stations Upon radio
SFOF the
to compute spacecraft thrust and
the midcourse turns through
correction. a series
command to align The
earth,
of angular
maneuvers vector.
the vernier required from
engine magnitude
axis
relative for
to the
spacecraft
velocity
direction
the midcourse
maneuver where
is transmitted it is received vernier and
the Goldstone Then
tracking
station command
to the spacecraft, causes three
stored. engines time. which the star
the execute
liquid-fueled for
rocket of
to operate This will action ultimately
at a specific provides bring
average
thrust
level
a specific spacecraft landing
period
a midcourse the vehicle
alteration to the the
of the
trajectory area. the sun After and the
selected spacecraft
lunar
midcourse Canopus
correction to maintain
is completed, its previous
reacquires
attitude.
Terminal
Descent
and
Soft Lunar 66 hours
Landing launch, tracking the spacecraft the approaches spacecraft the moon. changes attiA
Approximately Upon tude command to align the from
after
the Goldstone
station, with the
thrust
of its re'tro-rocket camera views
spacecraft surface area
velocity during
vector.
downward for the
looking purpose
television
the moon's
the approach As the
of transmitting the moon radar
pictures
of the landing speed
back 9000
to earth. feet per
spacecraft the altitude
approaches marking delay,
at a relative a signal of the
of about range
second, which,
generates
at a slant solid ejects
of 60 miles, main
after motor. from
a suitable This the
initiates and
ignition
propellant the altitude the
retro-rocket radar At out, an its At
ignition
subsequent and
burning
marking
retro-rocket of approximately
nozzle
begins feet,
to decelerate the main
spacecraft. burns
altitude empty this
40,000 from the
retro-rocket
and
case point
is separated the spacecraft signals system
spacecraft enoughto
approximately the and surface doppler of the
8 seconds moon
later.
is close from are
to receive system. to control
reliable Signals
control from
its altimeter processed by
velocity
radar
this
the fl_ght control
electronics
the throttle valves on the three vernier rocket engines to maintain the proper attitude and rate of descent. The spacecraft continues to decelerate until at an altitude of 14 feet, the vernier engines are turned off. At this time both horizontal short and vertical components to the gear and of velocity surface are small. The with the This spacecraft landing sequence falls the shock is illus-
remaining by
distance
of the moon
absorbed trated
the landing i-3.
the crushable
blocks.
in figure
Lunar
Operation After landing panel on the to the surface of the moon, line the and spacecraft the high-gain is commanded planar one array of the to the with the survival on to mode to
align antenna
the
solar to the
spacecraft-sun line.
spacecraft-earth to the planar array will solar
Subsequent and
commands
connect
transmitters (high power or low
antenna provide panel
switches
that transmitter consistent
power) from
which the
a usable and battery.
bandwidth After
available
initial touchdown can be turned
conformation, provide subsystem.
the various
payload
experiment
subsystems via
information
relative
to the lunar
environment
the telecommunications
FIGURE i-3.
SPACECRAFT
TERMINAL
DESCENT
pI:kECEDING
PAGE
BLANK
I_40T I:IL_ED.
II.
SPACECRAFT
DESCRIPTIVE
SUMMARY
GENERAL
ARRANGEMENT
Design launch vehicle,
of the the
Surveyor established and the
spacecraft DSIF nature and
is dictated Space-flight
by
the
configuration
of the
Operations mission. reliability, a high
Complexes, The use of
reliability
objectives, design
of the
spacecraft of proven toward
state-of-the-art ance succes of unproven s. Figure midcourse, The principal 2-i and
criteria, or
components
and
the avoidof
circuits
designs,
contribute
probability
is a basic post-landing
block
diagram
of the
spacecraft are
system. shown
The
stowed, 2-2.
configurations the system
of the A-ZIA are listed
in figure
elements
comprising
below.
Surveyor
Spacecraft
Subsystem
Elements
Structural Spaceframe--provides Landing Crushable mechanism Antenna/Solar toward the Leg Mechanism the basic -- absorbs part
and
Vehicular
Subsystem spacecraft. of the after shock of landing. large landing leg
structure major of landing
for the portion shock
Blocks
-- absorb
relatively
deflections. Panel earth and Positioner the solar (A/SPP}-panel toward actuate switches, spacecraft orients the the planar array antenna
sun. separation and nuts, tank
Pyrotechnic valves, Electrical Thermal thermally
Devices
-- mechanically power control
pin pullers, locking units.
electrical Cabling--
plungers
a detonator.
interconnects -- provide units.
Compartments sensitive
temperature-controlled
environment
for
SPACECRAFT
•O"T'CAL NER"'L i "N0 ' SE_,NG
LUNAR (RADARS) VELOCITY AND ALTITUDE_--_-Ib.J
l
•
I I:ATTITUDEVELOCITY _L,GH+ CONTROL ) 'NO I
CONTROL J
t I I
[
MIDCOuRSE PROPULSION MANEUVER RETRO MANEUVER
•
t
ELECTRIC • POWER CONTROL ENERGY STORAGE POWER SOURCES
_J
I • TRANSMIT AND RECEIVE • TELECOMMUNICATIONS COMMAND DISPERSAL • SIGNAL AND PROCESSING SWITCHING I • AND • I
• •
SCIENTIFIC
PAYLOAD
APPROACH
TELEVISION
ENGINEERING • MONITOR FUNCTIONS
INSTRUMENTATION BASIC BUS
INTEGRATING WIRING
STRUCTURE
MECHANISMS AND CONTROL VEHICULAR
: SIGNAL
AND
POWER
FLOW
• STRUCTURAL THERMAL
TO
DSIF
FIGURE
2-i.
SPACECRAFT
SYSTEM
BLOCK
DIAGRAM
Engineering Temperature frame status and and Acceleration performance. Sensors
Instrumentation -- provide for earth monitoring of space-
Propulsion Vernier trol Enginessupply reaction burning, forces and for midcourse and velocity correction, control attitude during con-
during
retro-rocket
attitude
terminal
descent. Main Retro-Rocket-decelerates spacecraft on approach to lunar surface prior to
final descent.
I0
®
ANTENNA SECONDARY SOLAR SENSOR
&
SOLAR
POSITIONER
PLANAR SOLAR PANEL ANTENNA
ARRAY
MICROMETEORITE SENSOR
LANDING MECHANISM
GEAR -3
RELEASE REG
TV
CAMERA (SURVEY)
NO.
S
TV
CAMERA (APPROACH)
NO.
4
/ CRUSHABLE BLOCK
ALTIMETER/VELOCITY SENSING ANTENNA
®
POST
LANDING
CONFIGURATION
fl-I
®
(I-2.
®
32.774 DIA
"l
SOLAR
PANEL
,
PLANAR
ARRAY
/
ANTENNA
RELEASE S REQD
MECHANISM
/CENTAUR INTERFACE
ALTITUDE MAIN ROCKET RETRO ENGINE
MARKING ANTENNA
SPACECRAFT INTERCONNECT STRUCTURE ADAPTER FIELD
JOINT
_ ", 104.704
MOUNTING FLANGE ATTACH STATION
/" CENTAUR
DIA
_ (_
THERMAL D,APHRAGM
STOWED
CONFIGURATION
FIGURE
2-2.
SPACECRAFT
GENERAL
ARRANGEMENT
Electrical
Power
Solar Main Battery solar Boost
Panel--
charges
battery
and
powers energy controls
spacecraft storage and for
during the
transit spacecraft.
and
lunar
day.
Battery--provides Charge panel. Regulator (BR)-Regulator
electrical (BCR)--
regulates
battery
charge
from
converts
unregulated
battery
power
to regulated
power
for
spacecraft.
T ele communications Transmitters and scientific (Z)transmit data engineering from lunar telemetry surface. earth-transmitted unit. Also commands two-way and route data in transit and engineering
payload
Receiver/Transponders these commands to the
(Z)-- receive central with decoder
provides
doppler
tracking Antennasreception landing Central mand control Signal
in conjunction two and and
transmitters. antennas antenna for for data transmission band data and command during
omnidirectional one planar phases. Unit-and five contains subsystem and array
wide
transmission
lunar
Decoder decoders on-off
a receiver-decoder decoders
selector, earth
two
central
comand
that process operations. and verification
commands
operations gather provide Auxiliary--
time-interval engineering signal
Processorsand Rate
the
signals telemetry.
from
various
subsystems Low Data
appropriate provides limited
conditioning
for at low
for transmission information
bit rates
for
use
with
low-power
transmitter
and
bandwidth.
Flight Inertial craft eration Primary once sun Reference Unitcontrol for provides of optical
Control rotational sensors. reference Also while an spaceaccel-
three-axis or radar
is not under reference Sun
provides
spacecraft provides is obtained.
flight. for accurate control of the spacecraft roll axis
Sensor--
acquisition
13
Inertia Switch-- closes at a nominal g level to predict retro rocket burnout for retro ejection timing. Canopus Sensor-- identifies attitude control reference. and tracks the star Canopus for accurate spacecraft
Flight Control Electronics-- processes guidance signals from the flight control sensors (inertial, optical, and radar) for stability and maneuvering. Secondary Sun Sensor- makes initial sun detection for gross alignment of spacecraft roll axis during transit, and for solar panel positioning toward the sun during lunar operation. Radar--the altitude marking radar (AMR) initiates firing of the retro-rocket on
approach to lunar surface. The radar altimeter and doppler velocity sensor (RADVS) measures slant range and three-axis velocity of spacecraft during descent phase, controlling the rate of descent and attitude via the vernier engines. Attitude Jet System-- provides reaction forces for spacecraft orientation maneuvers and attitude control during period from Centaur separation through preretro- rocket firing. Roll Actuator--provides Vernier Engine No. 1.
Approach Approach of I000 Television miles to about Camera-80 miles Television Subsystem of lunar surface from a range
roll control moments during vernier
engine thrust,
via
provides above
pictures lunar
surface.
Scientific Survey surface, Television free Experiment and of the Subsystem spacecraft
Payload -- provides after pictures of portions of lunar
space, Surface of the
landing. determines the mechanical
Soil Mechanics characteristics Alpha surface Scattering elemental
Samplerlunar
qualitatively
surface. Subsystem-gathers information to determine lunar
Experiment composition. Detection impacting
Micrometeorite micrometeorites
Experiment--measures the lunar surface.
lunar
ejecta
resulting
from
14
Seismometer moon,
Experiment
Subsystem -- measures physical disturbances
on the
BASIC DESIGN CONCEPTS The primary design objective has been to maximize the probability of successful spacecraft operation within the basic limitations imposed by launch vehicle capabilities, the extent of knowledge of transit and lunar environments, and the current technological state of the art. In keeping with this primary objective, design policies have been established which (1) minimize spacecraft complexity by placing responsibility for mission control and decision making on earth-based equipment wherever possible; (Z) provide the capability of transmitting a relatively large number of different data channels from the spacecraft; (3) include provisions for accommodating a relatively large number of individual commands from the earth; and (4) make all subsystems as autonomous and independent as practicable. These basic design policies complement each other and provide a large degree of flexibility in controlling the real-time operation of the spacecraft. Complete control of spacecraft operation is achieved through a loop that is closed through earth-based equipment and decision-making processes. The only portions of spacecraft operations that are not subject to this earth/spacecraft control loop are those associated with certain portions of the attitude stabilization and terminal descent phases, and solar panel deployment where earth control is complicated by requirements for critical timing. Although this design concept places greater demands on earth-based eq'uipment and facilities, it provides flexibility in control and data-transmission adaptability and growth potential. This concept enables the same basic spacecraft design to accommodate a wide range ofpossible payloads and missions. PAYLOAD INTEGRATION series of Surveyor spacecraft vehicles has been designed to experiment subsystems:
The A-ZIA
accommodate the following a. b. Survey television
experiment subsystem. subsystem.
Soil mechanics surface sampler experiment
c. Alpha scattering detector experiment _:_Instruments furnished by JPL
subsystem. _',_
15
d. e. Any primarily
Micrometeorite Seismometer combination by the total
detector experiment of experiment injection weight
experiment subsystem. subsystems capability the
subsystem. * may of the be
*
accommodated, vehicle. wide on the spectrum spacecraft is Hughes the
limited
launch
Flexibility different various through cally for electrical scientific the use
is provided and
to accommodate requirements in a typical auxiliary An the experiment spacecraft electrical units
potentially imposed
of by the
functional
instruments of electronic instrument. between normalizing experiment necessary
payload. designed auxiliary basic signals for the bus
This and unit and
flexibility built provides a particular the two. by
achieved specifi-
each
electrical] instrument
functional by
interface and the
conditioning When
between a specific
desired only
complement to remove which to adjust
spacecraft auxiliaries
has and
been other In of
determined, elements some cases
it is of the
appropriate are the being
experiment also be
subsystems necessary This
deleted
(off lateral
loaded). center
it may
spacecraft the exact
gravity by adding experiment
ballast.
approach
permits
definition of the
complement st cycle.
to be delayed
until relatively
late in the spacecraft
fabrication/te
*Instruments
furnished
by JPL
16
III.
STRUCTURAL
AND
VEHICULAR
SUBSYSTEM
The mal
structural
and electrical
vehicular
subsystem
provides mechanical
support, actuation, This blocks, subsystem
alignment, and
ther-
protection, for
interconnection, and leg its
touch-down includes the
stabilization basic antennas electrical SPACEFRAME The mounting structure, mal control trates tubing, around points provided spaceframe po s itioning LANDING The intermediate landing the forces
the
spacecraft landing
components. crushable positioner,
spaceframe, mechanisms, cabling,
mechanism, panel compartments.
omni-directional devices,
antenna/solar and thermal
pyrotechnic
spaceframe surfaces the main and
is the
basic
structure for the the vernier mast, and spaceframe interconnected points structure,
of the landing
spacecraft. gear, and control payload. up of thin the
It provides Centaur interconnect tanks, group, Figure wall 3-1 therdescent illus-
attachments rocket,
retro crushable
engines flight scientific is made to form for the and mast antenna is
associated sensor
compartments, radars, the basic the retro the at the and flight
blocks, sensors, The
control
spaceframe. frame rocket. Centaur corners supports m. members
aluminum
with the for
a modified rocket, leg hinge to the panel
hexagon attachment points top of the a are
Attachment interconnect of the the frame. planar
retro landing attached and solar
The array
through
m echanis LEG landing
MECHANISM leg mechansim the and shock (figure absorber, 3-2) is the maintain The landing control with absorver, made footpad, attitude legs jets. the up of the and the landing lock leg, strut. and long absorb radius leg is footand the The
A-frame, lep_ footpad, of impact points for
shock
absorbers
stability provide The landing
during the
touchdowr_ cold gas attitude spaceframe shock
attachment hinged pad to the attached
lower to the
corner outer
of the end. The
aluminum intermediate
honeycomb A-frame,
17
FIGURE 3-1.
BASIC SPACEFRAME
FIGURE 3 - 2 .
LANDING L E G AND FOOT PAD
18
lock strut are interconnected to permit folded stowage of the landing gear by telescopic action of the lock strut. Torsion springs at the leg hinge extend the legs when the squib-actuated pin pullers are operated on Centaur command. The legs can also be extended by earth command. The lock strut locks in the extended position and forms, with the shock absorber, a straight line member from footpad to the spaceframe upper corner. The hydraulic shock absorber compresses with landing load and absorbs the landing shock. Crushing of the footpads absorbs some of the impact energy. After landing, the shock absorbers are locked in place by squib-actuated pin pullers. CRUSHABLE BLOCKS The crushable blocks of aluminum honeycomb are attached to the bottom of the spaceframe at each corner to absorb part of the shock of large landing loads. The blocks contact the lunar surface upon occurrence of any relatively large landing leg deflections and absorb energy by crushing. OMNIDIRECTIONAL ANTENNAS MECHANISMS
The omnidirectional ("omni") antennas are mounted on the end of folding booms, hinged to the spaceframe, with omnidirectional antenna boom A near landing leg 1 and omnidirectional antenna boom B near landing leg 3. The omni antenna booms are stowed by folding against the spaceframe. Pins retain the booms in the stowed position and squib-actuated pin pullers release the booms on Centaur command. Torsion springs deploy the omni antenna after release. Omni antenna boom release is effected by a command from Centaur after the landing leg is extended and locked into position. Earth commands can also initiate omni antenna boom extension.
ANTENNA/SOLAR The the mast tion with one A/SPP PANEL connects POSITIONER (A/SPP) array and solar panel to the top of has three The axes of rota-
the high gain planar The planar
by hinge connections. respect to the mast:
array
antenna
roll, polar,
and elevation. array
solar panel has Stepping motors This the
axis of rotation with respect
to the planar in response
antenna.
rotate the axes freedom earth
in either direction permits
to commands array
from antenna
earth. toward
of movement
orienting the planar
and the solar panel toward
the sun simultaneously
after landing.
19
PYROTECHNIC The and and valves quantity no
DEVICES devices in table 3-1, All for mechanically which lists the devices safety actuate items, are design the their vased mechanisms, locations, on the switches, functions, "1 watt, A total of
pyrotechnic listed
required. fire
pyrotechnic range
1 ampere, 36 pyrotechnic
5 minutes" is used.
requirement.
devices CABLING electrical
ELECTRICAL The correct the cable tunnel heat in the
cabling power
interconnects flow. Cable the design installed
the
spacecraft permits wire is and routed To "tear contains unregulated pin puller aluminized shows the
components installation cable or connectors.
to provide removal The of
signal
and by
assemblies connecting to minimize losses wall
disconnecting the two heat the thermal loss lunar from night, A. pre from from the
compartments the compartments.
through further
a thermal minimize
during
a disconnect tear strip
strip" leads
is installed for RADVS and the 22-volt tear power, power strip,
of compartment squib On one flows power,
This
19-ampere return. removing tion after then
SS & A 22-volt the DSIF, a squib
power, separates Mylar
command part
compartment. the hole. from or Figure the
The 3-3 thermal
super-insulabefore and
in to close Cables inserts the tear Figure
disconnect contain wires
actuation. wire through lunar The night. scientific
exiting (nichrome strip) 3-4
compartments wire) loss from harness. designed installed or the or around in all the
thermally {except during
insulating those the
perma-nickle heat the
to minimize illustrates of the
compartments
wiring is be
payload each
spacecraft may
five
experiment individually bus, in the except vehicle
subsystem without for when {figures four weight an
installations, effect and on other
of which
removed basic left are B for to the can bus be
experiment To subsystem minimize is
subsystems the removed, is wiring cable reliability television disconnect installed from amount
spacecraft weight
balance.
of cable
experiment 3-4 and 3-5).
individual
harnesses
employed each experiment or assemalso use of
A connector All
in compartment the spacecraft that basic
experiment is without The
subsystems. contained
subsystem removed blies. external
in a separate the the survey as TV
assembly of the
installed
compromising for connectors
harness
harnesses bus
experiment points.
subsystem
basic
20
TABLE Type Pin Puller
3-1.
PYROTECHNIC
DEVICES L oc ation Quantity 2 3 panel locks 7 3 1 1 3 2 2 3 1 absorbers 3 4 auxiliary}
Omnidirectional Landing Planar Alpha leg locks
antennas
antenna/solar scattering
deployment sampler device
Soil mechanics-surface Electrical Separation Valve nut harness
disconnect
Retro-rocket Helium Nitrogen
attachments
operation
tank valves tank valves leg shock tank valve leg shock power absorbers
Locking
plunger
Landing Nitrogen
Locking
plunger electrical switch
Landing RADVS
Pyrotechnic
controls mechanism
(engineering Ig nito r Cable clips. Slack assemblies cabling Retro-rocket are attached
to the spaceframe mechanically
by
suitable brackets The
and
is provided is established harness legs
around
active points. connector
Centaur as
electrical
interface
through I. The
a 5Z-conductor connector mounts
assembled of the
part of basic spaceframe
bus wiring between
on the bottom
landing
1 and Z and mates
with a Centaur
connector
when
the spacecraft THERMAL Two items
is mounted.
COMPARTMENTS thermal compartments control (A and B) are provided is needed equipment throughout in these to house electronic (See figure is mounted 2-Z
for which
thermal
the mission. compartments
for compartment on a thermal A thermal surface tray
placement.
) The
(figure 3-6) which
distributes
heat throughout
the compartment. the lunar of
shell surrounds environments
the entire compartment An
to isolate it from composed
(figure 3-7).
insulating blanket,
Zl
a ) B e f o r e Actuation
b ) A f t e r Actuation
F I G U R E 3-3.
ELECTRICAL HARNESS DISCONNECT DEVICE
22
--__ L ....
1 [ .... 1 F-_.--1 J L __'___S L .... __J _
Fi_-_
JL___
T
-
•
!
rTwu-] I_NI I .A- I L. .J
I
.n. I
l Ii=_,i
,l TR _MITTER _RF _-TRAN_I, ilTTF.R SWITCH I _TCH
I
I I
I
"_"
I _UXlU_ I
II _L_.,_
IIII ; I I
w'_I'_ i,
F_
•
m
_2_. _
I _At. I lurxu L______J _-¢ I
2._ -/
qL_lLq_lLq &l_pt_y &N_/SOLAR
_
POQITIOqRR
&N O L(3CKS
......................
I .....................
_
so_ L.-__
_L
J
I
]
I
l ]
I L
COMPARTMENT
B
,z_ -E
SPACECRAFT SUBJ£CT SELECTION
TO
PAYLOAD PAYLOAD
WIRING
1 I
_,cK 0_C_ECT
I ® c_Ecm_ To K _s,_op _ a N_s
FIGURE
3-4.
TYPICAL
HARNESS
INTERCONNECTION
MICROMETEORITE SENSOR SEISMOMETER SENSOR i NICHROME WIRE M ICROMEEl ELECTRONICS SEI SMOMETER ELECTRON I CS
ALPHA SCATrERI G N DEPLOYMENTl N P PULLERS
SURFACE SAMPLER
HARNESS SUPPORT DETAIL
ALPHASCATTERI NG SENSOR
RTMENT B WALL MI CROMETEORI TE AUXILIARY SEISMOMETER AUXILIARY THERMAL SPLICE AREA AUTONOMOUS HARNESSES, ONE FOR EACH EXPERIMENT SUBSYSTEM
INTERFACE DISCONNECT, SCIENTIFIC PAYLOAD BASIC BUS (WITHIN COMPARTMENT B) CENTRAL S IGNAL / PROCESSOR
SURFACE SAMPLER AUXILIARY
ALPHASCATTERING AUXILIARY
BOOST / REGULATOR CENTRAL COMMAND DECODER /
/
SCATI'ERING ELECTRONICS
COMPARTMENT B POWER BUS
/ /
BASIC BUSCONTINUATION (TYPI CAL)
ALLSCIENTIFIC PAYLOAD WIRING THIS SIDEOF INTERFACE SCONNECT DI IS INTEGRAL WITH BASIC BUSWIRING ENGINEERING SIGNAL PROCESSOR
FIGURE
3-5.
HARNESS
REMOVAL
CONCEPT
25
I
I
MAIN " ENGINEERING MECHANISMS AUXILIARY
BATTERY
THERMAL CONTROL AND HEATER _ECEIVER TRANSPONDER B
BATTERY CHARGE REGULATOR RECEIVER TRANSPONDER A
TELEVISION AUXILIARY
E L ECTRON IC S _"_......._ MICROMETEORITE DETECTOR AUXILIARY _ \" _, . ,"_Z '__" / / / BOOST
,_. f _"_
/SIGNAL
PROCESSOR ALPHA SCATTERING
,/_e,,_ / _ ./')
ELECTRONICS THERMAL CONTROL -AND HEATER
_" X
_
ALPHA
SCATTERING
_ _ AUXILIARY LOW DATA RATE AUXILIARY
/ REGULATOR
FIGURE Z6
3-6.
THERMAL
TRAY
ASSEMBLY
,.,.q
I
t'-.-
A
c¢ 0 t-. ,,_ J 0 00 .J o., w 1Iw "Y'¢o
Z;
oO
,f,
f..4
Z; f..4
(,0
qr'_ ¸
z
0..,
U
b.--
I-I-- I1: ne n J :)J
0
L)
f..4
l"
\ a.I,OI-ar ar o I, I a. L 0._ I.--
,.4
L)
I o'3
r..4 o
I--I
[..q
o
Zr_ n_ i,.uJ_ n 1
D_
o0o
_d
\
\
,.,,. z,,,. _
o,.. =) -J
I
I
t_
"0
u
0
_
"_
,r.-t
i1)
_._
FIGURE 3 - 8 .
T H E R M A L SWITCH
29
8O
I u_ I-,,::I 3 60 COMPAR
6l o
43 °
_z
123.5 °
_
ao
13"5°
6 °
4o =1.--0 SOLAR LUNAR NOON
i
30 INCIDENCE ANGLE,
60 11, DEGREES
go
FIGURE
3-9.
COMPARTMENT
DISSIPATING
CAPABILITY
3O
TABLE
3-2.
THERMAL
COMPARTMENT
COMPONENT
INSTALLATION
Compartment Receivers Transmitters Rf SPDT (Z) (Z)
A Central Boost Central auxiliary
Compartment command regulator signal processor
B decoder
switch
Signal
processing switch
Engineering Low data
signal rate
processor
Rf transfer Battery Battery
auxiliary and heater
Thermal charge regulator auxiliary temperature a s s embly Resistor,
control
thermal sensor
calibrated
Engineering Television Thermal assembly Resistor, temperature Meter Wiring shunt harnes
mechanisms auxiliary'_
Wiring control and heater
harness,
compartment surface sampler
B
Soil mechanicsauxilia thermal sensor calibrated Seismometer Seismometer Micrometeorite auxiliary':-" Micrometeorite electronics-':`" Alpha Alpha scattering scattering r y::`"
auxiliary::-" electronics::' detector
s, compartment
A
detector
auxiliary::-" electronics::-"
":-"Part
of
scientific
payload.
31
PRECEDING
PAGE P_,LANK NOT
FtL_,',__.
IV.
ENGINEERING
INSTRUMENTATION
Temperature ing of spaceframe sensors, of platinum 5. 0 ma provided contained a basic
and status sensor,
acceleration and and
sensors
are There
provided are two both
for
telemetric
monitor-
performance. a high The accuracy
types
of temperature are made with are are up a
sensor,
of which are
resistance current
wire. source constant
basic
temperature accuracy The
sensors
provided sensors
constant with
and
the high
temperature constant
a 2.5ma
current
source.
current
sources
in the ESP. all of
There which These have
are the
sixty-three capability are
temperature of being monitored
sensors while
included still on units
in the the as
basic
bus, pad.
launching follows:
sensors
distributed control s
among
the spacecraft
Flight M
units
7 sensors
3 sensors 6 sensors
e chani sm units
Radar
Electrical T ransmitte Approach Vehicle Propulsion Survey Accelerometers, loading of the points leg 3. and displacement TV
power rs TV structure units
units
3 sensors Z sensors 1 sensor Z5 sensors
15 sensors ] sensor
switches, during are the FCSG.
andpotentiometers the thrust, near the transit,
are and
installed landing
to measure Three
phases.
accelerometers and one on
installed
retro-rocket/Centaur amplifiers system are
attachment on landing +15 g peak. points.
Accelerometer accelerometer by and
installed
Full-scale of the
range
of the legs
is approximately at the
Position Full
landing of the
is measured lock struts
a potentiometer omni antennas
leg hinge by
extension
landing
is indicated
33
mechanism-actuated to inhibit
switches.
Lock strut l has an "omni antenna extend" switch
omni antenna extension until landing legs are fully deployed.
and descriptive A, items documents iI thru 13. for the engineering instrumentation
Definitive are listed
in Appendix
34
V.
PROPULSION
SUBSYSTEM
The maneuvering phases and
propulsion the
subsystem
components
supply
the
reaction and
forces lunar
for landing
spacecraft
during 5-i
the midcourse illustrates the
correction elements
of the mission. control
Figure
of the propulsion
flight The
subsystems. subsystem consists of three-hypergolic-fueled vector correction engine The for and variablelanding the
propulsion engines
thrust, phase
vernier maneuvering
for midcourse a solid force by and
velocity
and
propellant
retro-rocket maneuver.
supplying
principal subsystem commands
deceleration is controlled from earth,
during flight
the landing control through initiated
propulsion maneuver sensor s.
preprogrammed by flight control
maneuvers
signals.
VERNIER The midcourse rocket descent chamber fuel tanks, and the
ENGINES vernier maneuver engine to the burning, lunar engine system (figure -5-2) supplies the reaction control forces during during for retroterminal thrust
velocity and surface. and oxidizer valves
vector
correction, vector vernier and
attitude attitude system system helium operation, over using
velocity The
control
engine The feed
consists
of three
assemblies three
a feed tanks, for
system.
is composed tank, and propellant
of three lines,
a high-pressure arming, be throttled valve Engine valves
necessary thrust by
system can
deactivation. of approximately signals supplied by flight conoperate cool30 by
The
of each
engine
a range control
a bipropellant Appendix
throttling ]5, item Z7).
flight
control
(see
firing
is accomplished by the
solenoid-controlled, trol. Throttling from to clear
helium-actuated valves a are controlled
on/off
controlled while the
individually valves The oxide
on/off
valves engine
collectively ing (NzO The jacket
single out
signal.
Vent
permit oxidizer (NO)
purging
of the
decomposed by volume
gases. nitric
is nitrogen the
tetroxide freezing ). Fuel point. and
4) with fuel
I0 percent
to depress
is monomethyl ignite
hydrazine when
.n_onohydrate mixed in the
(72 MMH. thrust
28 H20 chamber.
oxidizer
hypergolically
35
-Z
SUN SENSOR SECONDARY
00PPLER ANDALTIMETER VELOCITY RADAR ANTENNAS DOPPLER VELOCITY RADAR ANTENNA +Y
+X
l
30 d_ ROLL F_" / / _'_,/ L_.J YAW _L
F--_
J _ vERNIER ENG,NE
JETS
/L'-'J
O_RO
LEG
2
_
_A_
_
t"_
\, __
_\
_j
PR,MARY
_
SENSO
R
\
_
_n
P,,_
",,\
-y
U
C'I _ERN,ER _
_ _G'_ _
30 d
RETRO-ROCKET
ENGINE
LEG
3 NOTE:
I ALTITUDE MARKING RADAR
I_FOR
FIGURE
BEAM ORIENTATION,SEE 8-4
C)
I
÷Z
FIGURE
5-I.
ELEMENTS CONTROL
OF
PROPULSION SUBSYSTEM
AND
FLIGHT
36
_,__=3____._= .........
8EU_/= VALV_ 3U ]
d
I
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_E
OItlD/ZER
I
,_0¢¢.
_WlCO/.O
I
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I_'
Ip,
--
I I ii II
II
) _2OPELLAN'I"
,_ _[ ,'"
_ _
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,, I
,_ T Z<_ .....
:
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I
.
,:
....
z _]
......... _1_
I I
_a r_,_orr_c
VA L VLr
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,r_
_ ,,
u _O,_LLANT ,, #N-#,¢/: WLV.=
......
_ ....... _-_2V ,_ _
II II
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LEGEND
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O
-...... ..... O,t FUEl. /HECHAN/CAL _ R_DP_DTI_W. EL_'CT'R/C V_ILY_ _
_ re,,
CI41_C1_ _4LV_
FIGUt_E
5-2.
VERNIER
PROPULSION SCHEMA TIC
S-YSTEM, DIAGRAM
FUNCTIONAL
37
The of the three is hinge engine control specific thrust ing
thrust
chambers legs
(figures
5-3
and
5-4)
are
located
near
the Number
hinge
points
landing
on the bottom on an
of the main
spaceframe. motor
1 engine to rotate spacecraft The
(swivel) about moment impulse of each
mounted in the
electro-mechanical x-y plane for
roll actuator The
an axis arm and engine The
spacecraft engine
roll control. 36 inches thrust. installed The on
of each
is approximately vary by with engine gages
in length. approximate each and engine feed
total impulse is monitored control
strain
mount-
bracket. the
thermal
design
of the vernier of the system
engine between preventing
system 100°F freezutiliz-
maintains during ing ing or
temperature periods, by
of all portions from launch
0/° and
nonthrust overheating
to touchdown, and
propellant controls
a combination and electrical are
of active and solar
passive
thermal
surface Other
coatings system
heating. isolated from the source pair spaceframe nor heat to sink. near
components
thermally acts
ensure Fuel each
that the and
spacecraft are and each
structure contained tanks and
as neither tanks an
a heat with
oxidizer Fuel
in three each
one
of tanks
engine. line
oxidizer
have
interconnecting The arrangement and some
propellant of the segments temperature the helium data for tank, DSIF tanks of the
manifold on the
system
to all tanks is illustrated
all engines. 5-5.
spaceframe lines are
in figure heated sensors
Tanks,
propellant from and
electrically Thermal line
to condition on all tanks,
the propellant all engines,
Z0 ° to the three
100°F.
propellant Fuel and
segments tanks
permit each
telemetering positive complete Helium
thermal expulsion
monitoring. which assure valves sources. into the lation Appendix deflate
oxidizer the central under units tank
contain to permit
bladders and dump current
around
standpipe zero-g activated gas
expulsion release and
propellantcohesion are squib-operated The thrust helium chambers. and 27.
conditions. by 9.5
ampere
pulsed to force
constant
stores Valves
under
pressure
the propellants system, is given reguin
permit of residual
release helium.
of helium Tank
to the capacity
of pressure, B, item
dumping
MAIN
RETRO-ROCKET The main retro-rocket of the (figure 5-6), which lunar performs landing nozzle. the major portion of
the deceleration solid-propellant
spacecraft
during
maneuver,
is a spherical,
unit with
a partially
submerged
38
FIGURE 5-3.
VERNIER ENGINE ASSEMBLY
FIGURE 5-4.
VERNIER THRUST CHAMBER
39
1
P I
FIGURE 5-5.
VERNJER PROPULSION TANKS AND SPACEFRAME
40
14
IS 1 : 01
FROM
:E OF RE1 TRO GINE.
FIGURE 5-6.
MAIN RETRO-ROCKET ENGINE
41
The leg hinges, the
unit is attached with nozzle explosive flange gas
at three nut provide pressure
points
on
the main
spaceframe ejection. altitude
near
the
landing clips radar. the release Com-
disconnects attachment ejects
for post-firing points the for the
Friction marking when retro source.
around
Retro-rocket retro firing
igniter sequence operate by
altitude
marking squibs
radar and
is initiated. from a pulsed
Retro-rocket 19-ampere system. device
ignition
explosive mands are The Eastern rocket tion and
nuts
constant-current
initiated
the flight safety has dual
control arming
retro-rocket Test Range)
and
(required
by
the Air for safe are firing and
Force the retroactuaBoth
firing provisions
single for
bridgewire local and of the
squibs remote squibs
igniter. remote and
In addition, indication electrical
arm
of inadvertent isolation exists
firing
included. and pyrogen
mechanical igniter
between
squib
initiator
in the The
safe
condition. with 1332 the ib sec propellant, pounds. The range B, illustrated engine in cross may section vary The from by figure 8000 to 5-7,
retro-rocket
weighs I0,000 impulse A pressure
approximately pounds over
thrust
temperature (see Appendix on
of 50 ° to 70 °F. item 27). surface sensors ignition. engine blankets, above the
required
total
is 50,000 strain gage
is installed during
the motor Three
case thermal before
for
telemetering installed
case for
information retro-rocket thermal on
firing.
are
monitoring The depending to maintain ignition. engine
nozzle design
temperature of the
control thermal spot"
retro-rocket insulating
is completely and surface
passive, coatings of
its own "cold
capacity, propellant
the Because
temperature through will be
17 ° F
at the time
of the thermal the only.
gradient
engine
and at the
the prelaunch three grain engine will be
temperature, points
17°Ftemperature The bulk temperature
reached
attachment above
of the
propellant
+50 ° F. The retro-rocket nozzle engine is of spherical overall and length. conventional The design The utilizing engine grain case a partiallyan existing nozzle steel
submerged PBAA has
to minimize propellant and
utilizes
composite-type a graphite with throat an
geometry.
The
a plastic
exit cone. phenolic
is of high-strength case at a low
insulated ture level
asbestos-filled burning.
to maintain
the
tempera-
during
42
FLIGHT
WEIGHT
CASE
THRUST
SKIRT
STRUT
PROPELLANT
• INSULATION
COLLAR
PYROGEN
NOZZLE
ATTACHMENT
BOLTS
INSULATION
SAFE ARM
AND DEVICE
NOZZLE
THROAT
ALTITUDE
NOZZLE
ASSEMBLY
FIGURE
5-7.
RETRO-ROCKET
ASSEMBLY
Definitive listed in Appendix
and A,
descriptive items
documents 17.
for the
Propulsion
Subsystem
are
14 thru
43
PP',ECEDING
PAGE
E_L,_._'NK NOT
FILh_ED.
Vl.
ELECTRICAL
POWER SUBSYSTEM
Electrical tion can silver be
power
is supplied respect Three by
by
a solar sun dc by
cell array command,
(solar and
panel) a sealed
whose primary dc
posi-
oriented main bus
with
to the Z9-volt the
zinc
battery. are provided
regulated power
buses
and
a ZZ-volt to the
unregulated Figure A decoder actuated. control ments command 6-1
electrical
subsystem
spacecraft.
is a block
diagram bus
of the power is provided in the event bus
subsystem. so that power to the central circuit for the command is flight
Z9-volt will
"essential"
not be
interrupted
that the
overload
trip bus
The
Z9-volt
flight control The Z9-volt power. overload.
provides bus may
a separate satisfies be disabled
electronics. for Z9-volt in the
nonessential This bus
all other
requireor by
regulated event of an
automatically
The regulated units
ZZ-volt power
bus such
provides as own
unregulated solenoids,
power and
to those actuators,
circuits and
not
requiring
switches, regulation.
to electronic
providing
their
SOLAR
PANEL
The
solar
panel
assembl_, array approximately via
installed of 79Z
at the top
of the mast, arranged The
consists on
of a
series/parallel-connected honeycomb periodically the sun and substrate adjusted, will The The remain solar panel
solar
cell modules feet in area.
a planar be of a few the
9 square
panel
can
commands,
to compensate to the source incident
for the apparent solar radiation
motion within and
perpendicular is the prime
degrees. lunar day.
panel
of power a minimum
during of 81
transit watts.
is capable
of providing
BATTERY
A
14-cell A, of the
series-connected, provides when energy
silver-zinc storage
rechargeable spacecraft. watt-hours output voltage
battery, The
located
in
compartment capacity
for the is3375 The
minimum rate
battery (see
fully charged B, item 26).
at a discharge of the 14-cell
of 0. 5 ampere
Appendix
45
IBATTIrRy
C
HARG£ COMMANDS MECHANISMS AU X I L IARY ENGINEERING _ SWITCHED SWITCHED CONSTANT CURRENT ?gv Z?v PULSES
SOLAR
1
J REGUL
ATOR
PANEL 11
I [ OCR
I
BYPASS
!!
I
J I I J J BOOST REGULATOR VOLTAGE BOOST PREREGULATOR _. SENS=NG VOLTAG E
9._ am_ AND _9 o_t 20 m,
h
UNREGULATED BUS 4"4,0 V 22.0 -4,SV
t
COMMAND (N&Q(.E/O¢SAOLE
PRESSURE SENSING ANO CMARGE LOG,C
___ _OLTAOE"O I I
I I I I I I I
kJS
_ESSURE AND VOLTAGE
MAIN BATTERY
VOLTAGE DROPPING CIRCUIT
OVER
LOAD
TRIP
CIRCUIT
CURRENT SENSING TRIP CIRCUIT
zg v NO_I_Ii|NTIkL BUS
L_
COMMANO REGULATOR ON/OFF ByPASS ON NON EIIEN_AL BUS OFF REGULATOR CONTROL SENSING vOLTAGE TRIP I
t_
LOW INPUT vOLTAGE SENSING TRIP cIRCUIT
CIRCUIT
I
I I I
COMMAND
ON/OFF
I
I_ nV FLIGHT BUS CONTROL I RE MOT1[ VOLTAGE SENSI_
I
CONTROL REGULATOR
l
FIGURE
6-1.
ELECTRICAL
POWER
SUBSYSTEM
BLOCK
DIAGRAM
46
battery battery mental load. loads
as
measured
on the will be
load
side
of the mating -4.5) volts from
electrical
connector and no
of the environload to full
receptacle conditions, The and battery is charged
22
(+4.0,
for all operating 40 ° to 125 ° F peak loads from and
including is intended by the
temperature primarily panel.
to handle
lunar
night
solar
BATTERY
CHARGE
REGULATOR
UNIT
The rated the The
battery basic output charge
charge bus
regulator
unit, charging
located logic
in compartment and voltage
A,
is incorpoto enable
in the
to provide of the
conversion
varying battery logic
voltage
solar
panel
to be the
used
to charge charge allA/SPP is to couple
the battery. regulator, stepping the solar battery motors. panel
regulator and power
unit contains switching
optimum for
charge The
circuit, of the with from
circuits
function battery power solar
optimum maximum the solar
charge power panel
regulator transfer. at varying
circuit The
to the accepts mum
optimum
charge
regulator to maxiat the battery be con-
voltages
corresponding to the battery regulator can
panel
power The
output. operation from earth. logic
It delivers
this power charge
terminal trolled by The functions a.
voltage. commands battery necessary Provide pressure 27 volts).
of the optimum
charge to:
circuit
provides
sensing,
logic,
and
control
of
automatic reaches
charging 65 ±3 psi
of the
battery the
until the battery battery voltage
manifold is below
(provided
b.
Automatically pressure drops
restore below
battery
charging
when
the battery
manifold
60 ±5 psi. power for charging until the battery and reaches a tapered 27. 3
c.
Accept volts. charge
all available At this point, from
the battery charging
is potential-limited impedance.
results
battery which
d.
Respond charging
to earth functions. damage control
commands
override
all automatic
battery
e.
Prevent power
to the
charging event
logic of an
or open
to any or
other
component battery.
in the
unit in the
shorted
47
BOOST REGULATOR UNIT The boost regulator power. unit accepts unregulated power and delivers regulated
The boost regulator
receives an input voltage of 17.0___._jtoZ volts dc and 7. 3
provides 29.0 _i percent volts to the regulated output. This voltage specification is met only at the terminals of the regulator. The voltage available at a unit connector is less than this value by the line drop through the connecting wiring plus an additional switch drop of up to 0.5 volt, when transistor switching is required.
The fied maximum regulated The overload current voltage. trip for circuit, the boost loads. the located in the boost and regulator unit, provides to all to drop drop to output of the boost regulator is 7.0 amperes at the speci-
overload protection
regulator an
undervoltage causes the
protection output
nonessential to 27.75 zero and ±0.25
regulated volts,
When
overload
voltage will
voltage for
to the nonessential I000 milliseconds. on at the
regulated This to turn
loads period
remain load
at 0 volts switches
20 to
allows
the
individual If an
in all equipment after for the Z0
time
off automatically. recovered, the
overload will
is still present again drop to zero
overload ms,
trip and
circuit so on
has
voltage When
to i000
in a cyclic ±0.25 the volts,
manner. the output
the battery trip same
potential will The
drops remove
below
approximately loads
17.00 from
overload in this by earth
circuit manner.
all nonessential of the overload
regulated can be
operation
trip
circuit
controlled
commands. and A, descriptive items documents Z3. for the electrical power subsystem are
Definitive listed in Appendix
18 thru
48
VII.
TELECOMMUNICATIONS
SUBSYSTEM
INTRODU The which and
CTION telecommunications command A subsystem and consists decoding, provides of three and interconnected signal and groups
provide transmission.
reception data
telemetry
processing reception. commands. A
link group provides provides circuits
r-f transmission logic circuits analog
command A signal sub
decoding processing carrier
group group
decoding
for all earth to digital analog,
commutation, for processing
conversion, digital and
and video
modulation
of most
data
channels.
DATA
LINK
GROUP
The two
data
link group
comprises and
two a high
transmitters, gain planar
two array
receivers antenna. two-way
(figure A trans-
7-i),
omnidirectional mode
antennas, during
ponder
is employed The
the transit
phase
to permit
doppler The
shift
measurements. spacecraft-to-earth mode) The and total
earth-to-spacecraft link is operated PCM-FM-PM, bandwidth of the
link is a PCM-FM-PM PCM-FM-PM or system direct during FM transit
system.
(transponder operation. of system command oper-
PCM-FM-FM, information There These are
during
lunar
is dependent that can be
on the mode selected by
operation. from ating earth.
four
modes and as
of operation their usable
modes, are
information
bandwidths
while
at lunar a. Mode
distance, A-High
follows: antenna bandwidth antenna bandwidth with transmitter kcps. in low-power mode; in high-power mode;
gain
nominal b. Mode nominal c. Mode mode; d. Mode mode; C-
information B -- High gain
is ZZ0 with
transmitter
information
is Z kcps. with transmitter cps. in low-power in high-power
Omnidirectional information
antenna bandwidth antenna bandwidth
nominal D-
is 1000 with
Omnidirectional information
transmitter
nominal
is i0 cps.
49
1
FIGURE 7- 1.
COMMAND RECEIVER AND TRANSMITTER
F i g u r e 7 - 2 is a block d i a g r a m of the telecommunications s u b s y s t e m T r a n s m i t t e r R e c e i v e r , and T r a n s p o n d e r Interconnections. Design redundancy i s employed in dual t r a n s m i t t i n g and receiving s y s t e m s (with the exception of omni antenna coverage redundancy) t o e n s u r e r e c e p t i o n of c o m m a n d s f r o m e a r t h and e n s u r e t h a t the d e s i r e d data w i l l b e t r a n s m i t t e d b a c k to earth. Two identical t r a n s m i t t e r s , located in c o m p a r t m e n t A , a r e provided; e a c h t r a n s m i t t e r i n c o r p o r a t e s switching t o p r o v i d e e i t h e r a high- o r low-power output. T h i s is accomplished by switching a traveling-wave tube ( T W T ) amplifier i n t o , o r out o f , the c i r c u i t . Both t r a n s m i t t e r s i n c o r p o r a t e p r o v i s i o n s f o r f r e q u e n c y moduE a c h t r a n s m i t t e r c o n s u m e s about 70 w a t t s in highTransmitter performance lation and phase modulation. p a r a m e t e r s a r e a s follows: Nominal output frequency = 2 2 9 5 m c ( S e e Appendix B , Item 1 3 ) Nominal output power
= 4 0 dbm ( 1 0 w a t t s ) i n high-power m o d e
power m o d e and about 7 watts in t h e low-power m o d e .
( S e e Appendix B , I t e m 14)
50
ANTENNA A OMNIDIRECTIONAL
RF SWIT_,_ 0"--"
DIPLEXER
t--_
m_c 5
_
LOW PASS
FILTERH
RECEIVER
A
ANTENNA B OMNIDIRECTIONAL
r
TRANSPONDERJ
A TRANSPONDER
J
HIGH GAIN PLANAR
E_NA
2295 mc
LTRANSM'TTER :
B I T
2295 mc
TRANSMITTER
_
l
A
i_
,L
I I
I
I
3-/-/
6
COMMAND" CENTRAL DER COMMAND OECOI_E R A SUBSYSTEM DECODER NO. I
I o ,¢ m t o
i_ I I I _
NO. NO.I 2
[
CENTRAL COMMAND DECODER B SUBSYSTEM DECODER NO. 2
NO. 32
SUBSYSTEM DECODERS I
ASSIGNMENT
BASIC BUS IsTHROUGH 7 J TV 8,10 THROUGH 291 ADDED AS NECESSARY
SUBCARRIER OSCILLATOR
,SPA
1
L_
DECODER SUBSYSTEM NO. 29
3.9
kcps
_,
I
I
17.2 bp$ "_ ATTENUATOR H 560 cp$
I
SCIENTIFIC SCO'S PAYLOAD AND CHANNELS DATA ENGINEERING COMMUTATORS SIGNAL MODE I-MIDCOURSE MANEUVER tO0 WORD FRAME MODE 2-RETRO DESCENT I00 MODE 50 WORD FR_tME 3-VERNIER DESCENT WORD FRAME TRANSIT PROCESSING MECHANISMS PROPULSION ELECTRICAL RADAR FLIGHT CONTROL APPROACH TV POWER LINK ESP
137.5 960 cp$
bp$
csp ]
l 39 kcps I 550 bp$ ANALOG TO 1_,-! =
I
DATA
,
I
D,GITAL I=,
T
l
i--1%°u°
l
I L
MODE 4-MISCELLANEOUS AND LUNAR OPERATION I00 WORD FRAtME
"L35
kcps
t
I100 bps
SCIENTIFIC PAYLOAD LMODE 4) 33 kcps
t
4400
bpl
,
TELEVISION AUXILIARY AMPLIFIER VIDEO B l FRAME TELEVISION COMMUTATOR v V_DE0 VIDEO i_ i_ CAMERA 2 CAMERA 3 VIDEO APPROACH TELEVISION CAMERA ID .._ .SURVEY TELEVISION CAMEFtAS I I
I
PHASE AND FREQUENCY SUMMING AMPLIFIERS
'hl-
I SCO _ l a SCATTERING (
SUBCARRIER OSCILLATOR .A % COMMAND ENABLE/REJECT
I
¢1DETECTOR l SCO :t1:2 - SCATTERING i SCO 7.35 KC
'
r
2.3
kcps
If
l l l
J
5.4
kcps
_
GYRO SPEEDS
I 52.5 L kcps
ACCELEROMETER
NO. I, LEG NO. I, V8
SEISMOMETER 2.3 KC
'
VIDEO A
_MPLIFIER
]
-t
I t AMPLIFIER PRESUMMING It_ 90
o,c0,
kcps IlOkcps _'-_
ACCELEROMETER
NO.2,LEG NO.2, V9
/ J
SCO SURFACE SAMPLER ELEVATION 200 KC
I
IkCCELEROMETER
NO.3,LEG NO.3, VIO
SURFACE SAMPLER ACCELEROMETER DATA
ACCELEROMETER qp_
NO.4, F/C SENSOR GROUP, Vii
t
-_.,_
__
FIGURE
7-2.
TELECOMMUNICATIONS
SUBSYSTEM
BLOCK
DIAGRAM
_-/-7_
5Z
Nominal Two of the Appendix are when signal
output
power
=
Z0
dbm
(0. 1 watt) located
in low-power in compartment is 2115
mode A, mc are (see type and mode, received of the two a part
identical
receiver/transponders, Their These and nominal signal are
command B, Item
link. 18).
input
frequency
receivers
of the
double-conversion During the transponder with
frequency enabled and
modulated by command,
crystal-controlled. one of the for an receivers accompanying function.
is phase-locked transmitter. Each receiver thus
the
provides can perform
excitation
Either
receivers connected redundant signal about
this transponder
is permanently two the received require
to one
transmitter
for transponder the omnis
operation, are intact. Together for lunar
providing
transponders
provided at a ratio
In this mode the
is retransmitted 2.82 Three watts
of Z40
to 221. power are
receivers
continuous
unregulated antennas for command
operation. on the spacecraft. Two
telecommunications antennas array
provided
are and power
omnidirectional the third for
reception capable
and
transponder sufficient switching two
operation, effective function is
is a planar
antenna
of radiating An antenna of the antenna The
real-time use
television
transmission. antennas. boom. to receiver
included antennas nected switched
to allow is mounted to receiver to use
of alternate on an
Each One B.
omnidirectional conbe of the
extendable other
is permanently can Each
A either
and the
the
transmitters antenna.
omnidirectional consists
or the
high-gain
omnidirectional a turnstile Gain
antennas cone,
of a turnstile following than see
half-wavelength characteristics: (composite B, Item
dipole
exciting
slotted
with
the
nominal -I0 db
Greater antennas,
pattern, 17)
both
Appendix
Polarization Impedance Frequency The high-gain array Gain 3-db beamwidth antenna, with the
Right-hand 50 ohms
circular
S-band installed following with the solar panel on top of the mast, is a
planar
characteristics: Approximately 8.0 degrees 27 db E-plane, circular (see 6.5 Appendix degrees B, Item 16)
H-plane
Polarization
Right-hand
53
Impedance Frequency
50 ohms S-band using surface coatings to
Thermal control of the antennas is passive, maintain temperatures
COMMAND The from mand, provides process mands handle the DECODING command spacecraft each
within acceptable limits.
GROUP
decoding receivers, command
group
accepts
earth-transmitted sync and timing and
command signals from
messages each comand can comto
generates for correct the control
checks output direct (which a total
address addressed on-off
command
complements, The and system
signals commands control of 3Z4
to command (which time-interval
sybsystems. operations) The
quantitative is mechanized with
operations). and
spacecraft commands
direct
commands
quantitative
a resolu-
tion of one The ment one
part
in 10Z4. decoding decoding selector, five subsystem unit, located The identical in compartment command central B, is the basic unit consists decoders 7-3 for the and in the digital One comeleof
command command
of the
group. two
decoding command Figure command
receiver-decoder and
redundancy, relationship subsystem form words or more mand of two are
command
decoders. central
illustrates decoders,
of the receiver-decoder command types decoders. The
selector, command serial type and used
link transmits binary digital
information The
of standard-length command are
words. command with
of the direct commands
the
quantitative
type.
direct
always function
in conjunction flight control
a quantitative that
to select
the appropriate time may be
in the
subsystem
requires
a selectable words during
interval
operation. at all times command maintains link when no commands frequent bit sync, are use. but does being The not
Fill-in transmitted, fill-in word select
transmitted when the which
periods
requires and
is a direct
command
word
a subsystem Complement
decoder. checks are made on the address (and reject bits and their complements, on command
and bits.
similarly, Failure The
in the
case
of direct generates bits of the are
commands a message quantitative
fill-in words), signal that are
to "check" magnitude these
is telemetered not complement execution.
to earth. checked;
commands
however,
telemetered
to earth
for verificationbefore
54
<
o 0
0
_=_
0 L)
<
_4
I
t--
I
I
55
A receivers command decoder the four receivers output command receiver central rupted
receiver-decoder and one
selector central
is provided command earth.
to select decoders Two each
one for
of the two use
spacecraft
of the two transmitted form
in processing receiverto one of Both the
messages selector
from
flip-flops state
in the
a four-state of two be
counter, and One to be signal
corresponding command signal
combinations will always
receivers
two
central
decoders. will permit central
operating. receiver other select
receiver processed will
select
of the
corresponding The
in the the
operating corresponding The
decoder. output command for a period
inhibit
at the
input
of the central remains 500 off.
command Whenever
decoder. command the
nonselected is interselector of receiver comof the
decoder
modulation
exceeding
milliseconds, of the four
receiver-decoder combinations
automatically and decoder.
switches Thus,
to another if the receiver
possible a new
in use To
fails, avoid
receiver-decoder switching be
bination desired therefore, to be
will be
automatically between
selected. commands, which normal decoder entire
unintentional must always command,
combination
modulation is a dummy
present;
a fill-in word, between command
in essence command (figure command messages. incoming delivered
is required
transmitted The central the
transmissions. 7-4), which synchronizes subsystem, and and provides timing are the
controls major
operation
of the
decoding
processing are
of the derived
command from the and
Synchronization messages. Control
information generated decoders
signals command
in a timed to control types
sequence their
to all subsystem
operation. words an interlock are provided - a direct Each command, which address, command, command complement time durations a
Three quantitative consists and sync
of command and
command, address,
command.
direct command controls an The
of an
address The
complement, command, an
information.
quantitative also no
of operations and which sync
in flight control, but has an
contains
address
complement, command, inadvertent comple-
information, precede of such
command or
complement. critical command, of an sync the
interlock the
must
irreversible It also
prevents address
execution ment,
commands. command command complements.
consists and compares
address, information. address and
command, The central
complement, decoder
command
bits to rejection
their and
respective initiation
Failure reject signal
to check transmitted
results via
in message telemetry
of a message
to earth.
If
56
FIGURE 7-4.
I
CENTRAL COMMAND DECODER
both a d d r e s s and command complement checks a r e s u c c e s s f u l , a command enable s i g n a l i s generated and t r a n s m i t t e d t o e a r t h . T h e a d d r e s s and command information b i t s a r e delivered t o a l l s u b s y s t e m d e c o d e r s w h e r e they a r e decoded under the c o n t r o l of the a d d r e s s enable s i g n a l and the c o m m a n d enable signal, both of which a r e generated i n the c e n t r a l c o m mand decoder. The a d d r e s s that is assigned to all quantitative c o m m a n d s i s decoded i n the c e n t r a l command decoder. T h e function of a s u b s y s t e m command decoder i s t o supply actuating o r c o m m a n d signals to chosen locations by deciphering the digital information supplied b y the c e n t r a l command decoder under the control of the c e n t r a l c o m m a n d decoder. Subsystem decoders a r e available with t h r e e s i z e s of
57
matrices
so that 8 {3 modules),
Z0 (4 modules),
or 32 (5 modules) possible
command outputs are available. The maximum number of subsystem command decoders that may be addressed is 29.
At scientific The a. b. c. d. e. f. g. h. i. j. k. 1. present, payload twelve Data Signal the basic uses are 5. assigned as follows: approach camera (No. 4). bus uses 7 subsystem command decoders and the
link and
television
processing. power. mechanisms. mechanisms auxiliary.
Electrical Vehicle and
Engineering Flight
Control. survey survey camera camera {No. {No. sampler. 2). 3).
Television Television
Soil mechanics Alpha Particle
surface Scattering detector.
Micrometeorite Seismometer interface to a matrix subsystem
Command simultaneously the selected
signals and
from OR
the
central
command to the signal
decoder same OR
are gate
fed of central
to an
gate.
Also, enable
decoder, When
the address are
is fed from is fed out
the
command gate
decoder.
all signals flip-flops
present, redundancy). checks
a set pulse If the are turns both on
of the
to a pair address enable selected
of parallel and
(for
central
command a comamplifier of the of
decoder mand of the command the matrix
command will be
complement generated. This
successful, the power The states
signal
signal
subsystem signals to go
decoder,
energizing determine
its matrix. which
interface is allowed
at this time high.
command
output
58
SIGNAL PROCESSING GROUP The signal processing group gathers the engineering and verification signals from various subsystems and provides the appropriate signal conditioning. The signal processing subsystem comprises the engineering signal processor, central signal processor, auxiliary.
The data being the from engineering the Surveyor
low data rate auxiliary,
and signal processing
signal
processor, and
located puts
on
compartment form
B,
processes to
spacecraft This bus
it in a suitable handles data).
preparatory
transmitted
to earth. of the basic signal
processor (engineering
all data
required
to assess
performance The
engineering four
processor two current and
is composed sources four
of the following (thermal
major
comone
ponents: command The modes plus two
commutators, and reject signal
measurements), channels. capable of 9._8 words of
enable
channel_ processor
accelerometer a commutator 4 consist
engineering
contains
of four of data 48 words
of operation. sync plus words Z sync of data the
Commutator (a total of i00 words plus
modes words).
i, Z, and
Commutator Each 7-1 word
3 consists consists a sample The
of data bits, format on the
(a total of 50 words). one parity bit. Table
of iI digital frame positions the
I0 bits with frame I00 The
contains
first four
words
of the frame and sync) are
illustrated. also
first two To
(sync words sampling primarily
complement of the message formats by
explained. the scope and
illustrate
entire
frame
is beyond modes
of this document. 3 have been critical and correction,
of commutator
i, Z. for the
established periods commutator mode mode other Z; and 3.
telemetry
requirements descent: using
following
during
transit
or terminal
midcourse high gain
maneuver antenna,
mode
i; terminal descent format
descent using
commutator commutator established by all
terminal sampling
omnidirectional mode
antennas, 4 has been
The
of commutator excluding can offivebit be those
telemetry Each
requirements commutators at any The one
previously at any 17.2, by the
listed. time by 550, command II00, or (but
of these
operated rates:
only 4400
one bits
at a time) per
137.5,
second. located only by
bit rate
is controlled signal
operating
analog-toof the bit rate of
digital rate each
converter
in the bandwidth
central
processor. time. 7-2.
Selection The
is limited commutator
available is listed
at a given in table
frame
for
each
bit rate
59
TABLE
7-1.
PORTION
OF
ESP
COMMUTATOR
DATA
FRAME
Word
No.
Mode
No.
Signal
Commutated
O0
1 2 3 4
Sync Sync Sync Sync Sync Sync Sync Sync
Complement Complement Complement Complement
0
Primary on mode Doppler word 3 Doppler words 4 1 2 Omni Vernier 51)
Sun
Sensor
Pitch Zl, 41,
Error 61 and
(also 81) on
occurs
I, words Velocity
V x (also
occurs
mode
2 of
Velocity 21, No. 31 and
V
X
(also
occurs
on mode
3 of
41) A Power
1 Transmitted lines No. 2 Temp
Doppler word 52)
Velocity
Vy
(also
occurs
on mode
2 of
Doppler Velocity Vy (also words 22, 32 and 42) 4 Transmitter Primary on mode Sun A Temperature Yaw 23, V
occurs
on
mode
3 of
Sensor
Error 43, 63 occurs
(also and 83) on
occurs
l of words Velocity
Doppler word 53)
Z
(also
mode
2 of
6O
TABLE
7-1.
PORTION
OF
ESP
COMMUTATOR
DATA
FRAME
(Cont)
Word
No.
Mode
No.
Signal
Commutated
3 (Cont)
3
Doppler words
Velocity 23, Phase Z3, 33 and Error 43,
V
Z
(also
occurs
on mode
3 of
43) A (also 83) ZZ, Star Pitch Manual Mode PrecesLockon Z3, occurs on mode 4
4
Static
of words 4 Rate Cruise sion ZZ6, Mode
63 and Sun Mode
Z01,
Mode,
Inertia ZI9,
Switch, B05,
Enable Nominal
Signal
Thrust and Temp
Bias Doppler (also
Zl8,
Altimeter R-ADVS mode 4
Reliable 4 Main
RORA,
Reliable on
Battery 42)
occurs
of word
Sync
complement
is 0
0
0
1
1
1 sent
0
1
1
0
1 on
(this is the the next
barker
code
complement word
that is read sync
out and
to earth
commutator
preceding
to further
establish
commutator
frame
synchronization) Sync digits is 1 least l l 0 0 0 1 0 0 l 0 (this is a calculated the barker series code, of and
likely start eleven
to occur of each
at random, commutator T301 start
called frame. thru
indicates states and
the
This and
defines is read frame.
the out
of the
digit times, the
T311,
sent
to earth
to indicate
of a commutator
There commutators; and one digital signals
are two
five different analog voltage are from
types
of inputs two
available
in the
engineering types,
types,
temperature inputs A
measurement
type. which
There vary one
89 high-level 0 to +5 volts. and
available
for processing inputs provide
analog sampling
number than
of these one word
on more The ±0.2
than basic
commutator
on more high-level This
in a given inputs does will not be
commutator. approximately include errors
accuracy
of these
commutator however,
percent
of full scale. signal
tolerance, such as,
that exist
in the
sources,
transducers,
voltage
61
TABLE 7-Z.
TIME REQUIRED FOR ONE FRAME OF COMMUTATED Time Required (Seconds)
DATA
Sampling Rate Bits per Second 17. Z 137.5 550 Ii00 4400
Commutator 1 Words per (100 words Second per frame) 1.5 12.5 50 i00 400 64 8 2 l 0.25
Commutator 2 (i00 words per frame) 64 8 2 l 0.25
Commutator 3 (50 words per frame) 3Z 4 1 0.5 0.125
Commutator 4 ( 100 words per frame) 64 8
1 0.25
dividers, processing applied output absolute scale when within inputs
etc.
There
are
ii low-level varying from with
differential 0 to +i00 a nominal
inputs
available These The
for inputs amplifier The of fullis ±i percent generated by system. these The drop are
analog
signals
millivolts. gain of 50.
to a differential is then processed
amplifier as an
ordinary processed
0 to +5-volt on data. by three Most of is such these The
high-level is ±Z
signal. percent
accuracy the
of the data use of the
inputs
without using the will
calibration defined processor. measurements shunt
absolute
accuracy points processed control
a calibration engineering consist value
curve signal
calibration of the data
of current of each
the power that the
resistance never
current
expected
voltage
exceeds
i00 millivolts.
There ments. These
are
64
inputs
available are
for processing obtained by
basic
temperature a constant as
measurecurrent of of
measurements to temperature but do not exceed
supplying vary
5 milliamperes temperature measurements temperature There measurements.
sensors 1000
which The
in impedance accuracy
a function of these
ohms. on the
expected
is _:4° C, sensors. are seven These
depending
calibration
accuracy
of the particular
inputs
available
for processing are obtained by
high-accuracy supplying
temperature
measurements
a constant
62
current of Z. 5 milliamperes to temperature sensors which vary: in impedance as a function of temperature but do not exceed 2-000 ohms. The expected accuracy
of these measurements digital inputs is approximately handle certain of two Such _l. 0 ° C. quantities to be transmitted that consist
The only device off. of an or
indication electrical
of which quantity.
states a signal
prevails
regarding as being
a mechanical either than open on or
is defined
If it is on, or is off,
it is defined a voltage -3
as having between and ÷i
a current +5 and +i0
of not greater volts (or an
3 microIf a milli-
amperes, signal ampere. engineering
being
circuit}. 0.2
it is between are
volts
and words
is capable
of accepting available
There
a total of four
digital
(40 inputs)
in the
signal and
processor. enable signals from the central _ngineering on when command signal needed decoder modulate This bandwidth frequency is present, equal the subis when it will to the a
Reject 2. 3-kc carrier available. no enable
subcarrier oscillator The or the
oscillator will be
in the
processor. and when
commanded oscillator is present.
subcarrier signal
will remain When the
at its 2. 3-kc reject signal
reject subcarrier enable
deviate period signal for
oscillator signal
to a higher 21
frequency
for a period When to a lower
of the
(approximately deviate the
milliseconds}. oscillator signal
enable
is present, equal
it will
subcarrier reject
frequency 21 milli-
a period
to the period
of the
(approximately
seconds). The bines the central outputs signal from processor engineering are (figure 7-5), located in compartment and various basic or B, com-
the signals
signal
processor sent
bus phase-
subsystems. modulation analog data
These inputs and,
processed The converting
and
to the has
frequencythe data capacity for
of the where
transmitter. desirable,
subsystem
for handling
it to digital digits, and
subsequent signals The for con-
transmission. trolling signal
Synchronizing are contains summing r.
patterns, generated two
parity
timing processor. three an
commutators processor six
in the
central
signal converters,
central
analog-to-digital power
subcarrier
oscillators, is olation The signals sive nished
amplifiers,
switches,
and
analog-to-digital
amplifie
analog-to-digital 10-digit binary
converters numbers. voltage
accept The with
d-c
analog
signals by
and
convert
these
into
converter increments With
functions
making
succesfur-
comparisons by a binary
of the voltage
analog
of a reference comparison,
voltage a logical
weighter
network.
each
63
1
FIGURE 7 -5.
CENTRAL SIGNAL PROCESSOR
decision is m a d e , and the i n c r e m e n t of r e f e r e n c e voltage i s allowed t o r e m a i n o r i s s u b t r a c t e d out. This p r o c e s s continues until the e r r o r i n the b i n a r y r e p r e s e n T h e output of the c o m p a r a t o r w i l l These signals
tation is l e s s than one p a r t i n 2 1 ° at full s c a l e .
be the s e r i a l b i n a r y r e p r e s e n t a t i o n of the voltage being digitized. and sync w o r d s in a readout a m p l i f i e r .
a r e s e n t t o the analog-to-digital timing circuit;-y t o b e combined with a p a r i t y b i t ,
T h e signal p r o c e s s i n g s u b s y s t e m contains two analog-to-digital f o r redundancy and output c i r c u i t r y c o m m o n t o both.
converters
E i t h e r analog- to-digital
c o n v e r t e r will operate when it r e c e i v e s power on s i g n a l s f r o m the s u b s y s t e m T h e power on s i g n a l s f r o m the subsystem d e c o d e r s , available upon c o m m a n d f r o m e a r t h w i l l t u r n on e l e c t r o n i c conversion units i n the s e l e c t e d A / D c o n v e r t e r . T h e s e supply voltages d e c o d e r s and bias voltages f r o m the E S P c o m m u t a t o r ECU.
64
are applied to master switches which enable commutated analog signals to one converter and inhibit signal inputs to the converter that is off. Subcarrier oscillators are modulated by the outputs of the analog-to-digital converters. Five subcarrier oscillator frequencies are provided, each associated with a different rate of data transmission. Center frequencies of these subcarrier oscillators are shown in figure 7-?.
Output signals from the analog-to-digital converter subcarrier oscillators are combined in the summing amplifiers and are sent to either of the two transmitters. The transmitters may then be either phase modulated or frequency modulated. This system uses individual "final" summing amplifiers to provide signals for the four different modes: (1) frequency modulate transmitter A, (2) frequency modulate transmitter B, (3) phase modulate transmitter phase modulate transmitter ]3. Two additional summing amplifiers
to "pre-sum" number complexity through deviation output Centaur the outputs of various other sources. amplifiers The gain This and presumming thereby of inputs to the final for summing reduces
A, and (4)
are provided the circuit
reduces the channel frequency
required summing
amplification. is such gate
of each
individual transmitter
circuitry A series
that the which
desired
is produced.
provides This gate
isolation provides
is located a path data
at the
of the data
analog-to-digital link subsystem
converters. for transmission
to the
of engineering
before
separation. The transmission converters. dividing These carrier iary for The onmi the low low data rate auxiliary, located in compartment clock rates B, provides for data
at lower Data rates
bit rates
than
the lowest 137 pulse 1/2 from
of the are
analog-to-digital obtained processor. Two rate subauxilby
of 17 3/16
and
bits per the
second
550-bit-per-second are available and
clock
central upon
signal
bit rates
to all commutators 960 cps, are provided
command. low data
oscillators, use with the
at 560 low
in the
data are
rates. used primarily with the low-power must transmitter be limited to and
low
data
rates
directional adequate signal
antenna, operating processing
when
the information
bandwidth
maintain The kc
margins. auxiliary located the the in compartment r-f carrier ratio A, provides a 3.9-
subcarrier
oscillator
which By
modulates increasing
at a phase
modulation power
index
of 0. 3 radianpeak.
of carrier
to sideband
65
I
the low DSIF, mode
modulation while at the
index same
enhances time
the probability the
of carrier
acquisition
by data.
the This
permitting off when processor index.
reception
of engineering index
of operation time the
is commanded central signal
a higher can Power
modulation a
is desired,
at which oscillator iary
provide for the bus.
similar
subcarrier auxil-
at a higher from and
modulation the Z9-volt
signal
processing
is obtained Definitive
nonessential documents Z4 thru 34.
descriptive A items
for the
telecommunications
subsystem
are
listed
in Appendix
66
VIII.
FLIGHT
CONTROL
SUBSYSTEM
The during from down
flight control phase of the
subsystem of the
controls
the
spacecraft This
velocity covers
and
attitude period touch-
the transit separation on the lunar include:
Surveyor from basic
mission. the Centaur
phase
the
spacecraft The
vehicle
to spacecraft by the during radio for
surface.
functions and
performed orientation based descent on
flight control the entire data,
subsystem transit and
(I) attitude
stabilization
phase,
(2) midcourse phase upright retro
trajectory maneuver on
correction and vernier
command
(3) terminal in an
landing
of the forms control-of
spacecraft reference-are used.
position
the lunar
surface. and
Three descent
principal radar subsystem.
celestial Figure 8-1
reference, is a block
inertial diagram
reference, of the
flight
control
FLIGHT
CONTROL
SENSOR
GROUP
The sensors frame
flight and
control
sensor control
group
is made The
up
of a group
of optical is mounted
and on
inertial space-
the flight legs
electronics. the hinge
assembly of leg a three 3.
the
between The
1 and reference
3 near unit
point
inertial
(IRU) the
provides spacecraft
axis
rotational The gyros unit with
reference consists associated Each spaceSpacecraft at a of
and three
an
acceleration orthogonally control
reference mounted, circuitry with a
'along strapped and single rotation
roll axis. rate
down
integrating force balance
temperature IRG craft comprises rotation with rate pulses gyro
a linear degree angle.
accelerometer. which B, integrates item 9.)
a gyro rate
of freedom, (See Appendix by
to obtain
rotation precise control to track during mits
respect for
to inertial period
reference of time.
is obtained The flight
"torquing"
a gyro
a precise
control
electronics the
produce spacecraft
causing precession.
the gas The
jet attitude accelerometer correction
control
system
to rotate spacecraft descent
measures and lunar
acceleration and 26-volt trans-
midcourse the information power
velocity
vector
phases
to the flight control the gyro motors.
electronics. The gyro
Three-phase signal
400-cycle
operates
generators,
67
I
I I
0
_J
0 ,-1
1 '' IIII .,.I
8
,-1 0
I
I
0 c_
I
I
0
I.--I
S
I I
..4
I O0
kl
I I
_IZ
1'
w _
I I H
I I
I
I
68
temperature 2Z-volt
controllers, 29-volt dc
andaccelerometer power, respectively.
utilize Gyro
single-phase motor speed
10-volt signals
400-cycle, are
dc, and
telemetered
to earth.
Temperature restricted Internal range gyro
of the three in the vicinity
gyros
must
be
maintained
within
a closely
of 180 ° F and
for all spacecraft circuitry the and
flight attitudes. heaters are employed perthe gyros the
temperature the specified inertial
sensing
control To
to maintain turbations, rest and of the
temperatures. reference However, are
reduce
effect
of all external from heat, the has
the
unit is largely to dissipate to the IRU
thermally
decoupled generated The
spacecraft. circuitry
internally radiator. heat and within of the sun
associated
coupled
radiator
capability unit, sun
of radiating the gyro
the
total internal as well
electrical as direct
dissipation reflected the unit. spacecraft sensor also
of the IRU energy from the
excluding and
heaters,
spacecraft primary
to maintain sun line sensor during
a thermal detects coast
balance deviation
The the
roll axis supplies sun
from a signal sensor
sun-spacecraft the Canopus The
phases. when the
The sun
to open field
sensor primary
shutter sun
falls within of one lockon
a defined and four
of view.
sensor
consists The
directional for to
cadmium-sulphide transfer of control sun
photoconductive of the spacecraft The
cells.
lockon from the
cell provides secondary signals to within sun
a signal sensor
attitude
the primary control degree
sensor.
directional the spacecraft
cells
supply
enabling less than
the flight 0.Z +0.3
subsystem limit Thermal cycle
to control deadband,
roll axis
of the
spacecraft-sun results base from and the
line. combined outer the effect case support of (a) conductreatand the
control the case
of this unit and the
tion between ment, primary the the which sun
support
(b) the between
surface base
minimizes sensor
temperature case. cells The
gradients
surface
treatment to optimize
of the heat
inner
cavities
surrounding and
photoconductive cells. The inertia throw
is designed
flow
between
the housing
switch, switch,
a
spring-restrained at a nominal level
mass 3.5 g
that operates as retro-rocket decay
a single thrust decays.
pole This by
single g level
closes
corresponds
to a thrust
on a constant signal end
curve. at the
Therefore, instant the made.
timing
a command-to-initiate switch opens) a safe
retro-eject prediction of the
(initiated
inertia
of retro-burning
can
be
69
The
telemetering
accelerometer during
is an descent.
engineering The
instrumentation is a
sensor spring
that measures restrained, connected The This
acceleration seismic mass 29 volts sensor
accelerometer arm of a linear
that drives dc. detects, a fixed with
the pickoff
potentiometer
across Canopus
identifies, spacecraft the primary
and
locks
on
to the
star
Canopus. Canopus sensor sensor
function
establishes In combination required
roll attitude sun sensor,
relative the The
to the
line of sight. establishes uses
Canopus Canopus
the
celestial tube
three-axis B,
reference. item
a photomultiplier
(See Appendix in the 8-degree error
II) with
suitable
elements of
to
(i) establish Canopus the star
the presence,
nominal
field of view, to the angle star The systems power. The intensity Canopus
of a star of the as an
brightness; of view;
(Z) provide and
signals and
related
star aid
on to
roll field map-making
(3) measure Canopus and
telemeter
for positive electrical 400-cycle
identification. control dc
sensor from
contains 3-phase, to supply ment the
internal 26-volt,
mechanical and
operating unit
power
Z9-volt control cycle are
is designed Canopus align-
information +0.3 degree
enabling
the flight limit case
subsystem For so as
to hold thermal to have as well
to 0.2 external
including of the sensor
deadband. finished
control, the as capability the direct
surfaces
of continuously and reflected The decoder, the
radiating solar energy.
the internal
electrical
dissipation
flight and
control
electronics electrical that the
consists conversion
of control unit
circuits, The are
the programmerlocation not of this to of on
the ac/dc is such irradiation finished as well
circuits. surfaces The
spacecraft solar are
primary
radiating transit
subjected surfaces any
direct
in a normal to obtain as the
attitude.
external radiating
the case solar
capability
of continuously during items
reflected causing With level no
energy,
power control the unit period
dissipated of other is capable
coast within
phase, the
without
perturbations additional
to the thermal thermal control,
spacecraft. phase
of radiating
the thrust
of dissipation The control maneuver. flight control circuits for
for a maximum
of I0 minutes. guidance the signals and process or them for
flight control
circuits
accept
of propulsion The
systems
to achieve are The
desired on five include circuits
stability circuit
controlled located in the
control
circuits unit.
installed circuits (Z) logic
boards mode signal
electronics control mode
(I) control for input
switching processing;
selection;
7O
and
(3)
analog
circuits system.
for
converting
sensor
outputs
to
commands
for
the
propulsion The signals and power for
flight initiating
control and
programmer controlling units for of the
is
a
digital
unit within
that the
provides flight the
yes/no control programmer lockon
output electronic s
sequences spacecraft. control, provides and sun a timed generates maneuvers control from unit, the sensors, predetermined to generate sequence. serve buffering involving counter a dual between timed which time bus. the the
management the signals
Basically, lockon, and
generates to the time certain main delays Most commands system classified switches, signals, flight the
attitude
star of
according signals during variable correction.
command retro-staging for of or decoder as the when the
sequences; sequence; attitude to the
sequence precision and midcourse
output
radio-command velocity are transmitted programmer Other radar, switch. enable outputs for inputs, the
controlling inputs
flight
programmer flight respectively. altitude and marking the inertia sequences, specified control
magnitude or special spacecraft they arrive the
information central inputs, are separation in certain logic retro staging decoder from
sub-
legs-down These the input advanced
control flight Seven
programmer of latch and
controlling
operation output
amplifiers interface
purpose the delays provides delays.
of
providing
mode and by of timing: power is flight a 10-bit
memory control magnitude auto taken
capability electronics. shift time the power voltages circuits, sun electronics,
programmer are two performed types
Operations register/scalar delays dc and flight auto
external from The
internal
Electrical
Z9-volt supply required inertial sensors. and
control
electrical for
conversion operation unit, of the
units flight
develop control sensor, units the
and
control
the
d-c
and control
a-c
programmer-decoder and are primary installed dc and and in ZZ-volt the flight dc
reference The obtain electrical input
Canopus conversion
secondary control buses.
power
from
Z9-volt
SECONDARY The signals transit to and
SUN secondary enable
SENSOR sun gross sensor alignment of the solar effects of the the initial sun roll !unar detection axis operation. to and the supplies sunline The during
spacecraft during
positioning
panel
secondary
71
sun
sensor, on cell. of the its axis
an assembly the The solar solar axis panel
of five cadmium-sulphide Four of the cells sun are
photo-conductive directional is aligned cells with
cells, and one
is is a
mounted lockon axis and
panel. of the cells.
secondary The solar
sensor
the
principal position the view Centaur a
panel
is erected after
to the transit from sensor
aligned
to the
spacecraft
roll axis cells of the
separation sun
vehicle. complete maneuver Signals maneuvers primary sun
Since
the four
directional with each
secondary quadrant,
hemisphere, in any from the
cell viewing the sun
one
a yaw or
{pitch) more and cells. pitch of the when The which and operation. of to is the
direction secondary the
will bring sun sensor
into view
of one
produce
successive to within to the sun the
yaw
to orient sun sensor.
spacecraft
roll axis
field of view sun lockon sensor cell.
Control
is transferred of the from electrical electronics for output primary
primary
is within sun
the field of view sensor operates
sensor
secondary taken from from the
approximately conversion conversion from
5 volts unit unit
dc power, transit lunar
the flight control signal diodes processing are
during during
Protective "other earth.
provided The
operating levels
either
source are
in case telemetered
source"
failure.
of all five cells
Thermal conduction of the during case
control the
of this unit is obtained sensor case and the
from support
the
combined
effects surface -I00
of treatment
between
structure; between day.
is expected and-125
to maintain and +235 °F
the unit temperature during the lunar
and
+I60°F
transit
RADAR
S The altitude-marking from flight. measuring of in flange, When 52 the to 60 the The lunar radar surface generates to initiate radar a single Appendix nozzle washers ignition to is input eject powered commands, and between begins, the the output B, is an the is "altitude terminal a mark" descent signal phase pulse-type can The by be p._r_eset for at a preset of the
slant spacecraft
range
altitude-marking radar miles with {See
single-package sSgnal that
fixed-range slant radar the flange. develops The away ranges mounts nozzle
item retained
20).
altitude-marking clasps radar by the from through and ignitor the nozzle. a breaksignal, and around the
retro-rocket spring
friction
with retro-rocket
altitude-marking gas generated radar volts dc bus
sufficient
pressure radar also carries
altitude-marking from the the Z2
altitude-marking plug that
output
altitude-mark
72
telemetry 9. 3 kmc. The point
information. Figure thermal mission isolation 8-2
The
altitude-marking diagram for
radar
operates
at a frequency radar. only
of
is a block
of the
altitude-marking it operable mark
control
designed
this unit makes the 60-mile
to that Use
of the
in which of the
it generates
trigger
signal.
of thermal active
altitude-marking survival phase and
radar
electronics, of the
combined required
with operating radar
heating,
ensures During
transit the transit
attainment mission, 20°F by
temperature. components controlled are
of the
all altitude means
marking
maintained up to the
at approximately time of required
of thermostatically
heaters
operation.
The range lunar down. separate and
radar
altimeter
and
doppler
velocity of the from
sensor
(KADVS) with
measures respect
slant to the touchtwo
orthogonal during
three-axis the descent common
velocity phase
spacecraft
surface
retro-rocket and components
burnout
to near
the RADVS functions, four
utilizes altitude
circuitry and
to perform velocity mounted
determination (I) an i hinge r-f point;
three-axis (KPSM)
determination. on the omni
It comprises antenna beams velocity legs leg
assemblies: near leg under beams signal I and
section
1 structure I and
(2) the altimeter/velocity between under legs 1 and
antenna, 2; (3) the B between above the
4, mounted antenna, (4) the legs
compartmentA 2 and data 3. 3, mounted
sensing 3; and
compartment on the
I and 3 hinge,
converter The antenna
mounted
spaceframe, contain
between
assemblies
also
ANTENNA MARK SIGNAL_
PROCESSOR ENABLE SIGNAL,,_-o
v°°FlH H H H
SYNCHRONIZER TRANSMITTER AMPLIFIER AGC VIDEO O ETECTO R MIXER
CIRCULATOR
LOCAL OSCILLATOR
TR TUBE
FIGURE
8-2.
ALTITUDE
MARKING
RADAR
BLOCK
DIAGRAM
73
r e c e i v e r microwave components, c r y s t a l m i x e r , a n d p r e a m p l i f i e r f o r e a c h b e a m , a s i l l u s t r a t e d i n figure 8-3. The r - f section contains t h e two k l y s t r o n s a n d the a l t i m e t e r k l y s t r o n m o d u l a t o r . The a l t i m e t e r k l y s t r o n o p e r a t e s at a m e a n f r e quency of 12.9 kmc w h e r e a s t h e velocity k l y s t r o n o p e r a t e s a t 13.3 k m c . A waveguide connects the antenna a s s e m b l i e s to the r-f section. F i g u r e 8-4 i l l u s t r a t e s the o r i e n t a t i o n of the f o u r r a d a r b e a m s with r e s p e c t t o the s p a c e c r a f t coordinate s y s t e m . F i g u r e 8-5 i s a functional block d i a g r a m of t h e RADVS. R a d a r beam 4 i s p a r a l l e l to the s p a c e c r a f t Z a x i s a n d develops a l t i t u d e o r range information. B e a m s 1 , 2 , a n d 3 develop t h e t h r e e - a x i s velocity i n f o r m a t i o n with r e s p e c t to s p a c e c r a f t c o o r d i n a t e s . B e a m s 1, 2 , a n d 3 a r e o r i e n t e d about 2 5 d e g r e e s divergent to t h e s p a c e c r a f t t Z a x i s a n d p a s s through t h r e e c o r n e r s of a s q u a r e p a r a l l e l to t h e X - Y plane. Beams 1 a n d 2 f o r m a plane p a r a l l e l t o t h e
X a x i s . B e a m s 2 a n d 3 f o r m a plane p a r a l l e l t o the Y a x i s . Slant r a n g e along the Z a x i s i s d e t e r m i n e d in beam 4, whose t r a n s m i t t e d r - f i s swept i n f r e q u e n c y .
FIGURE 8-3.
74
RADVS ASSEMBLY WITH PREAMPLIFIER COVER REMOVED
L
KAD'q5
BEAM
OF_IIsNTATIOIq
_,_GUg._
8,_.
VELOCITY
SENSOR
ANTENNA
[
--
--
--
SIGNAL
DATA
CONVERTER
3_
_
TRACKER DOPPLER
NO 3
CONVERTER
l
'
v
__ 'I"
RF.L,IMILE OI_RATE
PREAMPLIFIERS
CRYSTAl.
R|,IAIILE OOPI_LI[R vELOCITY S(NSOA I35 mc
OCqERATI[ DOPP_(II III[LIAJL $(_SGm V(LOCITY Ir 00_ KATIE
_
( DI'03"D PROGRAMMER _
4 )
_&OA_ tu0( Q(LtAELI[ ALtl O_ IIIATE ( Ol" 021 ÷101' O31 ÷1D2"D31
RADAR ALTITUD£ SENS_t 12.g IIme _
RE L IAll_.l[ OI_KATI[ DOn't. £ II VE_.OCIT Y CON0_TIONAt. S[_SGm
J N0, I
VI CONVERTER
vI
so mv/tDs
:1
I RANGE _ IOOO FT.
MARK
_
14
FT,
t
v NO. 4 RADAR ALTITUDE RELIABLE OPERAT F
cO?:._,R/
?.RAN,,
VELOCITY
SENSOR
ANTENNA
FIGURE
8-5.
KADVS
BLOCK
DIAGRAM
76
A
sample
of the transmitted frequency, and their
frequency difference
in each extracted to the difference spacecraft
beam by
is mixed a crystal trip"
with
the
received The beam
energy
detector. time and
4 difference to range are
frequency the beam to the
is proportionate 4 path. velocity The of the
r-f
"round
therefore 3
along
frequencies along their
of beams respective
I, Z, and beam
proportional
paths.
Range the doppler along
is determined frequency the
by
compensating to produce Velocity
the beam an
4 difference signal by pair
frequency corresponding
for to
shift errors Z axis.
"audio"
range sensed beams signal along
spacecraft frequency
is determined of each
summing
the
doppler with data each
shift of the Three
reflections
of constant-frequency solved in the velocity axial or
divergent converter spacecraft velocity
paths.
simultaneous "audio" of summing each Y axes,
equations
are
to produce axis. The
a doppler method By
corresponding determines of velocity
to the whether beams,
transverse I-3, 2-3,
is computed. along the X,
using
pair
i.e. I-2,
the and B,
velocity altitude item 21.
Z, and and
respectively, for RADVS are
is determined. given in
Velocity Appendix
capabilities
accuracies
Thermal thermal completion from the capacity
control
of the RADVS acceptable operation. dc bus.
is passive, temperatures Power
relying from
on
surface the
coating to
and
to ensure
beginning
of its required unregulated
is supplied
to theRADVS
system
22-volt
ROLL
ACTUATOR
The and an
roll
actuator position vernier the range
consists transducer. engine
of
a two-phase This assembly
induction provides by parallel
motor,
a
gear
train,
angular during 1 over
roll-control swiveling to both vernier the Z-axis signals 26-vrms Position 10-vrms
moments engir_e and from 400-cps feedback 400-cps
thrust
phase
operation
of _+5_.5degrees engines
in a plane 2 and
to a line through the flight power control with
vernier
3 in response motor phase
to electrical operates from
subsystem. phase fixed
The and
actuator the
one
control
variable. from
is provided power.
by an
induction
potentiometer
operating
77
ATTITUDE
JETS
The attitude jets supply the reaction forces for spacecraft orientation maneuvers during the period from Centaur-Surveyor separation through preretrorocket firing. The attitude jet system consists of a spherical tank containing approximately 4.5 pounds of nitrogen under high pressure; regulating, and dumping valves for gas supply control; and three pairs of opposed gas jets with solenoid operated valves for each jet. (See Appendix B, item 12.) The gas jet pairs are installed at the ends of the three landing legs shown in figure 3-2. Number one jet pair lies in the X-Y or horizontal plane for roll maneuvers. Jet pairs 2 and 3 are approximately parallel to the vertical or Z axis. Cumulative thrust produces pitch rotation, and differential thrust produces yaw rotation. Execution of specific maneuver commands is accomplished by actuation of the required solenoid valves to release nitrogen gas to the designated nozzles. The six solenoids are connected to solid-state switches in the flight control electronics. Each gas__jet supplies a thrust of approximately 0.057 pound at a radius of 70 inches from the center of gravity for angular acceleration. The moment capability of the attitude jet system about each axis is: Roll axis
Pitch Yaw The measured The and its own axis axis of the nitrogen to earth. of this system is passive, the utilizing surface gas coatings supply and tank and +4.0 +4.25 +_7.0 the in-lb in-lb
/
in-lb 2 attitude jet are
temperature for tele,netry thermal thermal within and
number
control capacity
to maintain limits
temperatures
of the
the jet valves Definitive listed
acceptable descriptive items 35
throughout for the
the mission. flight control subsystem are
documents thru 44.
in AppendixA,
78
IX.
APPROACH
TELEVISION
SUBSYSTEM
The over During the
approach range from
television i000 to 80
subsystem _+20 miles
provides {slant
pictures range) above up
of the lunar the lunar to i00 earth.
surface surface.
the approach frames television camera
interval will be
covering
this altitude
range,
individual The approach
television approach television cables figure relation The craft and 9-1.
taken
in response is composed
to commands of the
from
subsystem and
downward-looking with
the television The approach
auxiliary television
unit, together camera
the appropriate in and 2-2. to the spaceA at all be its
connectors. The physical
is illustrated camera in figure relative by
location items
of the approach and structures field of view
television is shown is fixed
to other approach
spacecraft television system lens than 300 over and
camera
coordinate camera greater
is approximately which maximum
6.4 degrees will provide aperture The
6.4 degrees. pictures The
fixed-focus altitudes preset view
is employed feet at the the range
in-focus of f:4. center
iris may field of at the +0.5
before
launch
of f:4 to f:22. may be
of the
of the approach assembly block
television to be diagram approach heater elements
subsystem parallel
manually Z
adjusted axis within
time degree.
of spacecraft An in figure switch for
to the
spacecraft
overall 9-2.
of the approach camera blanket itself are
television {designated
subsystem TV camera
is shown 4) and the
Only
the
its electrical All and/or other
unique are
to the approach with the
television survey television camera and
subsystem. subsystem itself command point surface
illustrated
shared The only
other
spacecraft
subsystems. requiring Upon the command camera signals Such
approach
television power
is completely signals in the lunar for
self-contained, operation. phase,
primary earth
electrical
from
at the appropriate images of the lunar
approach
will convert that include
into complete and
composite-video blanking signals. the
all necessary received the from the
synchronization central logic, command and control
earth on and
commands, off, initiate
decoder the
turn
camera
picture
taking
camera
thermal
power.
79
,ELECTRICAL
CONNECTOR
VIDICON T B, UE
iLECTRONICS
Lfliwtna
LCND
~UJJSTMENT
FIGURE 9-1.
SURVEYOR TELEVISION APPROACH CAMERA
E l e c t r i c a l power f o r c a m e r a operation a n d f o r the a c t i v e t h e r m a l c o n t r o l i s delivered to the c a m e r a via the s p a c e c r a f t h a r n e s s f r o m the b a s i c bus c e n t r a l power control a n d distribution s y s t e m . T h i s e l e c t r i c a l power c o n s i s t s of t 2 9 vdc regulated voltage and t 2 2 vdc unregulated voltage. The composite s i n g l e - f r a m e video output f r o m the c a m e r a i s s e n t t o the t e l e v i s i o n auxiliary unit, p a s s e d through a s u m m i n g a m p l i f i e r , and then s e n t to the s p a c e c r a f t t r a n s m i t t e r . Two identical video outputs f r o m the c a m e r a a r e
conducted through s e p a r a t e c a b l e s to individual s u m m i n g a m p l i f i e r s i n the TV a u x i l i a r y f r o m which the signals a r e f e d through individual c a b l e s to e a c h s p a c e c r a f t t r a n s m i t t e r channel. C a m e r a condition i s o b s e r v e d by analog s i g n a l s which T h e r e they a r e
a r e sent f r o m the c a m e r a to t h e engineering s i g n a l p r o c e s s o r .
80
0
E_
0
0
!
o_
I
Z_
I--I
1
commutated in sequence mitter
and with
sent other picture
to the
central
signal data
processor
to be
digitized via the
and
combined transsent in
engineering sequences. consist
for transmission data outputs camera
spacecraft camera,
between
Analog
from
the
this manner, indication s.
typically
of temperature
and
operational
status
Closed-loop maintain camera permit before the
thermal
control
of the
camera within
vidicon appropriate circuitry and
faceplate operating have been
is provided limits designed
to during to
temperature The
of the vidicon camera and
operation. initial warmup actual The camera
associated electronics
of the camera operation. television of the camera camera
the camera
vidicon
faceplate
approach
provides is adjustable with
600
line-per-frame before launch
slow to cover
scan the as low as
operation. range I00 from
Sensitivity 800 to 3000
foot-lamberts,
acceptable
results
at levels
foot-lamberts. The approach camera is located and mounted on the spacecraft chart in a position and collimalaunch of the
that will allow tion lens vehicle approach constraints. surface reaching sequence. Figure and timing that at the
limited may be
observation provided of JPL.
of an the
illuminated spacecraft
calibration shroud
within This
of the
Centaur testing and or any point
option
arrangement assessment the in-shroud within the
allows limited
prelaunch by lighting chart
camera,
with
a resulting of either shutter
focus the lunar
Observation by a
calibration
is controlled the vidicon The
camera
that prevents
light from in the to sunlight.
faceplate also
until it is opened protects the vidlcon
at the appropriate from sequence direct
shutter
exposure by
9-3
illustrates of the
the picture-taking camera that may are both be
controlled upon and which upon command
the logic from earth. to camera
circuitry and
initiated running
Horizontal reduce
vertical circuit and
generators complexity. pulses, The
free
unsynchronized produce application the
camera
These
generators,
synchronizing power and run
blanking
start automatically nominal g00-millisecond for the the
of camera pulses When camera a
continuously. timing
vertical
blanking
represent start video frame logic
the principal command circuitry initiates a nominal
waveform from The
picture-taking command blanking triggers vidicon.
sequence. decoder, pulse the
is received is enabled. a picture-taking g0-millisecond
central
first vertical sequence exposure and
to appear shutter
thereafter to provide
the
camera
of the
82
2OO
rnsec-e_
1
I"_j-
I sec
O) VERTICAL PULSES
BLANKING
ii
II II
rl
n .
I
I
I
N+I
b)
"TAKE PICTURE" COMMAND
!111
II II
IJ'
VERTICAL
N
[1 I
I
c)
COMPOSITE OUTPUT
VtDEO
BLANK,NG AND/_'1
SYNC PULSES
HOR,ZONTAL / I
I
'1
PICTURE READOUT HORIZONTAL
I _ IIII [--AND SYNC
I
d)
OPEN
SHUTTER
20
rnsec
I n
II
[I
IJ
I _"!,. e) FUNCTIONAL SEQUENCE 36 sec =-I
_,'NcI SYNCl S,'NC SYNCI [ I II II I
f) VIDICON CONTROL THERMAL INHIBIT
SYNC
J 'NHIB'T iI
OFF OFF
I
,NH,B,T
FIGURE
9-3.
APPROACH
TELEVISION
PICTURE-TAKING
SEQUENCE
The vertical permits blanking picture the
first blanking
data
transmitted with display the
during
a picture-taking horizontal on earth to be
sequence synchronizing synchronized before
consists pulses. with receipt
of the This
pulse and with
superimposed monitors
receiving pulse video. and
a vertical of
horizontal picture
synchronizing readout takes
pulses place
Single
frame pulse.
in the interval video start closes the
following
vertical
blanking
Transmission and The
of the composite is terminated blanking and at the pulse resets
information of the next the video
requires (second) gate which
approximately vertical feeds blanking video
1.2 seconds pulse.
second
to the television
auxiliary
shutter.
83
Three subsequent after the
frame start
periods frame
are
required must
for proper not be
vidicon
erasure
so that the until three
command pulse
received
at the
spacecraft permits
third frames
blanking to occur An
of the
start frame being
sequence.
This
erasure
(picture sequence command the
readout
the first one) is possible second take erase a picture
in approximately in that scan and the camera
3.6 seconds. can receive being
alternate frame
of operation during the will
a start the
(picture respond erased. This picture
readout normally. alternate every
first), and the provide
TV
camera image
In this case method seconds. defining can
previous
will not be
completely
capability
for transmission
of a degraded
2.4 The
document
for the
television
subsystem
is listed
in Appendix
A
as
item
45.
84
X.
SCIENTIFIC
PAYLOAD
INTRODUCTION
The date various on
basic
design
policies experiment weight, Z,
involved subsystem center
in providing
the
capability and
to accommoimpli-
payload spacecraft
combinations and
the detailed harness
cations are craft
of gravity, ll. This
electrical defines
design, space-
described provisions
in Sections for
3 and
section
the other and
scientific
experiment with emphasis
subsystems on
in detail
describes and
the experiments operation. An those
themselves,
electrical
interfaces
experiment by
subsystem the basic bus,
is defined which control, surface. are
as
all spacecraft necessary
items, for
other
than
provided
uniquely
the installation, of a scien-
mounting,
deployment, on the
thermal lunar
and
functional an
performance experiment
tific experiment composed a. of the An
Typically,
subsystem
is
following
items: sensor assembly or mechanism or a combination
instrument two.
of the b. Mounting perform
provisions a secondary action.
and/or
deployment function
or
manipulative by initiating or
mechanisms terminating
to
mechanical
mechanical c. An and d. An as instrument operation instrument an
electronics
unit directly sensor.
associated
with
the
calibration
of the instrument auxiliary to match
unit to provide the electrical unit (if any) by
the
electrical
interface
and
act
"adapter" and
requirements to the power, bus.
of the instrument command and data
sensor
its electronics facilities cables
transmission e. Electrical other
provided
the basic
necessary elements.
to provide
interconnections
between
the
subsystem
85
An "instrument" is defined as a sensor and its directly associated electronics. The electronics conversion unit and the instrument auxiliary are not considered a part of the instrument. The scientific instruments and accompanying electronics units and/or mechanisms required are either built by Hughes or furnished by JPL. The instrument auxiliaries are provided by Hughes. The design approach outlined in Section Z is implemented to provide basic bus subsystems that are, to the greatest extent possible, independent of instrument complement. The spacecraft basic bus provides only two power forms, unregulated ZZ vdc battery power, and regulated Z9 vdc power. The telemetry system is designed to accept only input signals in the range from 0 to 5 volts. The central command decoder is part of the basic bus, with provision for connection to subsystem command decoders for each instrument. The auxiliary unit for each experiment subsystem then provides the interface circuitry between the standardized basic bus subsystems and each individual instru- . ment as defined above. Each auxiliary unit contains a standard subsystem command decoder capable of supplying 8, 20, or 32 commands, as required. Power switches for turning the instrument on and off, for controlling heater power, and mode changes, etc., are located in these units. The instrument signal outputs are conditioned by the auxiliary unit as necessary to bring them into the proper voltage and impedance ranges. Subcarrier oscillators are also included in the auxiliary units. Special power supplies (ECU's) can be included where power form requirements are different from the basic bus supplies (none are necessary for the present scientific payload). Thus the basic bus subsystems do not have to be changed appreciably when the payload instrument complement is changed, and each experiment is largely independent of other experiments. The A-21A Surveyor Spacecraft carries the following five experiment payload subsystems: 1. Survey Television Experiment Subsystem - transmits pictures of selected portions of the lunar surface and other scientific instruments on command from earth and is composed of two survey cameras and the television auxiliary. Soil Mechanics - Surface Sampling Experiment Subsystem - investigates lunar surface properties and is composed of an instrument, and an instrument auxiliary unit. Alpha Scattering Experiment Subsystem - performs elemental analysis of lunar surface material and is composed ofasensor, an instrument
Z.
3.
86
digital electronics mechanism. 4.
unit,
instrument auxiliary
unit, and a deployment
Micrometeorite Detector Experiment Subsystem - provides data of individual lunar ejecta and is composed of a sensor, an instrument electronics unit, and an instrument auxiliary unit. Seismometer Experiment Subsystem - measures physical disturbances of the moon and is composed of a sensor, an instrument electronics unit, and an instrument auxiliary.
5.
The following descriptions of the individual experiment subsystems include detailed discussions of the functional elements incorporated in each auxiliary. SURVEY TELEVISION EXPERIMENT SUBSYSTEM
The survey television experiment subsystem provides the capability of observing the lunar surface, portions of the spacecraft, and large sections of free space, on command from earth. This capability is provided through an experiment subsystem consisting of two cameras and a television auxiliary. Each of the cameras can be commanded to alter its angular field of view and to change the angular orientation of the center of the field of view with respect to the basic spacecraft coordinate system. Provisions are also made for inserting colored or polarizing filters into the camera optical system on command from earth. This arrangement permits colorimetric and polarimetric evaluation of individual television pictures. These features are augmented and made practical by the provisions for commanding the camera optical systems to alter focus distance and lens aperture to adjust for variation in object distance and light intensity. Provision is made to alter the lens opening (iris) either on direct command from earth or automatically, as desired. With both cameras viewing the same area, stereoscopic coverage may be obtained. The Survey Television Subsystem is also used to observe the operation of selected instruments carried by the spacecraft. A functional block diagram of the survey television subsystem is illustrated in figure i0-i. All of the items appearing on this diagram are shared with the approach television subsystem. The telecommunications subsystem is shared with other spacecraft subsystems. The survey television camera is illustrated in figure 10-Z. The location of the survey television camera on the spaceframe and its relation to other units is illustrated in the general arrangement drawing (figure Z-Z). 87
I
.I
u_ wl., p
|-
I
< r_ o
N U © m < Z ©
I---4
O Z
N
[
1 •_ii I
m Z ©
'l!
i
o!
N M
M
m
W
o
I as" I
S
88
I o_
_,,J Ow
I
I as= I l o" j
3;o Ow
_d
MIRROR
HOOD
/
/DRIVE
MIRROR TILT ASSEMBLY
MIRROR ROTATION DRIVE M O T O R 1
IF=-.:
FILTER WHEEL FILTER WHEEL/ ASSEMBLY
FOCUS POTENTIOMETER
I SHUTTER
ASSEMBLY
IRIS POTENTIOMETER
1 \
VIDICON TUBE /:
'
1~ /
/VIDICON RADIATOR
/
/
ELECTRONICS CONVERSION UNIT
FIGURE 1 0 - 2 .
S U R V E Y TELEVISION CAMERA
89
The to the The The sion CDC
relationship and the DSS
of the ground
spacecraft
survey
television
experiment
subsystem in figure spacecraft. for transmis10-3.
transmission provides modulates
system
is illustrated for the commands
CDC, DSS
in its command transmitter spacecraft. survey performs television
operation, and
commands the CDC
processes
to the The
experiment operations, back amplifies
subsystem and to the the
receives transmits
the video, DSS the
commands camera, is located spacecraft
from and at the and preby
the DSS, command DSIF). pares the
the desired
confirmation Upon reception,
information the DSS and
DSS. signal
(The from CDC.
it for demodulation the signal
decommutation
in the and in the is sent
After
demodulation in the System. system signals for may CDC
CDC,
is sent and
to discriminators processing information on 35 and mm
decommutators CDC to the Television television
Telemetry TV video
System frame and
to video
identification recording
display on
film.
Selected meter may
telemetry
be sys-
monitored tem, by
binary,
decimal,
analog displays the
displays. be obtained
In the shortly
television after
a permanent a Land Survey
record camera
of video
readout
using The
to photograph camera video
output
of the television images within
monitor. its field of synchronizacamera and is
television composite blanking
converts signals
optical which
view
into complete vertical
include
horizontal the
tion and completely received The
pulses. requiring power by the
In performing only
this function, power and decoded decoder command
self-contained, from the central
electrical system are
inputs
commands auxiliary.
distribution survey camera The and
the television by
commands in the via
received television the command and active
a subsystem coded All from +22 vdc earth
decoder commands
auxiliary. receiver thermal system as
subsystem central of the
receives decoder.
power cen-
for operation tral power lated
control +29 vdc
camera
is supplied voltage and
the
distribution
regulated
unregu-
voltage. The principal identical amplifier information gate. thermal, location signal output from the camera outputs is the are composite from signal-frame the cables. camera Frame amplifiers contains to to each all to
video. the
Two
signal-frame of the from Frame and and TV
video auxiliary the camera
supplied
summing
unit through is received data
individual by the the
identification through optical, describe 9O an
summing camera
enabling
identification angular
from
electronic, the relative
position
information
necessary pertinent
engineering
characteristics
1 I
I
I
FROM PCM DECOMMUTATOR
I I
FI
1
VERIFICATION DISPLAYS COMMAND DISPLAYS COMMAND PRINTER
I I I I
i I
I I
I I I I I I
I I
DSS GROUND TRANSMISSION SYSTEM
KEYBOARD
I'
I i I
I I
I I I I
i I
I
I i
I I
F
CONTROL SWITCHES TAPE READER a LOGIC CIRCUITS -____ 8 REGISTER CONVERTER
I I
I ]
I I
i]
S C O _ __
I I
I I I I I
I I I I
TELETYPE TO SFOF DIPLEXER __QUISITION AID NTENNA
EMERGENCY COMMAND
I
COMPUTER EDITING
II II
I
AUXILIARY SWITCH ENTRY
PROGRAMMER a LOGIC AUXILIARY CIRCUITS
I I I
I
II
TRACK I NG O_O_._ PHASE MODULATOR TRANSMITTER DIPLEXER NTENNA
CDC
COMMAND
AND
SYSTEM
I I I I I I
I I I I I I
I I I I I I
J I_
b_ M _ m m
SIGNAL t PROCESSOR ENGINEERING
I I I
53j
I I I
SIGNAL PROCESSOR ___ CENTRAL 45 [
I
I 1 I I
I
I
I
;
ENAB.LE GATE
I
I I
MASER ___ 2KC NBPF __ AMPLI FI ER ]'.F.
SUMMING AMPLIFIERS
22V 8k 29V lOWER
SURVEY CAMERA COMMUTATOR FOR FRAME 16 CHANNEL T.D. CONVERSION _._ UNIT ELECTRONIC
II
I
=
I I
DIPLEXER PARAMP L_ VARIABLE i;BANDWIDTH ___ DETECTOR PHASE H AMP VIDEO
I
\ \
/ /
I I I
SUBSYSTEM DECODER
|
ELECTRONIC HEATER BLANKET
I
L 179/180
SWITCH T V AUXILIARY
I I
I
VCO
I:
t
i,
DETECTOR PHASE
I
F
I I
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CENTRAL COMMAND DECODER B
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COMMAND RECEIVER 52
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SURVEY CAMERA SYSTEM
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DISCRIMINATORS (10)
PCM DE COMMUTATOR
ANALOG METER DISPLAYS (16)
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F.M. CALl BRATOR DEMODULATOR
!EPUT METER ___ DISPLAYS DECIMAL (16)
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TELEVISION
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_[ VIDEO MONITOR _--_..--..._-TIME -- _ _. -.._]_
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I I I I I J
FIGURE
10-3. SURVEY TV CAMERA/GROUND !INTERFACE, FUNCTIONAL BLOCK
EQUIPMENT DIAGRAM
q /'(_ of&.
televised been
scene.
The
frame
identification TV
signal and
consists routed the
of analog to the
data
that has signal
serially
commutated The central
in the signal
auxiliary
central data
processor. form. This as
processor to the
converts TV
analog and
to digital with compofrom of
signal pulse
is then coded
returned data in the
auxiliary amplifiers.
summed The
site video the summing composite thereafter The that vide can be
summing
total signal consisting
amplifiers video (figure survey
is sent
to the frame
transmitters
in a sequence data
information 10-4). television
with
identification
following
immediately
camera
is equipped its maximum nominal)
with
a variable focal length
focal
length
lens
commanded
to set at either (6.4 x 6.4 degrees to provide focus
position or
to pro-
a narrow-angle focal length
viewing (25.4 lens
capability x 25.4
at its nomito
minimum nal) viewing cover lar
position The 4 feet by
wide-angle of the camera
degrees
capability. from
distance The
is also
commandable views
the range to be axis
to infinity. means inside assembly
indirectly
the particuThe elevamotors in
scene
televised is mounted
of a gimbaled of the azimuth to position with be
mirror gimbal the mirror earth stepped
assembly. axis.
tion gimbal are provided
Stepping
in the mirror or elevation camera
surface
angularly The optical of approxi-
either
azimuth
in accordance may
commands. in increments The
centerline mately
of the
field of view and
3 degrees
in azimuth
5 degrees
in elevation.
mechanical 2-1/2 elevation motion
motion degrees optical of the in
of the mirror per path. mirror azimuth. around an step, A or
itself in elevation exactly half the
is in increments increment between rotation motion
of approximately of the the camera
angular
one-to-one in azimuth This axis
relationship and the
exists
mechanical
corresponding azimuth
of the camera
field of view
occurs very
because
corresponds centerline is provided small
to angular of the through heaters
rotation lens.
nearly control
parallel of the
to the optical survey camera and
camera
Temperature thermal the trol radiator
a passive located on
for
the vidicon within for the
faceplate camera.
electrical
electronic is also The
chassis provided
In addition, to achieve
closed-loop proper
thermal camera TV on
con-
the vidicon is capable emergency be
faceplate
operation. normal
survey
camera TV may
of operating mode. to the The
in either camera
a 600-1ine is turned mode
mode normal intent in a
or
in a 200-1ine mode and then mode
in the The pictures
switched
emergency
if required. quality for
of the latter signal
is to permit not as
the transmission wide as that normally
of degraded required
limited
bandwidth
television
93
! -_J H'I .67 ms I I l I J (NORMAL VER11CAL BLANKING PULSE j MODE) (EMERO MODE)
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V
(BLACK)
+0..5 V START 0 OF PICTURE SEQUENCE
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V.
4 -I .5 V
.200 (I_IORMAL
ms MODE)
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600 ml
-2.5 V (WHITE) -(EMERGENCY MODE)
FIGURE
10-4.
SURVEY
TELEVISION
COMPOSITE
VIDEO
OUTPUT
94
transmission. increase high-gain receipt priate and ground the
The probability or
basic of the of
objective successful powered
in
providing television transmitter television the
this
mode
of
transmission to earth when
is
to the
transmission is not usable.
antenna display
high
(Successful use of of approthe
emergency compatible
mode with
pictures
requires transmission
equipment mode. survey ) camera of 150
narrow-band
emergency The nominal opened This taking
is equipped milliseconds. period of time
with
a focal
plane
shutter
that provides of being command. for
a
exposure for an
The on
shutter
is also
capable earth exposure is about reflex
indefinite mode The
receipt
of proper
operating purposes.
is not directly maximum exposure shutter sampled shutter light
equivalent
to a "bulb" time A
picture-
equivalent beyond
exposure this interval.
1 second, splitter
regardless is provided passed a f:22, by the until
of continued in front the lens.
beam
of the If the driving
in the camera light is intensity of light in the camera viewed by is
to sample intense prevents At
the total light input and the f:22 in the shutter the iris is from manual mode may camera, direct or in not
sufficiently and removed.
circuitry the will a result
inhibited is
opening the occur (2) in shutter as the
excessive open of regardless changes:
intensity. area of field the
Light-intensity coverage of view, viewed or (3) by in of
changes the the
(1) of the
angular illumination The control on receipt camera
orientation of same the
reflected sun matic function The are remain fixed fore, approach In the angle. iris
scene
camera is also override
because employed the
a change in the autoinhibit
light-sensing Provision ,commands.
arrangement is made
circuitry. of ground
to
shutter
horizontal generated
synchronizing when off. between The the camera
pulses is
and turned
the on
vertical (power does
blanking applied) not require
pulses and a There-
independently until timing they are permits normal of mode the it is
turned
television the and horizontal independent in circuit pulse a
system and of
design vertical each
relationship both a
oscillators. other. and the This power principal is TV design
free-running significant the vertical sequence. and the that central
saving blanking
complexity represents
required. timing received auxiliary (via
waveform the command
picture
When command video after an sequence. logic
start
frame
command from is enabled. start frame pulse the
receivers decoder, pulse
decoder) circuitry
subsystem cal turn, blanking initiates
camera appears picture
The
first
vertiin the
earth-initiated This blanking
command, triggers
a complete
95
camera to be
shutter generated The first
to provide to turn signal the
vidicon
exposure on.
and,
simultaneously,
causes
a signal
transmitter during with
transmitted pulse
a typical superimposed ground
picture-taking horizontal monitors picture 10-5a). in the
sequence
con-
sists pulses. before place
of the vertical This receipt
blanking
synchronizing to be synchronized takes
waveform
is intended picture
to permit video.
of the actual
Single-frame (figure seconds
readout
in the interval
following signal
vertical requires
blanking about (second) 1.2
Transmission normal pulse. signal pulse mode The to the to close the and sec-
of the composite is terminated ond blanking
video at the
start
of the next the video
vertical provides acts as
blanking a video an enable TV
pulse resets
closes
gate
which and
transmitter, frame complete enable
the camera input gate
shutter, to the video identification
identification sequences pulse. The
summer
in the are
auxiliary.
Several during command the
of frame trailing
information initiates
transmitted
edge
of this pulse
a transmitter-off sequence. erasure. at the sequence. erasure nominally (figure during normally. TV
that terminates Three the subsequent the three third
transmission frame periods frame
for that particular are necessary must for
picture-taking proper vidicon
Therefore, until will per-
start
command pulse
not be start
received frame
spacecraft This
after mit
vertical
blanking
of the
complete
erasure The 61.8
frames three seconds to receive
to occur erasure in the
(picture frames emergency
readout require mode command respond the
occurs 3. 6 seconds 10-5b). the second In this image will It is
during
the first frame). mode the and and
in the normal possible erase case, not be for scan since
spacecraft the TV
a start frame a picture and
camera erasure from
to take has not
the third
yet occurred, Through
previous
completely the
erased capability
the vidicon.
this
specialized (imperfectly
operating erased}
sequence, picture
is provided every 2.4 camera earth
to transmit
a degraded
approximately The survey receipt
seconds. is equipped commands optical with to occupy filter. a filter wheel one of four that may different be posi-
television of proper use
rotated tions, before
on each time
providing of camera filters, with as
of a different
It is therefore of clear, of lunar scenes
possible, colored, through or
assembly, desired,
to select to enable
combinations
polarizing optical
observation
filters
known
characteristics.
96
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SOIL
MECHANICS-SURFACE The purpose of the
SAMPLING
EXPERIMENT sampling
SUBSYSTEM experiment subsystem sur-
soil mechanics-surface the mechanical
is: face;
(i) to qualitatively (2) to provide within
determine
characteristics and
of the lunar surface
means its area
for lunar
soil manipulation; The surface modulus
(3) to map to be and
elevations are: nature
of operation. shear
properties
investigated surface
of the
material,
strength,
of elasticity
contour. The lunar with soil mechanics-surface by digging, viewing) scraping, in a limited will measure above. sampling picking, area those near experiment and the contour Surveyor subsystem mapping will test the
surface television
(in conjunction The
spacecraft.
experiment the lunar
subsystem soil properties
parameters
necessary
to determine
listed
Experiment The an vide and Figure instrument,
Subsystem soil
Characteristics sampling auxiliary the unit, experiment electrical and the of the the subsystem cables instrument on the is necessary auxiliary spaceframe. composed to prounit, of
mechanics-surface an instrument between bracketrynecessary is a functional block
interconnections mounting 10-6
instrument to mount
instrument experiment
diagram
subsystem.
Instrument The device scoop attached electric fourth extension, provided tion is the forces forces to
Description basic soil attached are mechanics-surface to mounted the on end of an sampler extension and arm elevation Potentiometers of positions the consists arm vertically assembly and is extension are instrument. of output the (0 scoop to 50 used (see of a figure hinged clamshell-shaped 10-7). base The which by is three A
(scoop) and to arm the
a horizontally The extension azimuth, door. positions closed
spaceframe. that operates control the and the
manipulated motions. to Limit door. g of and the measure
motors motor
the scoop azimuth
elevation, indicate system, on the
switches An 0 to scoop accelera2000
are
open
and
measuring mounted pieking over over
providing to force 0. 0.1 to
a dual-range measure transducers 1 to 20.0 3.0 pounds pounds. vertical are
g),
instrument Two range range
deceleration provided; and the other one
during
action. the the
measures measures
vertical retraction
from of
98
T 2.2¥ EXTEND I EXTEN_ONDRIVE MOTOR RETURN EXTEND MECHANICS SURFACE SAMPLER AUXILIARY
RETRACT
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MOTOR I RETURN
SOIL
MECHANICS 22vd¢ UNREGULATED
SURFACE
SAMPLER
PRECISION POTENTIOMETERS
EXTENSION MOTOR RETRACT
T
SMS_ EXTENSION POSITION
Z
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ROTATE
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LEFT
AZIMUTH
DRIVE
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RETURN
LEFT
OPEN
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MOTOR RETURN OPEN
AZIMUTH MOTOR
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SMS$ AZIMUTH
CENTRAL POSITION SMSS RETRACTION CURRENT
SIGNAL
PROCESSOR
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AMPLIFIER SUMMING
SCOOP r_T_ -i CLOSE
DRIVE
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ELEVATE ELEVATION ELEVATE SWITCH DRIVE MOTOR RETURN ELEVATE
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ENGINEERING SIGNAL PR._._9 CE____S S_._OO R__
LIMIT 'S TRAIN STRAIN MEASURING ON _ POWER SMSS STRAIN MEASURING SWITCH AMPLIFIER POWER
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TO RADAR AND SQUIB RETURN
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AMPLIFIER ACCELEROMETER
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ENGINEERING MECHANISMS AUXILIARY
"_ACC ACCELERATION MEASURING OFF •
ACCELEROMETER ERATION MEASURING SWITCH _' Ir / t
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AND
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SMSS SMSS
ACCELERATION
DATA OUTPUT
SUBCARRIEROSCILLATOR
RELEASE
SMSS I
SQUIB HIGH-GAIN RANGE
POWER SWITCHED
LOW-GAIN
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SQUIB
FOWER
A
_/1 -Z_
FIGURE
10-6.
SOIL
MECHANICS
SURFACE
SAMPLEP_ BLOCK DIAGRAM
EXPERIMENT
SUBSYSTEM,
FUNCTIONAL
fro
FIGURE 10-7.
SOIL MECHANICS SURFACE SAMPLER INSTRUMENT PARTIALLY EXTENDED
A l t e r n a t e light and d a r k s t r i p e d m a r k i n g s on the inside and outside of the scoop m a y be viewed by the s u r v e y television to d e t e r m i n e penetration depth of the scoop, g r a i n s i z e , and the l e v e l of lunar m a t e r i a l contained within the scoop. Operation of the i n s t r u m e n t c o n s i s t s of manipulation in extension, elevation and a z i m u t h ; disengagement of the elevation d r i v e by a clutch device f o r picking action; r e t r a c t i o n of the extension mechanism along the s u r f a c e for digging; and l o w e r i n g the scoop to the s u r f a c e with the elevation d r i v e disengaged f o r point-topoint mapping of the l u n a r s u r f a c e . The i n s t r u m e n t c a n p e r f o r m l u n a r - s u r f a c e testing within a s e c t o r defined by i t s o p e r a t i n g l i m i t s in azimuth, extension, and r e t r a c t i o n . E i t h e r c o a r s e o r fine
101
incremental movements position the scoop within its sector of operation. fine movements occur in maximum increments of 0.6 inch in extension,
degrees in elevation, and 3.0 by degrees in azimuth. Motor control by and
The
2.5 or logic
coarse
fine increment and timing Three ment closes drive
selection in the
earth
command auxiliary
is accomplished unit. motors
appropriate
action
instrument
series-wound, elevation, door. An scoop the and
split-field, and
direct-current
move motor
the
instruand
in azimuth, the train The scoop
extension-retraction. clutch for picking directly solid-state clutch
Another disengages action. from the
opens
electromechanical can fall freely receive controlled one motor
the elevation
so that the motors and supply
clutch are Only
power by or
spacecraft located
22 in the one
vdc
unregulated
switches
instrument time. at the The scoop: a. b. c. d. e. The instrument transducer to indicate excitation sor (ESP).
auxiliary instrument
unit.
will be
operated
at any
mechanism
is capable
of developing
the following
forces
ExtensionRetractionAzimuthElevation Scoop jaw
l-lb
radial radial
push. pull. at maximum at maximum at edge. precision potentiometers It also system, The has and to indicate two two forcelimit switches receive procesESP. a static extension. extension.
20-1b l-lb
either up
way or
-- 30-1b -- 6-1bs
down
measured with
instrument position systems, fully from open
is equipped within an or the
envelope
of operation. transducer scoop positions. supply through position auxiliary to measure
accelerometer fully closed vdc are
potentiometers signal
the 4.85 outputs
potentiometer processed
in the
engineering mode
Their the
commutator potentiometer unit for the radial
4 in the
In addition, selector picking
output and SCO
from
the elevation instrument
is routed processing velocity
via during
switch action
in the
only.
This
is necessary
of the
scoop
at impact. Each signal vertical lated of the two force-transducer and an systems consist of a strain provided gage bridge, a
conditioning and
amplifier, forces. and
interconnecting signal conditioning
cable
to measure contain regulated a regupower
retraction power
The a d-c
amplifiers 29 vdc
bridge
supply
amplifier.
Spacecraft
102
is supplied to the signal conditioning amplifiers which generate a precise bridge excitation voltage. The bridge outputs are routed to a d-c amplifier the output of which is processed via the Mode 4 commutator. One strain gage bridge is located so that elevation forces (from 0 to 3 pounds) can be measured. Another strain gage bridge is mounted so that radial forces from 0 to 20 pounds applied to the scoop along the tape can be measured. An acceleration transducer system, consisting of a sensor, an amplifier,
and an interconnecting cable, measures scoop deceleration at impact during picking action. The sensor is mounted on the scoop near the blade. The output signal from the sensor is routed to the amplifier where it is simultaneously amplified through dual channels. The ratio of the two channel gains is approximately 40 to 1 with the amplitude of both channel outputs ranging from 0 to 5 volts dc. When the sensor output is low, the high-gain channel can be monitored. Conversely, when the sensor output is high, the low-gain channel can be monitored. Both channel outputs are routed to the soil mechanics surface sampler auxiliary where, upon earth command, a switch will select the desired gain. The selected output is sent from the instrument auxiliary unit to the telecommunications subsystem, where it is modulated directly upon the main carrier. Two limit switches, mounted on the scoop mechanism, indicate whether the scoop door is fully open or fully closed. The switches are excited by Z9 vdc regulated voltage. The switch outputs are routed to two destinations, (i) to digital channels on the Mode 4 commutator to be transmitted back to earth, and (g) to the instrument auxiliary unit to turn off the scoop motor when the scoop is either fully open or fully closed. The soil mechanics the rmal conditions : surface sampler instrument range is designed for the following
Operating temperature Survival temperature
-65 to Z57°F -300 to Z57°F
Prior to touchdown, the instrument mechanism is secured to the spacecraft by a clamp which may be released by a pyrotechnic device. The pyrotechnic device is energized from a 9. 5 ampere squib power supply in the engineering
103
mechanism A command the
auxiliary received
and from
controlled earth unit by
by the
a switch
in the
instrument
auxiliary
unit.
telecommunications decoder in the to the
subsystem engineering auxiliary in the
is routed mechanism unit. The
via
instrument and
auxiliary
subsystem
auxiliary command mechanism power
simultaneously
to a switch turns on the
instrument supply
simultaneously auxiliary turns and
squib
power
engineering unit. The squib The
the
switch
in the instrument after to prevent approximately inadvertent
auxiliary 20
supply
off automatically is interlocked
milliseconds. of the squib.
actuating
command
firing
Instrument The ing, and
Auxiliary instrument power
Unit auxiliary for are unit provides the command decoding, signal process-
management These power an unit. command pulse functions
soil mechanics-surface by a 20-command shunt,
sampling
experiment coma range the instru-
subsystem. mand decoder,
performed a current
subsystem a timer, within
switches, SCO, and
measuring logic
selector ment
switch,
appropriate
circuitry
located
auxiliary The
subsystem provides
decoder commands
processes required from the
the
inputs
received
from
the
CCD
and
output
in the subsystem
instrument decoder
auxiliary perform the
unit and functions
in the listed
instrument. in table sec 2 10-i.
Commands
The or clutch
2 sec/0.1 to either
timer
limits
the
operating or
time
of the
instrument (fine).
motors Normally, selectprevents fine and and
± 0.4 on
seconds in the During
(coarse) coarse fine
0. 1 _- 0.02 mode with
second the
the instrument able upon earth of the time
is turned command. elevation
timing mode
fine mode logic The
operating, to the
switching instrument.
release coarse coarse auxiliary current measuring
clutch for motor
to prevent and clutch
damage
intervals
operating of the
directly instrument.
govern The
the fine
incremental unit during contains scoop
movements, a current
respectively, shunt only. 4 that The
instrument motor
measures shunt output
extension-retraction is sent to a current
retraction in the Mode
channel
commutator.
ALPHA The of craft lunar is
SCATTERING alpha surface bombarded scattering materials. with
EXPERIMENT experiment A 6 mev portion alpha
SUBSYSTEM subsystem of the lunar performs surface Backward compositional adjacent scattered to the analysis space-
particles.
alpha
104
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106
particles, particles,
as well are
as
protons by
generated
within surface
the
sample
by
the incident
alpha
detected
solid-state
barrier
detectors.
Alpha
Scattering The alpha
Experiment scattering electronics electrical scattering except boron, and
Characteristics experiment unit, cables. experiment hydrogen nitrogen, sulfur. and subsystem helium. detects The sodium, proton scattered system alpha detects aluminum, measurement to be surface particles proan subsystem instrument is composed auxiliary of a sensor, unit, a deployment an
ir_strument mechanism, The from tons
digital and alpha
all elements from lithium, phosphorus, from one
fluorine, Although the
magnesium, of the
silicon, varies about
sensitivity threshold will be
element by weight. than
to another, Operation 0.1 micron.
the detection of the sensor
is expected impaired by
1 percent
contaminants The and
thicker important alpha
scientific particles,
measurements and
are
(I) the
energies and
of the protons scattered alpha
scattered as
(2) the intensity The The energy pulse
of protons detected
particles mined energy by
a function height detected
of energy. analysis. particles, 9-bit channels
of each
particle
is deterto the coded
pulse
height,
which
is proportional signal the which
of the
is converted words, or which energies parts the
to a time comprise
is then output,
into a 9-bit identify
word.
These height
instrument particles. electronics
the pulse
of the detected -- the command digital
The sensor. converter, sensor uses in the
instrument The digital
is packaged electronics logic
in two contains
and
the
memory width pulses The the
circuits, from the
power
and
digital
to convert form
the
variable
into
synchronous commands This with
binary
for transmission. controlling ifa
command outputs becomes
memory of the noisy detectors or
spacecraft sensor.
to set flip-flops proper operation
enables
detector
is contaminated
radioactive
material.
Sensor
Description The sensor, which source The are pulse will be deployed directly with surface to a pulse to the two lunar surface, conand four
tains proton
a
radioactive detectors. outputs data
of alpha
particles of the
alpha
detectors type. The
detectors
are and
barrier
alpha
detector The alpha
amplified output
applied
height-to-time is applied
converter. to a gated
of the height-to-time
converter
107
clock detectors a thin
in the are metal
alpha
scattering to the
instrument alpha to detectors stop
digital
electronics except that the
unit. proton To are prevent
The
proton have
similar foil in front
detectors registering The that from each the proton
of them a proton are pulse the
alpha guard and
particles. detectors applied
background puts inhibits command detector of the the
radiation guard proton memory with the
as detectors channel control associated
event, amplified
provided.
out-
to a discriminator Signals and
height-to-time operation of each
converter. alpha detector
guard
detector. the sensor must operate in a vacuum However, source close pressure. Tests large since temperature control the prevent thermal the is provided design temperature of the of the must be of and in required to the For suf-
When requires testing for lunar the an
measuring alpha instrument, operation source
composition, of 25 to 100
millicuries to use alpha are sources obtained may be
of Curium-242. the same can at be alpha mounted
it is not in vacuum.
necessary Weak
detector vacuum ficient
so that tests, length
reproducible a different set
spectra of sources
atmospheric
employed. to avoid
to allow The
accumulation instrument of low -40°C
of enough requires to +50°C special and
counts
statistical it has an range by operaof a
uncertainties. ting temperature to +75°C.
treatment a survival
range At
-166°C 5-watt case
temperatures, At high
temperature temperatures, finish
electricalheater. together from High with rising appropriate above
sensor in-
surface the survival sensors, at
strument
temperature. in the sensor and upon in the command The operating instrument from weight sequence earth of the for
accuracy unit, the will
temperature be monitored
electronics to maintain instrument this instrument a. b.
regular
intervals of this
temperature unit follows:
requirements is 3.2 +_ 0.2
instrument. A typical
electronics is as
pounds.
Standard Background position).
sample-count--
3 hours 3 hours
(sensor (sensor
located deployed
in stowed
position). count
to background
c.
Initial surface).
count
on
the
lunar
surface--
6 hours
(sensor
deployed
to lunar
d.
Data lunar
accumulation-surface).
minimum
of 24 hours
of data
(sensor
deployed
to
I08
In must between same not
order exceed the
to
obtain 30
reliable and step a
data, the of
the
total
interrupted time not in step
time c
in
steps
a and The for
b
minutes, of
interrupted step c must
1 hour. hours
time the
initiation
and
exceed
7-1/2
reason.
Instrument The figure 3-6), The are cps, gated into a identical while 900-kc 7-bit
Digital instrument weighs control except the proton clock binary the parity effective check by
Electronics digital 5.0 circuits for the & 0.2 for
Unit electronics pounds. the two instrument rates. is to word and then This and bit the a parity followed digital 29 vdc The 550 convert is shifted into results bit. by 8 an outputs, alpha cps. the readout The pulse into output in a 9-bit data is unit a logic from storage register word have alpha bit and rate proton is uses sensor to a sync of a pulses, 2200 a unit, located within compartment B (see
readout frequency counter 7-.bit
readout and word. dead are a 7-bit The
circuit the
register where consisting been received
minimize bit sync since and bit the The operating
time,
generated. bits a sync within the spacecraft
followed previous power voltages
7 data word,
If no zeroes
generated. generates instrument
converter from
electronics supply.
Instrument The instrument contains oscillators, The ment instrument auxiliary The instrument auxiliary with nificant "one" Source is a
Auxiliary instrument and the
Unit auxiliary spacecraft'. decoder power switches, unit in is figure and unit are a sync bit unit The capable and similar 10-15. the "Proton Data individual simultaneously by the a ohms. least 7 data Output" SCO's and bits, signals in are with bit. by 5 & 0.25 the from the provides instrument of providing the electrical auxiliary 20 interface unit commands, (see between figure two rectifiers to the seismometer 10-8) the
command two
subcarrier (SCP_s). instru-
three in
silicon-controlled appearance
auxiliary unit "Alpha digital shown Data
Output"
electronics These outputs of a parity by is
modulates present bit followed
instrument NRZ, most A digital volts. sig-
unit. a format bit first,
digital, the
consisting and
following volts 5000 and
significant "zero"
represented impedance
0 ± 0.25
digital
approximately
109
The two
outputs
of
the
SCOs
are in is
applied the phase central
to
a presumming signal processor. onto the
amplifier The main carrier
and output
then of of the
to the
phase-summing
amplifiers amplifiers
twophase-summing space craft The in the tran two smitte analog
modulated
r. outputs both in provided the range by of the high accuracy and temperature are processed sensors by the
experiment
are
0 to
5 volts
ESP. Electrical delivered subsystem controlled power to at by the Z9 power instrument vdc regulated switches from the for the via instrument the and within Z9 payload ZZ vdc operation harness unregulated. instrument Z2 v busses and from active the Power auxiliary are 1.4 aml thermal electrical for unit. the The control power experiment instrument is is
power
the
requirements
v and
5 watts,
re spe ctively.
Deployment The instrument sible position. Support by for and a mounting the sensor, sequence,
Me chanism sensor sensor from deployment prior the to mechanism spacecraft to the provides touchdown, background the and means deploying to the for it, stowing in an the irrever-
stowed;
count;
lunar
surface
for
the
sensor The the
in
the
stowed
position platform in the to standard steel band also
(figure serves
10-ga)
as a
is dust for to
provided cover calibration, provide In the mounting the stowed
platform. holds the sensor radiation is held
mounting sample
standard in proper view in
proper the
position spacecraft sample connected count. to
positions
relationship factor for by a
required position, platform.
thermal the sensor
place
the
Release (figure puller the sensor As the lunar as 10-gb) which
of is
the
sensor
from by the steel by a
the
stowed
to an
the
background squib of which
count actuated the
position pinto the position. band point to ro.tate position. the
accomplished
activating band. small the
explosive orientation band to the release, vertical causing the
disconnects is being provided
Proper orientation stowed out, past after the drum, toward
sensor
surface it is
rotates count orientation At this
deployed mechanism
from swings 15 on degrees the
background the axis. the background
deployment until begins to the the
unreels band as it
sensor
is
exert
a force arm
sensor out
sensor count
and
deployment
continue
II0
u?
Z
O©
I-.-I
I
O
P_
0
o >
Y
LQ ZU._ O--° ZC_
I q.
i
O @t a O
BAND
STOWED POSITION ORIENTATION BAND 32 deg
ULLER
I
STANDARD SAMPLE
s,oP
SA K/;O?NO
COMPLETE /
DEPLOYED_D ON SURFACE SURFACE ///////" I
i.'_'_'_. _
,
'
.'o ,A R CL-..'A'."
a) Deployment device b) Deployed from stowed position
FIGUKE
10-9. ALPHA DEP LO YMENT
SCATTER.ING ME CHANISM
SENSOR.
When arm
the and
sensor instrument
reaches sensor
the
correct
position, and from
a detent
retains out
the
deployment slot.
in position, sensor
the band
slips
of its disconnect to the lunar
Final surface pullers. position second the The from
deployment {figure
of the 10-9a) puller arm
the background by two explosive arm
count
position The
is controlled releases and the the
squib background
actuated count The
pin-
first pin allowing
deployment to settle
detent,
sensor spring
to the lunar which
surface.
a pin-puller allowing connection sensor.
releases the sensor
loaded port and
yolk face
connects
the arm
to surface. cable
sensor, only the Each
viewing sensor
to conform
to the lunar is the electrical
between
the
the
spacecraft
of the three switch power
pin pullers controlled by
utilized command
in.the from
deployment earth.
sequence The weight
are
ener-
gized
by
of the
113
deployment pounds. The so that the interference surface and
mechanism
and
support,
less
wiring
harness,
is approximately
5.5
deployment maximum as
mechanism practicable with
and view
sensor for the
are
positioned
on
the
spacecraft as little
sensor
is achieved of the
with
possible The
the operating between the
envelope the
soil mechanicssampler Z-2. The
sampler.
relationship sensor on
soil mechanics-surface can the be seen in figure the soil
the alpha areas
scattering
spacecraft between
common
of potential sampler
interference are
sensor I0-I0.
and
mechanics-surface
illustrated
in figure
MICROMETEOKITE The momentum, meteoroids
DETECTOR detector energy the lunar
EXPERIMENT experiment
SUBSYSTEM provides lunar ejecta data on the number, microspacecraft.
micrometeorite and kinetic
of individual surface
resulting
from
impacting
in the
vicinity
of the
Surveyor
Micrometeorite The and an
Detector
Characteristics detector experiment both contained cables. how the subsystem in one consists package, The experiment and the an of the sensor
micrometeorite electronics and (figure unit are tied associated 10-11)
instrument unit, diagram
unit
instrument subsystem instrument
auxiliary block electronics
interconnecting shows functionally
sensor
together.
Sensor
Description The sensor contains three detectors and bonded two thin signal to an film from impact capacitors the plate with to a the plate, of the unit and
detectors each 100 while side KC
consisting
of a microphone impact related plate.
bonded
of the common will be
The
microphone, striking energy sensor
crystal, the
to the momentum gross
of a particle and
capacitor Signals applied
signals from
indicate
trajectory are
kinetic in the
particle. then are A ture.
the three instrument
detectors
amplified unit for
to the
electronics in the from the heater,
digital
processing. unit temperaand its output
temperature
sensor, excitation
located voltage through
sensor
unit, within
monitors the ESP
It receives
a source Mode
of 0 to 5 volts A heating
is processed
4 commutator. in the the sensor day-night unit, provides terminator. the The
proportionally required if the
controlled experiment
located
is operated
near
114
Z
/
_r,3
/
/
/
m_ _m
s<
/
/
I
sensor receives turned includes circuit. buffer picked acoustical operation
is
designed power
to directly
survive from switch
the the located and
lunar
night
without 22 vdc
heating. unregulated auxiliary as unit are the part
The
heater and The is sensor
spacecraft in the
supply unit.
on
or
off by
a
instrument transducer
a buffer Pulses amplifier up by the
amplifier from on the the
calibration
of a calibration routed through The impact and any of the during the is
instrument
electronics to physically
transducer for The sensor
shock of the can
plate. sensors at
microphone
calibration assembly
capacitor be calibrated
transducer. upon Figure command. 10-1Z shows the one from side the from
time
the
sensor
and of the
its
basic is the by
field
of view. by
As
mounted
on 10 field
the
spacecraft, on
field by other TV
of view compartment side camera is
sensor A and
reduced support
approximately structure. 6 percent The because
percent of view
mast
reduced 2 and
approximately antenna
of the Instrument
shadowing
omnidirectional
B.
Electronics instrument analysis detector 10-11. unit counted
Unit electronics (PHi) logic unit contains microphone microphone and signals commutation from Every registers. does not cycle and the impact and and capacitor film film as are by clearshown sent a when are to
The pulse ing height circuits,
circuitry, circuitry,
capacitor circuitry sensor unit
calibration Energy for and occur, and pulse stored the
in figure the
momentum height analysis.
electronics is
registered During a series intervals
detector no fed ejecta
in accumulation commutator
impacts
of ones
continuously The output hit pulse
to the signal
telecommunications in this the case will
subsystem. resemble will circuits a 001 pulse bits from occurred. bits will the are cycle cycle and to identify height are the used a 100-cps one time the the analyzed for and square read wave. out the As When data the
a particle from the
occurs, height cycles,
commutator (PHi) generate of the Four signal impact and commutator from four
analyses it will for
from
registers. of a data momentum
commutator Three from height identify the bits the
start
word. signal pulse
are acoustical
used
readout
transducer. kinetic energy the
readout Two bits are out an
of the bits used the are
analyzed the side
capacitors. Three used once The
used
to out
on which accumulation
to read energy occurs, is
momentum
to read when
accuit may routed
mulation. also be
Although cycled by
the
impact output
a command
earth.
commutator
116
9t'-O_6-OZ
E_
-?
_E
m
.J< <{E u)I-H
__g.
O_z
_o_
) wo O_z _u w
OD
E o
_z
O_ _0
3
E
-,[
E
E
....
_-_ 1.3
_Z 0 _
u_
,--_ U?
®___
I
I
<
->
w iz:
I
I I I BO.LVIFI_II_OO _3V'Id 03
i [ii
I f I zlE_ mow L I
..................
]
........
r-
I",.
_30003Q
ONVI_IflO3
M3J.S.kSBns
I
o
I r,-
MICROPHONE
60
DEG
60
DEG
THIN AND PLATE
FILM MICROPHONE
0 ELECTRONICS PACKAGE
Lines limits factor; are and
denote angular of sensor view angles indicated both ends of sensor 30 DEG
typical of both sides
60
DEG
a s s embly
FIGURE
10-12.
VIEW
FACTOR INSTRUMENT
OF
MICROMETEORITE ASSEMBLY
DETECTOR
119
to the instrument monitored Upon tronics through earth
auxiliary the Mode command
for processing. 4 commutator. the capacitor power
The
output
of the
conversion
unit
is
clear
circuits
in the instrument burn out any
elecshorts in
unit will generate capacitors.
sufficient
to electrically
the thin film An generates ment craft
electronic B+ power
conversion for circuit This supply
unit
(ECU)
in the instrument in both unit the sensor
electronics and in the from
unit instruthe space-
elements
electronics 29 vdc
unit.
conversion through
receives
its power
regulated
a switch
in the instrument
auxiliary.
Instrument The located ment
Auxiliary instrument
Unit auxiliary B, unit for the The the micrometeorite interface auxiliary auxiliary decoding, The command detection between unit has unit experiment, the the instrusame in processing, components two
in compartment and the Surveyor as
provides
electrical instrument instrument command experiment. a subsystem
spacecraft.
physical figure and
appearance 10-15. This
the seismometer unit performs for the include
shown
auxiliary
signal
power
management these and
functions functions an SCO.
electronic decoder,
that accomplish power switches, The mand
subsystem decodes
command
decoder
receives addressed
signals to the
from
the
central
comdetector
decoder,
all commands and furnishes
micrometeorite to the
experiment electronics
subsystem unit and the
pulse
output
commands
instrument
sensor.
Signal lating
processing
is accomplished with the The
in the instrument micrometeorite subcarrier processor
auxiliary
unit by output
modu-
a subcarrier
oscillator
commutator output
from to
the instrument summing
electronics in the
unit. central
oscillator for
is applied via the
amplifiers
signal
transmission
telecommunications Two power, regulated other used solid-state
subsystem. switches One control the 29 vdc regulated from and and the 22 vdc spacecraft instrument. supply unregulated 29 vdc The
respectively. supply used
switch
controls
the power unit
in the the
instrument from
auxiliary the spacecraft
in the
switch in the
controls sensor
power
22 vdc
unregulated
heating
element.
IZ0
!
w
a single, amplifier advantages
low
frequency amplified
coupling by
between
the two. low-noise
The
d-c
drift from To
the preutilize the main
is not
the high-gain,
amplifier. the output by
of a balanced
arrangement
in the preamplifier, output. which and stages The
seismometer
coil is center-tapped fier is filtered push-pull through an and
to provide applied
a push-pull
of the preamplianother stage of
to the modulator The modulated
is followed signal
amplification. attenuator and
amplified of a-c
is then
applied
to the following
amplification.
It is then
demodulated The ments.
filtered. structure magnet of the sensor is composed an of four gap, with basic ele-
mechanical
(1) a cylindrical cylindrical on the same
assembly
forming within the
annular gap, main
(2) a fixed several turns
center-tapped of wire wound
coil centered form
magnet
but independent consisting 4.5 inches
of the of two
coil,
for
instrument and high.
calibration,
(3) the elastic housing
suspension
preformed and by
springs, 4.0 inches
(4) a cylindrical Thermal A control
approximately using super
in diameter augmented
is achieved sensor the
insulation
a special
heater.
temperature To
is included. instrument, instrument sensor. pulse a start calibrate command produce pulse is routed via
calibrate circuitry
calibrate in the stop
in the
electronics The step
unit to
a current off by output
step the is
auxiliary calibration Since
coil of the command
is subsequently transient
turned at the
after
a satisfactory the
obtained. stant
it is desirable of full-scale to vary
to maintain reading
calibration
output
pulse
at a cona step with the
percentage is used setting
independent of the
of the gain
setting,
attenuator the gain
the amplitude
calibration caging
pulse
inversely by
of the
amplifier. the
Mechanical support frame. cable. to the
is performed
clamping by actua-
seismometer ting a squib The maximum
mass device sensor
against to cut
Uncaging
is performed
the holding attached
is rigidly
spacecraft and The where location
at a location
where factor Z-g. on due
mechanical structure may
coupling
is afforded
the amplification is shown off vertical
to spacecraft The seismometer
is at a minimum. operate as much as
in figure depending
15 degrees
the landing The I. 2 _- 0.05
attitude instrument pounds.
of the
spacecraft. unit, which is located in compartment B, weighs
electronics
125
I
Instrument The face between
Auxiliary instrument the decoding, for capable the
Unit auxiliar instrument signal subsystem. of providing rectifier y unit and the (figur Surveyor and auxiliary commands, and instrument one e 10-15) provide s the This electrical unit functions subsystem power oscillator seismic ranging This auxiliary where output current is an signal unit. further analog data, from directly The command switches, (SCO). in the form to Z0 of cps a signal conditioning of the 0 to 5 an one performs inter -
spacecraft. power consists two management of electronic a
command required decoder
processing, The 20 (SCR), of the
silicon-controlled The analog with a principal of
subcarrier is the
output 0 to 5 volts, of within to the
signal source
containing about the 5000/ohms. instrument signal sensor
information
l/Z0
impedance oscillator
modulates modulated signal
subcarrier in is volts signal power payload power age. power turn is
routed
central
processor whose constant
provided. is excited processor for the
A temperature by a 5-milliampere (ESP). operation from of of watt 29 the of the volts This and
signal from
source within the
engineering Electrical via electrical voltthe
signal thermal bus regulated
is
conditioned of the power and
ESP. is
control electrical voltage
instrument subsystem. 22 volts dc
delivered This
harness consists Operation and The 0.5 defining through
basic dc
unregulated of regulated
instrument unregulated for
requires power. the scientific
approximately
3 watts
documents 61.
payload
are
listed
in
Appendix
A,
items
46
126
SEISMOMETER The and spatial
EXPERIMENT seismometer distribution (e.q. seismic such and
SUBSYSTEM subsystem moonquakes, noise measures (I)number, noise magnitude, level and
experiment of natural background as, effect
(2) background
spectrum or other
correlated, on
if possible, lunar surface and
with
thermal (3) (4) interversus of
sources properties
of temperature both near
materials) at depth;
elastic nal
structure damping of lunar and and
the lunar
surface depth,
constitution and type
- internal and state (number impacts
(Q),
density
versus
temperature
depth,
materials
versus
depth,
and upon
(5) distribution the ability
meteorite ferentiate
impacts between
energy moon
released quakes).
depends
to dif-
Seismometer The ment auxiliary
Experiment seismometer unit, instrument, experiment short suspended coil is period and
Subsystem experiment instrument and figure
Characteristics subsystem electronics 10-14 is co.asists unit. a functional of Figure block the sensor, 10-13 instrushows of the the
seismometer seismometer The gap of a
diagram
subsystem. seismometer magnet, top relative of the to velocity. quakes) acceleration transducer velocity over the itself, amplifier, output In much However, noise. one-tenth by if To and above specified in the for main the consists which coil, stationary The of a coil as mounted the inertial purposes. induces flat its a voltage response to frequency its natural motion, With experiment range. electronics required it earth, a high has and gain two been by the unit, an instruwithin the mass. flux An Motion
permanent wound magnet on
serves for
auxiliary of the
calibration main coil has above
suspended to displacement the
proportional ground and flat
differential (during to the to ground velocity ground is the low small flat sensor noise
subsystem
frequencies (rate output the of
natural
response Since is flat that that gain, very of
movement) differentiates
below ground
frequency. this ment response A increases output amplifier is high the
natural range,
frequency. the overall
above in up the than is not this
frequency the to the instrument levels
located power
sensor
telecommunications that possible could are the moon gain saturate available is is on on
subsystem. seismologically required. background command, is preceded
designing quieter this avoid
amplifier the true,
assumed
maximum amplifier gain gain. changes A with
possibility, maximum low noise
one-hundredth
chopper-amplifier
a direct-coupled,
preamplifier
121
i
CABLE
ACCESS
DOOR
OUTER
SHELL
SUPER NSULATION
MOUNTING (IOF3) DIA 7 in, 7.8 in.
FASTENER
HEIGHT
FIGURE
10-13.
SEISMOMETER
INSTRUMENT
122
FIGURE 10-15.
SEISMOMETER INSTRUMENT AUXILIARY UNIT
127
,
/
Xl.
SPACECRAFT
MASS PROPERTIES
WEIGHT
Total launch
spacecraft
weight
is limited 2150 pounds. II-I
by the A and months meets
capability weight The
of the Atlas/Centaur breakdown payload (as for
vehicle
to a nominal is given
detailed 11-2.
of 19 June each an mission
1964)
in tables
specific of the cg
will be payload
selected combination since shown,
several which
in advance weight of the and
flight to provide This all five
optimum
constraints. including
selection scientific Centaur. Fluctuations given
is necessary experiments Those
the
total weight exceed required
spacecraft, capabilities
may not
the boost for
of the Atlas/ will be of the may be deleted.
instruments details
a particular in periodic weight Status
mission revision status Report.
in exact
of design the current
result
figures
in this table; by referring
however, to the
spacecraft Weight
obtained
current
Surveyor
CENTER
OF
GRAVITY
AND
INEKTIAS radius-of-gyration, major and configurations: retromaneuvers The limits and moment-of-inertia for launch at
Spacecraft conditions are
center-of-gravity, specified deployed for for three
stowed
liftoff weight; and deployed
midcourse at landed defined along
at separated of the
weight;
for
touchdown plane are
weight.
center-of-gravity and the limits last of
travel the
in the X-Y
by the first two the Z-axis are
conditions, by
center-of-gravity s.
travel
defined
the first and
condition
Spacecraft-center-of-gravity weight are shown in figure Ii-I as
limits
in the
stowed
configuration
at launch and
established
by Atlas/Centaur
control
stability
capabilities. ravity are vernier limits after by Surveyor/Centaur the attitude separation for midcour of the The with se
Center-of-g and retro maneuvers and
limited engine
correction
capabilities
flight
control
subsystems
during
retro-rocket
burning. coincides
spacecraft
center
of gravity,
before
retro-rocket
installation,
129
TABLE
Ii-I.
SURVEYORA-ZIA
WEIGHT
SUMMARY Weight* (pounds) 698.81 1432.08 90.01 2221.90 624,72
Element Basic U sable Scientific Separated Landed bus propellant payload weight weight
*Based
on
Payload
Combination
1
TABLE
II-2.
SURVEYOR
A-ZlA
DETAILED
WEIGHT
STATUS Design (pounds)
Item Basic Flight Sensor Inertial Canopus Wiring Electronics Support Sensor Attitude Attitude N Z tank Actuator, Nitrogen Bus control group
Description
Current Weight
698.81 system flight control unit 49.0Z 35.39 8.13 4.9Z 0.95 17.89 3.50 secondary control jets and roll control solar system 0.35 13.28 1.62 9.08 1.08 1.50
reference sensor harness
130
TABLE
11-2.
SURVEYOR
A-Z1A
DETAILED
WEIGHT
STATUS Current Weight 103.86
(Cont) De sign (pounds)
Item Electronics Data link
Description
32.29 planar array A B 8.90 0.32 0.32 0.83 0.48 A B receiver A receiver B command signal decoder processor system 5.44 4.85 34.18 9.64 10.62 sensor 6.60 6.02 1.30 6.78 6.8O 5.73 3.35 and 3.88 and 6.84 6.84 3.88
Antenna, Antenna, Antenna, Rf Rf transfer spdt
omnidirectional omnidirectional switch switch
Transmitter Transmitter Command transponder Command transponder Central Central Doppler Signal Klystron Antenna, Antenna, Waveguide Boost Approach Engineering Engineering regulator TV
altitude/velocity data converter power altitude velocity assembly supply / velocity sensor
camera signal mechanism
4 processor auxiliary
131
TABLE
11-2.
SURVEYOR
A-21A
DETAILED
WEIGHT
STATUS Current Weight
(Cont) Design (pounds)
Item Thermal Altitude
Description assembly radar marking radar
control marking
0.24 8.43 0.92 0.34 3.30 O.6O 54.90 8.50 46.40 28.55
Insulation, Signal Battery Low data proce
altitude s sing
auxiliary
charge rate
regulator auxiliary
Electrical Solar Battery Mechanisms Positioner, Boom Boom
Power
panel
antenna/solar
panel antenna antenna arming A B
24.38 2.20 1.16 0.81 218.46
etc, omnidirectional etc, omnidirectional sensing vehicle basic hardware structure and
Separation Spacecraft Spaceframe, Installation The rmal
59.87 22.54 1.00 38.28 11.63 11.63 11.63
paint gear gear gear gear gr installation I 2 3 rel (3) blocks A
Landing Landing Landing Landing Pin
puller,
0.33 3.06 25.16
Auxiliary Thermal
crushable compartment
132
TABLE ll-Z.
Item Thermal Wire Pneumatic Release Engineering Latch, Heat
SURVEYORA-ZlA
Description B bus
DETAILED
WEIGHT STATUS (Cont)
Current Weight De sign (pounds) 18.06 44.20 0.74
compartment harness, line mechanism, measurement spacecraft collectors propulsion system, to basic s
main
retro sensors
1.50 4.45 0.33 0.78 223.16
Centaur
Spacecraft Propulsion Valve Tank, Tank, Tank, Thrust The rmal Lines Helium Propellant, Rocket Rocket
vernier helium
73.42 3.04 20.31
assembly, helium fuel (3)
9.96 (3) assembly (3) 9.96 18.44 5.62 fittings 6.09 2.50
oxidizer chamber control
and
miscellaneous
unusable main main retro retro retro
4.20 143.04 139.39 3.65 8.00
engine, engine,
Insulation, Contingency Propellant, Ve rnier
main
usable propellant propellant
1433.08 156.50 1236. Z0
Retro-Rocket
133
2.2 1.0
÷Y "I
-FX SEE INTERFACE CONTROL DRAWING
o° 'L
CG LIMITS FROM STATION 89 TO 118
--y VIEW (o) CG ENVELOPE
, /-:,,,,,,,,,,. /::7
+Y I _i'.':".".". C i05 53D + X 36 77D
SECTION PAYLOAD
A-A ENVELOP'E ONLY
I
I
o
1'
viEW I STA _B O0 CONICAL FLATTENED TO CLEAR .NOSE )E_ FAIRING (SEE SURFACE 2 75 NOSE B-B) ,n 2500 A I STA 6DO(REF) , (e) STA 157 22 SECTION LONGERON
I/llk
C SEE iNTERFACE DRAWING GDA FOR LOCAL RESTRICTIONS CONTROL 55-00050 ENVELOPE SECTION C-C DIMENSIONS IN INCHES SECTION6--8
FIGUKE
II-I.
ATLAS/CENTAUR
PAYLOAD
ENVELOPE
the
Z-axis
within
1.0 inch. must
The
spacecraft with the
center retro-rocket
of gravity thrust to coincide the
for the axis with
separated 0.18 center inch of
weight radial. gravity to meet
conditions The
coincide of the thrust
within the
position time
axis
is adjusted with
at the this The
of installation
to coinciae
spacecraft
center
of gravity
requirement. of travel are of the vertical center-of-gravity location in the touchin
limits
down Section
configuration XIII, so the
designed
to the landing will not topple
site assumptions, when landing.
described
spacecraft
134
EFFECTS OF PAYLOAD COMBINATIONS The concept of designing the spacecraft for autonomous operation and removal of one or more of the approved scientific instruments has been outlined in the previous sections. The approved list of scientific experiment for the A-21A series of spacecraft consists of the following: Survey television experiment subsystem (which can include either one or both of two television cameras) Soil mechanics-surface Alpha scattering Micrometeorite Seismometer sampler experiment subsystem subsystem subsystem subsystems
experiment
detector experiment experiment subsystem
The television experiment can be installed in any of three different combinations (i.e., TV camera 3 only, TV camera 2 only, or both TV cameras). When it is added to the other four experiments the total could be considered potentially as seven different possible experiment combinations. However, because of the relative locations of the television cameras on the spaceframe, the use of TV camera 2 by itself is not considered practical. Its capability to view the other scientific instruments is inferior to that of TV camera 3. Furthermore, its location is not as advantageous for center-of-gravity adjustment purposes as that of TV camera 3. Accordingly, the possible television combinations are restricted to I) TV camera 3 only, or (2)both TV cameras 3 and 2. The scheme employed in analyzing weight and balance effects consists of starting with a full complement of scientific experiments and then removing them in various combinations, to achieve particular configurations. In some cases removal of one or more experiment subsystems requires that ballast be added to either leg No. 2 or No. 3, or both, to keep the spacecraft lateral center of gravity within acceptable limits. Ballast is added at the footpad in increments of one pound up to a maximum weight of five pounds per leg. Use of ballast in excess of this figure potentially requires redesign of the leg and extension me chani sin. In some instances proper cg control cannot be achieved by adding 5 pounds at one or more footpads. In these cases it would be necessary to add supplemental ballast on or near the main spaceframe structure to achieve proper cg position. 135
An itemized weight summary for each scientific payload experiment is given in Table 11-3. Table ii-4 illustrates 13 different payload combinations considered. (Those combinations previously studied but which could not be brought into proper balance by adding footpad ballast have not been shown.) When the total dry landed weight of the spacecraft drops it will be necessary to add supplemental ballast (which can be around the geometric center of the spaceframe) to enable final made consistent with the minimum vernier engine thrust level chain ber. below 586 pounds, evenly distributed descent to be of 30 pounds per
As various payload combinations are selected for flight it will be necessary to provide the vernier fuel and oxidizer necessary to accommodate the total spacecraft/payload weight including ballast (if any). The total amount of main retro propellant loaded must also be adjusted for total spacecraft weight.
SPACECRAFT Spacecraft system is Z Z No. so as the axis axis, 1. to The formed ment by Ievel Since tooling plane is hoIes 23.75 whose height is the The make origin three between the on or axes COORDINATE components are denoted length top of the SYSTEM are X, of located Y, the and by Z. a right The Z A X-Y the hand Cartesian axis movement is line the perpendicular of spacecraft direction coordinate (figure along to 2-2) the the leg taken
coordinate negative plane
vertical the
spacecraft. the is to the right at The along
toward positive X axis the
spacecraft, axis
Y coordinate is perpendicular system the bali Surveyor system are the axes being but Tolerances in the X-Y may on plane, used X-Y holes.
center with
Y axis, handed. the X-Y and is geometric plane the no the
positive
coordinate lies tooling the in
plane
center is approximateiy
of
a triangle the adapter. attach-
vehicle origin to locate
Surveyor-Centaur physical origin. point The on origin the
coordinate the legs from roll
spacecraft, in the X-Y
inches yaw the and origin
tooling are as
holes. defined as above. with center of = -16 changes gravity to -19.5 the spacecraft The location in are component withina inches (above X, of Y and the Z axes center location one inch theX-Y of and radius
Pitch, respectively, gravity weight from plane). is
defined change the and
determined,
changes. the origin
Z
136
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137
TABLE
II-4.
PAYLOAD
COMBINATIONS
SUMMARY
Spacec Experiment Payload Combination Number TV Camera 3 Both TV Cameras Combinations PayZo&d Weightp pounds Resulting Center of Gravity, inches X Y Z
raft
Design Total Usable Propellant Weilht, pounds
Results
Surface Sampler
Micrometeorite Detector
BILllast Weight (ff amy), pounds Leg 2 Leg 3
Alpha Scattering
Seismometer
Spacecraft Gross Wolght, pounds
Dry Landed Weight, rounds
Propellant Weight - lb Ve ruierl Retro
I Z 3 4 5 6 7 8 9 10* 11 IZ
X X X X X X
X X X X X X X X
X
X X
X X
90.01 8Z. II
0.10 0. Z0 0.26 0.2,! 0.37 0.39 0.73 0.72 0.47 0.54 0.69 0.78 0.54
0. Z3 0.19 0.35 0.49 0.30 0.62 0.36 0.34 0.67 0, 55 0.4Z 0.23 0.64
0. Z5 0. Z8 0.44 0.54 0.48 0.73 0.81 0.80 0.81 0.77 0.81 0.81 0,84 4. Z0 Z. 50 0.70 Z. 60
1433.08 1416,84 1414.94 1399.50 1398.40 1381.26 1379.48 1394. 1368.9Z 136Z. 1367.80 1368.15 1353.71 Z6 17
ZZ21.90 Z197.76 2194.81 Z171.9Z ZlTO. 2144.73 Z142. Z163.96 Z 126. 2116.46 2124.69 2125.19 Zl0_, 67 38 lZ 37
624.72 616.82 615.77 606.32 607.87 599.37 598.54 605.69 593.36 590. 592.79 592.94 585.86 l0
151.38 149.74 149. 54 148.00 147.90 146.26 147.47 145.96 144.92 144, 144.80 144.85 143.41 26
1281, 1267. 1265.4
7 I
X X
X X X
81.06 73.61 73.16 64.66
IZSI. 5 I:_50.5 1Z35, 1246.7 1233.5 1224.0 1218.0 IZZ3.0 1223.3 1210.3 0
X X X X X X X .X X
61.33 70.28 56.05 5q. _9
X X X X X X X X
q_.88 q_.4_ 48.15
4.80 {. O0
*includes *:',_nciudes
I.
64
pounds pounds
ballast, ballast,
evenly evenly
distributed distributed
about about
the the
geometric geometric
center, center,
to obtain to obtain
minimum minimum
dry-laJrKled dryli&nded
weight weight
o1 of
586 586
pounds. pounds,
h, 00
Note:
Weight
and
balance
assumptions tank located at
are: x y = • 0.00 -3_._0
Nitrogen
Signal
data
converter
located
at x y
= =
I. 00 -Z;.O0
Weights
aqe
as
of
19
June
1964
weight
report
Locations
and
weights
of
scientific
payload
primary
instrument
or
sensor
are:
Weight Y Survey Survey Surface TV TV camera camera 2 3 -31. Z0 6.50 96 ZO ._. 8£.90 79.85 31. O0 {pound 15.96 13.96 g. 80 8)
15.00 23. O0
-Zq. -3_.
Sampler detector
Micrometeorite Seismometer Alpha scattering
7.50 .10,56 IO. 50
- 16. O0 -Zl.00 - 34.00
g6.34 5Z,16 70.87
2. O0 S,80 3. 20
138
XII.
SPACECRAFT
THERMAL
CONTROL
To passive design adequate while low and
obtain thermal
maximum control by
reliability techniques the requirement over the
with have
minimum ben used
complexity, wherever with
passive
and
semi-
practicable. a minimum long wide weight periods range
This penalty,
was
dictated
to achieve, for
thermal on
control a lunar
spacecraft having an
relatively
of time of high and
operating temperatures. lightweight
surface passive
exceptionally control
Although solution and
thermal operational
provides
a relatively
simple
to real-time
thermal
problems, and
it required surface control which of thermal is
the development treatments achieved control these control having through temperature control power
veriffication thermal
of special properties.
insulation
techniques
unique the use by
Semipassive thermal
thermal
of self-actuating varying the thermal active held
mechanical conductivity. heating
switches
Through
utilization and
techniques, consumption
electrical
requirements
are
to a minimum.
METHODS The or unit
OF
THERMAL thermal
CONTROL control system consists to provide of spacecraft of the individual subsystem environpassive surfaces and (2)
Surveyor
thermal for
control
systems, during
integrated all phases
acceptable operation.
thermal The
ments methods
all components include thermal
employed optimum
(I) special absorption to house heaters,
preparation and
of spacecraft
external
to achieve
emissive
characteristics, The paths compartment active
superJ_nsulated employed are
compartments (i) electrical operated acceptable
critical (2) thermal
equipment. conduction the
methods by
controlled
bimetallically tures within The a.
thermal limits. thermal will except
switches
to maintain
tempera-
spacecraft The
control remain during
design with the
is based
on
the following (Z) axis
assumptions: pointed
spacecraft the sun
the positive-thrust following periods:
toward
139
(I) (2) (3) (4)
h.
Before An
initial sun period
acquisition. of 1 hour maximum (full shadow).
eclipse
Midcourse Terminal the will normal be
correction descent transit pointed is 442 at
maneuver.
maneuver. attitude, the Btu/hr-ft of the from sun. z nominal illuminated zenith. of solar The angle (same lunar surface in as during is temperature accordance with transit). +Z60°F is figure at a the positive thrust axis of the space-
For craft
C°
Solar Maximum solar assumed iZ-l.
intensity
d.
temperature angle to of vary 0 degrees as
surface
a function
el
Radiative
properties
of lunar
surface: c = G = 0.875 0.93
Infrared Solar
emissivity
absorptivity
300
ASSUUPT_ONS u/2 ,. DUST :0 g3 B?_ w/fl I LUNAR OEPTl'l TEMPERATURE : $ ft : 40°F OvER HOMOGENEOUS l ROCK
OSO,.ao t,mm : 0 S:I$0 ADIABATIC ADIABATIC
MATF ROCK
RIAL
PROPERTIES OuST
-ISO
i _i) C
I 21 IBT 02
Btu/hr Ill/|l' Btu/IO
"OF'ftl
• 2SI10 ii2 Ih/fl
-4 |
Btu/hr-OF-ft
,
"OF
02
etu/IO
"°F
1
120 SOLAR 240 ANGLE, DEGREES 3GO LUNAqNO(_
FIGURE
12-1.
LUNAR
SURFACE
TEMPERATURE
140
f,
Minimum and
temperature at 65 degrees
of lunar
surface
at night
is -Z45°F
at the
equator
-260 °F
latitude.
141
PRECEDING;
PAGE
_LANK
NOT
FILCHED.
XIII.
OPERATIONAL
SEQUENCE
AND
FUNCTIONAL
PERFORMANCE
The mission craft
expected is described
performance
of the
spacecraft
during
a standard
operational on space-
in the following based The on design
paragraphs. capabilities phases and
Emphasis their
is placed to the are
performance phases. and lunar.
relation mission
mission
operational transit,
three
major
of a typical
prelaunch,
PRELAU
NCH
PHASE
Air
Force The
Eastern AFETR
Test imposes
Range
(AFETR) constraints are on the spacecraft, primarily The such conto
various measures AFMTCP
ensure "General straints
that adequate Range as are Safety
safety Plan,"
provided
to protect
personnel. for
80-?,
provides
the authority constraints
reflected are must
in the
spacecraft as
design.
Typical
affecting
spacecraft a.
design Squibs for
summarized be capable
follows: an energy of 1 watt at 1 ampere
of withstanding firing. system, destruction nitrogen capable
5 minutes explosive be
without destruct for
b.
An must
of actuation propellant tanks)
by
radio
command, by GD/A).
provided vessels
of solid and
(provided must meet
c.
Pressure proof
(i.e.,
helium
minimum
test requirements. and arming must mechanism be provided. such as as those launch listed azimuth above. conditions do not influence to ensure against inadvertent retro-
d.
A
safe
rocket Other the spacecraft AFETR
firing
restrictions as directly
design
Prelaunch The than a few
Operations final
at AFETR countdown and safety check circuits on the Surveyor that are spacecraft wired itself to meet (other
prelaunch power
critical
hard-line
143
AFETR requirements) is accomplished by means of an r-f link. Inherent in this checkout is the operational verification of the spacecraft transmitter, receiver, command decoding, and signal processing. During this countdown, the spacecraft pulse code modulation (PCM) data channels and Centaur telemetry can be monitored to provide the initial-condition values for these channels. Checks of the television and spacecraft transponder operation, as well as the ability of the spacecraft power system to properly supply the electrical load (i.e., with external power removed) can also be made. The success of the mission is dependent on the establishment of definite
initial conditions for some of the spacecraft subsystems. At the time of launch, the spacecraft battery must be fully charged, the marking range of the altitude marking radar selected, the Canopus sensor field of view adjusted properly for the particular launch data, the gyro temperatures within their operational range, and the temperatures of some spacecraft items such as the retro-rocket, vernier propellant, shock absorbers, compartments A and B, flight control electronics, and Canopus sensor at their prescribed launch values. Once the spacecraft is launched, external control of the spacecraft is not available until the initial DSIF acquisition is accomplished. Therefore, spacecraft operational conditions required during this interval must be established during the final countdown. These conditions are as follows:
a. Flight control coast system imparted can must be be phase must electronics be in the must be on, and mode of the the spacecraft so that the
flight control angular from b. The rates Centaur nitrogen
rate-stabilized as a result
to the nulled. inhibited
spacecraft
separation
from
flowing
to the jets
so
that it will be
conserved c. The to be d. One tion
during
countdown amplifiers via and the
and
boost, be
until Centaur on to permit provided
separation. accelerometer by the Centaur. data
accelerometer transmitted transmitter must be on
must
telemetry
channels
its associated time of launch
receiver/transponder to permit and
interconnecto facilitate the
at the
initial DSIF e. The
acquisition. traveling-wave will be ready tube for (TWT) high filaments power must be when on so the
transmitter
that the transmitter
operation
144
preseparation is received
f.
command from
for
spacecraft
transmitter
high
power
turnon
the Centaur. signal processor of PCM must data be operating in the proper and and the mode space-
The
spacecraft
to provide craft
the one
channel during
via
the Centaur boost
telemetry
links
the period
between
the
completion
of the initial DSIF
acquisition.
TRANSIT
PHASE
The a. b. c. d. e. f. g. h. i.
transit Launch DSIF Sun
phase through
of the Surveyor separation.
mission
includes
the following
events:
acquisition. acquisition. acquis_ion phase I. correction Z. descent de scent. retro descent. maneuver. maneuver. and verification.
Can.pus Coast
Midcourse Coast phase
Pre-terminal Terminal I. 2. 3. Main Vernier Touchdown.
descent.
Launch
thru
Separation are included via on the spacecraft telemetry simultaneously be available to permit system the telemetering during this phase via the the of PCM of the spaceof
Provisions and accelerometer This
data same system from the by PCM
the Centaur data will be
mission. craft the
telemetered following data
telemetry spacecraft
so
that it will The
also
separation the boost
Centaur. the
accelerometer
will indicate
vibration For certain
experienced the spacecraft must
spacecraft. properly after it separates from Signals the Centaur, that will
to perform be
operations
accomplished
before
separation.
145
cause these operations to occur are to be provided by the Centaur. operations are described in the following paragraphs.
Required
The landing gear and omnidirectional antennas are launched in their stowed (i.e., nonextended) positions so that the spacecraft can fit within the envelope of the shroud. The landing gear must be extended if the cold gas attitude control system is to operate properly since the attitude jets are installed near the ends of the legs and their moment control capability depends on the legs being extended. Consequently, gas jet actuation is inhibited until the legs are extended. The omnidirectional antennas must also be extended to provide the desired radiation pattern coverage for accomplishing the initial DSIF acquisition; the spacecraft high-power transmitter must be turned on to ensure that a signal of sufficient amplitude will be available for accomplishing the initial DSIF acquisition; and the solar panel must be aligned so that it can convert solar energy into electrical power. Although it is possible to command these functions from the ground in the event that the Centaur fails to deliver the signals required by the spacecraft, the ground-to-spacecraft communication link must be established before this can be done. The reduction of possible separation-induced angular rotation is accomplished by the coast phase attitude control system operating in a rate-stabilized mode. The coast phase attitude control system controls the spacecraft attitude by operating the three pairs of nitrogen gas jets located on the ends of the landing legs, and as indicated in table 13-I, it is the system used for attitude control throughout transit except during the midcourse and terminal descent thrusting phases. In the rate-stabilized mode, the system closes an electrical feedback loop around each of the gyros so that the spacecraft rotational rates about each of the spacecraft axes are sensed and reduced to approximately zero. To keep the gyro gimbals from hitting their stops and to limit the amount of nitrogen used to stabilize the vehicle (approximately l percent of the total amount carried), the rotational rates induced by separation are required to be less than 3 deg/sec. The spacecraft is mechanized so that two signals automatically generated during the preseparation and separation sequences enable the gas jet system. These signals are (1) a signal produced by all three landing legs extending in response to the Centaur command described previously, and (Z) a signal generated by the separation of the spacecraft from the Centaur. As a backup,
146
TABLE
13-1. SUMMARY BY COAST PHASE
OF OPERATIONAL ATTITUDE CONTROL
MODES PROVIDED SYSTEM Phase When Immediately separation re(within spacecraft C entau r of from of Mission Utilized after
Operational
Mode Closed around
Mechanization electrical gyros rates so are
Description loop that provided spacecraft and zero
R ate - s tabiliz ed mode
angular duced the Inertial mode
sensed
to approximately dead-band). so
system
System angular space the
operates position are sensed.
that changes in
in
During
midcourse pre-
of spacecraft The
correction, of retro retro phase, phase, burning
attitude
spacecraft
is controlled in space.
so that
it remains
fixed
during
endesand
tire vernier cent during stant portion and Maneuvers Mechanization mode, applie_l time This except is same as inertial current for the is yaw for
roll, con-
the
velocity for pitch
Before and
midcour
se
that a fixed torquer
terminal
to the gyro
descent
commanded current
from results
the ground. in a nominal for in a
spacecraft the commanded given Automatic acquisition sun
rate of 0. 5 deg/sec time, movement. resulting
angular
Spacecraft sec in yaw
maneuvered and sun pitch sensor
at 0. 5 deg/ in response logic sensor to
Sun
acquisition
secondary appears view.
until sun field of
in primary
147
TABLE BY
13-I. COAST
SUMMARY PHASE
OF ATTITUDE
OPERATIONAL CONTROL
MODES SYSTEM
PROVIDED (Cont) Phase When of Mission Utilized acquisi-
Ope
rational
Mode star
Mechanization Spacecraft sec maneuvered
De sc ription at 0.5 deg/
Automatic acquisition
Canopus tion
in roll until
Canopus
appears
within sensor. Optical c ele s tial) (or
field of view
of Canopus
Spacecraft controlled by primary
pitch by
and
yaw
attitude provided spaceby
Coast
phases
error
signals and
refe renc e
sun
sensor,
craft error Canopus
roll attitude signal
controlled by the
provided
sensor.
commands system nonstandard to
which be shut
either off at
permit any time
the
gas during
jet
system are
to
be
enabled to
or
permit
the
transit
provided
accommodate
situations. solar generated solar panel panel when launch-lock of the solar which motor. is relock by a third to deployment the spacecraft pin panel in When transferred limit limit the switch. switch automatic turn the to pullers axis is is initiated senses and the by two separation roll by axis a limit train position motor relock is is squib-firing from the pulses. Centaur, pin that pulses reached, through position disabled. is to
Automatic These initiate puller. initiates the the the solar pulse use of it DSIF pulses, the The
launch-lock switch of is
unlocking generator, stepping
sensed
a pulse panel generator a solar is
supplies solar the
a continuous panel roll axis the the pulse roll relock
output panel
stepping axis generator logic
When and
reached, After off.
sensed
acquisition,
power
deployment
commanded
DSIF
Acquisition The communication and and must on tracked locate to this the signal. maintained during link by the between the DSIF the spacecraft to the first and permit step and tune the the in ground spacecraft this the have sequence, ground been must to be be the receiver standard, con-
established trolled DSIF to 148 lock
stations As
mission. transmitter
spacecraft If the
signal separation
boost
and
phases
the spacecraft will come within view
mitters band data operating voltage-controlled over one to of the in pull-in the two in the high-power oscillator, spacecraft transmitter the of ground the
of mode,
the
tracking frequency
station controlled one The the
with
one by of
of its
its narrow
trans-
and omni
transmitting antennas. during is receiver. since frequency with the will be
channel spacecraft second until will part its
engineering receiver of the can acquisi-
phase-lock tion within the procedure the spacecraft
ground which range
signal transmitter
tuned This
signal result of by ground in
appears a shift in
spacecraft
transmitter the spacecraft frequency,
frequency, transmitter which in turn
achievement be controlled by the
transponder the space-
operation, craft receiver
will
controlled
transmitter.
With power The time mitter quency
the
spacecraft at the
transmitter antenna feed
operating will typically to change of the
in the high-power be from a minimum the time change
mode,
the
r-f
appearing transmitter
of Z. 7 watts. of launch which to the the transfreof the
frequency
is expected because
of acquisition, experiences will be
primarily during
temperature spacecraft The million
this period, at the time
so that the
transmitter
uncertain
of acquisition. ±Z0 parts per
frequency (ppm)
stability with
transmitter ture
is expected of less circularly
to be than
within
a tempera-
coefficient The two used
0.5
ppm/°F. omnidirectional reception the same but only antennas mounted on for the the space-
polarized for
craft sion.
are
simultaneously antennas do
one-at-a-time throughout They tend are
transmisentire on 471" the of
These
not have
level
of gain
steradians, spacecraft the other. over until tinent only
primarily so At one
because
of spacecraft
shadowing. antenna
installed
that the nulls the time
in the pattern
of one the
to cover will be over
the peaks
of acquisition, and cannot the be
since commanded
transmitter to transmit by one
transmitting antenna is per-
antenna
the other alone
acquisition
is achieved,
coverage Also,
provided since both
antenna
to spacecraft
transmission. the mission, reception. coverage -i0 a gain
receivers by
will be both
operating is of
continuously concern The provide possible upper for
throughout spacecraft
the coverage
provided
antennas
composite
of the two db for
antennas
used
for over
receiving 99 percent
will typically of the the polarizaan
a gain aspect hemisphere the
of at least angles, of the
single
polarization -6 db for dual
of at least and
polarization -7 db for must
over dual
spacecraft,
a gain
of at least however,
tion over
lower
hemisphere.
Transmission,
consider
149
appreciably Thus, for
lower
antenna gain
gain
whenever
the
spacecraft
attitude
is unknown. at least than -30 -30 db db over
the typical
of the possible
antenna aspect
at this time angles, with
is expected no nulls
to be deeper
99 percent
of the
10 degrees There can be
wide. is a constraint continuously on on the amount high power of time in the that the transit spacecraft environment high-power transmitter expected operation temperin
operated
the launch-to-acquisition will result in the
period. power
Continuous amplifier operation within
transmitter tube
transmitter
exceeding
its operating
ature.
Thus,
since must Under Z0
the high-power be accomplished standard
is initiated 1 hour the after
just before launch or
separation, the transmitter to be to high tem-
the acquisition will overheat. acquired power, perature Each within and
conditions, after commanded
spacecraft
is expected is switched
minutes
maximum can be
the transmitter to low power
the transmitter indicates
if telemetered
data
that the transmitter receivers receiver. (I) a noise 13.5 kc for A
is overheated. crystal mixer to have the
of the spacecraft FM
is a conventional typical no receiver greater
superheterodyne-type following width
is expected than 14 db, into the the
characteristics: more than
figure
(2) a bandreceiver, signal is
of no
getting
a ground
command
(3) a threshold centered least 40
signal-to-noise
ratio
in this bandwidth than 12 db, and
(assuming (4) a dynamic is the same
in the passband) db. The stability
of greater of the so that (to which
range as
of at
spacecraft
receivers in the
that of the spaceduring of transat the time
spacecraft craft
transmitters, frequency is the As be
the uncertainty the ground
knowledge must be
of the tuned
receiver
transmitter in the previous the
the acquisition) mitter stability.
same
as
that indicated
discussion
in the case reduced by
of the transmitters, measuring the
uncertainty frequency
of acquisition launch. The the mission static This one
can
receiver
just before
channel
of PGM signals error,
data which
which can
is being
transmitted acquisition transponder
during (such as
this
phase
of
includes phase data
aid in the agc, the
the transsignal,
ponder etc.). the
the
receiver
phase-lock
modulates carrier
a subcarrier The
signal
which,
in turn,
phase-modulates to modulate percentage enhancing the the of
transmitter signal
signal. a low
subcarrier index center
oscillator so
utilized high
carrier
provides power
modulation
that a very
the transmitted probability
is in the carrier receiver
frequency the carrier
thereby signal.
of the ground
acquiring
150
Sun
Acquisition The spacecraft relative next to during can in begin establishing descent in produced control sun sun be response by subsystem sensors. sensor stepped before deployment if the solar is mounted on the solar from panel the (figure stowed Under by spaceit can solar about by the to must the acquire sun transit convert on and which depends, solar lock the on passive (2) energy and accurate The a single jet align into to the sun to (1) establish of so most that it the
vehicle spacecraft faces power, the the
attitude
thermal the solar
control panel electrical reference and
components sun and (3) and aid and
required vehicle acquisition command The with gas signals
a known maneuvers. to
before lockon by a are sequence is by
midcourse
terminal
accomplished of spacecraft by
automatically maneuvers the and the solar flight primary secondary panel must
ground system.
gas in
jet
system provided
controlled the secondary Because 13-i), position normal craft be the
accordance
to its transit sun acquisition of the panel solar
position can panel be
in which conditions,
it is launched automatic
achieved. is initiated
Centaur
separation, by array means
is not
deployed To
automatically, this, the
accomplished
of ground must the
commands. be rotated then
accomplish
panel/planar the 85 spacecraft degrees.
combination and
approximately rotated by up
60 degrees from the
roll axis, These
solar
panel
roll axis will
rotations,
when
accomplished increments the active
ground per
command, When
result solar
in a stepping panel ular craft. At attitude
of the panel
in i/8-degree position, and be
command. of the panel of the the
the
is stepped to the The the
to its transit roll axis can
face
is perpendicspace-
spacecraft spacecraft time the
points commanded
in the direction to acquire command
top of the
then
sun. the sun, its
spacecraft random.
receives The
the
to acquire of the sun from
is expected involves
to be
acquisition
this random by the flight is
orientation control first mined axis
a sequence at a nominal the spacecraft yaw axes,
of rotations rate yaw and
(controlled deg/sec. sun
automatically In general,
programmer) about
of 0.5 axis
the vehicle deterpitch
rotated by until The the the
until the rotated sun
lies in the plane the spacecraft
roll and roll axis
is then
about
is within of the
the primary sun
sensor
field of view. hemisphere. Each hemisphere that quadrant. for the of of
field of view cells of this and
secondary has a view
sensor
is one
the four the
sensor
of one signal
quadrant the
of the upper sun is within rate
spacecraft output
provides are
an output processed
when 0.5
These
signals
to produce
deg/sec
commands
151
@
PLANAR
ARRAY ANTENNA
FLIGHT
CONTROL GROUP
SOLAR
PANI
PRIMARY SUN SENSOR PARALLEL TO S/C ROLL AXIS)
SECONDARY SUN SENSOR SOLAR PANEL) (PERPENDICULAR TO
FIGURE
13-I.
SUN
SENSORS,
LOCATIONS
AND
ORIENTATIONS
152
yaw no
and secondary
pitch
attitude sun sensor
control ceil will in moves signal the
loops is still sun's within is
as
indicated (i.e.,
in
figure sun is
13-Z. not in a yaw field primary to is is of
In the
the upper
event hemi-
that
illuminated be commanded
sphere), will
the eventuaily When
spacecraft result the sun
to in the
execute sensor of is sun the
maneuver view. sun the sensor, optical on parallel pitch and to
which
appearing the field and The and 13-1). has of
view control
a mode top to yaw the of
sun Jn the the error signals
lockonindicate both flight spacecraft angles provided Sun lockon that is the pitch control roll and
generated
switched sensor that nulling system
yaw
channels. group (figure by primary
primary a field of
mounted
sensor axis
view
aligned of in the
Simultaneous attitude sensor. control
accomplished by must certain the
the sun
response
be parts
achieved of the
within spacecraft
an
estimated do not
time suffer
of
1 hour
after damage
launch due
to to
ensure extremely
permanent
high
or
low
temperature.
Canopus With trolled. desired established. The which the to star the is
Acquisition the Before to thrust spacecraft the in
and
Verification locked on of direction the to the sun, only and space, the pitch-yaw descent, roll attitude attitude where must is conit be is
execution a given
midcourse in inertial
terminal the
spacecraft adjusted canopus before will attitude. in while to roll a
is
mechanized launch so within ground roll
to that its
accomplish with field the of to gas jet
this spacecraft view when
by
means locked the
of on
a
star to the is star rate spacecraft
sensor sun, rolled acquiof
appear The
spacecraft automatic the nominal The expected occurs, by
proper results
command via remains the
initiate system on to
the at the (i. e., When is
sition 0.5
spacecraft the spacecraft a star in the star
deg/sec
locked brightness field the roll of
sun. the this controlled
continues nes_ signal signai
until appears by by the the
of the
the
proper
brighta lockon the error
of Canopus) is generated
sensor and
view. attitude
sensor, sensor. sensor field that of the can
developed Although
Canopus sensor to verify
is view,
designed it also on
to
discriminate an the by sensor commanding output
against signal is actuaily the star sensor
stars which
that
may makes indeed to
appear it
in possible
the
provides which
object be roll
locked spacecraft output
is
Canopus. a
Verification complete 360-degree
accomplished while the
perform
telemetered
153
MANEUVER
/
', i
\
......
_.._, ......
.....
+,,,_,
SECONDARY SUN SENSOR AS VIEWED ALONG ROLL AXIS (CENTER OF CIRCLE SHOWN IS ROLL AXIS WITH TOP OF SPACECRAFT POINTING OUT OF PAPER)
LOGIC USED TO COMMAND SPACECRAFT MANEUVERS TO ACQUIRE SUN FROM SECONDARY SUN SENSOR CELL SIGNALS CELLS ILLUMINATED A D NONE A AND D B C B AND C A AND B C AND D COMMANDED MANEUVER
"_ YAW * YAW ._ YAW ._ YAW - YAW - YAW - YAW ._ PITCH - PITCH
FIGURE
13-2.
SECONDARY
SUN
SENSOR
ORIENTATION
LOGIC
154
(star sensitivity swept parison before
intensity) range out by between launch The for the
is
monitored. of sensor the the the Ganopus field positions particular is in in level before in the is not the of
Thus sensor view stars launch
a
map and as on date the
of
all
stars within
having the is
intensities 360-degree generated. on to a map be
in
the
failing vehicle map permits to perform
band A prepared com-
rotates and those
of
this
Ganopus in The a
verified. environment
Ganopus might be while
sensor expected it is
required Van Allen
radiation is become
such
as
belt.
sensor and to
expected operational correction to reference
to
be as is and
inoperative soon not verify required as planned lockon for the
radiation reduced. after to
environment, Since launch, the the first it
radiation to occur to the
midcourse will be possible roll
8 hours time before
acquire
Ganopus maneuvers
establish this
spacecraft
correction.
Coast
Phase
1
The ing with
coast
phases
of the mission to the sun mode
are and
characterized to the star
by Ganopus
the
spacecraft with the
coast-
its attitude on The
servoed
low-power
transmitter tracking. with signal
in the transponder transmitted power
to permit radiated PGM
continuous via
two-way
doppler antenna, at a rate as a coast
will be on
an omnidirectional engineering data
processing with the
equipment
to provide
consistent result phase
signal-to-noise conditions. midcourse is operating gas
ratio available Goast phase
at the ground i occurs
receiver and
of these Z occurs Since the amount this of
operating after, spacecraft of nitrogen the
before,
correction. under for greatest coast conditions the {typically to be used for for attitude 51 most of percent standard The For capability for a amount standard indiof the the misspacefor
sion, craft the
the during
required the of coast
maintaining portion expected attitude
phase the total
represents amount by on of the the gas is
3 ¢ case)
nitrogen phase is to plus
operations of nitrogen
accomplished gas this in table carried quantity 13-Z with
control 4.5 the
system. pounds.
spacecraft expected
nominally provide a reserve
mission, cated situations.
nominal quantity
a 3 0- probability
nonstandard
With sun and
the
spacecraft the
locked solar mission
on
to the
sun,
the on
solar
panel
will be
facing
the
will For
convert
energy
impinging for
it into necessary phase,
electrical space-
power. craft
the typical load
sequence
the transit which
the total
electrical
is in excess
of the power
will be
provided
to the loads
155
TABLE 13-Z.
COAST PHASE ATTITUDE
CONTROL CAPABILITY Time/Number of Maneuver s 30 minutes
Mode/Maneuver Rate-stabilized mode, including dissipation
of separation rates Inertial hold hold) 1 hour, 50 minutes 65 hours 1 1 4
6
5 ;:"
Optical mode (celestial Sun acquisition Star acquisition Star verification Roll maneuvers Yaw maneuvers
_Maneuvers corrections. required
under
standard
conditions
for
two
midcourse
by
the
solar being
panel
via
the optimum continuously demand Thus,
change by the
regulator. spacecraft and the
This battery power be
situation to make being
results up
in
power ence solar
supplied
the differby the time
between panel
the power (figure
of the loads the battery
supplied
13-3).
will not
fully
charged
at the
of landing. Of the total electrical of spacecraft load indicated in figure 13-3, the power A and required compartment The indicated dissipation A At has a typical in figure of 17 steady comshaded sun. Thus, for B
the operation must be within
components capability in the
within
compartment compartments.
the dissipation phase results phases,
of these
standard 13-4. watts
transit During
compartment B has
dissipations a steady
the coast
compartment to 37 watts, peaks than
with
transient
peaks with
of up
while of up
compartment to 75 watts. B and
dissipation partmentA during the
of Z5 watts has coast more phase,
transient area
present,
radiating while more
compartment B
is partially by the
compartment energy in the than
is fully illuminated B results 13-5. During during in the the
it is able The the
to dissipate power
compartment
this period. temperatures coast phase, of these
dissipation trays as
compartments in figure
compartment
indicated
156
Iooo
LAUNCH EVENTS LAUNCH SHROUD EJECT}ON TO SUN
T
ACQUISITION MIDCOURSE OPERATIONS BY CENTAUR -/ / EVENTS : CORRECTION PRESEPARATION COMMANDED MANEUVERS VELOCITY CORRECT_ONOF SUN
I
TERMINAL EVENTS: PRERETRO START TV MANEUVERS DESCENT
I
INTERMITTENT 940---_ TRIGGERS IGNITION _ _ R I I I
FRAMES RADAR ENGINE
MARKING VERNIER RETRO VERNIER TOUCHDOWN
goo
--
SPACECRAFT DSIF ,SUN
SEPARAT}ON
ACQUISITION ACQUISITION
I
'_--_ COA ST PI'-IASE I _"_
f_'_-
REACQUIStTION AND STAR
ENGINE IGNITION DESCENT
|
-_ -_ --COAST SE If
I
I
3O0
S/NADR tECRQIU:SITION
_
;I_GGHNE E RDATANR_T NETERROGATIO N!
A
2O0
O 0.
IO0
CONTINUOUS LOW RATE TRANSMISSION DATA NOMINAL ELECTRICAL LOAD
NOMINAL OCR OUTPUT
UrNA_IRATTERYLOAD
0 LAUNCH
Io
2o
30
TIME FROM LAUNCH,
40
HOURS
50
60
_
TOUCHDOWN
70
FIGURE
13-3.
TYPICAL AND
STANDARD TRANSIT POWER PROFILE
OPERATIONS
temperatures expected pha'ses
are
well
within
the
estimated
allowable
upper
limit during
of
IZ5°F. the coast
The
typical are shown the
performance in table coast in any signal
of other 13-3. the
spacecraft
equipment
During data
phases,
spacecraft data
can
be-commanded (i,e., types
to transmit commutator found on
PCM
continuously engineering are
of the four processor.
arrangements the
modes) these
of the modes
Briefly,
of data
described
below.
157
80
f-1
II
I _-: .......
:
0 I:00 5:00 600 7:00
__._L __Jl__ ._ _ _ __ _ _ ___ ._.i JJ__ _
m -_ TYPICAL 01SSIF_ATION --COMF_RTM(NT A _IN_LU_tES O_R AND STEPPING MOTOR SWITCHES) -CO_PARTMI[NT B
•
12:00 1300 1400 1500 TIME 1600' AFTER
.
2T30 LAUNCH, 3930 HOuRs
.
5230 5330 82:00
I
6300
I
6400 6500 66:00
FIGURE
13-4.
TYPICAL DURING
COMPARTMENT STANDARD TRANSIT
THERMAL PHASE
DISSIPATION
Mode
1 This mode contains most primarily coast and thrust phase attitude data used and for control a few data. propul-
It also sion
contains
of the
telemetered This mode
electrical is expected and
current to be
system
temperatures. sun and
monitoring correction. just the
attitude
maneuvers, be
star
acquisition, during
midcourse
velocity
It will also before proper the
of primary and attitude
interest terminal control
the high-data-rate since this data
interrogation can verify
midcourse of the
maneuvers system.
state
Mode
Z This mode appear used contains on mode the same coast phase and thrust phase attitude control since
data
which
I; however, terminal
the number when
of signal the data
samples
is less via
it is normally high-gain mode also
during
descent
transmission to be used.
the This
planar
array
permits
higher
transmission the retro burning
bit rates phase,
includes and
signals
covering power.
vernier
system
temperatures,
electrical
158
140
T
_COMPARTMENT _COMPARTMENT D COAST PHASE A B
120
TEMPERATURE -+5 ° AT _ w IE w 80 IOO "1"1_0 AT DStF TIME
TOLERANCE ACQUISITION _ 20 HOURS
6O
TO BE DETERMINED TEMPERATURE TOLERANCI
[
J
l
TRANSIT
PHASE
"_--LUNAR
PHASE--..--.----Ie_
I AND SUN
;TAR ACQUISITION VERIFICATION AND
MIDCOURSE CORRECTION
ACQUISITION
120 i:i:i:i:i:i:i:i:i:ii:i:i:i:i:i:ii:!: nW_
:::::::::::::::::::::::::::::::::::::
80
4O
CD
o
-4 0 4 TIME 12 AFTER LAUNCH., HOURS 20
!iiiii_iiiilililiiiiiiiiiiiiiiiiiiiii_iiiiiiiiililiiii!iiiii!
60 65 0 66 I TIME 2 AFTER 3 4 TOUCHDOWN, 5 6 HOURS
FIGURE Mode 3 This for
13-5.
COMPARTMENTS
AAND
B
TEMPERATURE
PROFILES
mode
contains the vernier data and
the signals descent altitude
used
for
lunar This
reflectivity data
measurements by
and
monitoring (RADVS) 4 This mode
sequence. control signals.
is provided
the doppler
radar Mode
contains and the data
primarily link panel
spacecraft
electrical
power,
thermal in the DSIF and
mechaacquisi-
nism tions status
positioning, to position during the
signals. and
It is intended the
for use
solar
monitor
spacecraft
power
thermal
mission.
159
TABLE 13-3.
TYPICAL
EXPECTED THERMAL DURING TRANSIT
Transit Temperatures
PERFORMANCE
(°F) Touchdown Coast Phase (Maximum Eclipse I/Z hour)
Survival Temperature Limits
Subsystem Vernier tem*
Painted pellant Blanketed propellant tanks Thrust chambers Plane valving Barrel/ extension cone Propellant transport line s Heated length Aluminized length Helium tank at
Operating Temperature Limits
Prelaunch Temperature
Acquisition (Maximum Eclipse 1 hour)
sys60 i00 80 to 90 ZO to 60 to 34 to I00
protanks
0 to I00
80
to 90
50
to 85
0 to I00
0 to I00
80
to 90
55
to 90
60 I00
to
35
to I00
0 to I00
80
to 90
45
to 90
ZO to 60
0 to 60
0 to I00
80
to 90
I0 to 50
I0 to 50
I0 to 50
0 to I00
80
to 90
30
to
90
50 I00
to
0 to IOG
0 to I00
80
to
90
Z0
to
90
55 to I00
Z5
to 10C
*Temperatures system
listed
under
touchdown
refer to a time
to those
temperatures
at time 6 minutes
of before
activation
which
correspond
of approximately
touchdown.
160
TABLE 13-3.
TYPICAL EXPECTED THERMAL DURING TRANSIT (Cont)
Transit Temperatures Acquisition (Maximum Eclipse i hour)
PERFORMANCE
(°F) Touchdown Coast Phase (Maximum Eclipse i / Z hour)
Survival Temper-' ature Limits
Subsystem RADVS system ;',_
Signal data
Ope rating Temperatur e Limits
P r e iaunc h Temperature
-40 257 -40
to
41
to 77
100 imum)
(max-
38 80
to
33 to 85
converter Klystron power ply Velocity sensing tenna Altimeter velocity sensing tenna Shock sorber Main retro abanansup-
to
-22 14
to
i00 imum)
(max-
- 24 16
to
-26
to 18
257
- 90 220
to
- 90 220
to
100 imum)
(max-
-123 253
to
-131
to 261
- 90 ZZO
to
-90 220
to
i00 imum)
(max-
-123 253
to
-135
to 265
-50 300 20 105
to
Z0 to 1Z5
85
(mini-
0to
15
80 120
to
15 to 125
mum) to
to
Z0 75
80 to 90
70
to
90
32 to 85
17
toTO
propellant
(30
hours after launc h) Solar panel -3Z0 235 *Temperatures system listed under to - ZOO 145 touchdown refer to a time to those to
80 to 90
-ZOO
to 140
117 140
to
- 140
to 140
temperatures
at time 6 minutes
of before
activation
which
correspond
of approximately
to uc hdo wn.
161
TABLE 13-3.
TYPICAL EXPECTED THERMAL DURING TRANSIT (Cont)
Transit Temperature
PERFORMANCE
s ('F)
Touchdown Coast Phase -166 -I00 to (Maximum Eclipse 1/Z hour) -234 to Z58
Survival Temperatu r e Limits -3Z0 300 systo
Subsystem Planar antenna Nitrogen tem Jets array
Operating Temperature Limits -255 Z75 to
Prelaunch Temperature 80 to 90
Acquisition (Maximum Eclipse 1 hour) -Z53 to 150
-30 160
to
-30 160
to
80
to
90
I00 160 30 ii0
to
T ank
-70 IZ5
to
-70 125
to
80
to
90
to
Spac line s Leg
eframe
-160
to
-160
to
80
to 90
90
lines -50 Z50 to
90
-50 Z50 under touchdown refer to a time to those to
80 to 90
150 Z50
to
,:,Temperatures system
listed
temperatures
at time 6 minutes
of before
activation
which
correspond
of approximately
touchdown.
Any rates: must be
of these 1100,
modes 550, on
can
be
transmitted or 17.Z bits
at any per
of the
available The choice
spacecraft of the at the
bit rate DSIF
4400,
137.5, the basis
second. of the
determined
of the
strength
received
signal
station. Because the transmitter (low power) the of losses power and in cabling available I0 watts minimum are and tolerances on the transmitter the nominal output. As the power I00 Table shown, coast output, milli13-4 at phases.
at the antenna (high power expected power)
varies at the
from
watts
transmitter antenna. during
indicates least 24
expected
available to be
at the available
milliwatts
of power
162
TABLE
13-4.
POWER
INTO
ANTENNAS Range of Values Nominal 45 milliwatt s
Typical Operational Low-power omnidirectional Low-power planar array transmitter, antenna transmitter, 4.06 watts 2.69 watts Mode transmitter, antenna transmitter, 36.2 milliwatts Minimum 24 milliwatts
58.8
milliwatts
High-power omnidirectional High-power planar array
4.
5 watts
5.88
watts
Based DSIF System
upon
the DSIF
system
parameters dated in table
documented 29 July 13-5. 1963,
in Surveyor the expected
Spacecraft/ quality of
Interface data obtained
Requirements is indicated
telemetry
Midcourse The sufficient primary As eters,
Correction soft landing time goal a to obtain
Maneuver of the Surveyor lunar data spacecraft via the Goldstone at a desired DSIF lunar location station with is the
tracking
of a Surveyor of expected
mission. dispersions in the launch be required vehicle injection the paramspacecraft miss from
result
a midcourse The
correction purpose spot
will normally
to correct the
trajectory. the desired and to adjust
of the correction on the moon, time
is to minimize
expected of
landing the
to optimize
the probability
soft landing, from the
spacecraft station
of flight so during
that the desired and subsequent to the
visibility to landing.
Goldstone
tracking
is achieved increment
In general, correction so-called
the velocity of two and
applied The primarily
spacecraft
to effect will be The
the
will consist critical plane
components.
first component for the miss.
in the
will correct
second
163
TABLE 13-5. Data PCM (at lunar distance) Accelerometer channels (Transmitted during retro burning)
EXPECTED QUALITY OF TELEMETRY Signal-to-Noise 8
+l db (rms signal at input
Ratio
to to r
Re sulting Bit error
Quality rate of
rms
noise)
approx.
3 x
10 -3
subcarrier >46 rms db (Peak
disc riminato signal at output discriminaproducing of the to of
<0.
5 percent due to
noise)
distortion noise
subcarrier tor for peak subcar
signal
deviation rie r (rms noise
Touchdown Strain Gage Channels
Z0 db rms
signal at output
to of
i0 percent tortion noise due
disto
subcarrier tor for peak
disc riminaproducing of the
signal
deviation
subcarrier)
TV
Z4
(Peak-to-peak noise level of
signal
to rms ance
at lumin100 foot
Lamberts)
component following a.
will be normal soft The landing main and retro
to the
critical
plane
and
will
be
sized
to optimize
the
time-of-arrival burnout velocity
constraints: must be between and limits imposed radars by and the the
operational mechanized b. A sufficient
ranges
of the doppler
velocity
altimeter
range/range-rate amount of vernier control thrust The for as the can can
descent fuel must be be exerted
curve. be reserved to ensure retro that the phase during of the termito ensure
required and
moment
during
the main until fuel and be
that the desired descent. used (defined
maintained of reserved correction must
touchdown in excess the nominal enough
the vernier amount nal
quantity midcourse
to be
descent
the fuel
margin)
large
164
that a s o f t landing c a n b e accomplished f o r the 30- v a r i a t i o n s in specific i m p u l s e , burnout velocity, and a l t i t u d e , e t c . c. The incidence angle of the unbraked velocity m u s t be l e s s than s o m e maximum value (nominally 4 5 d e g r e e s ) to e n s u r e that the m a r k i n g r a d a r
will operate properly.
This c o n s t r a i n t will a u t o m a t i c a l l y be s a t i s f i e d
b y s e l e c t i n g the landing s i t e p r o p e r l y (figure 13-6).
FIGURE 13-6.
POSSIBLE LANDING SITES
165
d.
The time at which the spacecraft arrives at the moon must be earlier than a minimum allowable time (nominally 3 hours) before the last lunar visibility at Goldstone to ensure that there is sufficient
visibility at Goldstone can be
time to
so that
conduct TV surveys after touchdown. than nominally Z hours after the first
the command from The because The actual Goldstone. midcourse velocity in the desired correction and execution of the
This time must also be later
descent accomplished
terminal
may
differ and been
from
that desired control systems. Specifi-
of uncertainties in the
spacecraft landing
attitude has
velocity
uncertainty 30240D, the limits
location Design
specified " and
in JPL this,
cation sets
"Surveyor on 13-7 the
Spacecraft
Specification, errors. at the moon
in turn,
midcourse
mechanization miss A
Figure of the cuted
shows
the estimated magnitude. results 15 minutes these assumed
(99 percent)
as
a function exeif the
midcourse 15 hours
correction after injection within for
30-meter-per-second in a miss the start of 47
maneuver kilometers
typically after
correction maneuver. miss
is applied Therefore,
of the initial pre-midcourse (which are typical), the
conditions illustrates
is within
60 kilometers.
This
figure
that for
a given
correction,
160 (n nr w Fw O .J GYRO DRIFT TIME
I
: 15 rain RADIUS ON OF OF CIRCULAR MISS DISPERSION 990/0 EXECUTION OF
LUNAR SURFACE ALL POSSIBLE
CONTAINING MIDCOURSE AS FUNCTION MAGNITUDE
120
--
i
ERRORS IS MIDCOURSE
PLOTTED MANEUVER
8
b80 TIME {n o3 OF CORRECTION '20 hr$ AFTER INJECTION (20 hr MANEUVER NOT PLANNED-SEE TEXT) 15 hrs 40 AFTER INJECT__
i
bZ tu (.) _r W 0.
30 MIDCOURSE VELOCITY INCREMENT,
60 METERS PER SECOND
90
FIGURE
13-7. MANEUVER
TYPICAL MISS MAGNITUDE
AS FUNCTION FOR NORMAL
OF MIDCOURSE IMPACT
166
the (e.g.,
expected Z0 hours
error
will versus
decrease 15 with hours) time.
if
the
correction the sensitivity for be larger for be since the
is
applied of same for the
at
a later
time to a given of at in Goldstone no longer 20
since However, miss will
trajectory the
correction correction hours 1965 tracking be visible than and
decreases required one 1966, station from applied
reason, a correction ascent from
amount applied
for at
a given 15 hours.
Further, cannot
direct commanded the
trajectories the will
a midcourse at 20 hours
maneuver after injection,
spacecraft
Goldstone.
The amount ing the
maximum
midcourse be
correction for
that can
be
executed
depends without
on
the
of fuel ability
that can of the
allotted
the midcourse
correction
jeopardizthe space-
spacecraft with
to achieve vernier and terminal
a soft landing. fuel to permit
Typically, up to a 30
craft second
will be
loaded
sufficient accomplished during
meter-perthat impact
correction
to be
to ensure descent
a 99-percent for a nominal up to 45
probability unbraked degrees The 30 (See
sufficient velocity Appendix mum second time For of
fuel is available Z690 item meters 5) and correction B, per
second,
incidence time
angles
B,
a nominal
of flight of 66 exceed
hours.
actual
maxi-
midcourse (See
capability item 3) value
could on
the nominal depending
meter-peron the actual
Appendix
any
mission, and
of flight, example,
the unbraked figure available 13-8 for
incidence illustrates midcourse burnout
angle,
the unbraked relationship and for a
impact between
velocity. the amount velocity and
the typical correction velocities on
of propellant for various
the unbraked specific time
impact
fuel margins incidence of impact correction descent of Z686 until B). the fuel and the Clearly, actual of for Z670 fuel angle.
and
of flight which
unbraked as for a function midcourse
Also
.indicated the amount satisfy of per burnout of all
this figure of propellant constraint For midcourse is
is a curve the of
depicts, can use
velocity, and 99 still percent meters margin the unbraked the for time all of
spacecraft
the cases.
reserving for
sufficient impact can straight is a be used
terminal velocities in lines function is also each A
example, propellant reached
second, velocity allowable velocity. and but incidence
case and of
(follow correction midcourse there is no
amount impact flight
midcourse Since the angle,
capability specific will
a function capability for To each
maximum exist
missions, on
a particular parameters. maneuver,
maximum
capability
mission,
depending
mission
perform to point
a midcourse in the proper
correction inertial
the
spacecraft
must
be
com-
manded
direction
arid to operate
its vernier
engine
167
8o I
6O 8O
o z
I
SPACECRAFT VERNIER VERNIER WEIGHT ENGINE ENGINE 2100 149 Ib Ib FUEL
Isp 300 sec MAIN RETRO BURNOUT FUEL VELOCITY
MARGIN
8
6O
-
/7"- zo
V
'>'.
"
/\
/
'./ 'y
\/ '
,5o
z w z >-
,Cd
2O
,///.,/.Z
(z
8
r_
/'/7 POST
./,,
z / / \>,\ ,%dT,o '_£TMot;:E:E,=_ 7 X/';' I
2660
MANEUVER IMPACT
2600
2630
MIDCOURSE
2690
VELOCITY, METERS PER
2720
SECOND
2 rSO
FIGURE
13-8.
TYPICAL
MIDCOURSE
MANEUVER
CAPABILITY
system maneuver engine for
to achieve
the
required the vehicle the
velocity
correction. so
The
first
step axis
in this of the vernier direction provides All of multiple used to the
is to orient system (i.e.,
angularly Z-axis) coast phase
that the
thrust with control pitch,
spacecraft The
is aligned attitude in yaw, rate one
the
desired
the velocity for
correction. maneuvering are
system and roll. and method
capability these angular command any one
in either
direction
maneuvers maneuvers the of the
performed performed to make is as about
at a nominal serially, an angular the for
of 0.5
deg/sec, The
are
at a time.
spacecraft three axes
change
of a desired
magnitude
about
follows: that axis
spacecraft the specific
is commanded time rate interval
to maneuver required for
in the desired it to maneuver A taining
direction through
the desired
angle by
at the fixed sending
of 0. 5 deg/sec. command This con-
maneuver the 10-bit
sequence binary
is initiated
a quantitative maneuver
equivalent
of the desired
duration.
168
magnitude and with is also the
is stored transmitted
in a register back
by
the
spacecraft receiver
flight control where When it may
programmer, be compared that the to
to the ground from
magnitude has the
commanded
the ground. magnitude, or two
it is verified then be commanded
spacecraft execute
received
the proper
it can yaw)
desired
maneuver programmer
(roll, pitch, contains
in the proper rates which
direction. permit it to
The count
spacecraft
clock
the number rate
of seconds provided Thus,
stored
in the programmer. storage angular
For
attitude
maneuvers, register com-
the clock is 409.6 manded 204.8
is such the
that the nominal maximum by single
capability change
of the be
seconds. and
that can
controlled Larger the two
automatically single angular time
the spacecraft maneuvers the earth. or roll-yaw) There can
programmer be executed
is nominally only by manu-
degrees,
ally controlling Normally, spacecraft storage magnitude serially. The plished commanded the thrust before
maneuver maneuvers
from
(roll-pitch
are
required
to orient command
the
applyin_ aboard
midcourse the
thrust. and
is no
multiple can be
capability
spacecraft, these
the programmer must
only
store
one
at a time.
Consequently,
maneuvers
accomplished
velocity
correction similar for by
applied
during for
the midcourse the maneuvers. As
sequence The
is accomis
in a manner to thrust is provided a constant magnitude
to that used period vernier
spacecraft in table
a desired the three
of time. engines
indicated the g for
13-6, to
to cause
spacecraft
experience time. verified The
acceleration of the desired similar
of nominally thrust time
0. II earth is sent the
the commanded and
to the maneuver
spacecraft time flight
in a manner case
to that used correction,
to verify the rate
magnitudes. control storage to
In the programmer capacity provide At control The the
of the velocity rate
second is such
available
clock of the
is employed. nominally of up
This
that the maximum time is sufficient
register
is 51. Z seconds. to 55 meters for per the
This second.
a nominal the end
velocity
of the commanded provides can also be
time the cutoff
midcourse for shutting
thrusting,
the flight engines. situtation,
programmer signal
signal from
off the vernier abnormal
cutoff
commanded
earth
if, in an
programmer
should
fail to provide
this signal.
169
TABLE 13-6. Operational Midcour se velocity correction Mode
VERNIER ENGINE THRUST CONTROL MODES Desc ription of Mechanization Vernier engine thrust servoed to provide a fixed vehicle acceleration for the commanded time interval. The vernier engines are throttled differentially to correct for main retro thrust misalignment. Two commandable thrust values (ZOOand 150 ibs) are available. However, the flight control electronics will override these values when the main retro misalignment requires compensation thrust from the vernier engines different from these values. Phase of Mission When Utilized Midcourse maneuver
Main retro
Retro burn period
Main retro separation
Constant total vernier
thrust
Retro separation
of near maximum value (typically >Z80 pounds) is provided for an interval controlled automatically by the flight control programmer. Thrust to mass ratio maintained constant at nominally 0.9 lunar g. Vernier descent Vernier descent
Acc ele ration control
Velocity control by range reference
Thrust is controlled
by the
doppler radar signals to approximate an optimum (minimum fuel consumption) descent trajectory.
170
TABLE 13-6.
Operational Constant control Mode velocity
VERNIER ENGINE THRUST CONTROL MODES (Cont)
Description Thrust constant nominally of Mechanization so that a of Phase When Final descent of Mission Utilized vernier
is controlled spacecraft
velocity
5 fps is obtained.
Coast
Phase Following
2 execution of the midcourse the sun and star, velocity thereby for the correction, returning reacquisition of the the the spacecraft is
commanded optical to be those in the star as
to reacquire control mode.
spacecraft sun and
to the star as and and
It will be either by
possible
of the same
accomplished performed reverse it did
(i) performing
maneuvers but
magnitude direction the sun
before order, or
the midcourse (2) by sun
thrusting, the
in the opposite to acquire
commanding and star
spacecraft
in the original
acquisition.
Pre-Terminal
Descent
Planar
Array The
Positioning planar result The array antenna is stepped by after earth commands to the maneuvers commands,
high-gain that will
position are with
in its pointing
to the earth
the pre-retro by ground
performed. each command
stepping resulting
of the antenna in a i/8-degree the planar maneuvers
is accomplished step. array so positioning amount
It is desirable execution which reducing completed. ily as the of the system the number The
to accomplish descent
prior
to the during
terminal is placed of TV
that the is kept taken must and
of time
in the inertial pictures
mode be
to a minimum after be
without are primarbetween
that can
the maneuvers changes to vary
angle of the
to which unbraked 68 degrees the array
the planar incidence for would
array angle,
stepped
a function 38
is expected of ±45
approximately (For vertical
and
incidence have )
angles
degrees
off vertical. 5Z degrees,
impact,
to be
stepped
approximately
involving At (nominally angle
approximately 48.5 degrees,
416
commands. array cone) array
the planar half-angle the planar
will begin of the Canopus
to obscure sensor a cone
the field of view sun channel. At an
a 5-degree degrees,
of 68
will permit
only
of approximately
171
0. 75 degree half-angle clear field of view. With a restricted field of view, the Canopus sensor may loose the sun reference, resulting in the roll attitude of the spacecraft reverting to the inertial mode. The approximate time required to step from 48 to 68 degrees is 80 seconds, and the amount of time required to send and verify the commands required to initiate the first roll maneuver is estimated to be approximately 155 seconds.
Thus the system could be on inertial for almost 4 minutes in addition to the presently planned time if the complete stepping sequence were accomplished for a mission where 68 degrees of stepping were required. It may be necessary to step the array out in two sequences. The array could first be stepped out to an angle which would ensure no shadowing of the Canopus sun channel (e.g., 45 degrees) before the maneuvers are performed. Then, any additional stepping that is required in excess of the initial stepping could be accomplished subsequent to the maneuvers. The time required for this additional stepping (a maximum of 9Z seconds) could reduce the time available for obtaining approach TV pictures prior to retro-rocket
Maneuvers Before are performed. the retro-thrusting The (nominally main (a retro-engine roll) causes descent a roll period, sequence and a pitch is or with array bandwidth missions PCM are can A to the on the thrust where and obtained. be accomplished in command and from by the proper spacecraft gas jets the same representing receipt of telemetry. at a nominal manner the this as to three shown yaw the point spacecraft in figure maneuvers 1-3. align vector. the earth for lighting the so The the The that normally first thrust last the necesof permit) the two axis
ignition.
maneuvers of the
maneuver) velocity to
approximately the planar
maneuver sary approach and high-rate
telecommunication TV pictures (4400 and Each of retro these bits burning three (for
information those per
required lunar accelerometer
transmission conditions
second) phases
data
during
pre-retro
maneuvers maneuvers. is is sent verified by
the
midcourse maneuver by the
correction duration spacecraft is of 0.5 then
quantitative spacecraft, the ground provided
desired command The angular
maneuver rate
executed
deg/sec.
172
Approach The approach the last the time The miles first from Since transmitting mutator
TeLevision spacecraft TV camera can be commanded interval to take between the up to I00 that array TV the pictures spacecraft with the
in the
time
executes and sequence. than I000
roll maneuver that the pictures
to point marking taken
the high-gain radar when the
planar
toward the retro
the earth ignition no less
altitude will be
signal
initiates
spacecraft
is at a range
the lunar the present TV video the These
surface. design signals picture blocks of the spacecraft does with will be will be status are picture PCM not provide data capability of
simultaneously sequencing of pictures
of one normally
of the comconsisting in which be moni-
modes,
TV
in blocks
of l0 pictures. engineering tored, will be predicted but not data
separated of the being
by
periods can
is telemetered when necessary
so that the commands last TV
spacecraft sent. as The
particularly commanded time so close
spacecraft to the to occur, can verify taken surface. of taken
to transmit at which as
the
at a time trigger of PCM
close
the altitude-marking-radar the transmission It is estimated range
is expected data which
to jeopardize
that this trigger before retro The the at approach 1000 miles by
signal
is generated.
that the last picture miles 13-9. from The the lunar
ignition
will be TV
at a nominal is shown 6.4
of 80
approach TV
geometry
in figure by 6.4
field of view
camera
is nominally than 180
degrees on each
so that a picture side. impact The
will be
greater
kilometers
coverage The smear control
provided pictures because system
the field of view in the pre-retro angular rates
is a function period are
of the unbraked expected while to have
angle.
taken of the
a negligible attitude
of the vehicle
the cold
gas
is still controlling In addition
the vehicle
attitude. it may be possible Any because and to obtain pictures of the are also a few during of to
to the pre-retro when phase the
pictures, sequence
additional the the have Also, retro main
pictures burning
landing
permits.
taken effects
will be and TV
subject
to degradation exhaust because and gases gases
retro-engine
vernier lines
engine of smear retro
expected thrusting. subject to have
a minimum any picture caused lines
of five taken by
of effects separation and
of retro will be
between vernier
burnout exhaust
to one
degradation to four TV
engine
is expected
of smear.
173
OFFSET
ANGLE AT I000
mi
.0 OF VIEW 5.4 deo
vELOCITY VECTOR AT RETRO
RETRO ALTITUDE
3UMERA AXiS AT IO00 mi SLANT RANGE
LANDING SITE
LUNAR SURFACE
FIGURE
13-9.
APPROACH
TELEVISION
GEOMETRY
It will probably when the spacecraft be pointing those
not
be
possible
to obtain
TV
pictures control
after since
retro the
separation array
attitude
is switched
to doppler
planar
will not
to the earth. resulting by the amount is set by can by be in night moon landings during as well or as some day terminal can landings,
For the
launches will be The
spacecraft phase.
shaded
part
all of the the
descent be shaded
maximum
of time two
during
which
spacecraft
during that the is limited
this phase spacecraft primarily not installed phase
different on inertial amount controlled may be drift
considerations: hold (i.e.,
(I) the amount relinquish and sun
of time lockon) craft on The the
placed
the expected
of gyro
drift,
(Z) spacedepend
equipment spacecraft
in thermally orientation
compartments and become 30
which
coast
damaged
inoperative. Coinis
limitation
attributable
to allowable due to thermal
gyro
is nominally
minutes. equipment
cidentally,
the constraint
limitations
of spacecraft
174
also approximately 30 minutes. Table 13-7 shows the spacecraft equipment whose performance would be affected by being shaded during the terminal descent phase. This includes the vernier system, flight control system, landing leg shock
absorbers, signal data and the solar and panel. the The most critical items are the doppler radar converter shock absorbers.
Lunar
Approach To reduce
Altitude
Measurements of false alarm, the altitude slant (Z615 marking of IZ0 radar ±45 will miles.
the probability until the
not be
enabled For
spacecraft of impact
is at a nominal velocities
range
the design
range
to Z69Z before latest assuming
meters/second), the radar marks time) the to
the number will nominally 257
of measurements v_ry from 36 and
provided (for highest earliest
in the interval velocity and
enabling that marking
(for lowest exceeds
velocity the agc
enabling
time), that
received radar is
signal enabled.
threshold
at the time
the altitude
Measurements descent the four ments phase doppler by
of lunar measuring radar beams. descent,
reflectivity the received There assuming
can
also
be
obtained of the be
during radar
the
vernier in
signal should that
strength
return
nominally the radars
a total of 5Z0 locked on
measure-
during
vernier
are
at the time
of burnout.
Terminal The and
Descent terminal descent phase comprises main retro descent, vernier descent,
touchdown.
TABLE
13-7.
MAXIMUM
SHADE
TIME
FOR
SPACECRAFT Maximum Time
EQUIPMENT Allowable in Shade hours)
Spacecraft Vernier RADVS Helium Shock Solar engines signal gas
Equipment (3) data converter
(approximate 0.8 0.5 0.8 0.5 0.8
supply (3)
absorbers panel
175
Main
Retro The
Descent events The that constitute sequence radar when surface, After expected are the automatic by retro sequence signal are illustrated by the of 60 in
figure
13-10.
is initiated the
the trigger is nominally
provided
altitude-marking miles from
spacecraft
at a slant by
range
the lunar
and
is controlled delay and main
automatically and
the flight conwell in
trol programmer. advance of the
a prescribed marking ignited. engine is turned time The
(commanded
verified
stored retro
in the flight control engine is ignited After
programmer), a 1 second delay of
the vernier delay 0. 55
engines
after
to permit second,
the vernier the RADVS
thrust on.
to stabilize.
a further
i
o
W d t.) O
4
p-
_A 1.)
i
TIME
(NOT
TO SCALE)
FIGURE
13-I0.
TYPICAL
RETRO
SEQUENCE
OF
EVENTS
176
The total thrust provided by the vernier engines during the retro burning phase can be commanded in advance of the retro period to one of two values, typically 150 or ZOOpounds. The lower level is provided for use on those trajectories where the impact (lunar approach) velocity is low and hence the burnout velocity will be low. By reducing the vernier thrust, the burnout velocity can be increased to a more acceptable value. The descent trajectory design is based on the assumption that the doppler velocity and altitude radars cannot be used reliably while the main retro-engine is burning. Thus, the spacecraft attitude will be controlled so that the attitude will remain fixed inertially throughout the retro phase. The main function of the vernier necessary controlled engines during this period is to provide the
moment to accomplish this control. Where possible they will also be to maintain the total vernier thrust at a constant (i.e., 150 or ZOO
pounds) value. The thrust phase attitude control system is mechanized so that the moment demand overrides the thrust level command when the two are not
compatible. Moment vernier engine engines can The pounds As and be control about swiveled about the roll axis is provided by swiveling one of the The
a radial
line perpendicular ±5.5 by
to the vehicle
roll axis.
approximately provided removal
degrees. retro-engine of the incoming by an is nominally spacecraft inertia 9000 velocity. switch. 3.5 sepaThe g
average results burns when
thrust in the out, the
the main
of the bulk in thrust acceleration control
the grain senses
the decrease spacecraft
is sensed has
switch and
decreased
to nominally the retro
provides
a signal This
to the flight signal
programmer
to initiate engine thrust spacecraft
ration
sequence.
results
in the vernier level at the so
level
being
increased experiencing the
to the maximum the maximum In addition, delay
programmed deceleration
that the
will be thereby down. aiding After
time timer the
of separation, begins main to count retro causing
separation. time
the flight control duration
a fixed
of sufficient the blown
to allow delivers
thrust the
to reduce
to a negligible separation The sufficient signal from nuts
value, to be
programmer apart.
a signal
retro
flight control to permit permits the
programmer retro-engine spacecraft velocity
continues to clear pitch and
to count the yaw
down
and,
after
a delay arming
spacecraft, attitude the doppler
provides control to be
an
which inertial
the
switched
to doppler
reference
when
velocity
reliable
177
signal signal
is is
generated. generated,
If
this
signal will be
is
already switched
present immediately
at
the to
time doppler
that
the control.
arming
control
Burnout The by the
Conditions spacecraft retro level for As velocity and altitude the at burnout are velocity and the determined at main primarily retro ignition, the
the main thrust
ignition
altitude, by the main
spacecraft
provided
retro-engine and burnout increment
the vernier weight
engines, at retro sum
duration ignition.
which shown
this thrust in figure velocity, gravity retro
is provided, 13-11, the
spacecraft
velocity
is the from
vector
of the retro
initial spacecraft phase, and Since mately applied retro
the
velocity
resulting
the main
the lunar the main
term. phase velocity by increment the main is supplied retro and to an vernier at the approxiengines) time of main is a
constant
total impulse nearly
(provided opposite value
in a direction ignition,
to the spacecraft retro spacecraft
velocity
the absolute only
of the main of the
phase
velocity start
increment of terminal
function
primarily
of the weight
at the
MAIN
RETRO
PHASE
VELOCITY GLE, 8
INCREMENT,
_V
VELOCITY
AT
IGNITION,
GRAVITY
TERM
t b qm
t b : DURATION OF MAIN RETRO BURN BURNOUT VELOCITY, _b qm: ACCELERATION TO MOON DUE
FIGURE
13-Ii.
DETERMINATION
OF
MAIN
RETRO
BURNOUT
VELOCITY
178
descent. course determined primarily midcourse. figure
This correction. by
weight
is Since
determined the spacecraft impact the unbraked
by
the
amount at
of
fuel retro the and
used ignition burnout the fuel
in
making is
the
mid-
velocity velocity impact vector, velocity occurs
essentially is thus at shown in
the
unbraked of retro
velocity consumed locus
a function Main 13-12. The main descent spacecraft 13-12. If the main engines landing there Along values Control of This retro
burnout
nominally
on
the
burnout
burnout phase to follow which
constraints follow. optimum to at a point
are During
determined this phase,
by
the it is
requirements desired to indicated decelerathrust the too effect a curve, far soft
of conin
the trol figure tion. the and the
vernier the
the
slant-range-velocity a gravity below to other supply locus, amount the vernier of return hand, of turn the the if with curve, spacecraft burnout fuel
contour a constant maximum to occurs to
curve retro will will
corresponds burns not result. be out
of
vernier a hard curve,
sufficient On the
above
will with of the and
not the
be
an
adequate burnout and during
vernier
landing. different
nominal velocity
typical midcourse phases the These operate
3 _ dispersion fuel is used
ellipses are shown. by must are also accuracy altimeter must should by vertical together the
for
impact spacecraft altitude ranges The than doppler
accomplished conditions
the be shown at will funcbe
doppler within in figure the
velocity operational 13-12. greater at the
radars. of these
Therefore sensors. will 700 fps. not In
burnout
constraints within the the desired
radar
velocities not tion below ment velocity assumption burning. the ratio that a lunar in the certain the function before these that
nominally above descent'curve The which not minimum the exceed cannot 850
addition, Since the burnout
doppler radars velocity is with the set
velocities vernier
fps is
(nominal). reached,
these
burnout velocity
values. the angles must that The surface doppler the former and
allowable roll 75 used
requireand with the the is
spacecraft 45 and be
axis
makes
vector
degrees, reliably that the a
respectively, while three the doppler high from beams can be main
radars
retro-engine beams intercept
restriction guarantees channels. The along so retro is that
ensures that there latter at the burnout. with degrees the of least
is
sufficiently arises the roll doppler axis
signal-to-noise the be aligned requirement greater with the than
restriction one of
velocity minimum vector
component value following
spacecraft
velocity
Since velocity must
the
roll also
axis be
aligned 45
velocity the vertical
vector and
at 75
burnout, degrees
the of the
burnout roll
within
179
6O
/ I I
5O DOPPLER LIMIT "_" NOMINAL BURNOUT LOCus
I
I I I I
I
99% DISPERSION ELLIPSE NO MIDCOURSE CORRECTION 4O ALTIMETER LIMIT FOR / / / _ _ / _(
/
/
/\
f
_ MAX,MO. M,DCOORSE
/ 99% DISPERSION ELLIPSE FOR /
/
--" / \'
\
\ /
CORRECT,O._ \ _
/
,/
z (I 3O
/
/
Ii
F-
I
\\1 /
/x\
\_
ii
/
2O
i I\\
\
DOPPLER
/ /
/ /
/COMMAND
DESCENT TRAJECTORY
4 VELOCITY, HUNDREDS OF FEET
6 PER SECOND
8
I0
FIGURE
13-1Z.
VERNIER
DESCENT
PHASE
180
axis
at burnout. can be
The
minimum by
nominal proper descent unbraked
burnout choice 6 = 0). impact nominal and greatest
velocity
consistent offset
with angle
the above for the
constraint main retro
minimized
of the thrust This yields and
thrust
(for vertical the lowest The impact
a nominal smallest for
burnout spacecraft the combinamust velocity be is
velocity weight
of 270 at retro
fps for
speed burnout
ignition. unbraked 700-fps
maximum speed
velocity
tion of highest chosen typically
spacecraft This
weight
so that the 525 fps.
constraint
is not violated.
nominal
Vernier
Descent Following the separation reliable of the (range main retro-engine range-rate) engine from the spacecraft and reaching and the
before optimum
obtaining descent
radar curve
and/or
signals thrust
(fuel-wise), ratio
the vernier equivalent
is servoed 0.9 lunar g.
to mainIf the is
tain a constant doppler switched by velocity
thrust-to-mass reliable signal velocity
to nominally and yaw
is present, reference,
the pitch and
attitude attitude
control
to the doppler
the vehicle to the vehicle
is controlled (V x and Vy)
servoing zero. When
the components
of velocity
normal
roll axis
toward
the optimum
descent down i000
trajectory the desired feet,
curve
is reached,
the thrust
is concurve. to
trolled At an
to bring altitude the
the vehicle of nominally factor is 5 fps to the
preprogrammed altimeter At
range-velocity will provide
the radar system.
a signal
change the
scale control
of the doppler
a velocity reference. and
of nominally A constant
10 fps, velocis
thrust
switched is then inertial
to the doppler commanded, hold mode. phase to be the
velocity and
ity of nominally switched back
the pitch
yaw
attitude
control
During permitting ing the
the vernier spacecraft
descent
spacecraft and
will transmit providing data
PCM for
data, determin-
operations
monitored
lunar
reflectivity. the shut spacecraft reaches an by altitude a signal stirring of nominally from up the a dust radar cloud 14 feet, the vernier to mini-
When engines mize the are
off automatically of the
altimeter at landing.
possibility
spacecraft
T o uc hdo wn With remaining the vernier engine lunar system surface. cut off, the As the spacecraft will free-fall touches the the
13 feet to the
spacecraft
surface,
181
the tripod landing gear system (with the spring-damper
legs) provides stability
and, in conjunction with the crushable blocks on the spacecraft body, helps to absorb the kinetic energy of the landing. The spacecraft is designed so that it will not topple on landing (i. e., will settle in an upright position) on the assumed lunar terrain when (1) the vertical velocity is Z0 fps or less (See Appendix B, item 4) (2) the limits of lateral velocity are as shown in figure 13-13, and (3) the centerof-gravity location (shown in figure 13-14) is not exceeded. These limits are based on a vehicle radius of gyration of 3Z inches or less. Lunar operations will not be impaired by the structural loads imposed under the conditions indicated in figure 13-13 within the landed weight range of figure 13-14, provided (I) the vertical center of gravity is not below spacecraft station 63.48 and (Z) the vehicle radius of gyration is 28 inches or more, under the same vertical velocity, slope,
and friction The terrain coefficient noted below. terrain at the landing as: (I) a slope site is unknown. not exceeding strength and Surface proper The assumed
nature
of the lunar
at the landing
site is defined or less,
15 degrees, from 50 to of fricare of the strength
(2) protuberances 25,000 psi
I0 cm
(3) terrain and design
compressive purposes), 0 to 1.0. to ensure
(for landing landing
dynamics
(4) coefficient dust
tion with uncertain, spacecraft may be
the
gear
foot pads means
of from provided
conditions
but and
reasonable components from
are
functioning
in the presence but is not
of dust. included
Surface in the
compressive design
different
that noted
requirements.
LUNAR
PHASE
Postlanding The descent The ments high
Engineering initial period subsystems power
Sequence after and landing will be not utilized required the to turn during off all terminal lunar operations. compartwhen
phase
functions
electrical
dissipation will lunar
within require noon
thermally
controlled
during
a normal occurs under
landing near
immediate conditions.
turn-off, Engineering
particularly data
the landing sions by
transmis-
to verify
the spacecraft
response antennas to be
to turn-off with turned the
commands spacecraft
will be
accomplished in the flight and be
means
of the omnidirectional mode. power; line, The functions
transmitter RADVS power;
high-power control vernier
off include: and
approach fuel, and
television oxidizer
camera thermal
4 power control.
temperature Commands
control; will then
tank
182
'I5
SPACECRAFT WEIGHT = 59_) TO 7i_ lb VERT1CAL VELOCITY' __ 20 fps A cg = BETWEEN SPACECRAFT STATIONS 63.48 AN{) 66,48 RADIUS OF GYR_.TION ; 28 T0 _2 ;n FRICTION COEFFICIENT __ 1.0
z O (...) LIJ u') n,.u.J Q. I-I,.i.I I.,i,J LI.,.
10
SLOPE
_>I_) deg
1
>.F-,(O O _1 ill >
n." I-,J I./J _J nn O .J ._J <_ :3
-5 ',-.,-,....__._...
X
-tO
-5 TOUCHDOWN
0 INCIDENCE
5 ANGLE, DEGREES
tO
t5
FIGURE
13-13.
MAXIMUM OF
ALLOWABLE TOUCHDOWN
LATERAL INCIDENCE
VELOCITY ANGLE
AS
FUNCTION
183
69
f,
SHOCK ABSORBER
6e
TEMPERATURE "TOUCHDOWN,
AT OF = 20 40 6O 70 8O I00 I20,
66 oLD U_ 0 n_ W Z tu o ,_1 65
ALLOWABLE LANDED MINIMUM WEIGHT
I12: IJJ 64 550 600 LANDED VEHICLE 650 WEIGHT, 700 POUNDS 750
FIGURE
13-14.
MAXIMUM AS
ALLOWABLE FUNCTION OF
VERTICAL LANDED WEIGHT
CENTER
OF
GRAVITY
184
transmitted operations The ducted
to lock are
the landing to be of the
gears completed spacecraft design,
and
dump
helium.
In a typical 4 minutes. landing shock data may
case,
these
expected
in approximately survival and of the
assessment
will be for
con-
to acquire
reliability, The functions array
subsystem of touchdown of the
operating survival second
postlanding
operational verification nal
planning. of such
assessment as and
include sigoutput thermal
operation solar and panel
transmitter, solar panel of the
processing, spacecraft system The extent
planar
positioning, and
power, control
temperatures electrical the on
the condition system. assessment
response
and
power
to which
postlanding solar lighting
of the
spacecraft
is possible time landing
or
desirable
is dependent after the be touchdown, solar
conditions,
Goldstone requirements.
visibility A survey night
remaining will which noon
and
payload
operational (as well acquisition as
preclude must landing for
panel
assessment earth
that of the the
television, antenna). A
delayed reduce
un_til after the extent
with
high-gain by
may
of assessment because
possible
restricting
the time
available encountered Goldstone craft
continuous at that time.
operation
of the high with of the
spacecraft minimum
temperatures of 3 hours of
If touchdown
occurs
the
visibility
remaining, se, may as
the evaluation opposed
relative
importance
of spacedata by
assessment
per station,
to the acquisition for the optimum
of initial scientific allocation
the Goldstone time.
be
desirable
of available
The be
postlanding for
positioning day landings for
of the and
solar
panel
and
planar for night
array
antenna
will The teleof landing, sun
accomplished
is also
desirable
landings. necessary
positioning communication power for
is required bandwidth lunar day
the case for
of day
landings
to assure and for
the
television and solar the and
transmissions charging. to face
the generation of a night morning
operations the
battery panel earth time and,
In the case
it is .desirable and the planar The
to position array
the direction wide-band for the
of the
to acquire utilized
to provide required
telemetery
capadepend noon, at conductwill be
bility. on on
procedures environmental
initial positioning occurs near lunar
the lunar the
conditions controlled acquisition the planar
if the landing
spacecraft The
thermally angular with day
compartment of the earth array. The by
temperatures will be search accomplished pattern
existing by
touchdown. ing a search
pattern for near
required
restricted
terminator
landings
determining
spacecraft
roll attitude
185
from of
the a most
solar
panel
data roll
and
for
night
or
near
noon
landings
by
the
assumption
probable
attitude.
After ing from solar lunar
initial incidence libration twice
positioning, angle must per and be earth
repositioning the planar array
of
the to
solar
panel pointing intervals. is
to
follow errors
the
chang-
correct
resulting
accomplished day the during
at the at
periodic lunar day
Repositioning although prior to it tele-
approximately may vision be desirable operation,
anticipated, such ratio. as
to position to obtain
antenna video
specific
times,
optimum
signal-to-noise
When two celestial
the
landing
site
is in
known, spacecraft earth
spacecraft coordinates. positions and star in
attitude For spacecraft
may day
be
determined the from
from
bodies will and
known sun array
landings,
spacecraft solar frames ing, atures array the panel with
provide planar
and
coordinates from survey case low of
the
positions mirror be be
positions For of
television a night landtemperthe planar
their
corresponding may may
positions. because only
the their
survey and earth
cameras position
inoperable available
ambient by
if earth
acquisition
is
successfully
accomplished.
Spacecraft
Lunar
Day
Ope
rational
Capability
The shown solar mode may
lunar
day
operational The the
capability nominal case of the solar the
of
the
Surveyor output charge power of as
A-ZIA
configuration of in the time normal or
is
in Figure angle of be is
13-15. shown and to the for
panel
a function
optimum
regulator level which actually
operation available by the
represents spacecraft. of optimum panel by the
maximum The amount
at various generated to The dissipate degree to
times is the
power
limited heat which
capability by the
thermally charge power the
controlled regulator is utilized of the the
compartment compartment. depends thermally charge in both energy by on the
generated the
available and is
solar limited the heat
mission
sequence compartin A comand B.
of operation ments partment In general, all
capability by
controlled regulator compartments is available
to dissipate A and with desired the
generated subsystems landing, within be a be the landing
optimum located
spacecraft day
a lunar operations
sufficient constraints near the
solar
to con-
duct The case start is
imposed day/night charge
the
compartments. in before which the
single
exception time
would may not The
terminator, the battery
sufficient of the in
available of may
to fully the be
lunar terms
night. of watts
capability heat that
thermally dissipated
controlled under
compartments thermal tray
shown
of
186
I00 SOLAR
I PANEL OUTPUT POWER
_ 75
_
OCR
IN NOiMAL
MOOE
(NOMINAL]
5o
NOMINAL COMPARTMENT DISSIPATING CAPABILITY XIMUM 25 THERMAL TRAY THERMAL TEMPERATURE, +125°F COMPARTMENT_
0
1.0(,
d
>u _o 0.75
I I
\
_, _,_ J % '_
-_,CROMETEOR,TE // OETECTOR.
SEISMOMETER, ALPHA SCATTERING OR NARROW BAND ENGINEERING DATA T ANSMISSION ' _/ /SURFACE
z _O I,-
I I
_TELEVISION TEMPERATURE LOW LIMIT
i
,
SAMPLER
'7
_,_,_ SURVEY. FRAME _ _ J PER
./
I I I
I
j SUR_FACE SAMPLER LOW TEMPERATURE LIMIT
0.50 _. _ _ _.
TEPPING _
OPERATION TELEVISION 3.6 sec
._" ._"
0 =E 2E 0.25
I
LIMIT O 50 IOO TIME [ 90 I 60 ] 30 SOLAR FROM 150 DAY/NIGHT I 0 ANGLE, DEGREES 200 TERMINATOR, HOURS I 30 250
IE
I i
I I
350 I 90
300
I 60
_SPACECRAFT EXTERNAL
THERMALLY SENSORS NOT
CONTROLLED COMPARTMENT INCLUDED EXCEPT AS SPEClFED
CAPABILITY'ON_;
FIGURE
13-15.
LUNAR DAY FOR EQUATORIAL
OPERATIONAL LANDING
CAPABILITY
temperature under assumed ment the
conditions compartment includes
of +125 ° F standard
and
as
such
represents conditions.
the
maximum
capability
environmental as and well
The
environment on the and compart-
the lunar by
environment solar panel
as
the influence array
radiator
surface
the
planar
positions
temperatures. The (a) (b) (C) complement Survey Soil Alpha of scientific (2) surface instrument sampler instruments consists of the following:
television mechanics scattering
187
(d) (e) The maximum by the heat
Micrometeorite Seismometer operational steady-state dissipating allowable by to the
detector
capability cycling capability thermal the tray
of which of
the is the
above permitted thermally
instruments within controlled is 1Z5 ° maximum thermal F;
is the
shown constraints
in
terms
of imposed
the
compartments. therefore, operation tray temperature maximum of compoof
The
maximum heat nents
temperature as occurs well at
dissipation internal
compartments
as a
compartments
125 ° F. During the normal (a) (b) (c) (d) (e) the On mode Optimum Transmitter Transmitter Command Engineering required (f) (g) During Ig5 ° F mand thermal At Boost regulator on and auxiliary thermal period units tray of the on as required are cycle slightly the below comperiod of the instrument cycle the spacecraft is assumed in
of operation charge on
as follows: regulator in normal band or mode narrow band as required)
{either
wide
connected receivers signal on
to planar (Z) and
array
processor
central
signal
processor
on
as
Instrument the
electronic compartment
this period and rising. are
temperatures instrument the
During assumed
the Off to be
only
receivers tray lunar
operating less
with than
resulting
decay
in the
temperatures noon, per
to slightly
I75 ° F. television in the mapping mode
the operation typically
of the results
survey
(3. 6 seconds mately Off or partment B.
frame) and
in thermal A and
dissipations B,
of approxiDuring in comthe
36 watts compartment A by
Ii watts cooling
in compartments period approximately and average no
respectively. are dissipated
3 watts dissipations
the command in order
receivers to limit B the
occur
in compartment A is
Therefore, and
heat
dissipation
in compartment mapping mode
to 6 watts limited
in compartment cycle of about
to 4 watts I0 percent
the operation or 6 minutes
of the per
to a duty
hour.
188
The operation of the micrometeorite detector, alpha scattering instrument, and seismometer is shown in figure 13-15. The above instruments may be operated either individually or simultaneously at essentially the same duty cycle. The principal operating mode of the surface sampler, the continuous stepping of one of the four drive axes, is shown in figure 13-15. The operation in the accelerometer mode although requiring ard is therefore not shown. The spacecraft capability wide-band telemetry is only of short duration
shown in figure
13-15 is subject to a number of
variables. The first of these is the operating mode of the optimum charge regulator. It was previously stated that the regulator was operated in the normal mode during the On period and turned off during the instrument Off period. A tradeoff may be made, however, in battery charging versus data transmission. If the charge regulator is turned off or operated in the bypass mode during transmission, the reduction in the heat generated in compartment A would, if compartment A were the limiting compartment, allow increased transmission time. If battery charging is important during the period when the indicated duty cycle is less than i00 percent, the battery may be charged during portions of the off cycle (thereby causing added thermal dissipation within the compartment) with a resulting decrease in the allowable transmission time. Other variables affecting the illustrated capability include optimum charge regulator variations due to the dispersions associated with the solar panel output power, variations in the dissipations by spacecraft electronics, the optimum charge regulator and the boost regulator due to variations in the battery terminal voltage, and variations in the actual compartment dissipating capability due to lunar environment and compartment thermal parameters. The above variables are potentially capable of significantly affecting the illustrated capability during the noon interval. Additional tion.
than will until with vision At the thus the allowable provide maximum landing, may be
capability
maximum. a limited tray
is available to optimize
temperatures The transient compartment capability of IZ5 ° F engineering thermally During allow the the
the lunar sequence of operatypical case will be due less capacity to its mass operation case,
touchdown,
compartment
in the thermal for
continuous
spacecraft In the normal and
temperature
is reached. assessment restricted region
a noon survey
the postlanding completed
the initial teleas operation, transient time.
before
operation
indicated the
in figure
13-15
is required. would for
of restricted
nonoperation which
of equipment may be used
accumulation
of additional at some desired
capability
extended
transmissions
189
A
tradeoff the panel
may above
also
be
made heat could it may panel output
in
the sources
utilization during in an increase to to the
of the
the
planar
array
or the
solar reduction heat
panel. of
Since the
represent
noon in
interval, compartment planar lower its
temperatures Although of the solar
result be
radiating the and dissi-
capability. positioning yet pation retain
undesirable
alter sun
the to
array
position,
off-normal could provide
temperature thermal
a minimum if
additional
compartment
capability
desired.
Spacecraft The expected During l0 of at ° F,
Lunar transit battery lunar night
Night energy state
Operational deficit, of charge
Capability nominally of 2800 1000 watt-hours A losses. In 6.8 B. than operation that the The order and and Table to watt-hours at B are 13-8 maintain 3.9 watts will a temperature controlled presents the of at will 2375 is of fully at a result of in an
70 ° F. and
operation to lunar minimize night
compartments heat losses.
50 ° F summary
respectively,
compartment the above in and an
heat
compartments heat must be of available If the before available the for
temperatures, compartments average for during a load lunar the
a nominal A of and less
watts battery
dissipated 50 ° F energy landing day/night the night The tures point, engineering is presented of may by
discharge ampere
a temperature the
one-half to
reduce
remaining occurs
night day such
about
watt-hours. charged will be
battery
terminator operations. heat be the dissipation provided operation data. in the figure survey may The is followed survival decay For to the 0 ° F, case energy be The
nominal
3375
watt-hours
energy
required by of heaters the or,
for
the more
maintenance efficiently system under of various
of from to
compartment an acquire operating or operational scientific
temperastandor
telecommunication night If the is by not taking of of limited allowable landing assessment of the it is of spacecraft capability maintenance required, advantage the battery
lunar 13-16. television obtained
conditions the night thermal temperaThe for maximum survival survi-
communication spacecraft compartment above control
heating val
additional of at the the
capability
capacities. tures extended A to
complete by the
discharge slow should minimum a night decay be
compartment to the battery desirable the interval
temperatures. required temperature. to spacecraft touchdown allocate a
time the of
compartment
small
portion survival.
of A
the nearly
available complete
for
the
touchdown survival is
assessment
expected.
190
TABLE 13-8.
AZI-A
COMPARTMENT
Model
LUNAR NIGHT HEAT LOSS SUMMARY
AZI-A A* Heat Loss Watts B**
Heat Path
Basic less compartment wiring harness
Compartment at 50 ° F
Compartment at 10°F
Supports Thermal Mylar Thermal Wiring Basic Low inserts 3 coax cables low switches super tunnel harness bus conductance insulation
1.52 1.13 2.32 0.76
O. 75 0. 67 1.38 0
0.34
0.68
0.26
0
conductance Tear strip (34 payload strips) -,--,--,0.16 0
Scientific
TV
cameras
2 and
3
0.16 0 0 negligible
0.06 0.16 0.06 O. O9 0
Surface
sampler
Seismometer
Alpha
scattering
Micrometeorite detector Total compartment on on design design low losses shown shown conductance on on drawing drawing wiring
0.16
6.8±0.5
3.9 revision revision E. E.
± 0.25
*Based **Based ***Assumes
261214 361240 harness.
191
5DO0 _t BATTERY DEPLETED, COMPARTMENT TEMPERATURE DECAYS TO OeF A
POST LANDING ENERGY ALLOCATION OF 5ow-hr REQUIRED FOR ENGINEERING ASSESSMENT AND PLANAR ARRAY/SOL.AR
4000 "K-')(" 0 z i < 3375 ±i75 _'')('_ 3000 0.SAMP LOAO = 3373w-hr "_ , --'_'-I" ESTIMITED _"_'UULLY BATTERYLCHARGE DISPERSION, CHARGED BATTERY AT 50"F 4- 30" AND PANEL_ --POSITIONIN i
_._ w z w 2373 ±seo 2000 ....
r
_
•
I-_'_ E-N-E'I_G¥-'A"V'AILABLE _ AT TOUCHDOWN WITH BATTERY _TEMPERATURE REDUCED TO
_ _ _ "_
I
NOTV SURVIVAL_
I
_
\
I000
TVSURV_VAL. _TA OPERATION
SEISMOMETER 30% OPERATION 20.2 w REQUIRED 0 4OO ,00 TIME FROM
f 30",. I
I I l 200
_ _
_.'.-;._-X,;,_-_, .......... _ I
_ _ I _-_"
_
\ _ \ _. _ \
_. _ _. -_ _1
iO0 TERMINATOR HOURS
0
NIGHT/DAY
A
DAY/NIGHT T E R M I NATOR
'
TERMI
A
OAY/NIGHT NATOR
FIGURE
13-16.
LUNAR
NIGHT
SURVIVAL
CAPABILITY
It of
is the
desirable wide-band after lunar
to
acquire transmission
the
earth
with
the
high The
gain survey low
antenna television
for
the may
assessment be inoperaof
capability. because of
tive the
touchdown, night.
however,
the
environmental
temperatures
Hypothetical A shown
Sequence
of
Operation of operation touchdown and planar completed for the case of an equatorial the sun landing and the earth is
hypothetical
sequence The panel
in figure by
13-17. the solar
survival array
evaluation are completed second
and
acquisition with the
during
first hour the heat steadyof the initial
initial television in the thermally
survey
in the
hour. the
Although maximum capabilities
dissipated state
controlled
compartments the
exceeds thermal
compartment
dissipating in the nominal within
capability, case are
transient to permit temperature
compartments 2 hours
expected tray
completion constraint
of the of
of operations
the thermal
IZ5 ° F.
192
STATION
TRACKING
TOUCHDOWN SUN/EARTH
SURVIVAL ACQUISITION
ASSESSMENT
PLANAR ARRAY/SOLAR REPOSITIONING TELEVISION NARROW COLOR SURFACE ALPHA OPERATION ANGLE SURVEY SAMPLER SCATTERING
PANEL
MAPPING
SURVEY
OPERATION INSTRUMENT
SEISMOMETER MICROMETEORITE ENGINEERING DATA DETECTOR TRANSMISSION
-%
SOLAR APPROXIMATE TIME OVER OPERATIONAL INDICATED DUTY CYCLE ASSUMING OR PERCENT NOMINAL ON SPACECRAFT
ANGLE,
DEGREES
r--]
INTERVAL
COMPARTMENT NON-LIMITING CONTINUOUS TELEVISION CONTINUOUS ENGINEERING (_) SURFACE PICKING NOT
THERMAL DISSIPATING BY EXTERNAL SENSORS INSTRUMENT INTERRUPTED ENGINEERING INTERROGATIONS, OPERATION 5 rain EACH DATA S
CAPABILITY. ASSUMED OVER HALF INDICATED HOUR INTERVAL.
TRANSMISSION rain EACH HALF WITH HOUR, MINIMUM FRAMES.
SAMPLER OPERATION IN
STEPPING ALTERNATED REQUIRING WIDE-BAND INDICATED DUTY CYCLE
TELEVISION TELEMETRY
INCLUDED
FIGURE
13-17.
HYPOTHETICAL EQUATORIAL
LUNAR LANDING
PHASE ASSUMED
SEQUENCE--
Operations compartment
during
the
remainder capability television exceeds
of the lunar indicated camera
day
are
shown 13-15.
as constrained
by
the
dissipating that the
in figure operating
Assuming thermal mally mapping may veys be
capability, imposed
established by the and per mapping mapping thermal
by ther-
its
characteristics, controlled surveys conducted each
the thermal during 1000 the near frames
constraints noon each, tape.
compartments of approximately by predetermined color
interval, at 3.6 Two The and
black seconds
white frame, suroperacooldown
command are
complete television to allow
with
of three
filters
included. status
tion is interrupted of the compartments. If sufficient micrometeorite
to monitor
the spacecraft
telecommunication detector will be
bandwidth operated
is available
the
seismometer only for
and
continuously,
interrupted
193
compartment thermal cooling as required and during the wide-band transmissions required by other scientific instruments. It is assumed that the sensors do not constrain operation during the noon interval. The operation of the soil mechanics surface sampler is conducted in two modes. Scraping, digging, and mapping of the lunar surface is conducted in a narrow-band telecommunications mode with the stepping alternated with television frames in the wide-band mode as required. The picking operation is conducted in the wide-band telecommunications mode with television viewing used for remote control. The 167° F maximum nonoperating temperature limit of the alpha scattering instrument sensor, together with its limited available view factor to space, could result under worst case conditions in potential permanent damage to the sensor if allowed to remain in the stowed position after a touchdown at lunar noon. Since the 3-hour standard sample count is conducted in the stowed position and is limited to a maximum interruption of 30 minutes the deployment of the instrument will preclude the acquisition of standard sample data. It is assumed that the above worst case situation must be alleviated by thermal design if additional thermal studies of the sensor head and its environment confirm the existence of an actual problem. Therefore this worst case condition is not considered in the illustrated typical sequence of operation. In order to conform to the instrument maximum interruption constraint, the alpha scattering experiment is not initiated until continuous spacecraft operation is permitted by the thermally controlled compartments. During the lunar day and night additional operations are required to support
the requirements of the scientific payload. The assessment of touchdown survival will provide data necessary for the operational planning associated with the scientific payload. The positions of the sun, earth, and stars in spacecraft coordinates obtained during the initial acquisition period and television survey will provide data necessary to compute spacecraft attitude. Engineering interrogations are provided each half hour to monitor spacecraft status. Repositioning of the solar panel and planar array should be performed at intervals of about IZ hours, although if the rate of lunar libration is less than maximum, repositioning of the planar array may be conducted at less frequent intervals. The solar incidence angle on the solar panel is thereby maintained at less than 3.5 degrees and the planar array positioning error may be maintained at less than I degree. In an 194
optimum sequence the final positioning of the solar panel prior to the start of lunar night causes it to face in the direction of the norning sun, with a full battery
charge achieved at the time of final positioning.
195
PRECEDING
PAGE _'_ K,_';_-,,, NOT FILMED. ,.,.
APPENDICES
Appendix ration,
A
lists specifications and function
that establish Surveyor namely,
design Spacecraft, the basic
requirements, ModelA-ZIA. bus and which the the
configuThe scientific specifica-
performance, into two are
of the
list is divided payload. tions are These
major
categories, divided
further
into the
subsystems
under
listed. B lists characteristics under only, C gives and subsystems e_cept for data to which those and for Surveyor apply. in the Spacecraft, All items Model are A-21A.
Appendix These tional are listed
they
for informa-
reference Appendix
that appear performance
preceding
sections. which exist
the physical
differences
between
the A-21
A-ZIA
spacecraft.
197
pRECEDING
PAGE
BLAN_
NOT..,I_|I-I_.-_P'
APPENDIX
A.
DOCUMENT
AND
SPECIFICATION
LIST
BASIC
BUS:
Structural
and
Vehicular
Subsystem
Item
Document DS
Number 230029
Document Spaceframe Configuration
Name Design
Vehicle
DS
230003
Spaceframe Integrating
and
Unit
Structure Gear Gear Requirements Shock Absorber
3 4
DS DS
Z38900 Z38901
Landing Landing Column
5
DS
238904
Landing Blocks
Gear
Crushable
Body
DS GS
Z36130 226100
Antenna/Solar Electromagnetic Specification, Spacecraft Equipment and
Panel
Positioner
Interface Surveyor Associated
CS
Z39506
Spacecraft Grounding, Shielding
Design Bonding,
Parameters, and
DS
Z38704
Spacecraft Compartment
Thermal A Thermal B
Control,
I0
DS
Z38706
Spacecraft Compartment
Control,
199
Engineering
Instrumentation
Subsystem
Item ll
Document CS CS
Number 988653-2 988654-I0 239004 239004
Document Temperature
Name
Sensors
12
13
DS DS
Accelerometers Engineering Sensors, Amplifier Measurement AccelerometerSystem
Propulsion
Subsystem
14
PS
238611
and
Main Thiokol Rocket
Retro
Rocket Main
Engine Retro-
E155-62
Spec, Engine
15
DS PS
238666 238610 262526
and
Surveyor System Safety Main and
Vernier
Propulsion
16
PS
Arming Rocket
Assembly, Engine
Retro Actuator
17
DS
234675
Roll
Electrical 18 239513
Power
Subsystem Power by and Thermal Event Assembly Dissipation
Functional Panel
19
DS PS
237601 237787 237790 237901 231631 231632
and
Solar
20 21 22 23
PS PS DS DS
Solar Battery Battery Boost
Panel
Solar
Cell
Module
Charge Regulator
Regulator
200
L.
Telecommunications
Subsystem
Item 24
Document 239511
Number Spacecraft
Document Command for
Name
Assignments Payload Z5 Z31681 Decibel Summary 26 DS 231706 Spacecraft Transmitters 27 28 DS DS Z31707 231718 Receiver
Scientific
Allocation
and Margin
Telecommunications
Transponders Antennas
Telecommunication and R-F Switching Decoding Decoder} Signal Signal Data Rate
29
DS
231641
Central Command
Unit
(Central
30 31 3Z 33 34
DS DS DS DS
Z31711 Z31710 Z316!6 Z31613
Engineering Central Low Signal Spacecraft Control for
Processor
Processor Auxiliary Auxiliary Channel Payload
Processing Data
Z3951Z
Scientific
Flight Control
Subsystem
35 36 37 38
DS DS DS DS PS
234600 234630 234622 234636 Z 34661 234621 234610 and
Flight Control
Sensor
Group
Inertial Reference Primary Sun Sensor
Unit
Inertia - Burnout
Sensor
Switch
39 40
PS DS
Canopus Flight
Sensor Control Electronics Unit
201
Flight
Control
Subsystem
(Cont)
Item 41 42 43
Document DS DS PS
Number 234623 232601 232902 Secondary Altitude Radar Velocity
Document Sun Marking Altimeter Sensor
Name Sensor Radar and Doppler
(RADVS) - Gas Tank
44
PS PS
234641 234640 231621
Attitude Gas
Jet System
Jet Attitude
Controls Camera
45
DS
Approach
Television
SCIENTIFIC
PAYLOAD:
46
DS
231622
Survey
Television
Camera-
Spacecraft Television 47 FR Z31611 Alpha
Telecommunication
Scattering
Instrument
Auxiliary 48 FR Z31612 Micrometeorite Instrument Ejecta Auxiliary Auxiliary Instrument Detector
49
5O
FR FR
231715 231716
Sample
Processor
Soil Mechanics Auxiliary
51 52
FR FR
239319 239320
Alpha
Scattering
Experiment Ejecta Detection
Micrometeorite Experiment
53 54 55 56
FR FR FR FR
239323 239334 239335 239389
Soil Mechanics Seismological Seismometer Soil
Experiment Experiment Auxiliary Sampler
Mechanics-Surface
Auxiliary
2O2
Television
Subsystem
(Cont)
Item 57 58 59 6O 61
Document FR
Number 239390
Document
Name
Soil Mechanics-Surface Sampling Experiment Ejecta Detector
IS 239302 IS 239303 IS 239330 IS 239381
Micrometeorite Alpha Single Scattering Axis
Experiment
Seismometer Sampler
Soil Mechanics-Surface
203
_-7
pfKECEDING
pAcE
_LA'_"!" b_O'_£_L_ED.
APPENDIX
B. CHARACTERISTICS DATA
System
Item Number /1 Spacecraft Transit
/
Item Injected Time AV (Nominal) at 20 hours Weight 2221.27
Data
/2
66 hr
30meters/sec 15 fps (vertical); 5 fps (lateral)
Midcourse
114
Soft
Landing
Capability
_6o tb paytoa20 HOURS
,_
/
A-21 A-2IA A-21 AND A-21A
_,,=,
bJn =E=E I--LU 80
O. =E 0 L)
TEMPERATURE
TOLERANCES
60
TO
BE
DETERMINED
t
I
AND SUN
TRANSIT
PHASE _I_LUNAR PHASE_
i
ACQUISITION VERIFICATION AND ACQUISITION
MIDCOURSE CORRECTION
_
c
w
120
rr b J03
8O
Z z :::::::::::::::::::::::::::::::::::::::: ...... :::::::::::::::::::::::::::::::::::::::::::::
,,,o_ 4O
.....
0
_:i:_:_:!:_:_iiiililililililiiiiiiii_:_:_:_:_:i:_:i _
0
4 TIME
12 AFTER LAUNCH, HOURS
20
60
65
0 66
I rIME
2 AFTER
3
4 TOUCHDOWN,
5
HOURS
FIGURE
C-2. COMPARTMENTS TEMPERATURE PROFILE
A AND B THERMAL DIFFERENCES
TRAY
221
60
_
I 1
I
COMPARTMENT COMPARTMENT A B LOAD OF 3w COMPARTMENT IN COMMAND ALL LOW A CONSTANT RECEIVERS
COPPER WIRE HARNESS ON A-21 CONDUCTACE WIRING HARNESS ON
A-21A
LANDING
SITE:
LUNAR
EQiATOR
¢n
40
l
b. 0
///
I"125 II1
IZ
rr L_ O. =E I-"
_
_1 < ¢J W -r
20
Z U.I
A-ZIA
Ilg G. 0 0
°
I.Z
/
/
0
,.,
,s
Q.
,s o
130
_
' 135
--
-2O
0 SOLAR
30 iNCIDENCE ANGLE,'r/,
60 DEGREES
90
FIGURE
C-3.
COMPARTMENT DIFFERENCES,
DISSIPATING LUNAR DAY
CAPABILITY
222
60
\
5O
DOPPLER
LIMJT
/
ELLIPSE CORRECTION
99% FOR
DISPERSION NO MIDCOURSE
I / / / I I
ALTIMETER 40
LIMIT
\
z 5O
A-21A
/
I,-
t t \
99% FOR DISPERSION MAXIMUM ELLIPSE MIOCOURSE \ % J ,_ CORRECTION
IN / I
2O
A-21
ANO
A-21A CINEAR DOPPLER
/
/
/ /
IO
COMMAND DESCENT TRAJECTORY
o
2
VELOCITY, HUNDREDS OF FEET
6 PER SECOND
io
FIGURE
C-4.
VERNIER
DESCENT
PHASE
DIFFERENCES
223