THE LUNAR ORBITER

Prepared for

National Aeronautics and Space AdministrationLangley Research Center

Prepared by

Space DivisionTHE BOEING COMPANY

Seattle, Washington

RevisedApril 1966

iii

CONTENTS

INTRODUCTION

UNMANNED LUNAR EXPLORATION

The Lunar Orbiter Project

The Lunar Orbiter Spacecraft

Ground Support Network

THE MISSION

Prelaunch

Launch and Boost

Translunar Flight

Initial Lunar Orbit

The Photographic Mission

THE SPACECRAFT

Photography

Electrical Power

Attitude Control

Velocity Control

Communications

Temperature Control

GROUND DATA HANDLING

AFETR Equipment

Deep Space Station Equipment

Lunar Data Processing Laboratory

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vi

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6

7

10

12

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14

16

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20

22

26

28

29

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32

34

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ILLUSTRATIONS

Figure Page

Lunar Orbiter Spacecraft Frontispiece1 Lunar Exploration Program 22 Comparative Resolution Photographs 33 Area of Interest 44 Potential Photographic Coverage 55 Lunar Orbiter Project Organization 66 Spacecraft Environmental Testing 87 Atlas/Agena Launch Vehicle 98 Location of AFETR, DSIF, and SFOF 109 Deep Space Station Tracking Antenna 11

10 The Mission 1311 Ground Tracks for Two Launch Azimuths 1412 Typical Flight Profile 1513 Lunar Orbits 1714 Photographic Modes 1915 Lunar Orbiter Spacecraft 2116 Photographic Subsystem 2217 Velocity/Height Sensor 2318 Film Format 2519 Film Scanner 2620 Electrical Power 2721 Attitude Control Modes 2822 Velocity Control 2923 Communications Modes 3124 Ternperature Control 3325 Deep Space Station Data Flow - Operational 3526 Ground Reconstruction Electronics and Film Recording Equipment 3627 Photo Data Acquisition, Reconstruction, and Reassembly 3728 Photo Reassembly 38

Note: Lunar photographs on the cover and in Figures 2(a) and 2(b) are reproduced bypermission of the director of the Lick Observatory, University of California,Mt. Ham ilton, California.

v

INTRODUCTION

The Lunar Orbiter project is a necessary and valuable contributor to man's knowledgeof the Moon and its environment. The Lunar Orbiter will provide extensive photographicexploration of the lunar surface to aid in selection of possible landing areas for theProject Apollo manned landing mission.

This document was prePared to provide familiarization with the purpose and objectivesof the Lunar Orbiter project for the information and orientation of persons not closelyassociated with the program. A complete but general description of the spacecraft,ground equipment, and operations activity shows important interfaces and currentplanning. A description of a typical Lunar Orbiter mission is presented to indicatethe approach to acquisition of essential lunar data.

The document generally reflects spacecraft design and mission concepts as of March1966.

Note: Readers unfamiliar with the metric system will find the following conversionfactors helpful:

1 kilometer 0.6214 mile

1 square kilometer = 0.3861 square mile

1 meter = 3.2808 feet or 39.37 inches

1 inch = 25.40 millimeters.

vi

UNMANNED LUNAR EXPLORATION

The most detailed examination of the Moon, using the latest Earth-based instrumentsand methods, does not adequately define the lunar terrain and environment for a mannedlunar landing. Precisely where on the lunar surface are there suitable landing sites fora manned vehicle that are not laced with crevices or jagged rocks? How long can mansurvive when exposed to the radiation and micrometeoroid bombardment in the lunarenvironment? How accurate are present calculations of the size, shape, and gravita­tional field of the Moon? These and many other questions must be answered to achievemaximum safety for the first lunar explorers.

The Lunar Orbiter is one of three unmanned program s devised to provide this vitalinformation (see Figure 1):

• Ranger -Rangers VII, VIII, and IX have given man his first close views of thelunar surface. The Ranger sPacecraft were guided to selected areas on the lunarsurface. The multiple television camera payload was activated minutes prior toimpact. Television signals were transmitted to Earth, and the final display showedthe impact area in detail. Although the Ranger program has provided some ex­tremely valuable high-resolution lunar photographs, the coverage is too limited toprovide sufficient information for a manned lunar landing.

• Surveyor - Surveyor will make a soft landing at a predetermined point on the lunarsurface. During landing, touchdown dynamics and bearing strength will be meas­ured. After landing, additional data of local surface conditions will be collectedwhile eye-level television cameras scan the landscaPe. These data will be trans­mitted to Earth as they are gathered.

• Lunar Orbiter -The Lunar Orbiter Program will provide extensive photographiccoverage of large specified areas of the lunar surface. Surface detail will be com­parable to the sampling by Ranger at impact and far better than is possible to ob­tain from Earth. Maximum photographic resolution that can be achieved with Earthtelescopes is on the order of 800 meters (1. e., objects less than 1/2 mile in diam­eter cannot be recognized). Enlargement of telescopic photographs would not im­prove resolution, and would only result in unrecognizable blur (Figures 2a, b, andc). Many high-resolution photographs similar to the Ranger photographs can beprovided by the Lunar Orbiter with a resolution of 1 meter (Figure 2d), each, how­ever, covering a much greater area. Photographs of medium resolution (8 meters)of the surrounding area will be taken simultaneously with each high-resolutionphotograph. The photographs to be taken by Lunar Orbiter will allow extrapolationof the SUrveyor data to the greater areas covered by the Lunar Orbiter.

A typical Lunar Orbiter mission begins at Cape Kennedy with initiation of prelaunchactivities. Following launch and separation of boosters, the Llinar Orbiter orientsitself in space, performs one or two midcourse corrections, and injects itself into an

"->

RANGER

COMMAND & SERVICEMODULE (APOLLO)

19ro • 19m

Figure t Lunar Exploration Program

52METERS

~ I-----t-

,i£',~+~~~~.:;,~,~,.,;'.~'

,'~~

(D) RANGER PHOTORESOLUTION = 1 METER

Figure 2 Com"parative Resolution Photos

(C) PORTION OF LICKPHOTO ENLARGEDRESOLUTION =790 METERS

SCALE ....... 880 METERS

..

(A) LICK OBSERVATORYPHOTO OF MOON

SCALE ......... 190 KM

AREA COVERAGE BY ONE LUNAR ORBIT~': THIGH RESOLUTION (1 METER) EXPOSURE.. METERS

~ 1-9-4-S-UC-H-EX-P-O-S-U-RE-S-C-A-N-BE-M-A-D-E_W_IT_H ~ I_ THE LUNAR ORBITER -.l-

(B) LICK PH~TO 1~.~~~~~~~~-16,600 METERS~~~~~~~~~.IMARE IMBRIUM

SCALE~75 KM

initial orbit around the Moon. When target lighting conditions are proper) the space­craft is inserted into an elliptical orbit and passes over the target area at low attitudeto obtain the lunar surface photographs.

From the design altitude of 46 kilometers) each high-resolution photograph covers 69square kilometers of lunar surface) and each medium -resolution photograph has a 1182square-kilometer field of view.

Figure 3 shows the primary area of photographic interest and target sites to be photo­graphed on the initial mission. Actual area covered is dependent on the concentrationof photographs necessary to fulfill specific mission requirements. The design mission.consists of sampling each of the 10 target sites with 16 consecutive exposures; sitesearch and site examination are also illustrated in Figure 4. Photographs will be de­veloped in flight and transmitted to Earth. Analysis of Lunar Orbiter data) in conjunc­tion with that gathered by Ranger and Surveyor, will provide answers to most of thevariables affecting safety of the first lunar explorers, as well as invaluable scientificdata about the Moon.

MOON'S ROTAtIONON ITS AXIS

'111111/11111111111111111111

MOON'S TRAVELAROUND EARTHAS OBSERVEDFROM EARTH

- TARGET SITES

Figure 3 Area of Interest

4

• 0 250 500Iw • wi I

SCALE (KM)

o 250 500

.. - - ISCALE (KM)

AREA COVERED INDESIGN MISSION INCLUDESPHOTOGRAPHING 10 TARGETSITES WITH ONE SPACECRAFT,16 EXPOSURES PER SITE

REPRESENTATIVE __-oN

AREAS COVERED BYONE SPACECRAFT INA SITE EXAMINATIONTYPE MISSION, 96EXPOSURES PER AREA

30°I W 60° 50° 40°

'\: ~10° S .........----.--;l\r----+-~~~t_:;W_-------_\_---_+_-----~----1

\

REPRESENTATIVEAREAS COVERED BYONE SPACECRAFTIN A SITE SEARCHTYPE MISSION, 64EXPOSURES PER AREA

NOTE: THE ILLUSTRATED AREAS SHOWN FOR "SITE EXAMINATION" AND "SITE SEARCH"HAVE BEEN SUBDIVIDED AND DISPLACED FROM EACH OTHER BY APPROXIMATELY20°. THIS IS NECESSARY SINCE LIGHTING CONDITIONS OVER ONE AREAREMAIN FAVORABLE FOR ABOUT 7 ORBITS. AS ILLUSTRATED EACH SITE EXAMIN­ATION AREA REQUIRES 6 ORBITS AND EACH SITE SEARCH AREA REQUIRES 7 ORBITSTO COMPLETE.

ORBITAL PARAMETERS ARE: PHOTO ALTITUDE (PERILUNE) - 46 KILOMETERS,APOLUNE (HIGH POINT OF ORBIT) - 1850 KILOMETERS,INCLINATION - 12°ORBITAL PERIOD - 3.5 HOURS

Figure 4 Potential Photo Coverage

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THE LUNAR ORBITER PROJECT

The Lunar Orbiter project, one of the lunar and planetary programs directed by NASAHeadquarters Office of Space Sciences and Applications, is managed by Langley Re­search Center. Project organization is shown in Figure 5. The project is essentialto the overall program of lunar exploratiol). Project implementation, from conceptionthrough mission completion, has been geared to ensure maximum probability of success.

A fundamental approach to designing the spacecraft-supporting operational systems isthe use of as much proven hardware as possible so that a minimum of system develop­ment is required. Reliability has been emphasized throughout project planning by thisuse of proven hardware design and by safety engineering in all aspects of subsystemdesign. Quality control, which began during spacecraft design, is being continued asan integral part of the test program; failure data is rapidly fed back to design andfabrication areas.

Three sets of spacecraft components and eight assembled spacecraft are being furnishedfor the program. Three of the spacecraft are for ground testing; five are designated forflight operations. Spacecraft component sets have been subjected to thorough flightacceptance, qualification, reliability demonstrations, and subsystem design verifica­tion tests. A flight-acceptance test is performed on each of the three ground-testspacecraft to functionally check its operation. Each ground-test spacecraft is then

PROGRAM DIRECTION

NASA HeadquartersOffice of Space Sciences

•PROJECT DIRECTION

NASA Langley Research CenterLunar Orbiter Project Office

I

+ ~LAUNCH LAUNCH PRIME DEEP SPACEVEHICLE SITE CONTRACTOR NETWORK

NASA Eastern Test The Boei ng Co Jet PropulsionLewis Research Lunar OrbiterCenter Range

Project Laboratory

I+ +

SPACECRAFT AGE

Boeing RCA Eastman Boeing RCA Eastman

Figure 5 Lunar Orbiter Project Organization

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used for individual testing. One is taken to Goldstone, California, for the performancedemonstration test, to determine compatibility of the spacecraft with ground controland instrumentation facilities.

After the Goldstone test, the spacecraft is moved to Cape Kennedy for Eastern TestRange launch operation integration test to confirm its physical and functional compati­bility with the booster and range instrumentation. The second ground-test spacecraftis used for qualification testing to prove design adequacy or locate potential failurepoints caused by over-stresses, tolerances, combinations of operational sequences,and environmental conditions. The third ground-test spacecraft is used for missionsimulation testing in a space chamber (Figure 6) to evaluate the spacecraft operationunder realistic space vacuum and temperature conditions for periods equivalent tomission time. The five flight spacecraft receive functional and environmental flight­acceptance tests prior to their shipment to the Air Force Eastern Test Range (AFETR)for launch operations. Boeing, using a test-team concept, has assigned a group of testengineers to each spacecraft. These men accompany the spacecraft from manufacturethrough launch, performing tests, correcting faults, and reporting problem areas. Acontinuity of lmowledge is thus maintained throughout the life of each spacecraft, en­suring a complete understanding of its test history.

THE LUNAR ORBITER SPACECRAFT

The spacecraft (see frontispiece) is an assembly of high-reliability components withthe flexibility to fulfill several mission assignments. An Atlas/Agena launch vehicle(Figure 7) will place the Orbiter in translunar trajectory, where an internal rocket­engine propulsion system will provide midcourse trajectory correction and, later,inject the spacecraft into lunar orbit. Full three -axis stabilization and attitude con­trol is provided by ejecting nitrogen gas through small jets. Solar sensors are usedto give pitch and yaw reference, and a star tracker (Canopus) is used for roll reference.Solar panels furnish electrical power and charge the nickel-cadmium battery, exceptwhen the spacecraft is behind the Moon or maneuvering; during these times batterypower is used. A flight programmer sequences all commands, which may be eitherstored before launch or transmitted during flight. Commands to be transmitted dur-ing flight are stored, verified by retransmission, and executed at the proper time.

The camera has two lenses -a long-focal-Iength lens for high-resolution photographsand a wide-angle lens that simultaneously records a medium-resolution frame of thesurrounding area. Camera frame, film supply, and numerous mechanisms are thesame for the two lenses. Exposed film is developed in flight, by a damp monochemi­cal process, into a high-quality photographic negative. A minute beam of light scansthe completed negative and, on passing through it, is converted to electrical signalsfor radio transmission to Earth.

In addition to photographic data, lunar-environmental and spacecraft-performancemeasurements are made by the sensors of an instrumentation subsystem. The dataare transmitted over the radio link to Earth-based receivers. Designed photographicmission life is 30 days, but transmission of environmental and performance data willcontinue for another 11 months.

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8

Figure 6 Spacecraft Environmental Testing

Figure 7

·;."'::-'·1~1

Atlas/Agena Launch Vehicle

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GROUND SUPPORT NETWORK

Launch operations for the Orbiter mission are performed by AFETR. Mission con-trol is maintained by AFETR until after separation of the Agena boost vehicle. Con­trol is then transferred to the Space Flight Operations Facility (SFOF) at Pasadena,California, where all spacecraft flight trajectory computations are made and commandsgenerated. Actual communication with the spacecraft is maintained by the Deep SpaceInstrumentation Facility (DSIF) , which consists of three deep space stations: Goldstone,California; Madrid, Spain; and Woomera, Australia (see Figure 8).

Figure 8 Location of AFETR, DSIF and SFOF

These stations are geographically located so that at all times at least one is in contactwith the spacecraft to monitor and track the craft, transmit all ground commands, andreceive and record data transmitted by the spacecraft. A deep space station trackingantenna is shown in Figure 9. Only a minimum of Lunar Orbiter-peculiar ground equip­ment is required to supplement the existing facilities.

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Figure 9 Deep Space StationTracking Antenna

1]

THE MISSION

Beginning in 1966, five Lunar Orbiter spacecraft will be launched to obtain informationvital to the Apollo program. Although the spacecraft are identical, each mission isunique in timing and trajectory, in orbit and target, and in objective - expansion ofman's knowledge of the Moon.

The spacecraft, in order to achieve its photographic mission, must be placed in orbitaround the Moon - thence its name "Lunar Orbiter." While a satellite can be placedinto a relatively accurate Earth orbit from a known point on Earth's surface, theLunarOrbiter must be injected into a precise, low-altitude orbit around the Moon, whosesize and shape, as well as center of gravity and mass, are not precisely known. Also,the critical maneuver of retrofiring to place the spacecraft into a specified lunar orbitmust be controlled from Earth after a 240,000-mile journey through space.

In addition to the principal operational complexities of lunar orbit acquisition, theLunar Orbiter photographic mission poses other unique requirements. In the follow­ing sections, these needs and their fulfillment will be seen through various phases ofthe typical mission (Figure 10).

PRELAUNCH

Members of the test team work with the spacecraft from its assembly at Boeing throughits many test phases, following it from Seattle to Cape Kennedy. At the launch site,final checkout of the Orbiter completes the extensive test program. Film is loaded inthe camera, squibs installed, fuel and oxidizer pumPed into the propellant tanks, andnitrogen bottle pressurized. With great care, an aluminized-mylar thermal barrieris placed around the craft and the nose shroud installed.

Simultaneously, the Atlas/Agena launch vehicle undergoes final launch preparation.The spacecraft is then transported to the launch pad and mated to the Agena, and finalprecountdown tests performed.

LAUNCH AND BOOST

The countdown begins, initiating a checkout of the Atlas, Agena, and spacecraft com­bined as one functional unit - the Lunar Orbiter space vehicle. Personnel at AFETRand in the deep space network are informed continually of the countdown progress. Ifa delay occurs, ground computers determine a new launch azimuth and a new parking­orbit duration (Figure 11). Delays longer than 2 hours may require a rescheduling ofthe countdown -a day, or, possibly, a month later.

At the completion of the countdown, the Atlas engines fire, boosting the space vehicleto an altitude of 85 miles before Atlas separation (Figure 12). The spacecraft noseshroud, no longer required for atmospheric protection, is jettisoned.

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rREASSEMBLEDPHOTOS,SUPPORTINGMISSION DATA

I)~NASA

PHOTOGRAPHLUNAR SURFACE.TRANSMITDATA TO DSIF

DATA

SPACECRAFT CONTROLTRACKING ICOMMAND TRANSMISSIONDATA RECORDING /'PHOTO RECONSTRUCTION

~ /'

MIDCOURSECORRECTION IAGENA ... ~

ESTABLISHES .... -- ,LUNAR ~ f " MIDCOURSE

TRAJECTORY ~•••••••~" CORRECTION

" AGENAAGENA / I SEPARATION

ESTABLISHESEARTH ORBITtI

I

NOSE-SHROUD ---~SEPARATION i~.....~

ATLASSEPARATION

~~~~TER I.....~\SEPARATION ,.

LAUNCH IVEHICLE

ASSEMBLY COUNTDOW\

SPACECRAFT 1 & LAUNCH I LAUNCH MISSIONLAUNCH. ~EHICLE CONTROLPREPARATION .. TRACKING ~""'''' y " < r ,

, f ,ttL \ ,_._, , · · NE. . c I , - I· . " ••' GOLDSTO~; , Ii""""[ rc:!: AFETR···••TELETYPE••• SFOF '•• ~~~~YPE WOOMERA

& VOICE & VOICE MADRID

Figure 10 The Missionw

TYPICAL LOCATION )OF TRANSLUNARINJECTION (INSERTIONINTO MOON TRAJECTORY)

111111111111110 IRECTION OF EARTH'S ROTATIONIIIII\IIIII"

A translunar injection point remains nearly stationary in space over a short time interval; however,this point with respect to the Earth's surface moves from east to west due to the Earth's rotation onits axis. This injection point can be reached at any time during a short interval by varying thedirection (azimuth) of launch and the parking orbit time. By scheduling a 900 launch, delays arecompensated for by gradually increasing the azimuth to 1140 , providing a "launch window" ofnominally two hours. To avoid launch over populated areas and remain within the trackingrange the launch direction is restricted to the angles mentioned.

Figure 11 Ground Tracks for Two Launch Azimuths

Next, the Agena engine fires, accelerating the Agena/Orbiter combination into a 100­mile-altitude "parking" orbit. After coasting to the translunar injection point, theAgena engine reignites and pushes the spacecraft to escape velocity. The Lunar Orbiterspacecraft then separates from the Agena and begins its long coast toward Moon rendez­vous, which occurs 3 days later.

TRANSLUNAR FLIGHT

After laWlch, mission control is transferred from Cape Kennedy to the SFOF atPasadena. For the Atlas/Agena, the job is done; for the Orbiter, the task begins.

Signals from the flight programmer explode squibs to release two pairs of solar panels,which spring out from the base of the spacecraft. The omnidirectional antenna and the3-foot-diameter, dish-shaped directional antenna are similarly extended (Figure 12).

The spacecraft now maneuvers until the solar panels face the Sun. Sun sensors controlthe ejection of nitrogen from small nozzles, causing the craft to pitch and yaw until thecorrect pitch and yaw attitude is achieved.

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Other nitrogen jets roll the spacecraft until the star-tracker faces Canopus, a southern­hemisphere star identified by its brightness. To verify that the star-tracker is "see­ing" the correct star, the spacecraft may be commanded to roll through a completerevolution while star-tracker measurements of light intensity are telemetered to Earth.From the telemetry, a "star map" can be prepared and compared with a known starmap for that trajectory. Coincidence of the star patterns provides verification. Con­tinued tracking by the deep space stations enables SFOF to compute spacecraft range,velocity, and flight path.

Slight deviations from the intended trajectory are probable, and correction must bemade. About 10 hours after launch, guidance corrections are transmitted to the craft,verified, and stored. The Lunar Orbiter breaks Sun/Canopus orientation as the gasjets turn the craft until the rocket engine points in the desired direction. Valves open,allowing fuel and oxidizer to flow into the engine and ignite. The 100-pound thrustchanges the spacecraft's speed and direction until an accelerometer, measuring thevelocity change, cuts off the engine. The spacecraft gains Sun/Canopus orientationas it continues on its adjusted course toward the target. A second correction, if re­quired, is made 50 hours or more after launch. Fuel limitations allow only smallguidance corrections; the spacecraft relies mainly on the initial guidance accuracy ofthe Atlas/Agena boost vehicle.

Since the Moon orbits the Earth in a slightly elliptical path, the transit time fromEarth to Moon will depend on launch date as well as launch velocity. While a high­velocity, short-time transit provides greater accuracy, it also requires an increasedbooster capability and additional spacecraft propellants to achieve slowdown on arrival.A compromise has been selected, with a transit time of about 3 days.

INITIAL LUNAR ORBIT

During translunar flight, trajectory information provided by deep space station track­ing is used by SFOF to compute the velocity and direction changes required to achieveinitial lunar orbit. This infor~ation, transmitted to the spacecraft flight programmerand stored, will activate the nitrogen jets to turn the spacecraft and point the rocketengine against the direction of flight. As the Lunar Orbiter penetrates deeper into thelunar gravitational field, the rocket engine ignites at the prescribed time. The slowedspacecraft, approaching to a minimum altitude of 200 kilometers above the lunar sur­face, no longer has sufficient velocity to proceed outward against the attraction of lunargravity. The Orbiter becomes a satellite of the Moon, as SuniCanopus orientation isreestablished. Initial orbit parameters are: perilune, 200 kilometers, and apolune,1850 kilometers.

Because exact lunar size, shape, and gravity are unknown, the actual orbit may be dif­ferent from the planned orbit. Tracking and computation by the deep space networkwill accurately establish the actual orbit within three to five passages around the Moon.The exact location must be known to successfully inject the Orbiter into an ellipticalorbit in which the spacecraft will approach within 46 kilometers. of the lunar surface.

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READOUTCONTINUES

MOON'S AXIALROTATION

TEST PHOTOS& GRAVITYANALYSIS

INITIALORBITINJECTION

STARTPHOTOS

~7DAY,

15 DAYS 5 DAYS

\ ~ J20 DAYS ',ZERO

"--- - 28 ~~DAYS "

",\

PHOTOS

START PHOTOREADOUT

DIRECTION OFMOON ROTATION•__________6~OOE

THE SPACECRAFT LUNAR ORBIT IS FIXED IN SPACE, SUCH THAT FOR THE DURATION OF THEMISSION, THE ORBIT WILL CROSS THE LUNAR EQUATOR IN THE SUNRISE ZONE. AS THEMOON ROTATES ABOUT THE EARTH, IT~ AXIAL ROTATION BRINGS THE FIRST TARGET SITEINTO THE SUNRISE ZONE, AND SIMULTANEOUSLY, UNDER THE SPACECRAFT ORBIT. DURINGTHE FOLLOWING EIGHT DAYS, THE SPACECRAFT TOGETHER WITH THE SUNRISE ZONE WORKTHEIR WAY ACROSS THE AREA OF INTEREST AS THE LUNAR ORBITER PHOTOGRAPHS EACHOF THE TARGET SITES.

Figure 13 Lunar Orbits

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THE PHOTOGRAPHIC MISSION

To determine height and slope of lunar mountains and craters by photometric tech­niques, photography must be done shortly after the lunar sunrise, when shadows arenear maximtun. Since, during the life of the photographic mission, the spacecraftorbit will vary only slightly with respect to the Sun, the perilune of the initial orbit isplaced over the lunar sunrise zone, not over the target area. As the spacecraft con­tinues orbiting over the lunar sunrise, the rotation of the Moon on its axis brings thetarget area under the orbit and, simultaneously, into the sunrise zone (Figure 13).

As the target area moves into the sunrise zone, the rocket engine again is pointedagainst the direction of flight and fired, changing the initial orbit perilune (200 Idlom­eters) to the design perilune altitude of 46 kilometers over the target sites. Photo­graphs are taken at this altitude as the Moon's rotation continually presents a newlandscape on each pass over the target area.

Each exposure of the dual-lens camera simultaneously exposes one high-resolutionframe and one medium -resolution frame (Figure 14a). The camera can be commandedto take 1, 4, 8, or 16 consecutive exposures on an orbital pass over a target site witheither a short-time interval (2.2 seconds) or a long-time interval (8.8 seconds) betweenexposures. Figure 14b illustrates the design mission coverage of one target sitewhichrequires 16 consecutive exposures at the 2. 2-second interval. As can be noted, thisprovides continuous high-resolution coverage with forward edge overlap. Contiguoushigh-resolution coverage (Figure 14c) is similarly obtained by repeating the photo­graphic pattern on successive orbits.

Overlap of 50 percent or more of the medium-resolution photographs allows stereoviewing, revealing surface contours not perceptible from single photographs. Thistype of coverage (Figure 14d) is obtained by utilizing the long interval of time betweenconsecutive exposures and by repeating the photographic pattern on alternate orbits.

The flexibility of the photographic subsystem, coupled with orbit variables such as in­clination, perilune altitude, and orbital period, provide great flexibility in selectingthe size, shape, and location of the area to be photographed. For example, by increas­ing the perilune altitude, the area coverage can be extensively increased - at the ex­pense, however, of reduced photographic resolution.

The exposed film is developed into a film negative and stored on a take-up reel. Toreturn the photogt'aphs to Earth, a beam of light 1/20 the thickness of a human hairscans the image on the negative. The resulting electrical signal, varying with theamount of light shining through the negative onto the face of a photomultiplier tube, istransmitted to a visible deep space station via the directional antenna.

The information can be transmitted only when the spacecraft is visible to the Sun forsolar energy and the Earth for line-of-sight communication. These conditions occurmore than half the time during a typical mission. Several frames of information may

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MULTIPLE OVERLAPMEDIUM RESOLUTION~----OVERLAP HIGH

RESOLUTION

2

30

4;g

S!-t

5 Zc~

6 c:Dm;g

7ORBITAL PERIOD3.5 HOURS 8

FORWARDPERILUNE ALTITUDE OVERLAP 946 KILOMETERS HIGH DIRECTION

RESOLUTION OF FLIGHT 10

HIGH RESOLUTION

MEDIUM RESOLUTION

o. SINGLE EXPOSURE AT THE 46 KILOMETER ALTITUDE;HIGH RESOLUTION COVERAGE IS 16.6 KM X 4.15 KM,MEDIUM RESOLUTION COVERAGE IS 37.4 KM X 31 .6 KM.

b. DESIGN MISSION TARGET SITE COVERAGE, 16 CONSECUTIVEEXPOSURES DURING ONE ORBITAL PASS OVER THE TARGETSITE. THE INTERVAL BETWEEN EXPOSURES (APPROXIMATELY2.2 SECONDS) IS TIMED TO PROVIDE OVERLAP OF THE HIGHRESOLUTION FRAMES.

C. EIGHT TYPICAL FRAMES FROM A SITE EXAMINATION TYPEMISSION - CONTIGUOUS HIGH RESOLUTION COVERAGE ISPROVIDED BY RAPID EXPOSURE RATE AS IN (b) TO GIVE HIGHRESOLUTION FORWARD OVERLAP AND BY PHOTOGRAPHINGON CONSECUTIVE ORBITS (9 & 10) TO GIVE HIGH RESOLUTIONSIDE OVERLAP.

d. EIGHT TYPICAL EXPOSURES FROM A SITE SEARCH TYPE MISSION ­STEREO MEDIUM RESOLUTION COVERAGE IS PROVIDED BY IN­CREASING THE TIME INTERVAL BETWEEN EXPOSURES (APPROX­IMATELY 8.8 SECONDS) TO OBTAIN 50% MEDIUM RESOLUTIONFORWARD OVERLAP AND BY PHOTOGRAPHING ON ALTERNATEORBITS (7 & 9) TO OBTAIN MEDIUM RESOLUTION SIDE OVERLAP.

Figure 14 Photographic Modes

be transmitted before the completion of photography, but, since transmission andphotography cannot occur simultaneously, most frames are transmitted after allpictures are taken.

At the receiving deep space station, the information is displayed line by line on theface of a kinescope, a tube similar to that used in television. Cameras photographthe kinescope display on 35 -mm film.

To simplify analysis of the reconstructed photos, the 35-mm film is flown to a lunar­data-processing laboratory where the 35-mm film strips are reassembled into 9- by14-inch photographs. A composite photograph of the lunar surface, photographed inone high-resolution frame, measures 14 by 55 inches, with a scale of 1 inch equal toapproximately 1000 feet.

Data recorded on the film in flight and magnetically taped telemetry d~ta provide time,height, location, lighting conditions, and related information required for completephotographic analysis.

After transmitting all photographic data, the Lunar Orbiter continues to transmit in­formation on micrometeoroids, solar flares, and magnetically trapped particles.Periodic tracking will refine the information on the Moon's gravity and shape. Whennitrogen gas for the reaction control thrusters is exhausted; the spacecraft cannotmaintain Sun/Canopus orientation. The solar panels no longer face the Sun to gen­erate electrical power, and the spacecraft becomes silent, but continues its circuitof the Moon.

THE SPACECRAFT

The 850-pound Lunar Orbiter spacecraft (Figure 15) is 5 feet in diameter and 5.5 feettall, excluding the solar panels and antennas. During launch, the solar panels arefolded up under the spacecraft base, and the antennas are held against the side of thestructure. A nose shroud only 5 feet 5 inches in diameter houses the entire spacecraft.With the solar panels and antennas deployed, the maximum span is increased to 18.5feet along the antenna boom sand 12 feet 2 inches across the solar panels.

The spacecraft's structure consists of a main equiPment mounting deck and an uppermodule supported by trusses and an arch. The module supports the gimbaled velocity­control engine and propellant tanks and may be removed as a unit for engine testing.It also carries, directly under the engine, the high-pressure nitrogen tank which pro­vides pressurization for the engine feed system and the attitude-control thrusters.

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OXIDIZER TANK

CANOPUSSTARTRACKER

INERTIALREFERENCEUNIT

MICROMETEOROIDDnECTOR

FLIGHTPROGRAMMER

ATTITUDE CONTROL THRUSTERS

HEAT SHielD

VELOCITY UPPE S CCONTROL ENGINE~--_-~.t--.J--"1or- MOD~lrRU TURAl

ATTITUDE HEAT SHIELDCONTROLTHRUSTERS COARSE

FUEL SUN SENSOR

TANK OXIDIZERTANK

MICROMETEOROI DDETECTOR

EQUIPMENT MOUNTING DECK

PHOTOGRAPHICSUBSYSTEM

lOW-GAIN(if;ANTENNA

Figure 15 Lunar Orbiter Spacecraft

21

PHOTOGRAPHY

The photographic subsystem (Figure 16) is housed in a pressurized, temperature­controlled container. The precision equipment can expose film, develop high-qualitynegatives, and convert the images into electrical signals for transmission to Earth.

The dual-lens camera uses a 24-inch focal-length lens for high-resolution photographsand a 3-inch focal-length, wide-angle lens for medium-resolution photographs. Shutter,platen, and image-motion compensation (IMC) are provided for each lens, but the film,film advance, and shutter operation are common to both. The 70-mm high-definitionfilm is relatively immune to radiation fogging, but requires slow shutter speed.

A velocity/height (V IH) sensor (Figure 17) is used in conjunction with the 1MC to reduceimage smear that would normally result from high spacecraft velocity and slow shutterspeed. The camera mode (number and rate of exposures) is selected by ground com­mand and controlled by camera electronics.

,

FilmTake-up

READOUT

Video__--. Signals to

Communi cationsSystem

PROCESSOR

Dryer

ReadoutLooper

HeatSource

CAMERA,...----~,.-----"" r,.----------,

Figure 16 Photographic Subsystem

22

'" rI "-

I '.I

I/

Di recti on " Iof Flight 1/ ............, '$

"'IIIIIIIIIII~II~I / )//L Jl

....... ...J I/ .......r--'

/-.1

//

II I,_r-/

- ... '

'-.

Distance Traveled by SpacecraftDuring Single Exposure

-_ ...

--::::.

The velocity/height (V/H) sensor provides the spacecraft with velocityand height information by sampling a ring of lunar surface through thehigh resolution lens and comparing successive scans. The angular positionchange of consecutive scans with respect to time gives the velocity/heightchange ratio. This ratio is used for image motion compensation to"lock on" a ground target during each exposure to reduce image smearor blur. The V/H data is also used to control time between exposures,to control spacecraft yaw angle during photography, and is telemetered toaid in photograph analysis.

Figure 17 Velocity Height Sensor

23

The automatic camera sequence proceeds as follows: lenses are uncovered and theV/H sensor is activated; the film is clamped to the platen and flattened by vacuum; the!Me clutch is engaged and moves the platen to compensate for ground speed; and theshutter is opened to expose the film. Mter exposure, the IMC clutch is disengaged,the platen is returned to the original position, and the film is released and advancedin preparation for the next film exposure.

Each shutter cycle exposes one medium-resolution (square) frame and one high­resolution (rectangular) frame centered around the same surface area (see Figure 18).Because of the spacing between the lenses, the two images are not adjacent on the film,but subsequent exposures are interspersed on the film to avoid film waste. The time­of-exposure data, sandwiched between photograph frames, and edge data, preexposedon the film, aids in the reconstruction and evaluation of finished photographs.

Film processing is accomplished by temporary lamination of the exposed film with aprocessing web that has been presoaked in a developing solution. The damp negative,dried by a thermostatically controlled heater, is stored pending readout.

24

(A) GROUND FORMAT

I""-----37.4km-----.......·I (C) EDGE DATA STRIP - Includes9 Level Gray Scale andResolving Power Charts

...-----2.29 mm-----~

Time 12

No. 12

Fi 1m Travel '"Duri ng Readout y

No. 13

mm

TIme 11

No. 1165

7.65 mm

No. 12

Time ofExposure No. 10

3mm2 mm (Blank) (8) FILM FORMAT(TIme)

Di recti on ofSpacecraftMovement

Figure 18 Film Format

Note:Ground Format Dimensionsare for Exposure from a25 Mi Ie Altitude

j-16.6km-jt4.15 km__....__-l

Exposure No. 11 No. 10

It4IIi---219 .18 mm---......... 3 mm (Blank)

31.6 km

~ Film Travel,. During Photography

The film scanner/readout process (Figure 19) is accomplished in small increments bya very narrow light beam projected through the film negative onto a photo cell that con­verts light energy into a video signal. Nearly 17,000 horizontal scans of this 5-micronbeam are used to scan each 1/10-inch segment of the film before the film is advancedand the process repeated. The readout of photographic intelligence gathered by asingle camera operation requires 45 minutes.

For mission flexibility, the camera, film processor, and scanner/readout mechanismoperate independently. During a single orbit, up to 16 exposures may be made andprocessed. Selected exposures can be scanned and transmitted between photographicsequences, but most exposures will be stored for scanning and transmission at thecompletion of photography.

ELECTRICAL POWER

Solar panels develop all necessary electrical power during the long spacecraft mission.When solar energy is not striking the solar cells, such as during spacecraft maneuversor occultation of the Sun, electrical power is supplied by a 28-volt battery (Figure 20).Power regulators and controllers protect the solar panels, battery, and spacecraftsystem s from unusual power fluctuations. The number of solar cells provided allowsfor the possibility that some cells may be damaged by micrometeoroicls during themission.

Video Signalto Transmi tter

Photomultiplier Tube

Figure 19 Film Scanner

26

....z-:!:~-~

Ot: -

50­U")

SunlightCondition

Powor toOther SiCSubt)'lteml

OUTPU.TVOLTAGEREGULATOR

Spacecraft Maneuver(Attitude DeterminesWhether Solar Panelsare Receivi n9 SolarRadiation or ore Occulted)

BATTERVCHARGERCONTIOLLEI

Figure 20 Electrical Power

ATTITUDE CONTROL

Throughout all phases of the mission, the spacecraft attitude must b~ accurately con­trolled. The Sun/Canopus reference attitude provides optimum utilization of solarradiation for spacecraft power and serves as the reference attitude for spacecraftmaneuvers. In establishing and maintaining this attitude, Sun sensors control rotationabout the pitch and yaw axes, and a star tracker controls rotation about the roll axis.Signals from these sensing devices control nitrogen v;as ejection from the appropriatethrusters (Figure 21) to acquire and maintain the necessary sPacecraft orientation.

GYros in the inertial reference unit maintain the reference attitude whenever the Sunor Canopus is hidden. When other attitudes are required, a command from Earth pro­vides the exact rotation required around each axis, and rate outputs from the gYrosenable the flight programmer to determine when the new attitude has been reached.The gyros then control the new attitude.

yz

FLIGHT ELECTRONICSCONTROL ASSEMBLY MODE

SWITCHING(FLIGHT

..} PROGRAMMER)

..: ( SUN-cANOPUS. REFERENCE MODE

COMPARISON... (CLOSED LOOPELECTRONICS)

(COMMANDED•• ~ MANEUVER

•• MODE

.

=} fNERTIALREFERENCE

•• MODE

~

f PHOTOGRAPHY.. - MODE"

Attitude Control ModesFigure 21

Crab Error ISill'!Cl fram ._._.V/H Sensor

INERTIAL REFERENCE UNIT

CAN0 PUS SENSOR

PHOTOCELLS

Ccmmanded } _._._._AttitudeReorientation - - - - -(From Decoder) .

28

VELOCITY CONTROL

The initial spacecraft trajectory is established by the Atlas/Agena launch vehicle, butminor changes to the trajectory and the velocity changes required for lunar orbit in­sertion must be accomplished by the spacecraft. An attitude maneuver establishes thedirection of thrust; velocity change, detected by an accelerometer and compared withrequirements stored in the flight programmer, determines thrust duration.

When the flight programmer commands the velocity-control engine valves to open (Fig­ure 22), gas pressure, acting on propellant tank bladders, forces the fuel and oxidizerinto the lOO-pound-thrust engine. No ignition system is required because the propel­lants ignite when mixed, and thrust continues until the engine valves are closed.

~~

To Sun

~~To Canopus

No. Velocity Maneuver Maneuver Sequence

1- First Midcourse a. Attitude control system "polntst' theCorrection rocket engine in accordance with

ground computed data.b. Val ves are opened and fuel, under

nitrogen pressure, flows Into enginewhere combustion occurs. Accelerometersenses the commanded velocity changeand closes valves to shut down engine.

c. Sun-Canopus orientation Is gained

2. Second Midcourse a.,b.,c. SameCorrection

3. Initial Orbit a.,b. SameInjection c. The slowed spacecraft, unable to escape the

gravitational field of the moon, becomesa satellite of the moon-canopusorientation Is regained

4. Photographic Orbl t a.,b. SameInjection c. Spacecraft velocity Is reduced, lowering

perilune altitude to 46 kllometen, asSun-Canopus orientation Is regained.

//

Figure 22 Velocity Control

29

COMMUNICATIONS

Simultaneous reception of command messages and transmission of photographic, per­formance, and lunar environmental data is accomplished by the spacecraft communi­cations system (Figure 23). The communications system also provides velocity andranging signals used by the deep space network tracking system.

All incoming signals are received by the low-gain antenna. The transponder automati­cally responds to the rf carrier and any range code to assist the DSIF in tracking,velocity, and range determination. Commands from the DSIF are routed to the com­mand decoder and stored. The command, as received, is transmitted to Earth whereit is checked for accuracy. If verified, an execute command is transmitted to thespacecraft, and the information stored in the decoder is advanced to the flightprogrammer.

Performance telemetry data and data gathered by radiation and micrometeoroid sensorswill be encoded, multiplexed, and transmitted to Earth continually by means of the low­gain antenna and low transmitter power. Whenever photographic data are to be trans­mitted' the photographic system readout mechanism and traveling-wave-tube amplifiermust be turned on by command from Earth. The photographic data, mixed with theperformance and environmental telemetry data, are transmitted via the high-gainantenna.

30

Carrier &Pseudo Noise

Range Codefrom DSIF

TRACKING AND RANGING MODE COMMAND MODE

Performance Telemetry &lunar Environmental

Data Telemetry to DSIF

LOW POWER MODE

Vide.'_ / /

Ph.tog,.phy /c

PerformanceTlM Samples "

MICROMETEOROI DDETECTOR

HIGH POWER MODE

Figure23 Communications Modes

31

TEMPERATURE CONTROL

Temperature control is provided because many of the spacecraft components could bedamaged by the temperature extremes encountered in space (Figure 24). An alumi­nized mylar thermal barrier encloses the area between the upper heat shield andequipment mounting deck with only the Canopus star-tracker lens exposed. In additionto this flexible insulating shroud, the engine heat shield is insulated, leaving only theequipment mounting deck for heat transfer. The equipment mounting deck is coatedwith a substance that has a high heat emission-absorption ratio, which helps keep thespacecraft temperature between 85°F (when the craft is in direct sunlight) and 35°F(when it is not). When the temperature drops too low, small electric heaters willsupplement the heat generated by operating equipment.

32

Figure 24

Almost All Heat TransferThrough This Surface

Surface Pai nted to Assure HighRadiation to Absorption Ratio toPrevent Overheati ng (MaximumExterior Surface Temperature + 250°F)---~-""'IL"

,,~

76% of SolarEnergy Refl ected

Temperature Control

Alternating layers of

aluminized Mylar&

Dacron scri m cI oth

Thermal Barrier Detai I

InsulatedHeat Shield

GROUND DATA HANDLING

The Space Flight Operations Facility (SFOF) at Pasadena, California, has primary re­sponsibility for Lunar Orbiter mission control. A telephone-teletype link and a high­speed data link (HSDL) maintain continuous communication between the SFOF, the AirForce Eastern Test Range (AFE TR), and the deep space stations making up the DeepSpace Instrumentation Facility (DSIF). Tracking and performance telemetry data arerelayed to the SFOF from AFETR during the launch phase and from deep space stationsthroughout the spacecraft mission. Commands are teletyped from the SFOF and veri­fied at the deep space stations before transmission to the spacecraft.

AFETR EQillPMENT

Existing equipment at the AFETR is supplemented by a blockhouse spacecraft console,umbilical cabling, and rf repeater antenna system, and by the mobile checkout vanduring launch operations. This special equipment is used to monitor the spacecraftperformance during countdown. Postlaunch command, tracking, ranging, and telem­etry data recording is accomplished by existing AFETR equipment.

DEEP SPACE STATION EQUIPMENT

Deep space stations are located so that at least one is always in contact with the space­craft after its insertion into translunar trajectory. For the first 30 days of the mission,the spacecraft is monitored continually; thereafter, 30 minutes per day are set asidefor spacecraft monitoring and tracking.

Spacecraft tracking and ranging is accomplished by existing DSIF equipment (Figure25). Command equipment is installed at each deep space station for receiving, de­coding, and verifying commands from the SFOF before these commands are relayedby transmitter to the spacecraft.

Most of the Lunar Orbiter mission-dependent equipment installed in the DSIF is forprocessing and displaying data telemetered from the spacecraft. The rf carrier fromthe spacecraft is received by the nss antenna-receiver system. The 30-kc subcarrier,containing spacecraft performance and lunar environmental data, is routed to perform­ance telemetry equipment and recorded on magnetic tape. This equipment demodulates,sYnchronizes, and decommutates the data for computerized transmission to the SFOFvia the HSDL or via teletype. The performance telemetry equipment also providesprintout of all 120 telemetry channels, and displays anyone of these at the deep spacestations.

34

SPACECRAFTTLM - TELEMETRY

TTY - TELETYPE

HSDL - HIGH SPEEDDATA LINE

1······················1 EXISTING~::::::::::::::::::::: DSS EQUIPMENT

lOMC

VIDEO30 KC DEMODULATOR

30 KCGROUNDRECON STRUCTIONELECTRONICS

QUALITYEVALUATIONVIEWER

H~AGi,fEffc::::::TAPE :::.:::RECORDER ~:~.:••:.:.:.:.:~•••••••••• I ••••••••••••:.

BITSYNCHRO­NIZER

30 KCDEMODULATOR

MANUAL

COMMANDS

COMMANDDRAWERNO.2

COMMANDCONDITION­ING UNIT

~:~pAPER' TAPE' .':~:::READER AND::.~~~~.~~~9~~Q~.~~:

COMMANDDRAWERNO.1

Vlo

~

~u

DECOMMU­TATORDRAWERNO.1

I

DECOMMUTATORDRAWERNO.2

Figure 25 Deep Space Station Data Flow--Operational

35

The photographic data lO-Mc carrier is demodulated and routed from the receiver tothe ground reconstruction equipment (Figure 26) where it is displayed on kinescopes.Motion-picture cameras expose 35-mm film to the ldnescope presentation to provide

Figure 26 Ground Reconstruction Electronicsand Film Recording Equipment

a permanent film record of the photographic data. Two film records are made at thedeep space station. Small portions of the film are processed and viewed on a qualityevaluation viewer so that picture quality can be determined. The viewing of early filmallows determination of the necessary remotely controlled adjustments to the space­craft camera or readout mechanism that will improve quality of subsequent pictures.

Tape recordings of spacecraft performance, lunar environmental telemetry data, andfilm record of photographic data are air transported to the lunar-data-processing lab­oratory. Because of Earth rotation and deep space station locations, the Lunar OrbitersPacecraft, when visible to Earth, can be monitored continuously by the deep spacenetwork, but each deep sPace station can collect only a portion of the informationgathered during the mission. Tracking, telemetry, and photographic data are consoli­dated at the lunar-data-processing laboratory (Figure 27).

36

To NASALangley

Process

~Q.7 Q;)

70mm Negative

9

35mm Strips

•nnn~\\-REASSEMBLY

LUNAR DATA PROCESSING LABORATORY.......................................................................................

ACQUISITION

:::.•. <::.::-:''' ~=:t==.._:::::~_~=.1..=6.~=···:::I·1=r;..<'=.:;=..6='~!~·~ Edge Data....-----~ ,.'~ 1..:> L.:::-:..J ...-' . - "901

Kinescope Camera

-----------.........~

RECONSTRUCTION

~L SiC ANTENNA

SPACECRAFT ,~

~i·~~·~;~C~.;;..~~~~~~ ~~;~;;.;~~;~;~..•......•.•••......•....•..•...•..•••.....•••i1/ Film 0 i

Record

L-_--..__~--lIII!Is~y~nc__~.c(]Video

Figure 27 Photographic Data Acquisition, Reconstruction, and Reassembly

LUNAR-DATA-PROCESSING LABORATORY

At the lunar-data-processing laboratory t tracking data and telemetry data are integratedinto a master mission tape. Also, the film record is reassembled into the necessaryformat for analysis.

Each readout scan in the spacecraft is recorded on the 35-mm ground film as a segment,approximately 0.75 by 16 inches (Figure 28). A set of 14 of these segments are edge­matched to form a composite that is photographically reduced to a 9- by 14-inch re­assembled photograph. A high-resolution exposure composite requires seven of thesereassembled photographs and measures 14 by 55 inches. The high-resolution photo­graphs gathered during the design mission would form a composite 18- by 45-footphotograph, which would represent a 63- by 166-kilometer lunar surface area from a46 -kilometer altitude.

Langley Research Center receives reassembled photographs, the mission master taperecording, and preliminary analysis reports within 30 days after transmission to Earth.

---.55.50..--------·1

(D) COMPOSITE OF ONE HIGH RESOLUTION FRAME(6.2 REASSEMBLED PHOTOS)

(A) SPACECRAFT FILM FORMATr ~~70 mm RESOLUTION

L FRAME

~::::;t'~=l======~~:::::::======.::=================:i

...35 mm

.L---:" ---::~-----_----_:__--

Figure 28 Photo Reassembly

38