Apollo Experience Report Simulation of Manned Space Flight for Crew Training

8/8/2019 Apollo Experience Report Simulation of Manned Space Flight for Crew Training

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N A S A T E C H N IC A L N O T E NASA TN D-7112

Uv,uz CASE F I L ECOPY

APOLLO EXPERIENCE REPORT -SIMULATION OF M A N N E D SPACE FLIGHTFOR CREW T R A I N I N Gty C. H . WodGPzg, Studey Faber, John J . Vun Buckel,Cbades C. Olasky, Wayae K. Williums, J u h Le C. Mire,and Jumes R. HomerMunazed Spucecrafi CenterHozcstoiz, Texas 77058 ? _dNAT ION AL AERONAUTICS AND SPACE ADMINISTR ATION WA SHIN 6TU# , 0. C. MARCH 1973


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I . R e m No. I 2. Government Accession No. I 3. Recipient's Catalog No.NASA TN D-7112 IAPOLLO EXPERIENCE REPORTSIMULATION O F MANNED SPACE FLIGHT FOR CReWTRAINING

C. H. Woodling, Stanley Faber, John J. Van Bockel, Charl es C.Olasky, Wayne K. Williams, John L. C. M i r e , and Ja mes R.Homer, MSC

7. Authorls)

3. Performing Organization Name and AddressManned Spacecr aft CenterHouston, Texas 77058

2. Sponsoring Agency Name and AddressNational Aeronautics and Space AdministrationWashington, D. C. 20546

17. Key Words (Suggested by AuthorW)' Project Mercury- Gemini Progra m* Apollo Program- Crew Training Simulators

March 19736. Performing Organization Code

18. Distribution Statement

8. Performing Organization Report No.MSC S-346

10. Work Unit No.076-00-00-00-72

19. Security Classif. (of this report) 20. Security Claaif . (of this page) 21. No. of PagesNone None 60

11. Contract or Grant No.

22. Price$3.00

13 . Type of Report and Period CoveredTechnical Note

14. Sponsoring Agency coder5. Supplementary NotesThe MSC Dire cto r waived the use of the International System of Units (SI) for this Apollo Experi-

ence Report, because, in his judgment, the use of SI Units would impair the usefu lness of thereport or result in excessive cost.6. Abstract

Through space-flight expe rience and the development of simulat ors to meet the assoc iated train ingrequire ments, se ve ra l fact ors have been established as fundamental for providing adequate flightsimu lato rs for c rew training. The development of flight sim ulators from Proj ect Mercury througkthe Apollo 15 mission is described in this report. The functional uses , c hara cter ist ics, and de-velopment problems of the var ious simulato rs are discussed fo r the benefit of future program s.


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CONTENTS

SectionSUMMARY . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .INTRODUCTION . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .BACKGROUND . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Classification . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .Discussion . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

CREW-STATION FIDELITY . . . . . . . . . . . . . . . . . . . . . . . . . . .Controls and Displays . . . . . . . . . . . . . . . . . . . . . . . . . . . . .Stowage . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .Lighting . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .Aural Cues . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .Markings and Nomenclature . . . . . . . . . . . . . . . . . . . . . . . . . .Space Suit and Cabin Environment Requirement . . . . . . . . . . . . . . .Crew Couches . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .Crew-Station Hardware . . . . . . . . . . . . . . . . . . . . . . . . . . . .Concluding R ema rks . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

VISUAL DISPLAY SIMULATION . . . . . . . . . . . . . . . . . . . . . . . . .Display System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .Celestial Simulation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .FarBodies . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .Target Vehicle f or Rendezvous . . . . . . . . . . . . . . . . . . . . . . . .Target Vehicle f or Stationkeeping and Docking . . . . . . . . . . . . . . . .Near-Body Scenes . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .Landing Scenes . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

page11256111112121212131313141414161819202227

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SectionConcluding Remarks . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

MOVING-BASE SIMULATIONS . . . . . . . . . . . . . . . . . . . . . . . . . .Application of Moving-Base Simulators . . . . . . . . . . . . . . . . . . . .Simulation Experience . . . . . . . . . . . . . . . . . . . . . . . . . . . . .Concluding Remarks . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

SIMULATOR CONFIGURATION MANAGEMENT . . . . . . . . . . . . . . . . .Configuration Tracking . . . . . . . . . . . . . . . . . . . . . . . . . . . .Configuration Control . . . . . . . . . . . . . . . . . . . . . . . . . . . . .Configuration Accountability . . . . . . . . . . . . . . . . . . . . . . . . .Concluding Re ma rk s . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

CONCLUSIONS . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .REFERENCE . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .APPENDIX-SIMULATORDESCRIPTION . . . . . . . . . . . . . . . . . . .

Page29293134363636404345464647

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TABLES

Table PageI FLIGHT CREW SIMULATORS. . . . . . . . . . . . . . . . . . . . . . 3II SIMULATOR TRAINING SUMMARY. . . . . . . . . . . . . . . . . . . 4

111 SIMULATOR USE FOR FLIGHT CREW TRAINING . . . . . . . . . . . 4rv SIMULATED NETWORK SIMULATIONS

Figure

56789101112131415

(a) Gemini Program . . . . . . . . . . . . . . . . . . . . . . . . . . 8(b) Apollo Program . . . . . . . . . . . . . . . . . . . . . . . . . . . 8FIGURES

PageMercury procedures simulator . . . . . . . . . . . . . . . . . . . . . 5Gemini mission simulator . . . . . . . . . . . . . . . . . . . . . . . . 5Command module simulator . . . . . . . . . . . . . . . . . . . . . . . 5Lunar module simulator. . . . . . . . . . . . . . . . . . . . . . . . . 5Lunar landing training vehicle . . . . . . . . . . . . . . . . . . . . . 6Command module procedures simulator . . . . . . . . . . . . . . . . 9Lunar module procedures simulator . . . . . . . . . . . . . . . . . . 9Translation and docking simulator (Apollo) . . . . . . . . . . . . . . .Dynamic crew procedures simulator (Apollo) . . . . . . . . . . . . . 10

10

Typical reflective infinity display system . . . . . . . . . . . . . . . 15Docking model of CSM n lunar docking simula tor . . . . . . . . . . . 21Air-lubricated free-attitude trainer . . . . . . . . . . . . . . . . . . 23

28Partial-gravity simulator . . . . . . . . . . . . . . . . . . . . . . . . 30Mobile partial-gravity simulator . . . . . . . . . . . . . . . . . . . . . 30Lunar-surface model (scale 1:2000) of lunar module simulator . . . .

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Figure Page16 Lunar landing res ea rc h facility (Langley Resea rch Cen ter) . . . . . . . 3117 Multiple-axis spin test inertial facility (Lewis Resea rch Center) . . . . 3118 Centrifuge (Naval Air Development Center. Johnsville.Pennsylvania) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3119 Rendezvous and docking simulator (Langley Resea rch Center) . . . . . 3320 Tran slati on and docking simulator (Gemini) . . . . . . . . . . . . . . . . 321 Dynamic crew procedures simulator (Gemini) . . . . . . . . . . . . . . 3422 Lunar landing training vehicle simulator . . . . . . . . . . . . . . . . . 3423 Simulator modification flow . . . . . . . . . . . . . . . . . . . . . . . . 4124 Mercury part-task train er . . . . . . . . . . . . . . . . . . . . . . . . 4825 Centrifuge (Manned Spacecraft Cent er) . . . . . . . . . . . . . . . . . . 51

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ALFACCACCBCCPCFECMCMPSCMSCRTCSMDCPSE C PEIGEOFO VGFEGMSKSCL&A

LCRLLRFLLTVLMLMPS

ACRONYMS

air-lubricated free-attitudecontract change authorizationConfiguration Control BoardConfiguration Control Pane lcontractor -f urnished equipmentcommand modulecommand module procedures simulatorcommand module simulatorcathode-ray tubecommand and service moduledynamic cr ew procedures simulatorengineering change proposalelectronic image generatorengineering o rde rfield of viewGovernment-furnished equipmentGemini mission simulatorKennedy Space Centerlanding and ascent

lunar module change requestlunar landing re se ar ch facilitylunar landing training vehiclelunar modulelunar module procedures simulator

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LMSMCCMCRMEPMPSMRMSCPDRRDSRECPSCPSCRTDSTV3 -D

lunar module simulatorMission Control Centermanufacturing change requestmission effects projectorMercury procedures simulatormodification requestManned Spacecraf t Cente rpreliminary design reviewrendezvous docking simulatorreque st f or engineering change proposalSimulator Control Panelsoftware change requesttrans lation and docking simulato rtelevisionthree dimensional

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APOLLO EXPERIENCE REPORTSIMULA TION OF MANNED SPACE FL IGHT FOR C R W T RA IN IN G

B y C . H. Woodling, Stanley Faber, Joh n J. V a n Bockel,C h a r l e s C . Olasky, Wayne K. Wil l iams,J ohn L . C . Mire , and James R. HomerManned Spacecraft CenterS U M M A R Y

Fr om the ear'ly phas es of Projec t Mercury through the Gemini and Apollo Pro-gra ms, flight simu lator s have been the key elements in the astronaut training programs.As the miss ions pro gress ed in complexity, the sophistication, number, and varie ty ofsimul ator s employed fo r astronau t training were incr eased correspondingly. Throughspace-flight experience and evolution of the sim ulato rs to meet associated t rain ing re-quirements, sev era l fac tor s have been established as critical and basic f or providingadequate flight simulato rs fo r crew training. Included in these factors are high-fidelitycrew stations , especially in the area of con trols and displays; accu rate simulation ofthe spacecraft syste ms, including the guidance computer and navigation sys tem; com-plete visua l display s ys te ms for simula ted out-the-window sce nes ; and cert ain moving-base simu la to rs fo r high-fidelity training in parti cula r portions of the missions. Thesignificance of the se fac to rs fo r new pro gra ms wi l l depend to a large degree on the mis-sion objectives and requirements. Nevertheless, flight simul ators incorporating someof thes e it ems in their design and operation will be vital in future astronaut trainingprograms.

INTRODUCTIONIn this report, "simulation" refers to th e operation of train ers f or instructingflight crew s in the various con trol and monitoring procedures of manned spacec raft.The purpose of simulation for crew training is high-fidelity duplication of a wide rangeof inflight conditions and vari able s to obtain precise flight cre w respo nse to sophisti-

cated and critical mission events. Repeated simulation exposure allows the crew tobecome proficient in interfacing with the flight hardware and ground-support e lemen ts,thus enhancing mission success and safety. In this context, a simulator is defined asa complex set of hardware (including computers, visual display systems, and simulatedcrew stations) that presents, with a high degree of accu racy, the total flight character-istics of the actual spacecra ft and mission. Excluded fr om this classification are train-ers that are full-scale spacecraf t mockups designed pri mar ily to refine crew tasksrelated to the handling of equipment in performing such activi ties as stowage, ingressand egr es s, maneuvering in zero -g and 1/6-g (lunar surf ace) conditions, handling


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photographic equipment, and genera l housekeeping within the co nstraints imposed byflight hardware and the space-flight environment. Although the r ema inder of t his paperis concerned primarily with simulators, it is only prope r to give recognition t o thesemockups that played such a significant and necessary role in complementing the simu-lator training for all major mission phases. A typical example was the extensive trai n-ing program c ar ri ed out fo r the Apollo lunar landing mission using full-scale mockupsof the lunar module (LM) for lunar -su rface t raining and of the command and se rv ic emodule (CSM) for ingress and egress and scientific instrument module training.The topics of thi s r epo rt include some of the mor e significant asp ec ts of space-flight simulation: crew-sta tion fidelity, visual display req uir ements , moving-base si m-ulations, and configuration management. The various topics relate primarily theexperience gained in the se a reas of simulation since the beginning of manned spaceflight. It is hoped that future programs might benefit through discussion of thisexperience.Credit for the se par ate section s of this repor t is given to those who, through the irdi re ct involvement with manned-space-flight simulation and training, were able to re-por t firsthand the data contained herein. These individuals and the ir respect ive sec tionsare as follows: C. H. Woodling and John J. Van Bockel, background and discussion;James R . Homer and John L . C. Mire, crew-station fidelity; Stanley Fabe r, visualdisplay requirements ; Wayne K. Williams, moving-base simulations; Cha rle s C. Olasky ,simulator configuration management; and C. H. Woodling, over al l compilation andediting.

BACKGROUNDIt is pertinent to preface these discussions with a listing of the crew sim ula tor s

employed in support of manned spa ce flight, to note the s imu lat or use for tra ining duringeach of the flight progra ms, and to disc uss briefly the history of simul ator s fro m thebeginning of Project Mercury.The crew-training simula tors used fo r P roject Mercury and for the Gemini andApollo Progr ams a r e listed in table I. These simul ator s a r e described in the appendix.The crew-training use of the various s imu lat ors throughout the flight program is pre-sented in tables 11 and III. During Project Mercury and the Gemini and Apollo Pro-gr am s, each crewman (on an average) spent one-third o r more of the tota l trainingprogram t im e in simulations (table 11). The crews of the lunar landing missions (Apollomissions 11 to 15) spent slightly mor e than 50 percen t of t he ir tot al training time insimul ator training. In addition to the actual time spent in the various simulators, other

training activities were accomplished in support of the sim ula tor training. Fo r ex-ample, the Apollo crews averaged more than 150 hours of sy st em s briefings as a pre-requisite to simulator training.A breakdown of the s imulator us e by pro gram and simulation facility is presentedin table III. Progressing from Proje ct Mercury through the Gemini Progr am to theApollo Progr am, the number and complexity of the simulators inc reased. Also, it isinter esting to note the increased emphasis on the mission sim ula tor s. The 708 hoursspent by the crew s in the Project Mercury procedures tra ine rs r epres ent

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approximately 53 percent of the total simulation time of 1330 hours. In the GeminiProgr am, the hours logged in the mission simulators are 67 percent of the total; and,in the Apollo Program, the command module simulator (CMS) nd lunar module simu-lator (LMS) hours combined are 80 percent of the total of 29 967 hours spent in allApollo simulations.

TABLE I. - FTJGHT CREW SDIULA'KIIIS

Mercury procedures simulator (3)Centrifuge (Naval Am Development Center,J c h ~ v lU e ,Pa.)

P r i m spacecraftIrd miseaon ~imulatorMovurg-bPae airnulator fo r launch, launch abort.and entry with pssocioted accelerationenvironment

Air-lubricated free-attitude trainer

&mini mission simulator (2 )Dynamic crew procedures simula tor

Moving-base simulator for multiplr-axis pilotcontrol tanka such as on-orbit attitude control,r e t r d i r e , and entry

Primary spacecrafta d mlssim simulatorMoving-bpse simulator for launch. launch aimrt,

and =*Translation ud docking simu lato r Moving-bpse simu lator fo r formatim flying a d1

Movin!g-baBe simulator for launch, launch abort.a d ntry with associated leceleratimI eovirOllmentCentrifuge (Naval Air Development Center,JohnsviUe. Pa.)Apollo

Command module simulator (3)Lunv module simulator (2 )Command module procedures simulator

Lunar module procedures simulator

Dynamic c r e w procedures simulator

Translation and docldng simulator

Centrifuge (NASAMannedSpacecraft Cen ter)

I*my arding t r r i r r i vehicle (LLTV)

Lunar larding training vehicle sim ulatwhagley lunar ading research facility

Primary CSM mission simulatorPrimary Lu missim simulatorP u t - t a s k procedures simulator for CSMreadarmshrt-task procedures simulator for LM unardescent and ament iacludlllg rendezvous

a d ntryfWiw docking

Moving-base imulator for launch, launch Pborts,

Moving-base imulator for LM-active formation

Moving-base sim ula tor f or launch, laun:h abort,a d em with associated accelerationenvimnmentFree-flight trpining vehicle for f i na l phase of lunu

M i n t ?Wmil ia r iu t im with LLTV myatemsDynamic simulato r (tetbered)fo r final phpac d

Dynamic imulator for W i n g in tk V6 -g lunar-l W nding

surface environment

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TABLE n. SIMULATOR TRAINING SUMMARY

Simulator portio]of total trainingpercent

Number ofcrewmenSimulator Simulator timetime, hr per crewman

(a ) (average), hrTotal training

program time, hr program time,ProgramIMercuryGemini 17 991Apollo (through 29 967 69 248

Total 38 261 91 277mission 15 )

aExclusive of simula tor preb riefin gs and postbriefings.

TABLE nI. SIMULATOR USE FOR FLIGHT CREW TRAINING1imulator

Mercury procedures simulator (2 )Air-lubricated free attitude trainerMultiple-axis spin test inert ial facility

Time per program, hr~~ ~

Mercury Gemini

Part-task trainer (2 )CentrdugeGemini mission simulator (2 )Dynamic crew procedures simulatorTranslation and docking simulatorPart- task trainerRendezvous simulato rCommand module simulator (3 )Lunar module simulator (2 )Command module procedure s simulato rLunar module procedures simulatorLunar landing training vehicle andsimulat or, lunar landing res earc h

facilitya

evaluatorsManned Spacecralt Center mission

Translation and docking simula tor(Langley Research Center)Contractor mission evaluators

Massach usetts Ins titute of Technologyevaluators (2 )Full-mission engineering simulatorSimulator prebriefing and postbriefingPartial-gravity simulator, mobilepartial-gravity simulatorTotal

aTime computed on the basis of 2 hours logged for each LLTV flight.bNot included in total simulator hours.

Apollo(throughmission 15 )- -

_ -- ---

58- -55 7

64--_ _

14 58410 6721 00 8

75 394 9

146

87

85 956

17 4b703b27

29 967

Totalt i n e , h r

70 88227

17 5496

4 68298534 0

611 4 1 7

14 58410 6721 00 815 394 9

146

87

85 956

174b l 343b t 20

38 261

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C assi f icationSimulators can be grouped into se ve ra l classifications. For the purpose of thisre po rt , the following definitions will apply.Full-mission simulator. - A full-mission simulator is a device in which all crew-

rela ted phases of the mission are consolidated into one complex. Included in this cate-gorv are the Mercury procedures simulator (MPS)fig. l), the Gemini missionV Isimulator (GMS) (fig. 2), the CMS (fig. 3 ) , and the LMS fig. 4).

Figure 1. - Mercury proceduressimulator.

Figure 3 . - Command module simulator.

Figure 2. - Gemini mission simulator.

Figure 4. - Lunar module simulator.

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Part-task simulator. - A part-task simulator is a device in which only a portionof the mission o r se vera l major crew ta sk s are simulated. Fidelity in these tasks maybe gr ea te r than that achieved with the mission simu lato r.Moving-base simulator. - A moving-base simulator is a device in which the cr ewstation o r crewmemb er (or both) is subjected to some physical motion. This physicalmotion could be intended to give the pilot rea lis tic cu es of acceleration,velocity , o r

position o r could be an undesirable art if ac t of the simulation technique. One exampleof a part-task and moving-base simulator is the LLTV (fig. 5) , which provides a high-fidelity, six-degree-of-freedom duplication of the final lunar descent fr om an altitudeof approximately 500 feet to touchdown.

LA

Figure 5 . - Lunar landing trainingvehicle.

Fixed-base simulator. - A fixed-basesimulator is a device in which no physicalmotion is transmitted eith er to the crew sta-tion or to the pilot. All the dynamics of thesimulation are provided through displays onthe crew-station panels and by movement ofout-the-window scenes.D iscussion

The value of high-fidelity simulationwas well known through a ir cr af t flight ex-perience before Proj ect Mercury. The de-pendence on simulation fo r space missionsuc ces s and crew safety generally is grea terthan the dependence on simulation for air-cr af t test ing because of the nat ur es of thetwo flight pro gra ms. Space-flight cre wsare fully committed at lift-off to an entiremission , in which a broad envelope of opera-tional variab les is exercised. Aircraft tes tres ear ch usually allows for a more gradualexe rci sin g of the total flight envelopethrough a series of "buildup" flights.Whereas the airc raf t test pilot can obtainmuch of h is train ing coincident with actualflight testing, the crew s fo r space missionsmust receive all their training and be highlyproficient in all flight tas ks before the mis -sion. Prim aril y for this reason, the space-flight sim ula tors requir e the highest deg ree of fidelity of spacecr aft and missionsimulation. Aircraf t experience was used extensively in the development and operationof the first space-flight s imul ator , the MPS. Fo r the follow-on space programs,Gemini and Apollo, the maj or developmental impetus fo r the s imul ator s was derivedfr om space program experience,

Early in Project Mercury, the effect of space -flight environmental fa ct or s uponcr ew performance caused considerable concern. Consequently, training emphasized

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cr ew exposure to such conditions as high acceleration forces, zero-g conditions, heat,noise, and spacecra ft tumbling motion. The Mercury ast ronaut s received many tra in-ing exercis es in an attempt to duplicate these environmental conditions. Par tic ula rconcern wa s expressed about crew capability to manually control the spacecra ft duringthe high acceleration loads imposed during launch and entry. As a result, the Mercuryastronauts participated in four centrifuge programs at the Naval Air Development Cen-ter, Johnsville, Pennsylvania. Project Mercury flight re su lt s verified that the condi-tions of space flight had no adverse effects on crew performance for missions as longas 22 hours in duration. Consequently, crew-training programs for the Gemini andApollo missions deemphasized such considerations and concentrated to a grea t extenton the many and complex operational considerations. Project Mercury experience didpoint out the importance of simulat ing out-the-window scenes , which became even moreimportant fo r the Gemini and Apollo missions. The Mercury sim ula tor s did not haveadequate out-the-window displays, and a major effort to remedy this situation fo r theGemini si mulato rs wa s begun.The value of high-fidelity simula tors was verified throughout Project Mercury,

thus establishing for the Gemini and Apollo Programs the requirement for a f u l l -mission-simulator inventory both at the NASA Manned Spacecraft Center (MSC) and atthe NASA John F. Kennedy Space Center (KSC) launch site. Operation of simulators atKSC was deemed necessary primarily because of crew participation in spec ific checkouttests of the spacecraft and launch vehicle at the launch site.For each Mercury flight, f u l l dr es s rehearsals (referred to as simulated networksimulations) of the most significant flight phases integrating the crew, the flight plan,and the ground-support elemen ts wer e accomplished as pa rt of the preflight prep ara -tions. These simulations proved to be extremely valuable for the flight and groundcrew s and, consequently, were fur ther developed and expanded fo r the .Gemini andApollo Programs. A simulated network simulation for the Apollo lunar landing mission

involving the LMS and CMS tied in with the Mission Control Center (MCC) in Houston,Texas, required the precise coordination and synchronization of 10 arge-scale digitalcomputers . The magnitude of this phase of the simulation training program is indicatedin table IV in t e r ms of the number of days spent running full-mission simulations duringthe Gemini and Apollo Programs.

The Gemini Prog ram required mor e sophisticated simu lat ors than did Proje ctMercury, principally because of the Gemini rendezvous miss ion objective. Crew capa-bility to effect the rendezvous by using primarily out-the-window information was essen-tial to the mission and dictated that an elaborate visual system be incorporated into theGemini missi on simula tors. The development and use of an advanced sta te-of-the-artinfinity optical display system added considerably to the r ea li sm and value of s imulat ortraining for the Gemini crews.

Resu lts of the Gemini Progr am indicated quite clea rly that a well-defined and ef-fective configuration management syst em was needed to maintain the simul ators in aconfiguration that corresponded closely with the continuously changing spacecraft.Basically, this meant a quick-response systemoperated by personnel cognizant of allspacecra ft changes and capable of deciding on and contracting fo r the necessa ry modi-fications and incorporating these into the various simu lator s. A configuration controlpanel commi ttee , with anci llary working groups, was formed for the Gemini and ApolloPr og ra ms with good resul ts.

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TABLE IV . - SIMULATED NETWORK SIMULATIONS~(a) Gemini Program

~~

MissionI11IVVVI

VI11IxXX IXI1

VII/VI

~

Simulation sessions, days4771399- 9767

(b) Apollo Program

Miss ion~~

789101112131415

CMS/MCC18141011710131519

Simulation session s, daysLMS/MCC CMS/LMS/MCC

00877129137

Total181420181825273531

aThis summary includes only the miss ion simulations involving the use of mis-sion si mula tors with flight crews and does not include the many more additional daysof "math model" simulations (without the miss ion si mul ato rs) fo r flight cont roll ertraining.

Forecas ts of the simulator training requirement s for the Apollo Pr og ram showedthat a large inventory of various types (full miss ion, par t tas k, moving bas e) of stra-tegical ly located simu la tors would be needed. The Apollo Pr og ra m imposed seve reschedule and simulator-complexity constraints; a requiremen t was dictated to trai n alar ge number of three-man cre ws within a relatively short time in sophisticated and

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critical mission simulators (three command module s imul ator s and two lunar modulesi mula to rs ) with complete visual systems. The capability fo r integral operation of thesim ula tor s with each other and with the MCC n Houston al so was required. The greaternumber of command module simula tor s was needed to accommodate the g rea ter numberof command-module-only flights that comprised e ar ly pr ogr am concepts. In addition,cer tai n special-purpose part-task simulators were considered necessary to providesupplementary training, to afford overall training flexibility, and to support the crewprocedures development effort (an ntegral part of crew training). Among these special-purpose sim ula tor s wer e the command module procedures simulator (CMPS) (fig. 6),the lunar module procedures simulator (LMPS) (fig. 71, the Apollo translation and dock-ing simulator (TDS) fig. 8), and the Apollo dynamic cre w procedures simula tor (DCPS)(fig. 9). Pe rh aps the most notable special-purpose simulator was the LLTV. Becauseof the limitations of accurately simulating this critical phase of the lunar mission onground-based simula tor s, the free-flying LLTV was employed to provide an ext remelyrealistic dynamic simulation of the lunar approach and touchdown.

Concerning the simu lators themselves, se ve ra l basic decisions wer e made tosati sfy the fidelity requirements. One of these decis ions was that all spacecraft suh-sy st em s would be simulated by the com-puters; that is , no actual (hardware) space-handled in a special way was simulation ofthe Apollo onboard guidance and navigationcomputer. The simulation successfullyused was based on an "interpreter" concept

cra ft subsystems would be used. One area

Figure 6. - Command module proceduressimulator. Figure 7. - Lunar module procedur essimulator.9


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Figure 8. - Translation and dockingsimulator (Apollo).in which a general-purpose dig ital com-puter was programed to accept the sam eprogram as the flight computer and to re-spond to spacecraft sys tem s precisely asthe flight computer would. The interpre-ter superseded a barely adequate, costly,and always late functional approach fo rsimulation of the onboard computer. Figure 9. - Dynamic crew proceduressimulator (Apollo).

Another area that requi red considerab le updating and redes ign was the difficultvisual simulation of the lunar landing. As will be discussed in the section on visualdisplay simulation, the initial concept of the LMS fell sh ort of the rea lis m required.An intensive development effort was undertaken to improve the visual simula tion fo r thelunar landing and at the same time to permit timely use of the LMS fo r the ea rl y orbi taltraining. This effort was successfully concluded to provide high-fidelity landing andascent training for the first lunar landing and subsequent flights.

Although the remainder of this report is oriented toward the difficultie s and ex-perience gained through the operation of the simulato r hardware, it by no means impliesa lesser importance of the software systems. Indeed, a large percentage of the overallsimulation effort throughout all flight programs was associated with the simulation soft-ware . Fo r example, the Apollo lunar landing mission simulation for the CMS consistedof 750 000 words of memory residing in four large digital computers . Simiiarly, theLM simulation in the LMS required 600 000words in th ree digital computers. At thepeak of preparation for the first lunar landing mission, approximately 175 support con-tr ac to r personnel we re assigned to the development and control of the software s ys temsof thr ee command module simulators and two lunar module simulato rs. Another200 contractor personnel were assigned to the hardware operations and maintenance.Simulator changes required to keep pace with the changing spacecr aft and mission num-bered more than 30 per month in this same time frame. Configuration management ofthese hardware and software sys te ms is discussed in the section entitled "SimulatorConfiguration Management. 'I10


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CREW-STATION FIDELITYVarious considera tions fo r the design of the simulator crew station are describedin this section. The Mercury, Gemini, and Apollo simu lato r cre w stat ions , particu-larly those for the mission simulators , were designed and constructed with a high de-

gree of fidelity to simulate closely both spacecraft performance and int erio r appearance.

Controls and DisplaysThe controls and display panels provide the necess ary inte rface between the crewand the spacecra ft. Only through the medium of properly designed pane ls and functionalintegration with simulated sy ste ms software can the training be performed efficientlyand adequately. The controls are used to provide intelligence to the supporting com-puters in the form of analog voltage or disc rete signal inputs. To meet crew-trainingrequirements, it becomes necessa ry to simulate all known char acte rist ics of a controlswitch, such as lock and spring-loaded positions, as well as the forc es required tochange positions. Commercia l switches that look and operate like flight units usuallycan be procured at a fraction of the cost of the spacecraft hardware . Valves to controlthe flow of fluids in the actua l spacec raft usually ar e not required in an electrically sen-sitive simulator . Basically, two types of valve operations requi re simulation. Thetype of valve requi red fo r adjusting the flow of a fluid is referred to as a proportional-flow valve, and the discr ete detent-hold operation u s e s a shifting-port type of valve.The function of the proportional-flow valve is simulated by the attachment of a potentiom-eter to the control shaft for position sensing. Switch activation by cam and detents isused to simulat e shifting-port valve action. The torque for ce profile can be simulatedaccurately by placing a frictiona l, adjustable sleeve around the shaft being turned.When cre w training req uir es a close duplication of hardware operat ing character-istics, the use of flight-configured hardware has provided the best results. Such com-ponents as hand controllers must duplicate as nearly as possible force profiles, deadbands, feel of multiple detents (soft stops), handle fr ee -r et ur n characte ris tic s, mechan-ical damping, and signal generation for both discrete and proportional displacement.A wide range of comm erc ial variable controls (such as potentiometers) that havecharacteristics acceptable for use without modification are available to simu late equiv-alent flight hardware . However, speciali zed assembli es consisting of a transformerelement and a potentiometer or a rheostat (both of which may have certain cha rac teri s-tic profiles t o interface delicately with an instrument or system in a balanced relation-ship) would be difficult to obtain unless procured from the manufacturer of the flight

article. Such units usually are procured for the si mulat ors without the flight qualifica-tion test nece ssa ry fo r spacecraft hardware, thereby reducing the cost of the units andshortening procurement time. Circuit breakers also require some special considera-tion. In simulating wiring malfunctions (resulting in high curr en t flows) that t ri p re-spective circuit brea kers , it is not practical to use spacecraft hardware that trip s atrelatively high currents (e.g. , 100 amperes). Instead, a standardized low curr ent(0.01 ampere on a 26-volt dc sys tem ) circuit breaker has been used throughout thesimulators.

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StowageSince the early Mercury flights, the proper location of onboard data and looseequipment has been a necessary inflight task and, therefore, a training requirement.The need for assigning every item a specific stowage location and the procedure forhandling loose items could not be neglected. Ideally, flight-type equipment for stowagetraining is pref er re d. However, because of the disadvantages of the high co st of the

equipment and the frequent handling under one-g conditions, actual flight equipment isseldom used for simulator stowage.well i f it functionally fulfi lls the intended use.For training purposes, a nonflight item s er ve s as

LightingLighting of the cr ew compartment was not considered c rit ica l in t e rms of intensi-ties and spectrum composition. Experience has shown that light levels sufficient fo rgen era l illumination are satisfactory as long as the simulated intensity does not deviateappreciably (+20 ercent) from the actual.Because of the use of optical eyep ieces (te lescope and sextant) for navigationalsighting in the Apollo spacecraft, the floodlight intensity settings were usually low,which caused some difficulty in reading the met er s. As a consequence, all simulator

luminescent elements, as in the spacecraft.me te rs were lighted integral ly, and panel nomenclature was blacklighted with ele ctr o-

Aural CuesAudible cues are simulated and presented t o the flight crew s whenever applicableduring the training exer cise . On the Apollo missio n simulat ors (CMS and LMS), suchcues as booster thr ust , cabin decompression, and the firi ng of pyrotechnic dev ices andattitude-control jets are simulated. Noise leve ls in the communication channels as wellas ambient noise are simulated and trans mitted to the flight crews through heads ets andthrough concealed loudspeakers within the cr ew stat ions.Aural cues are generated by various means, such as a low-frequency reverbera-tion that feeds into a summing ampl ifie r. The signal output (simulated au ra l cue) of thesumming amplifie r then is fed through a voltage-controlled att enuator to a summing am -plifier and then to the loudspeaker in the mission simu lato r cockpit. The ins tru ctor athis station is included in thi s loop to enable him to control the amplitude of the soundmanually with a decibel-control potentiometer.

Markings and NomenclatureAll controls and displays , stowage areas, and other hardware must be properlymarked with the same nomenclature used in the flight spacecra ft. Nomenclature is

tions. However, in time, the crewmem bers rare ly re so rt to reading nomenclature be-cause they become s o fami li ar with the location and function of each contr ol o r displaythat reading is unnecessary.

especia lly important during the ea rl y phases of cr ew training to aid in system opera-

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Space S u i t and Cab in Env i ronment Requi rementIn the various s imula tors , suited operation throughout all phases of flight, yetwith the capability for the c rew t o t rai n unsuited for comfort, has been emphasized.The crew- station environment is maintained at ambient pre ss ur e with circulated,

temperature-regulated air. All the cont rols fo r the crew-s tation environment are man-ual, and the systems are purely simulator oriented. The crew-station-environmentparame te rs of temper atu re , humidity, and noise level are electrical ly monitored at theinstructor-oper ator station. The crew-station cabin-temperature and cabin-pressureme te rs display simula ted values in accordance with the flight profile; that i s, while onthe launch pad, the simulated cabin press ure is 14.5 psi and decreases proportionatelyduring launch, sett ling at 5.0 ps i fo r or bit conditions.A realistic simulation of suited conditions is provided. The us e of fully pr es su r-ized su its provides the proper constraints for controlling the simulator with rigid-suitconditions. Great car e always has been taken to prevent the possibility of a rapid pres -su re buildup in the simulator s u i t loops. The CMS suit loop has a motor-driven pres-

sure control that vari es the airflow at a fixed ra te of 2 psi/min, reaching 3.75 psig forhard- suit and 0.25 psig for soft-suit conditions with constant airflows under ei the r man-ua l or automatic (computer) control.

Crew CouchesExac t rep lica s of the c rew couches have been used in the simulators to provideproper body restraint, correct reac h patt erns , and accu ra te couch-manipulationcapability.

Crew-Station HardwareSimulator crew-station hardware is discussed in two categories : contractor-

furni shed equipment (CFE) and Government-furnished equipment (GFE).Contractor-furnished equipment. - The contractor in this cas e usually was theprime s pacecraft contractor. The spacecraf t contractor has sev era l avenues open toprovide hardware for the s imulators. If it is not practicable t o simulate the s pacecraftcomponents, actual spacecraft hardware is supplied. Such components as attitude con-tro lle r ass emblies, thrust and translation controller assemblies, attitude indicators,and polycarbonate cover assemblies are examples of ac tual spacecr aft items procureddirectly from the prime contractor. Much of the CFE for the simulators is manufac-

tured on the production line at the same time the spacecraft pa rts are fabricated. Sim-ulator pa rts made in thi s manner are identical to spacecra ft hardware.In many cas es , simulator p art s do not require the sophisticated design o r therigid tolerances demanded for the spacecraft. The frequent use of rejected spacecraf tpa rt s on the simula tors resu lts in cost savings. During the Apollo Progr am, the space-craft contractor maintained another manufacturing capability commonly referred to asthe "model" shop. The personnel in th is shop usually work from engineering designsketches to fabricate prototype units fo r spacecraf t qualification tests. The model shopis often used to fabr icate simulator hardware. Components made in th is shop fo r the

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simulator s do not have to meet the stringent spacecr aft specifications. Occasionally,hardware f o r the simulators is made from initial design sketches to meet the desiredleadtime nece ssary to support crew-training exercises. Various other contr actor.sour ces of simulator ha rdwar e are available, including the simul ator manufacturer.Government-furnished equipment. - Simulator part s can be made at MSC. Thissou rce of fabricated ha rdware is generally the preferred route, on a cost basis. Fed-eral stock provides a pri me sourc e fo r standard off -the-shelf it ems fo r installation andmaintenance work.A substantial amount of loose hardware is provided to the spacec raft simulator asGFE. Optical sights , s u i t hoses, backpacks, and most of the stowed it ems are pro-vided by the supplying organization at MSC. Some uni ts are included in the simulatorson a permanent basis, but other units are furnished by the suppliers as required for thetraining exercises. Other hardware, such as hand controllers and attitude indicators,is provided f o r the simu lator s by the c ontr acto rs who supply these ite ms f or the

spacecraft.

Concluding R e m a r k sA high deg ree of crew-station fidelity, part icularly in the full -miss ion si mulato rs,is necessary fo r accur ate simulation of space craf t performance and mi ssion char act er-istics. In addition to a complete simulation of the s pac ecr aft controls and displays ,such items as stowage provisions, lighting, aur al cue s, and cr ew hardware are ex-tre mely important in providing thi s required fidelity.

VISU AL DISPLAY SIMULATIONEarly in the evolution of manned space flight, the requi rement for a window in thespacecraft w a s identified. Although the window served seve ra l purposes, the most im-portant was as a backup attitude reference. Experience in airc raf t control ta sk s andduring early space flight had established that training was necessa ry to enable the crew-member to cor rel ate his out-the-window ref ere nce s with his inst rumen ts. The varioustechniques used to produce the simulated out-the-window scenes throughout Pro je ctMercury and the Gemini and Apollo Programs are described in this section.

D i s p l a y S y s t e mThe first element of a visual simulation syste m is the display system, the elementinto which an image is projected for viewing by the crew. The determining fac to rs ofthis element are the apparent distance to the image being viewed and the field of view(FOV) of the window. A genera l requir ement is that all images appear at infinity asthey do in reality. For docking, the simulation requirement was that the image appearto approach the cre wmembe r as in the actual situation. Field-of -view requ ir ementswere based on the pa rti cul ar spacecraf t window being simulated, with s ome allowancefor pilot eye o r head movement. The Mercury spacec raft requi rement was a 60" FOV,although that of the Apollo LM (the largest requirement) was a 110" FOV as measuredat the eye.

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A ll mission simulators used an infinity display system. Figure 10 s a schematicof a typical window sys tem using ref lective optics techniques. To satisfy the dockingdisplay requiremen t, the display or output tube was moved from its infinity position.Approximately 4 inches of movement of the display tube made the image appear to movefr om infinity to 8 feet from the viewer.

The infinity image syst em has thedesirable characteris tic that, as the crew-man moves his head, the i tem being viewedappears to move cor rectl y. Conversely,because of parallax, viewer head motiongrossly affect s what is seen with direc t-view, screen-type sy stem s; therefore, theviewer must hold his head relatively stillthe sou rce of the ob-

Specifications fo r the FOV could nothave been met without som e design com-promises. In the Mercury and Gemini sim-ulators, the FO V requirements (60" and88 " , respectively) fo r completely filled win-dow scenes were met as long as the pilotmaintained hi s e yes within an area definedas the exit pupil. This was not a severeconstraint in these simu lato rs, because ofCMS, ocation of the exit pupil at the window

Sphericalminueyepi-

&asplittersWindm

qt positim the relatively limited head freedom. In theallowed unlimited viewing-except as con-of the spacecraft. The design FO V w a s 90".Fi@re lo . micd strained by the physical char acte rist icsdisplay syst em.The exit pupil fo r the sid e windows was not

quite as large as the windows; however, the small d ar k bands outside the exit pupilwe re hardly noticeable and were not a ser iou s problem.In the LMS, he large FOV, oupled with a l ar ge window (17by 22 inches to ex-t reme apexes) and almost total freedom of the crewman to position his head, requ iredthe greatest compromise. The first compromise was to place the exit pupil i n the areaof the crewman's design eyepoint (i. e. , the normal eye position). Fro m this design

eyepoint, the crewm an could move approximately 4 inches in any direction with no lossof display. The second comp romis e was to point the axis of the forward-window displaydown to fe atu re the landing scene at the expense of the upward scene. In the actualspacecraft, the crewman can look up approximately 20" fro m the design eyepoint, andeven more by hunching down. In the simulator, the crewman was limited to less than10". Even with this compromise, the bottom apex of the LM window was blank asviewed fr om forward eyepoints. Because these adjustment s we re chosen very carefullyagainst anticipated u se of the window, the limi tations have had no adve rse effect ontraining.

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The infinity display s ys tems used with the mission simul ator s produced the de-sired effects as long as proper consideration was given to mission simulation require-ments. These syst ems did have some drawbacks. Relatively high cost (30 to 40 percentof total mission simulato r cost), larg e physical size s, ghost image s, extremely lowlight transmission , and backlight scatter were the most significant.application of the infinity optics technique to simulation, additional development has re-sulted in reductions of cos t and s ize . These new systems are of both refractive andreflective types. The ref rac tiv e types tend to be bright er, although with some loss inimage quality, than the reflective types.

Since the initial

C e l e s t i a l S i mu l a t i o nSimulation of the stars was predicted on both navigational and a ttitu de-refe rencerequirement s. In orbital flight, these requirements are intermixed because th e pilotmust know hi s location to use the stars as an attitude refe renc e. In simulation of thestar field, considera tion was given to the number and rel ative brightness of the starsselected and to the static and dynamic accurac ies req uired. The number of stars chosenfor all projects to date is approximately 1000, consisting of all those brighter than a+5 magnitude. The logic of selecting th is number was that the navigation stars and theconstellations used for identifying them are all brighter than the fifth magnitude. Oneothe r factor in this selection was that fifth magnitude stars are the dimmest generallyvisible t o the unaided eye from the ground.In training crews for constellation recognition, variations i n brightness, at leastt o the whole magnitude, must be simulated. The specifications required at least sevendiscrete levels (+ 5 to - l ) , with cor re ct relat ive brightness between levels desirable butnot mandatory. The specifications al so requ ired that t h i s variation be in the t ruebrightness of the star, not as produced by merely increasing light-source sizes.Static positional accuracy specificat ions were dictated by the need t o identify con-stellations. It w a s determined that the appearance of conste llat ions changes signifi-cantly with even minor positional e rr or s . Furt hermo re, if the s tar s were to be usedin a navigational task, an instrument such as a sextant would be necessar y and an accu-rac y compatible with the instrument wa s required. To reduce cost, star positional re-quirements were moderated to 1 milliradian (approximately 0.06") for all navigationstars and 0 .5 " for all other stars. Specifications als o requi red that if the constellationappeared incorrect with a 0.5" er ro r in a given star location, the positional accuracyfor that star must be tightened to t h a t of the navigation star.The specified dynamic positional accuracy also required compromise, as the re-quirement was complicated by the la rg e range of vehicle angular motions possible in a

spacecraft. Pilot-controlled rates range up to 20 deg/sec , with uncontrolled vehicleangular rates of 50 to 60 deg/sec poss ible . Although a perfectly stationary spacecraftis not possible, rates can be as low as 0.1 deg/min fo r sma ll spacecraft motions.From his stable and vibrati onless viewing position, the crewman could detect such smallmotion. The final specifications derived f or all spacecraft simulation programs re-quired as smooth a low-speed drive as could be produced and permitted lag s in positionat angul ar rat es above 15 deg/sec. Stepping motions that were detectable by the crew-member s were deemed unacceptable.

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In all simulators, the same basic technique was used t o produce the star displays.Approximately 1000 individual stars were simulated by small, highly reflective steelballs set into the sur face of a (cel est ial) sphere . When the highly polished balls wereilluminated by a point sou rce of light, they reflected the point of light, the rela tivebrightness being a function of the size and coating of the balls. Balls as small as0.1 inch and as large as 0.8 inch in diamete r have been used. Mi rr or su rfa ces concen-tr i c with the ce lest ial sp here we re used t o keep the individual balls in focus over theFOV of the window. This technique produced excellent representat ions of the starfield. The only anomaly noted on the display was a halo created by some of the coatedballs. Unfortunately, in the LMS he halo stars were all navigation stars, and the halomade recognition relatively easy. The refore, the use of coatings to simulate variousmagnitudes is not recommended.A chronic difficulty in the simulat ion of stars was in the dynamic drives of thecelestial spheres. Smooth, ste pless servomechanism dr ives that hold positional cali-brat ions have not been obtainable, and only continuous maintenance and adjustment haveproduced acceptable operation. This problem has not been limited to the servos andtorque motors but has been noted in the enti re chain, from the equations of motion,through the digital conversion equipment, to the sphere s. The problem has becomemore acute with each new program, because the range of dynamic motion produced bycrew action and under crew contro l has increased from approximately 10 deg/sec forthe Mercury spacecraft to approximately 20 deg/sec f or the LM. Although, in the past ,continual maintenance and frequent ca librat ion have tended to minimize the impact ofthe servomechanism problems, additional researc h in product improvement as well asimproved mathematical and computational techniques should be continued.In addition to the window and the unity-power telescope, the command module (CM)required simulation of a 28-power, 2" FOV sextant. This optica l ins trument directl yinterfaced with the onboard guidance computer f o r navigational procedures . The re-quirement fo r the sextant simulation w a s to produce the Apollo navigation stars (56 innumber) with co rr ec t background stars, as well as to provide earth or moon limb atvarious altitudes. Static accurac ies were the same as specified for the spacecraft,that i s , 10 arc seconds. The dynamic accuracy requirements of the sextant simulationwere not high, because all tracking and marking ta sk s were performed at very low scenedynamic rates. The sextant simulation was produced by using a single extremely accu-rately positioned navigation star in one optical leg and a selection of as many as90 sl ides of star fields, limbs, and landmarks in a second leg. A pattern genera torand filters were used to produce the background fields fo r the navigation star. Pairsof rhombic pr is ms were used to produce the motions in both legs. The two legs werecombined into a single image by a beamsplitter assembly. For most Apollo training,five sli des covering the most-used navigation sta rs were used in the second leg. Inoperation, if either leg were more than 1" off t h e intended line of sight , the sextantsimulation was turned off.Difficulties encountered in the sextant simulation were similar to those with thecelestial sphere. That is , producing smooth motion at the extremely low rate used intaking navigation marks was not always obtainable in the simula tor opera tion. Anotherdifficulty was encountered with the initial simulator design requir ement fo r updating thebackground field. Various changes in the Apollo navigation stars occurred s ince thebeginning of the program. The method of generation of background stars made it tooexpensive t o update to new background star fields, and the background-star generator

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was not used. As a result, it was not possible for the c rew member to identify the par-ticular navigation star in the sextant. Fortunately, th is turned out to be only a minorlimitation. The accur acy and stability of the actua l spa cec raf t guidance, navigation,and stabilization syst ems we re quite good, and the inflight pro ced ure s were developedto make most effective use of thi s spacecr aft capability. That is, the spacecraft crew-member was not required to identify a navigation star with the sextan t, because his on-board guidance computer programs ensured that the navigation star was always within0.5 O of predicted value.

Far BodiesFo r the purpose of th is discuss ion, far bodies are identified as the di stant moon,

ear th, sun, and planets, that is, all those bodies in the solar system that subtend anangle of approximately 0.5" fr om the viewpoint. These far bodies wer e simulated fornavigational reference and for added realism. In the Apollo simulation program, theywe re deemed important enough to be included. Fo r the sun, the specifications we re fo rposition, size, and high rel ative br igh tness in the windows when the individual windowwas pointed toward the sun. The requi rement s fo r distant ear th and moon were fo r po-sition, brightness compatible with star simulation, and apparent size change as a func-tion of range to each body. In addition, simulation of the sun te rminator on the fa r bodywas desirable. The toler ances on these requ iremen ts we re not tight. Simulation of thela rg er planets, although desi rable fo r navigational ref ere nce , was not mandatory.

As noted previously, these effects were not included in the Mercury and Geminisimula tor s. In the Gemini simulator, the sun in the window or , ra th er , the total wash-out of the scene caused by the sun in the window was simula ted by video flooding thetelevision (TV) cathode-ray tube (CRT).The sun simulation in the CMS used a rc lamps to produce a 0.5" spot of l igh t ineach of the windows. Although thes e lamps did not dupl icate the exact sun brightness,

the simulation was sufficient to wash out the stars and other dim images and served toremind the crew to maneuver t o some other attitude if they desire d to perf orm any out-the -window visual t asks .With the LMS, a high leve l of sun brightnes s was not required, and the sun wassimulated by a sm al l disk attached to the celestial spher e. By this means, the desir edeffect wa s provided at much s ma ll er computation and hardw are c ost than in the CMS.Sun lighting (or sun shafting) in the crew s tation wa s not provided in eit her the CMS orthe LMS. Although thi s omission in the simulation has r esult ed in the cr ew s' usingspace craft lights in the simula tors that are not normally requ ired in flight, the problemhas not been a major one.The simulation of the dis tan t moon and ear th in the Apollo mission si mulato rs als ofell shor t of the original spec ifica tions. In the CMS, a direct-view film sy stem with28 discrete phases of the dis tant moon was defined. Because these image s wer e toosmall to be reproduced on film, the technique h as been abandoned. For most CM simu-lation work, the sun simulato r was used t o rep re sent the distant body. In the LMS, atechnique similar to the LMS sun simulation w as used, but the dis k was somewhat lessreflective than the sola r disk. By shaping the disk, the phases of the moon o r ear thal so we re presented. The distant ea rt h and distant moon wer e moved o r manually

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adjusted to provide for movement of the distant body relative to the stars, changes insun ter minato r on the body, and changes in relative size with variation in distance.In summary, the far-body simulation requirements have been modified to be con-sis ten t with crew-training needs and state-of -the-& hardware. The requir ements fo rfar-body simulation fr om the eyepoint are summarized as follows:1. Position toleranc e- olerance should be 17" or general realism, 4" ornavigational reference.2. Size tolerance- olerance should be *O. 25".3. Brightness -Order of rela tive brightness should be sun, moon, and ea rth;all should appe ar brighter than the brightest stars.4. Phase (solar erminator)- hase is desirable, but not mandatory.5. Dynamics - ar body should have no apparent motion rel at ive to the star for

spacecraft-attitude motion simulation. Far body should move with respect to the starsfo r long-term emp hemeri s effects; however, step changes are acceptable.6. Washout and sun shaf ting- ashout and sun shafting are desirable.The CM sextant (previously described) contained a set of s lides of the dis tan tear th and of the distant moon. These slides reproduced the effect of the terminator onthe moon for all 28 days of the lunar month and on the ear th for those days the CM wouldbe in lunar orbit. This slide technique proved acceptable.

Target Vehicle for RendezvousThe target-vehicle simulation was spl it into two parts: long distance, whe r e thetarget-vehicle attitude is not important (e. g., rendezvous); and shor t distance, wherethe vehicle attitude in addition to the range becomes important (e. g. , stationkeeping).The switchover point was defined as 2 miles, or where the t arget vehicle would subtendan angle less than 0.2".The tar get vehicle f or distant rendezvous does not requ ire attitude information.The two requirements are position in the window and relative brightness with range.Fo r tracking in darkness, the spot of light was required to blink to simulate a flashingbeacon. For daylight tracking, the spot should be steady to simulate reflected sunlight.Other facto rs a r e an accura cy of 4" on position, minimum brightness equivalent to afifth-magnitude star, and a maximum range of 250 miles for the unaided eye. Maximum

brightness was not specified; however, the intent was to make the minimum-range tar-ge t as bright as possible. The spot of light should be smal l enough tha t it does notstand out against the background stars unrealistically.In all simula tor s with rendezvous capability, the window simulation was producedby electronic techniques on a CRT. This technique met the requ irements. The majorproblem wa s in positional stability of the display. In the early simulations, frequentalinements were required; however, over the span of the Gemini and Apollo Progr ams,improved electr onic circui t design has reduced this problem t o a minimum.

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In the Apollo CM, the unity-power telescope and a 28-power sextant wer e usedduring rendezvous. The simulation requ irem ents for these inst ruments are similarto that of the basic window except fo r positional accu racy . The telescope w as used topoint the sextant to a specified accuracy of 0.5" rela tive to the sextant line of sight.For compatibility with navigational requirements, the sextant accuracy was specifiedas 30 arc seconds.The telescope simulation fo r distant bodies was a later addition to the o riginalsimu lato r. Because of space cons traints and the stability problem discuss ed previously,a CRT display w a s not prac tica l. Instead, the point of light representing the tar get ve-hicle was produced by a simp le light bulb. Displacements in the FOV and si ze we re

produced by mechanical means.The CM sextant simulation al so was added to the origina l sim ula tor design. Therequirement of the distant ta rge t was sati sfied by using the sextant navigation star sim-ulation (described in the pa ragraphs on celestial simulation) to repre sent the target-vehicle position. The resultant positional accuracy was much bett er than 30 arcseconds. There wer e two limitations with thi s sextant simulation. First, with

28-power magnification, the size , shape, and attitude of the targ et vehicle we re appar-ent fo r relatively long ranges. Second, the omission behind the LM of a lunar back-ground made spotting the LM much easier in the simul ator than i n flight. Of the twolimitations , the lack of a lunar background was considered more critical. To fullysimulate the actua l situation, a moving scene beneath the CM and behind the LM wouldbe required. However, a full simulation is not mandatory. A static scene of the lunardisk, o r of the lunar limb and termi nato r, fulfills th e minimum requirements.T a rge t V e h i c l e f o r S t a t ionk eep i ng and D oc k i ng

At ranges up to 2 miles o r when the target vehicle subtends an angle g reat er than0.2", simulation of the target-vehicle attitude, shape, size, and ext ernal feature s ismandatory. The ta rget vehicle should be in proper orientation relative t o th e activespacecraft and should be illuminated in accordance with the re lat ive position of the sun,earth, moon, and any lighting sou rce on the active spacecraf t. The angular accuracyof these vehicle features should be +2" at ranges beyond 40 feet and 4 " t ran ges within40 feet. Linear paramete rs such as size should be accurat e to within +4 percent. Tol-er an ces on lighting functions are large, approximately rt30 percent. All dynamic mo-tions should be smooth, with no obvious stepping actions. Additional specifications area proper background behind the target vehicle, with no apparent bleedthrough of stars,eart h, o r moon terr ain.

Two general techniques have been used for target-vehicle simuiation: a directanalog system of clos ed- cir cui t T V and models, and an electronically generated (drawn)image. In both sys tem s, the input to the display sys tem was through a CRT in the in-finity optics systems. The elec troni c image genera tor (EIG) was used successfully inone of the Gemini mission simul ator s, in the Gemini part -tas k tr ai ne r, and in theLMPS. In the EIG system, the ta rget vehic le was drawn on the face of the CRT. Theoutline o r envelope of the target was drawn at a 60-hertz ra te; however, t he surf acewas filled i n at a 15.75-kilohertz rate. The image generation contained nine de gree sof fr ee dom and produced such phenomena as line -of -sight blanking, illumination

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shadowing, and perspective distor tion. Simple target shapes (cylinders, cones , andothers) as well as combinations of these shapes were readily simulated with simplesurface markings and details.In the more conventional system such as that used on the Apollo mission simula-to r s , a scale model of the ta rget was viewed by means of a closed-circuit TV system.

The model was mounted in a three-gimbal system to reproduce target-vehicle rotations.Generally, the innermost gimbals are in the model. The gimbal syste m should be off-set relative t o the c am er a optical axis to avoid gimbal lock at docking attitude. The TVcameras o r lens s ys tem s attached to the TV came ras als o are gimbaled in this c as e torepre sent the active-vehicle rotations. In the CMS, however, ra the r than cam era rota-tion, active-vehicle motion was introduced by displacement of the video raster on thedisplay CRT.

The relat ive range was introduced through a combination of techniques. In all TVsystems, a tr ac k and carr iag e assembly was used to move the ca me ra away fr om themodel. Ca me ra motion produced an apparent model-size reduction. For the ApolloCMS and the Gemini mission simulat ors, this size reduction was supplemented by ara ste r-s hri nk technique to provide total changes in range. Raster shr ink of 67 to 1 wassuccessfully used in the CMS. The L M S used two models, a one-eightieth and aone-twentieth scale. The total required range or apparent size change was obtained bymoving away fr om one model, then switching to the other and moving from it. An ex-ample of the type of model used is shown in figure 11. Th is model, which is from theLMS, contains the lighted docking target in one of the CM docking windows. The dock-ing probe and the CM conical portion wer e made of rubb er to protect the p robe andca me ra s in the event of inadvertent collision.

Figure 11. - Docking model of CSMin lunar docking s imulato r.

A ll TV systems are black-and-white,high-resolution ( gr ea te r than 1000 TV lines)systems. Each system repres ented the beststate of the art at the time of procurement.Since that time, however, these sy ste mshave been undergoing almost continuous re-work to improve basic performance andreliability.

Sun, earth, moon, spotlight, and otherlighting effects we re introduced in theTV/model sy st em s by high-intensity l am pssurrounding the models. Both mechanicallycontrolled lamps and switched banks oflamps we re used with reasonable success.Both th e TV/model and electronic-image techniques have produced satisfactorydisplays for stationkeeping and docking. Inth e EIG technique, complex shapes cannotbe drawn; therefore, re al ism is significantlyless. Conversely, the EIG is a muchsimpler system t o maintain and operat e.

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With the TV/model technique, ext ernal fe at ures such as windows, antennas, and dock-ing aids can be portrayed very accurately. These featu res are extremely important,especially at the close-in docking distances . The picture fidelity obtainable with theTV/model sys tem must be weighed against the complexity of the e lect romechanical de-vices required by this system. The TV/model technique has been animated so success-fully in the missi on simul ato rs that full-scale -vehicle simulation fo r docking is nolonger considered necessary, as it was in e ar ly Apollo training.

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Near-Body Scenes

.1

Scenes of the near body while in or bi ta l flight cover the a ltitude range above theea rt h o r the moon from 8 to approximately 1000 nautical miles . At these a ltitudes, thesur fac e was a ssumed to be reasonably smooth, and te rr ai n that appeared three dimen-sional (3-D) as not required. Near-body sc en es wer e necessar y for two purposes.The first was as a gener al reference fro m which the pilot could evaluate hi s attitudeand estimate his position over the te rr ai n. To meet these needs, horizon position,groundtrack movement, and gene ral continental fea tur es wer e required. The secondpurpose was fo r landmark identification and tracking. To meet these requir ement s, amore detailed t err ain scene was neces sary, and the spacecraft position relative to thescene had to be more exact. For the earth, color was a highly desirable feature in ter-rai n identification. For eit her function, sev era l additional feat ures were desira ble,such as a day -night terminato r and the highlight brightness that would approximate theout-the-window scene. Regarding the accur aci es, the following pa ra me te rs are noted :

ScenePosition of the horizon nadir, o r other

reference in the field of view, deg . . .Groundtrack movement (azimuth angle),deg . . . . . . . . . . . . . . . . . . .Size of landmasses, percent . . . . . . .Horizon curvature, percent . . . . . . . .

esolution- eneral, deg . . . . . . . ..andmark areas, deg . . . . . . . . . . Attitude function Landmark function+-j-5I00.5 .5Landmark area was defined generally as a 10-mile-radius cir cle about the spe-cific l a n dma r k . The transition from the accuracy of the landmark area to the generalscene should occur over a 100-mile radius.

a

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For the attitude function, the allowable distortion should permit identification of grosscontinental fe at ur es . Fo r the landmark function, the allowable distortion should permi teasy identification of landmass fea ture s, capes, bays, peninsulas, ri ve r basins, majo rcraters on the moon, and s o forth.In Pro jec t Mercury, the view through the periscope and through the window was

simulated in what might be termed "moving-base devices." The periscope was simu-lated in the air-lubricated free-attitude (ALFA) trainer (fig. 12) by viewing a 12-foot-diameter rear-projection screen. The nadir scene was projected on the sc reen fro ma hand-painted strip film. Considering the l a r g e distortion of the spacec raft and tr ai ne rperiscopes, the basic requirements of attitude reference and recognition of g ross con-tinental fe at ur es we re met. The requirement for window simulation in Pro jec t Mercurywas uniquely tied to yaw-angle identification while i n ret ro fi re attitude. This need arosef rom an inflight problem during one of the e a r l y Mercury flights. Before the followingflight, a yaw recognition device had been conceived and fabr icated. The simulation con-sisted of a 32-foot-diameter sc re en curved t o represent a portion of a globe 63 feet inradius. The shape wa s obtained by inflating an airtight envelope consisting of onetranslucent and one transparent Mylar sheet. A fi lmst ri p depicting moving cloud coverwas projected onto the translucent screen. The crewmember standing at the centerand approximately 2 feet fro m the dome was at the proper scaled altitude. A simple,hand-held box outlined hi s window. Using this simulation, the crew member could ob-serv e any yaw angle and readily le ar n toidentify the yaw angles of i nt er es t whilelooking toward the horizon.

Figure 12. - Air-lubricated free-attitude tra iner .

During the Gemini Prog ram, the near-earth s cenes were provided in the missionsimulators. Two different techniques, bothof which met the basic attitude-referencerequirement, were used. At the comple-tion of the Gemini Pr og ra m, th e equipmentwas updated and instal led in the CMPS andthe LMPS.The CMPS system (from the MSC-

based GMS) used a 6-foot-diameter globe,an articula ted probe, and a closed-circuitTV system. Modifications fr om the GMSincluded a new eart h globe and a new sup-port system for the globe and probe. Theprobe exit pupil was positioned over theglobe as a function of altitude and latitude.Rotation of the globe produced the p ro pe rlongitude. Spacecraft rotations we re intro-duced by motion of optical elements withinthe probe. The updated globe was built toaccurate sphericity requirements and artis-tically rendered with high detail in selectedlandmark areas. The sphericity plus astable probe mount provided good horizon

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positional accuracy and curvature up to an altit ude of 4000 nautical miles . An artifactin the globe and probe sys tem was the need t o nutate the probe in azimu th as a functionof latitude, an effec t that complicated the probe d riv e equations.The LMPS used a flying spot scanner sys tem (from the KSC-based GMS) with fi lm

as the data -sto rage element. Latitude and longitude were produced by film translation,and spacecraf t motions were introduced by electr onic manipulation of the flying spotscann er output. The scanner patt ern w a s spi ra l and centered at the nadir to producea clean, sharp horizon. To provide equal element spacing as measured f rom the view-point in the spac ecra ft, the sp ira l scan spacing was nonlinear. The design resolutionwa s equivalent t o the best high-resolution TV raster systems. In the GMS, two parallelsyst em s were used to produce a semblance of col or. However, in the modification toadapt th is system to the LMPS, thi s provision was eliminated. The modification alsoallowed simulation of an altitude range f ro m 2500 feet to 100 m i l e s by using zoom optics,zoom electronics, and a series of var iou s film scale s. The rem ain der of the LMPSmodification was direct ed toward improving overal l quality of the visua l disp lay by usinga larger f i lm forma t and improved electronics. Even with thes e improvements, theLMPS system has not proved as stable and as flexible as the more conventional mechan-ical and optical systems.

The LMS used a syst em si mi la r in cer tain as pect s to both of the previously de-scri bed systems. In the LMS, an articulated probe and closed-cir cuit TV viewed afi lm st ri p of the near body. The LMS fil ms and the LMPS fi lms we re actually the sameimage material printed at slightly different sca les. The major difference in the sys -te ms was that, in the LMS, four windows were active and four probes were requ ired,whereas, in the LMPS, the display wa s provided t o one for ward window and to the ove r-head window. Fur the rmo re, the LMS scene requi remen ts included landmark identifica-tion and tracking in addition to the vehicle attitude-reference task . In the LMS, acommon filmstrip was projected through zoom optics onto four s cr eens . The altituderange fro m near zer o to orbit was simulated by a combination of opt ical zoom and fi lmchanges. The probes were mounted at a fixed distance above the s cr ee ns and arti cu-lated with the ability to scan to any point on the scr ee n. At the higher altitudes, sph er -ical distortion was introduced by moving additional optical el eme nts into the projec tionchain. The horizon was produced at the higher altit udes by illuminating an area smallerthan full-screen size. At alti tudes below 100 000 feet, a servomechanism-operatedmi r ro r system surrounding the sc reen produced the horizon. The lunar scen e fi lmsused were, in chronological order, artistic rendition of the front s ide coupled withartistic imagination fo r rendition of the back, artistic rendition updated with LunarOrbiter data, Lunar Orbiter photographs photomosaicked into a filmstrip, and LunarOrbi te r photographs photomosaicked with lunar photographs fr om Apollo flights. Finalconfiguration of the film consisted of Lunar Orbit er s tr ip s fo r the high-altitude, full-orb it scene s and Lunar O rbit er s tr ip s mosaicked with Apollo photographs f or the low-alti tude scenes in the vicinity of the lunar landing site.

Several problems we re associated with the sys tem that ma terially reduced theuse fulnes s of the LMS. A problem result ed from the sp litting of the film image to fou rsc re en s. The illumination on each vidicon was one-third to one-half of the desi red min-imum. Another difficul ty involved the zoom opt ics. The design requ ired no movementof the optical axis with zoom. The dynamic wander of the original hardware was suchthat the usable zoom ratio was rest rict ed to 3 fro m the design ra tio of 10. The wanderal so made registra tion difficult to maintain when switching from one to a second film

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scale. A problem was also associated with the film graphics. It was not prac tical toobtain continuous fi lms tri ps with the information content desired. The Lunar Orbiterstrips with Apollo photographs were available only for small areas of the moon. Also,fra mel et lines in the Lunar Orbiter s tr ips were reproduced j ust as vividly as the cratersand rilles. Furthe rmore , in the Lunar Orbit er films, it was not possible to simulatethe effect of the sun elevation angle. Hand-painted f ilms could have alleviated cer tainof these problems; however, the cost in manpower and time was prohibitive.

The overall effect of these problems w a s a degradation in the LMS fidelity, whichnecessita ted use of other techniques and facilities to complement the LMS. For in-stance, for the landmark identification and tracking tas k, detailed briefings and lunarmap revi ews wer e held with the crew. The accuracy of the spacecraft navigation sys-tem also helped, because the crew rare ly had to search for a landmark during flighty.The desired target was usually where mission planning said it should be.

The CMS near-body simulation , fr om the outset, was mainly for the purpose oflandmark tracking. The syst em consisted of direc t view of color film projections ofnear-body imagery. Only such a direct-view system could provide the resolution spec-ified. This syste m, known as the mission effec ts projector (MEP), w a s used in thefour cabin windows and in the unity-power telescope. In the MEP, the image was pro-jected through a series of optical devices onto a spherical screen. This scre en servedas the input of the infinity image system. The optics were driven by servomechanismsto simulate spacecraft rotations, limited altitude effects, day-night term ina tor , and soforth. Na d i r position was introduced by driving the fil m, and large altitude changeswer e provided by dissolving to fi lms with different scale factors. The MEP representedan extreme ly complex elec tronic , mechanical, and optical device. For instance, eachMEP had over 30 computer-controlled servomotors and over 40 other computer-generated signals. Some compromises wer e required in the final operating configura-tion: the maximum detai l sceni c area, as measured from the spacecr aft, was limitedto a 90" cone centered about the nadir; the minimum altitude simulated was 100 nauticalmiles: and t h e film centerline was placed approximately along the groundtrack. Thefirst of these was the greate st limitation, but it was partially alleviated by using a cloudcover tha t moved with the ea rth to f i l l the sce ne to the horizon.

The fil ms covered relatively narrow s tri ps centered on the groundtrack; and, be-cause of r et rogression of the orbi t (weste rly movement of the ascending node), 17 con-tinuous stri ps were required to cover the Apollo earth-orbit mission. To resolve theround-to-flat mapping problem and the orbit-to-orbit scene repeatability, an extremelyaccura te globe, used as the dat a source, was photographed by a special modified-slitcamera. The preparation of the filmst rip from this globe is described in reference 1.The altitude variation fo r near-body orbit s was specified as 100 o 1000 nautical

mile s and was obtained by a combination of three film scales with a 2.15:l zoom capa-bility. Over 250 eet of film stored in four film cassett es were required fo r each MEPsystem.The CM S MEP system, like the LMS, did not meet all the initial design objectives.The complexity of the mechanical-optical device caused problems, and the number ofMEP system s required to f i l l the windows of the three mission sim ula tor s made im -provements extremely costly. The first of the difficulties concerned the illumination

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system, in t e rms of both the leve l of light and the uniformity of illumination ac ross thefu ll FOV. The relatively low leve l of illumination manifested itself in many other prob-lem areas, The syst em design required an extremely high luminous f l u on the su rfaceof the f i lm . Even with thi s f l ux , the final brightness as meas ured fro m the window wasapproximately one-fourth of the original ly specif ied 4 . 5 ft- L and approximatelyone -eighth of the des ired illumination to s imulat e a highlight brightness representativeof a tr ue out-the-window scene. With the high light flux , the film would be destroyedalmos t immediately by infr ared and ultraviolet radiation. To limit this damage, heatabsorption and rejection filters we re used between the arc lam p and the film. Addi-tional f i lm cooling was achieved by a ve ry high, continuous airflow. This high airfl ow,in turn, caused flu tte r and eventual des tructi on of the film. In the final mode of ope ra-tion, a compromise was selected that accepted som e film deterioration caused by theheating and flu tter and a scene illumination that r equi red a relatively dar k cre w station.In the simulator , the cr ew made use of the panel lighting and the s ma ll floodlights,whereas i n the spacecra ft the sun illumination through th e windows was generallysufficient.

Fo r other deficiencies in the sy stem, prac tical solutions were n ever found, and,as a result, procedural workarounds were developed to alleviat e the impac t on training.One such deficiency involved the resolution of t he imagery that was pre sent ed t o thecrewmember in the cre w station. Although the inherent resolution of the direct-viewfilm system satisfied the specified requi remen ts, copies of the film could not be prol-duced consistently with the sa me high image quality. The pro ces sing of 100-foot lengthsof f ilm does not produce r es ul ts of the qual ity that c an be achieved by copying individualphotographs. A 30-percent degrada tion was not uncommon. Another problem involvedthe impossibility of maintaining the prop er film positioning accuracy ove r the 80 feet offilm in each of the cass ettes . Minor film stretching and warpage introduced e r r o r s ofsignificant magnitude. In addition, the dr ive s yst em of the film did not have nearly astight a tolerance as required. Finally, a problem existed with the colo rs in the film.While the fi lms themselves contain v ery saturated colo rs, the projector tended to muteand mutilate the colors as displayed in windows - or instance, t h e blue oceansappeared almost black in the windows.

In both the CMS and the LMS, a defect existed in that the star simulation wouldbleed through the near-body sce nes . The equipment, as originally designed, containedhardware to occult the stars in a cir cul ar pattern of the sam e diame ter as the nea r bodywould subtend. This hardw are proved difficult t o maintain, was nonlinear with positionand shape of the body in the FOV, and , as a resu lt, did not successfully occult all thestars. Finally, a shor tcoming of all near-body simul ation s was the low scene bright-ness . This condition requir ed the crewmen to da rk adapt to discern the necessary de-tail out the window.Significant funds have been spen t in developing the near-body simulations fo r thevario us space-flight vehicles. It is recommended that, in the future, the requirementsfo r out-the-window sc en es be very carefully analyzed and that extreme care be takento limit the simulator requi reme nts to those obtainable with reasonably simple hard -ware . For example, in the CMS, only the unity-power telescope in rea lity requir edstrin gent tracking accuraci es. The other four active windows could have been animatedwith much less sophisticated hardware . Experience has als o shown that window-scenegenerators should be more integrated as in the LMS and the CMPS than the system-per-window s u ch as in the CMS. Such design considerations in future applications shouldboth reduce initial co st s and resu lt in lower maintenance ef fort and costs .

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Landing ScenesLanding scene s imulation was required only in the L M S or training in the lunarapproach and landing phases. The altitude range des ired was rom a high-gate orbreakout alti tude down to and including touchdown. The scene was required to providethe c rew member with attitude and altitude information plus pertinent surface-feature

details. In all other pr og rams (for instance, during ea rt h entry and landing), the re wasa minimum of c rew activity re lative to the window disp lays; the ref ore , little or no ex-ternal scene simulation was supplied. In the LMS, the requirement w a s defined as acontinuous scene fro m an altitude of approximately 15 000 feet to touchdown. The or ig -inal requirement at the time of simulator procurement was fo r a generalized lunar sur-face containing represen ta tive features of the moon. Thi s requi rement was expandedsubsequently to include modeling of the actual landing sites. Three-dimensional sur-face deta il was required when surface irregularities became visible to the crewman.For the lunar landing simulation, the attitude at which surface features became impor-tant w as defined as 2000 feet. An additional requirement was the cast ing of shadowssuch as would be caused by loca l surface ir regul arit ies when illuminated with collimatedsunlight.The visual acuity requirement was to identify objects subtending approximately0 . 5 " at the pilot's eye. This a

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Apollo Experience Report Simulation of Manned Space Flight for Crew Training

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