apollo experience report crew station integration. volume v lighting considerations

8/8/2019 Apollo Experience Report Crew Station Integration. Volume V Lighting Considerations

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N A S A TN 0 -729* >

APOLLO EXPERIENCE REPORT -CREW STATION INTEGRATIONVolume V = Lighting Considerationsby Charles D. WheelwrightManned Sacecrafl CenterHouston, Texm 77058N A T I O N A L A E R O N A U T I C S A N D S PA CE A D M I N I S TR A T I O N W A S H I N G T O N , D. C. JUNE 1 9


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1 . Re r t N o. 2. Government Accession No.NAgA TN D-7290APOLLO EXPERIENCE REPOR T4. Title and Subti tleCREW STATION INTEGRATIONVOLUME V - LIGHTING CONSIDERATIONS

7. Author(s)Ch ar l es D . Wheelwright , MSC~

9. Perform ing Organiz ation Name and AddressManned Space craf t CenterHouston, Texas 77058

3. Recipient's Catalog No.

5. Report Date6. Performing Organization Code

June 1973

8. Performing Organization Report No.MSC S-360

10. Work Uni t No.924-23-D7-82-72

11. Contract or Grant No

2. Sponsoring Agency Name and Address13 . Type of Repor t and Period Covered

Technical NoteNat ional Aerona ut i cs and Sp ace Admini s t ra tionWashington, D. C. 20546

7. Key Words (Suggested by Author(s))' E l ec t ro l um i nescence* F l u o r e s c e n c e* Incandescence '* X enon L am ps* Da rk A daptat ion

14 . Sponsoring Agency Code

18. Distribution Statement

I5. Supplementary NotesThe MSC D ire cto r waived the us e of the Internat ional S ys tem s of U ni ts (SI) fo r this Apol loExpe r i ence Repor t because , i n h i s judgment , t he use of SI Uni ts would im pai r t he usefu lne ssof t he r ep o r t o r r e s u l t i n exces s i ve cos t .

6. AbstractT he l igh ti ng r equ i r e m en t s fo r t he A po ll o spacec ra f t a r e p re sen t ed . T he na t u ra l li gh ti ng f ac t o r sare d i scus se d i n t e rm s of m a j o r cons t r a i n t s . A gene ra l descr ip t ion of the ex tern a l and in t e rna ll igh t ing sys t em s for the comma nd and lunar modules is pre sen t ed w i th a di scus s ion of the pr im aryapproach and des ign cr i t e r i a fo llowed dur ing development. Some of th e mo re d i f fi cu lt p ro blem sencountered dur ing the implem enta t ion of a new l ight ing system are rev iewed.

19. Security Classif. (of this report) 20 . Security C lassif. (of this page) 21 . No. of PagesNon e None 24 22 . Price$3.00

* For sale b y the Nat i ona l Techn i ca l In fo rmat ion Serv ice , Spr i ng f i e l d , V i rg i n i a 22151


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FOREWORDThis technical note documents experience gained in the area of spacecraft crewstation design and operations during the Apollo Program. Emphasis is given to thetime period ranging from early 1964 up to, and including, the Apollo lunar landingmission of July 1969. This time period cove rs three important phases of the ApolloProgram: the design phase, hardware construction, and mission operations.Th is technical note con sis ts of five volumes. Volume I, "Crew Station Designand Development, 9 t g i v e s an overview of the total crew s tation integration task. Vol-umes 11, 111, IV, and V a r e specialized volumes, each of which is devoted to a basicfunctional area within the Apollo crew station. The subject of each volume is indicated

by its title, as follows.Volume 11, "Crew Station Displays and Controls"Volume 111, "Spacecraft Hand Controller Development"Volume N , 'Stowage and the Support Team Concept''Volume V , "Lighting Considerations"

Louis D. AllenManned Spacecraft Center

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APOLLO EXPERIENCE REPORTCREW STATION INTEGRATION

VOLUME V - L I G H T I N G C O N S I D E R A T IO N SBy C h a r l e s D . W h e e l w r i g h tM a n n e d S p ac e cr af t C e n t e r

S U M M A R Y

The interna l and external lighting fo r the Apollo spacecra ft required a high levelof quality and reliability to ensure excellent readability of the displays and control sduring the mission. The lighting hardware was designed to withstand the high vibrationand acceleration during launch and to function for the duration of the mission in theextreme environments of space.The lighting system requirements were based on many studies and lighting evalu-ations. The Apollo lighting system consisted of two major categories : the inte rnallighting system, which included all integral displays and controls lighting and generalin te ri or floodlighting; and the exte rnal spacecraft lighting, which included illumination

ai ds f o r rendezvous and docking, extravehicular activity, and work stations. Integra1lighting r e fe r s to the methods of lighting display and control panels and instrumentsfr om within.

The general illumination in the command module was provided by fluorescentlamps and in the lunar module by incandescent lamps. Fo r the integral displays light-ing in both vehicles, electroluminescent lamps were used, which was the first applica-tion of backlighting of the complete ins trument p a n e l in spacecraft and the first rxtensiveus e of electroluminescent lighting by the aerospace industry. The external spacecra ftlighting used on both the command and lunar modules consisted of orientat ion lights,extravehicular activity lights, and a docking floodlight located on the command module.These lights wer e incandescent. To aid in visual tracking, a xenon flashing light wasprovided on both vehicles. The docking tar get on the lunar module and the extravehic-ular activity handrails on the command module were illuminated by radioluminescentdisks . Fo r the docking target on the command module, a combination of electrolumi-nescent and incandescent lamps was used.

Many lighting studies and mockup evaluations we re conducted to as su re that t helighting conditions we re appropriate fo r adequate monitoring of subsystem performance.The type of lighting fix ture, the locations, and the light intensity to be used in the twoApollo vehic les wer e estab lished by the lighting studies and mockup evaluations.*


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The major problem encountered was with the use of electroluminescent lamps,a relatively new lighting source. Because these lamps were installed directly into thepanels, sizing and shaping of the lamps to specific areas required very c lose manu-facturing tolerances to achieve optimum light output fr om the displays.The Apollo lighting system was highly successful as a result of continuous re -

views and mockup evaluations of new state-of- the-art design and the capabili ty for con-trolling the design interface between lamp manufacturers and hardware contracto rs.Also of importance was the early utilization of lighting laboratory calibra tions andstandardization of the light spectrum distribution within both the command and lunarmodules.

I N T R O D U C T I O NSuccessful execution of the Apollo lunar landing mission requi red that the crewproperly perform a variety of visual tasks. Many of these t asks depended heavily on

the performance of spacecra ft illumination subsystems or devices. The lighting hard-wa re had to function adequately for the duration of the mission in ext reme environmentranging fr om sea- level conditions to the vacuum of deep space.Spacecraft lighting presented problems unique in the illumination indus try. Fo rexample, each pound of equipment required approximately 500 pounds of fuel f o r around trip to the moon. Mate rial s had to be qualified by numerous te st s to determinethat they could withstand the severity of the space-flight environment, parti cularly ac-celeration and vibration levels that obviously exceed those i n most presen t lightingenvironments. Reliability and redundancy were of major importance when the level ofartificia l light had to range fr om high candlepower (for a la rm lights, which had to beseen with periphera l vision during periods of high acceleration) to total darkness dur-

ing space observations (requiring the elimination of sunshafting inside the vehicle).These crit eria had a pr im ar y effect on the choice of lighting sys te ms fo r the Apollocommand and lunar modules.

APOLLO LI G H T l N G REQUI REMENTSThe lighting requirements for the command module (CM) and lunar module ( L M )were established to maintain the eye adaptation necessary fo r crew tas ks . Eye adaptation re fe rs to the visual adaptation of the eye when exposed to different ambient lightlevels. The lighting system pa ra me te rs we re based on the following cr it er ia .1. Task analysis to define cri tic al visual ta sks2 . Vehicle design and performance3 . Internal/external illumination relationships

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4 . Considerations of operato r fatigue5. Optimum system performance and maximum reliabilityAfter studies were conducted at the NASA Manned Spacecraft Cen ter and by pe r-

sonnel of two cont rac tors , and afte r lighting evaluations had occur red in mockups atthe con tractor facilities , the following it em s and lighting requirement s were recom-mended and inco rporated into the lighting system of the Apollo command and lunarmodules.

*

1. White lighting, for us e in ca se s when cone vision and color perception w er eneeded and when the lighting power requirements had to be low2 . Integral lighting, fo r lighting uniformity and minimal g la re3 . White electroluminescent (EL) lighting (at 0.5 f 0 .2 ft-L), for nomenclatureand meter s4 . Green EL lighting (at 15 t 3 ft-L), for improved contrast in numeric read-out displays5. Incandescent mas te r al ar m warning lights (at 150 f 50 t-L)

a. Caution and warning (50 t-L)b. Component caution (15 t-L)

6. Continuous dimming, to provide dark adaptation during guidance and naviga-tion operations and to maintain legible indicators for continuous monitoring7. Floodlighting, as orienta tion and backup to integral lighting

The ma st er al ar m system and the caution and warning syst em used incandescentlighting because of the higher light leve ls requi red.The test evaluations and resul ts determined the type of lighting to be used for theApollo spacecraft. The inte rnal and external lighting requirement s, the component

type, the lighting used , the luminance intensity of the lighting, and the color specifica-tion a r e presented in table I.

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TABLE I. - ILLUMINATION COLOR AND BRIGHTNESS

Primary lightingmethodomponent Color underincident illuminationTransillumination Brightness at Brightnessrated capacity, adjustment,

it-L it-Lcolor

Integral EL pr eferred White, lu nar white,CIEa coordinatesX = 0 . 3 3 0 + 0.030Y = 0 .330 t 0 .030

2ontinuous fro m 0(0.01 minimum)Pushbuttons -panel nomenclaturedisplays and con-t rols ( D K )

0.5 f 0.2

100

50 t 10

50 f 10

Control panelsPushbuttonsGray 36231

Background: black 37038,Characters: white 37875gray 36076

Background: translucentCharacters: black 37038Background: translucentCharacters: black 37038Background: translucentCharacters: black 37038

gray 36076, white 37875

gray 36076, white 37875

gray 36076, white 37875

Master alarm Integral (incandescent) Aviation red ( perMIL-CZ5050A) Fixed

Warning annunciators Integral (incandescent) Aviation red (perMIL-CZ5050A)

Aviation yellow (perMIL-CZ5050A)

:ontinuous from1.5 f 0.5

:ontinuous from1.5 f 0.5Caution annunciators Integral (incandescent)

Aviation yellow ( pe rMIL-CZ5050A) 1 5 t 3omponent caution Integral (incandescent) :ontinuous from0.05 t 0.02

:ontinuous from0.02 t 0.01

Background: translucentgray/whiteBackground: translucentLegend: black 37038Background: translucentLegend: black 37038

gray 36076, white 37875

gray 36076, white 37875

Status annunciator Integral (EL o rincandescent) Aviation white~

l o + 3

10:;dvisory annunciator Aviation white o rgreen :ontinuous from0 .02 t 0 . 0 1Integral (EL o rincandescent)Integral (EL )lags, two-position White,X = 0.330 f 0 .030 0.5 f 0.2 :ontinuous from 0 Energized: alternateblack 37038 andwhite 37875 stripingDeenereized: erav 36231

Flags, three-position Integral (EL) White,X = 0 .330 + 0 .030 0.5 f 0.2 :ontinuous from 0 Malfunction: re d velva-gloo r equivalent; if labeled;letters: black 37038 ongray 36231Energized: alternateblack 37038 andwhite 37875 stripingDeenergized: gray 36231

Meter (color coding oflighted D K )Pointers Silhouette or ELfloodlight (integral)

Integral (EL )

Black 37038, yellow 33538,red (rocket)Black 37038, white 37875

- -0 . 5 t 0.2

0.05 f 0.2

8 minimum

20 f 5

- -Jontinuous from 0

:ontinurns f rom 0

:ontinuous from 0

:ontinurns from 0

Indexes White,X = Y = 0 . 330 +0 .030White,X = Y = 0 .330 f0 .030Green, wavelength:

5000 to 5300 AAviation green

Characters Integral (EL) Black 37038, white 37875

Time-shared labelsand multipliersRange markings

Integral (EL)

Integral (incandescent)~-inplay malfunctionindicator lightsCircuit breaker

Integral (incandescent) Aviation red ( perMIL-CZ5050A) 2ot 5 :ontinuous from0.05 + 0.02 Translucent gray/white

Background: black 37038Characters: white 37875Flood White,X = Y = 0 .300 t0 . 0 3

0.5 f 0.2~

:ontinurns from 0

7 to 18 Background: gray 36076lphanumeric read-outs Integral EL Green , dominantwavelength :4900 to 5300 A:ontinurns from0.02 0.01

PInternational Commission on Illumination.

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TABLE I. - ILLUMINATION COLOR A N D BRIGHTNESS - Concluded

Component Color underincident illuminationTransillumination Brightness at Brightnessrated capacity, adjustment,rimary lighting ft-L ft-Lethod color

Internal

CM utility light

LM utility light

0. 1 f 0.02I At launch.Self-luminous devices, Green, wavelength: Initial,switchtip

Por tab le, incandescent White, unfiltered bo. 2 minimum Fixed - -Portab le, incandescent White, unfiltered b5.2 at 3 ft 25 to 11, 15 to I - -

at 3 ft

Fixed

FixedFixedFixed

Pale green

------

I ontinuous fro m 0 Illuminance A o r betterI 5 . 0 + Iow-level floodlighting Incandesce nt or IWhite, unfilteredI fluorescent

Aviation re dAviation gree nAviation whiteAviation yellowAviation yellow/whiteWhite, unfilteredWhite

~~ _ --- Iigh-level floodlighting Incandescen t or IWhite, unfilteredI fluorescent

15 minimum'. 15 minimum'. 25 minimum' 25 minimum'. 25 minimumd1500 minimumDependent ondetectionrange re-quiremente

Fixed

Orientation lights(running):Port (left)Starboard (right)AftBottomForward

- _Docking floodlightTracking light

Green,5150 T;i: 6I

~ ~~~

Incandescent

0.6 f 0. 1at launch

IncandescentIncandescentIncandescentIncandescentcncandescent

CM docking targe t

Xenon

Background EL Gree n High, 17 to22Cr os s incandescen t Red LOW, I to 10

~

I I

Two evels - -

Extravehicular activityGreen

lighting b 0 . 6 f 0 .2at 3 ft0.2 f 0 . 1I R LLM docking targe t

FixedFixedFixed

eThe CM light is 160 beam c-sec, giving a detection range of 50 nautical miles; the LM light is 1000 beam c-sec, giving adetection rang e of 120 nautical miles.

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APOLLO LIG H T l N G S Y S T E MThe Apollo lighting system consisted of int ernal cr ew station lighting and externalspa cec raf t lighting. The internal lighting syste m provided ambient light for activityin the couch and lower equipment bay; f or reading panel nomenclature, indicators , andswitch positions; and fo r tunnel activities. The spacecraft external lighting sys tem

furni shed artifici al light fo r extravehicu lar activity (EVA), rendezvous, docking, opti-ca l tracking, and recovery aid after splashdown.

Internal LightingTwo types of in terna l lights were used on the CM and LM, and the s am e sy ste mswe re used on both space craft , except where differences are noted in the following dis-cussion. The pr ima ry lighting of the display and control ( D W ) a rea of the Apollospacecraft w a s by transillumination, which is a type of panel lighting in which the light

is emitted from behind the panel (figs. 1 and 2 ) . Apollo repre sen ts the fi rs t applicationin spacec raf t of back illumination of thecomplete instrument panel (nomenclature,instruments, dials, etc. ). The primar ycol ors used on both spacecraft w ere whitefo r nomenclature and instru ments , greenfor alphanumerics, red for warning, andyellow for caution. Three methods oflighting were used within the spacec raf t :self-illumination; incident-direct flood-lighting, including wedge lighting ofmeter faces; and transillumination. Ingener al, these three methods wer e usedto provide lighting of indic ato rs, con tro ls,read-outs, displays, system switches,nomenclature, annunciator (signal devicecommanding attention usually by visualand auditory means ) pushbuttons, andsigna l lights. The white nomenclature,instru ment, and control lights had amaximum brightness of 0 . 3 to 0.7 ft-L;the gr een alphanumeric read-outs had amaximum brightness of 8 to 18 ft-L. Thecrew could dim the integral lighting sys-tem fro m the maximum to near zer o(0 .01 ft-L).

The secondary lighting system onthe Apollo spacecraft consis ted of flood-lighting system used incandescent lightand was primarily a redundant, secondarylighting sys tem in ca se of panel and in-strume nt integral lighting failu re. The

. lights (figs. 2 , 3 , and 4). The LM flood-

6

Figure 1. - Command module controland display panel lighting.


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Figure 2. - Lunar module controland display pane l lighting andfloodlight .

Figure 4. - Command modulefloodlighting system.

Figure 3 . - Command moduledisplay panel, illuminatedby floodlights.illumination intensity of the display panelswas relatively low but was not less than0 . 2 ft-c on the main display panel. TheCM floodlighting syst em was pa rt of thetotal pr ima ry lighting system; therefo re,the floodlight intensity for the CM wasconsiderably higher than that for t he LM,having a nominal brightness of 30 ft-L.

Before the adoption of the rela tivelynew EL lighting system fo r the Apollospacec raft, the relative mer it s of the sys-tem had to be determined. Color match-ing and balance had to be achieved, andminimal brightness levels had to bemaintained throughout a mission.

Integr al lighting. - Transillumination on D&C panels and instrumen t lighting inthe Apollo veh icle s was provided by EL lighting. Of chief concern we re the behavior ofthe EL l amps under the effect of burning time , temp era ture , frequency, and voltage;the effect of manufacturing proc es s reproducibility on such factors as lifetime, bright-ne ss, and colo r; and an accurate sy stem with which to meas ure the effects.

The burn history of a green EL lamp used in the LM-3 vehicle is shown in fig-u r e 5. Specifying brightness, color, and minimum life is not enough; a burn-in periodis required. This initial period depends on the part icu lar manufacturer of the EL lampand the range of lifetime curve s that can be expected as a re su lt of the degree of re-producibility of the EL-lamp manufacturing proces ses . A burn-in time is chosen at thepoint on the cu rve at which the slope begins t o level off s o that brightness stability canbe expected during the mission . Because of brightness decrea se, calculations are madeS O that, by th e end of the mis sion, the brightness will not fall below the minimum0.2 ft-L nec essa ry for visual acquisition.

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After 50-hr burn (minimum)Lamps installed in instrument

Instrument installed ih spacecraftLM-3 crew compartment fit and function

Present launch specification

.-m 5 - Acceptable, launc h

Maximum and minimum specificationsx - brightness measurementI I I I I I I I I0 2M1 400 600 800 1MM 1200 1400 16OJ 18M)Operation time, h r

Figure 5 . - Brightness agingcha rac ter ist ics of greenEL lamps.

If the brightness values of the lamppermit, the rated voltage on the lamp maybe set at some value lower than that rec-ommended by the manufacturer. Thevoltage ma yth en be increased as the lampefficiency decreases, thereby retainingoriginal brightness values and effectivelyincreasing the life of the lamp. Apollo ELlighting was operated a t 75 volts r m s and400 hertz, but the voltage could be in-creased to 115 volts.

The effect of t emperature on ELlifetime was of concern in the Apollo Pro-gra m because possible temperatur e l imitsfor the EL lightingrange fro m 45" to145" F . The dete rioration of EL lightingat 145" F is very rapid i f the lamps areon. However, by supplementing the EL lighting with floodlighting during extreme en-vironmental conditions, the EL lifetime can be extended throughout the mission.The requir ed brightness on the EL panels and inst rumen ts generally was 0 . 5 f0 .2 ft-L at 7 5 volts rm s and 4 0 0 hertz after 5 0 to 100 hour s of burn-in time. Thesevalues were not difficult to attain, although difference s between brightness measuringme te rs varied considerably. The brightness variations of prototype panels had rangedi3 0 percent of nominal brightness over the panel. Measurements in the LM-3 toLM-5 flight vehicles reveri fied this range. The Apollo cr ews commented that the ELintegral lighting w a s very good.Dimming all the EL lighting with one control was neither possible nor desi rabl eon the CM o r LM. The gre en EL alphanumer ics, which requ ired much higher bright-ne ss fo r proper contrast with floodlight illumination, had to be dimmed separately.The dimming characte rist ics of the alphanumerics control shaft ar e shown in figure 6;this shaft also controlled the brightness of the annunciators through a special circuit.Difficulty w a s experienced in providing pro pe r lighting f or toggle switches to beused under dark adaptation, fo r which most of the E L integra l lighting was intended.On the CM, it was determined that edge spillover fr om the EL panels was sufficientfo r th is purpose. On the LM, radioluminescent (RL) disks wer e mounted in the toggleswitches. The radioactive sour ce f o r the radioluminescence w a s promethium-1 4 7 .The intensity of the RL lighting was 0.09 to 0.1 ft-L at launch. The half life of the RLdisks was 18 months. The promethium was encapsulated inside the acryl ic toggle

switches. When the acr yli c w a s changed to Kel-F, a fire-resist ant material, a reac-tion between the Kel-F and the radioactive so ur ce developed which resu lted in radio-active leakage. This condition was cor re ct ed by encapsulating the radioactive sour cein gl as s capsules and sealing the caps ules into the Kel-F toggles.The driving frequency on the EL lamp affects color and brightness. The bright-ne ss effect caused by frequency changes has been published in the manufacturers' bro-chures and has been proved to be li near . A t 7 5 volts, a green EL lamp w i l l change

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Annunciator voltage1111113 4 5 x l d 1 hertz in driving frequency. Frequency

Main display console alphanumeric voltage

approximately 0.01 ft-L for a variation ofwas controlled to f 7 hert z in the Apollospacecraft. Brightness could vary ini-20 40 60 80 100 120 14030

20 Linear -Incandescent annunc iators tially *O. 05 ft-L, assuming a fil ter trans- .:< mission of 70 percent. The effect offrequency var iation upon the color of the

facturer s' brochures. With a k7-hertzEL lighting w a s expected to change byfO.001, and the X coordinate was expectedto change by much le ss . In this case , Xand Y coordinates ref er to the interactionof a point on the International Commissionon Illumination (CIE) chromaticity chartthat specifies the dominant wavelength ofthe light (no units used).

?'IlorFreenquipmentL alphanumerics,ay lower EL lamps has been published in the manu-' ;L Range of green EL alphanumerics, maing 1* I display console variation, the Y coordinate for the greenI- II Y C u s t o m i z e d , 2-step potentiometerI *. l o 20 '40 60 80 100 120Shaft rotation, percentuustomized potentiometeruinear potentiometer

1 5 1 0 x l d o h m s

2 4 6 8 1 0 x l d o h m s

Figure 6. - Dimming characte rist ics ofC M alphanumerics control shaft. Color dri ft occ urs during the lamplifetime. The extent of dr ift by one lamp

is shown in figure 7. Although it is evi-dent that most lamps will drif t beyondspecified tolerances, the drift is expectedto be uniform and toward the sa me wave-length region. Also indicated in figure 7is the fact that, although X and Y toler-ances on the white EL lighting a r e rea -sonably close to the present st at e of theart, visually matching the lighting hues ofthe different instrument vendors is nearlyimpossible because the human eye is ex-tremely sensitive to color differences.

s of 100 noticeable hues fromhue reference point

The effect of voltage varia tion on ELote Dots in the square0 .1 .2 . 3 . 4 .5 .6 .7 .8 represent the shiftX in color with time. lamp brightness is an order of magnitudegr eat er than the effect of frequency.Brightness changes at the ra te of 0.025 ft-Lfor each volt of driving-voltage change atthe 75-volt region of operation. Thi s factwas significant primar ily in the qualifica-tion and acceptance test because the accuracy of measuring the driving voltage had tobe controlled careful ly. Otherwise, with a measurement accuracy of f0.02 ft-L, a

A-vol t tolerance at 75 volts would vary brightness by 0. 5 f 0.10 ft-L and thereby de-feat the kO.2-f t -L tolerance.

Figure 7 . - Eye sensitivi ty to Apollowhite lighting.

Inherent inefficiency of edge lighting presented some problems with brigh tnessqualification on the face of the m ete rs when EL lighting was used. White EL bright-ness at 75 volt s and 400 hertz aver ages approximately 2 ft-L afte r burn-in. A t least1 2 sq ua re inches of EL lighting were required to obtain candlepower comparable to thatof a grain-of-wheat lamp, which is a miniaturized incandescent lamp about the si ze ofa gr ai n of wheat.9


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The efficiency of E L lighting as a transilluminant on the panels averaged2 215 mW/in of lamp at 7 5 volts r m s and 40 0 hertz, or 4 0 mW/in at 115 volts. Thesefig ure s were obtained fro m statisti cal data fr om all the l amps on the panels.The most difficult problems i n the use of E L panels were the shape and s izing ofthe E L lamps, the tolerances required by the lamp manufacturer, and the manufactur-

er's ability to produce lamps with unusual shapes. Generally, the lamps fell into th efollowing four major groups.1 . The group 1 lamps were used for circuit-breaker panels and were made aslong as possible and to the width requi red to light the legends on the panels. In somecases, the length of the lighted area exceeded the lamp manufacturer's production capa-bility. In these ca se s, two o r more lamps were used, and the spaces between lampswere placed to correspond with the open areas of the panel (areas without letters).2 . The group 2 lamps contained holes o r cutouts to provide a cc es s areas forswitches, potentiometers, and rotar y devices. These lamps had to be designed to lightall legends and also to allow the lamp manufacturer enough unlighted areas to hermet-

ically seal the lamp.3 . The group 3 lamps were unusually shaped, with thin sections, thick sections,cr os s shapes, and L shapes, and had squar e o r round cutouts within them.4 . The group 4 lamps were small , individual lamps, 3/16 by 7/16 inch to 1/4 by3 inches in size.The group 1 lamps did not present any significant problems in the design stage.Group 2 and 3 lamps required great ca re in locating the circuit areas or unlightedareas as well as in selecting the tolerances allowed to the lamp manufacturer. It w a sdiscovered that no attempt should be made to design lamps of l e ss than 3/4 square inch

in size i n lighted areas; smaller lamps would not meet the illumination-intensityspecification.The lamp termi nals presented problems; "standard" termi nals were not usablefo r flight application because the reliability was le ss than desi rable. The wire meshbroke quite easily and the grommet s became loose , resulting i n a poor electrical con-nection. Phosphor bronze wi re mesh embedded in the lamination w a s used. A plastictophat was placed on the lamination to accept the number 26 wiring, and silve r threadswere embedded in the phosphor to reduce line drop. A s a resu lt of EL lighting degradation with temperature i nc reas e, another problem developed with the use of E L withinflight instruments. The specification on flight instruments *as 160" F, but the olderE L lamps were reliable only to 140" F. The lamp manufacturer modified the lamps,

and they then were qualified for the 160" F tempe ratu re environment. These newlamps were used for the first time on the L M - 5 flight vehicl e.The us e of EL lighting for wedge lighting on inst rumen ts presented some prob-lems . For example, problems were encountered i n the lighting used in the flight direc-to r attitude indicator . When E L lighting was u sed to illuminate a sphere from the side,the center of the sphere remained da rk , and uniform illumination was not obtained. Thproblem w a s resolved by using an E L light with a thick center and a narrowing area o rwedge toward the s ide .

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Other prob lems were encoun tered on the Apollo spacecra ft, but they were si mil arto the problems previously discussed and, therefore, will not be mentioned in thisreport .Floodlighting. - Floodlighting in the Apollo spacecraf t began with the conventionalincandescent sou rce. However, subsequent vibrational and heat-dissipation te st s dis- .

couraged the us e of this type of lighting in the CM in favor of a more efficient andrugged type. The incandescent source w a s maintained in the LM, and an isolationmount was use d to produce the qualification of the unit s under the Apollo launch vibra-tion (fig. 8). The LM floodlighting syste m was composed of white incandescent lamps,as follows.1. Overhead lights, one each above panels 1 and 2 (fig. 2)2. Forward lights, one each above panels 5 and 6 (fig. 2)3. Side-panel lights, a total of 31 lights above the rows of side-pane l D&C

The overhead and forward lights had dimming capability; the side-panel lights did not.

D

A tubular fluorescent lamp (fig. 4)was used in the CM. This lamp is f re-quency sensitive; but, through specialcircuits, is operable fro m a 28-volt dcsource . A conve rter within the light fix-tur e alter ed the 28 volts dc to ac. Theelectronic components were solid state tominimize weight. Noise proble ms (bothaural and electronic) ar e inherent butwere eliminated satisfactorily by themanufacturer.L. - The brigh tness leve l of the lamp wasincrea sed favorably during the 2 yea rs ofdevelopment. The initial value was great -er than 2100 ft-L. This brightness was

inc reased to 5000 ft-L by increasing ef-ficiency and by changing tube diamete rFigure 8. - Lunar module floodlight.

and length, which produced not only a higher level of to tal luminous flux output but alsohigher brig htness levels. Specified color coordinates for the floodlights were X = 0.365to X = 0.425 and Y = 0.365 to Y = 0.400, chosen to simulate incandescent colo r, whichgives a natura l appearance to the color of the skin.

The CM floodlights provided illumination pr ima ri ly f or the D&C panel s and lower .equipment bay and fo r gene ral activity within the spacec raft . The lights were also re-dundant with the in teg ral lighting. A secondary s yst em w a s integrated within the pri -ma ry floodlighting sy stem fo r redundancy a n d fo r added illumination during high-gravity.conditions. Required illumination levels approached 60 ft-c at the highlight ar eas ofthe panels, with the use of both primar y and secondary syste ms.

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During t h e Apollo 9 flight, the crew objected to the high temperature of the flood-light lens. When both lamps were operating in a 5-psia zero -g environment, the lenstemperature reached 170" F. An evaluation was conducted of the lens temperature ef-fe ct on the crewman's touch. The re su lt s of the evaluation a r e summar ized as follows.

*Temperature Result130" F Not objectionable.

170" F

145" F

Very hot, burning sensation . Can be touchedfor 1 to 3 seconds. -

Distracting ther mal shock. Can be touchedfo r 5 to 8 seconds.

Subsequent to the touch t emperatur e evaluation, the procedure was modified f o rsingle lamp operation, using the dua l lamps only when high level illumination was r e -quired. With a single lamp operating, the lens temperature reached a maximum of130" F. If the secondary lamp was energized while the pri mar y was operating, thetim e required for the lens to reach 170" F was approximately 30 minutes. On futurepr og rams , the operational temperatu re of equipment that inter faces with the crewshould be closely examined during prel iminary design pha ses and followed throughouthardware development.Dimming controls were provided for the pr imar y floodlighting system, while anon-off control w a s provided for the secondary sy ste m. Dimming cha rac ter ist ics of afluorescent lamp, as compared to an incandescent lamp, are shown in figure 9. Thecurves indicate the g re at er efficiency and a satura tion leve l of light output fo r thefluorescent lamp. Some hy st er es is in light output was experienced but was minimizedby additional cir cui try . The lower light level attained before extinction was approxi-mately 0.20 percent of maximum.Because of se ve ra l spec ial pro jec ts during the mission, floodlighting i n the CMhad to serv e more purposes than general illumination. Because it was desirable toobserve the astronauts in a zero-g environment, illumination levels for television andfilm ca mera s had to be satisfied. During television ca me ra testing i n the CM mockup,the floodlights near the fac e of the cente r astrona ut "blinded" the exposure control ofthe television cam er a, and satisfactory pictu res wer e not obtained. An automatic ex-posure control that w a s sensitive to overall scene illumination (instead of sm al l, brightareas of the scene) w a s used , but a t the expense of a le ss sensitive vidicon tube.

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Regulator fluorescent = 50 percentRegulator incandescent = 150percentR-Lmax - ft-Lmin looReg = It-L .min

lamp output28 V = 100 percent

.1L I I I I I I 10 5. 10 15 20 25 30 35V dc

Rheostat control system, 2 8 .OVconstant inpu t

0 4 8 12 16 MPower supply load, W

(a ) Dimming comparisons. (b) Dimming efficiency.Figure 9. - Dimming chara cte ris tic s of fluorescent and incandescent lamps.Tunnel lighting. - Additional floodlighting w a s provided to fcrnish ambient light-

ing for crew activity in the tunnel region between the docked CM and LM. The pr imar yactivity within either tunnel was hard-docked crew tr an sf er , and removal and replace-ment of the docking drogue and probe.Thr ee light fixtures , mounted on opposite si de s of the tunnel approximately at eyelevel, were used in the CM. A pair of ruggedized lamps, which were the sa me lampsused in the annunciators located on the main D&C panels, we re used in the light fixtures.The LM tunnel lighting w a s provided with the LM'utility light.Utility light. - Floodlighting would not be complete without a utility light or flash-light (fig. 10).jective tests indicated tha t only 0.05 ft- c was needed to perform such ta sks as findingan extra pair of socks. It also was indicated that 0.3 to 0.5 ft-c is sufficient fo r read-ing panel nomenclature and that 0.1 ft-c is sufficient fo r equipment stowage and re-':rieval and for removal of sc rews fr om panels. It was assum ed that the astronautsa r e dark adapted and want to remain that way. Too much light could destroy darkadaptation. However, reading can be accomplished with the utility light by holding thelight closer to the nomenclature.

Candlepower requi rements depend on the intended use of the light. Sub-

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It was decided that vari able inten-sity should not be used on the CM utilitylight because of heat generation and energywas te by the rheostat. The intensity of thelight was 1. 0 ft-c at 1 foot. Only one lightpe r crewman was provided i n the CM, withan 8-foot cable that connected to a 28-voltdc outlet.

:. .' .The LM general floodlighting wastoo low in intensity to provide sufficientillumination for reading flight plans, st arch ar ts , navigational data, and lunar land-ing maps. Theref ore , the LM utility lightwas of gr ea te r intensity than the CM utilitylight. The maximum intensity of the util-

Figure lo '- Commandmodule utilit y light and penlight.ity light was 5 ft-c at 3 feet. Again, itwas decided that variable intensity should not be used. Also, it w a s recognized that

dimming would be mandatory; therefore, a two-step dis cre te dimming control wasadded. There were two utility lights in the LM, one fo r each crewman. The intensityof the command pilot's light (on the left s ide ) was 5. 0 ft-c (high) and 2.0 ft-c (low) at3 feet, while the intensity for the lunar module pilot (on the right side) was 1. 5 ft -c(high) and 0.5 ft-c (low) at 3 feet. Both lights had an 8-foot cable that connected to a28-volt dc outlet. The base of the light had a univer sal clamp and ball. The light couldbe clamped to any int eri or LM struc tur e 0.5 to 1 .5 inches in diame ter. During LMtunnel activity, one or both of these lights could be mounted on the tunnel str uc tu re toprovide tunnel illumination.Special lighting. - Each astronaut w a s furnished a penlight for activity in hard-to-see areas when one o r two crewmember s we re maintaining dark adaptation. The dis-

tribut ion pat tern of thi s light at a distance of 2 fee t consisted of an uneven hotspot (8 to4 0 ft-c, average 15 ft-c) approximately 4 inches in mean diamete r, surrounded by adimmer area ( 0 . 1 5 ft-c) at 1 foot radius that extended to approximately 8 feet i ndiameter.It was discovered that the l ife and reliability of the penlights were not good andthat the operational life would vary from a few hours to sev era l weeks. Theref ore,

th ree lights per crewman were stowed on board f or redundancy and backup in ca se offai lur e. It would not be desirable to us e this sa me penlight in future manned missions,although the concept of providing the crewmen individual flashlights is valid. A reliablemedium-intensity flashlight should be part of the cr ew lighting inventory on all spacemissions to facilitate auxiliary lighting in hardlto-illuminate areas.Sunshafting. - The effect of sunlight e nteri ng the windows pre sen ts a problem. Atfirst, it may see m reasonable to use the sunlight for illumination. However, fur the rana lys is revea ls that sunshafting through the windows is not des irable . Sunlight is

nearly parallel, simila r to a spotlight. Consequently, whatever is illuminated by sun-light within the crew compartment will be illuminated to the point that everything el sewill be nearly silhouetted. As the spacecr aft tur ns, the position of the illuminated Spotwill move, perhaps into the face of the ast rona ut. Dark adaptation, which is necessary

.

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in seve ra l tasks, would be impossible . Also, the heat energy introduced into the space-craf t by the sunlight would have to be dissipated. Therefo re, i t was decided that , undernormal conditions, sunlight would be eliminated completely fr om the crew compartment. .The elimination of sunlight was accomplished by the u s e of opaque window shades.

Besides being opaque, the window shades also had to be highly reflect ive to pre -vent an y additional heat energy fro m overloading the environmental control sy stem fo rthe crew compartment. A highly specular reflective mat eri al is usually a materia l oflow emissivity, so that energy that is absorbed by the shade must be emitted principallyf rom a nonreflective surface. Thus, the equilibrium temperature of the shade shouldbe low enough that an astronau t can touch it without ha rm but not s o low that the shadematerial is rendered ineffective.The basic window-shade concept was a 1/32-inch- thick aluminum sheet config-ured to the shape of the CM window. The first 0.5 inch around the periphery had aVelcro sea l fo r attachment to the window.

Exte rna l Light ingExternal lighting was provided on the Apollo spacecraft for rendezvous , station-keeping, docking, optical tracking, EVA, and contingency extravehicular t ra ns fe r.Rendezvous and docking lighting. - The rendezvous and docking maneuver re-quired int erf ace lighting between the LM and the CM.and 12) was us ed fo r detection, illumination, and attitude orientat ion. All of the lightingaids had been proved successful fo r rendezvous and docking during the Gemini Pr og ra m.Both the LM and CM were equipped with a flashing xenon light. The light on the CMw a s mounted on the se rv ic e module (SM) in the positive Z axis and 12" toward the posi-

tive Y axis (fig. 11). The cone of radiation w a s 2 60" (120"), which placed the upperedge of the cone paralle l to the SM mold line. The intensity w a s 160 beam candle-second (c- sec)/flash. The unit, beam c-sec/flash, re fe rs to the integrated intensityof a flashing-type lamp. A t 60 nautical miles, the light is equivalent i n brightness toa third-magnitude star. It could be detected with optical a ids at 160 nautical miles.This light was a modification of the light used on the Gemini Agena Target Vehicle. Toprovide the longer visual range necessa ry, the light intensity was increased by chang-ing the cone of radiation from +80 (160') to +60' (120') and increasing the voltageinput to the light f ro m 32 to 55 volts.

External lighting (figs. 11

The LM tracking light was mounted between the two forward windows (fig. 12).This light must be acquired at a range of 400 naut ical miles using the sextant. Thetracking-light flash rate was 60 flashes/min, with a pulse width of 20 mill iseconds.The light intensity was rated at 1000 beam c-sec/flash. Thr ee range-detection evalua-tions wer e conducted on a prototype light to verify the maximum range. Two evaluations.we re conducted on the ground using neutral-density goggles to simulate ranges f ro m 1 to150 nautica l mile s. The third evaluation was conducted in an ai rc ra ft flight evaluationat an altitude of 25 000 feet. These evaluations confirmed that the light would be de-tected at 130 nautical mi les visually and at 420 nautical mil es with the aid of the CMsextant. On two Apollo flights, fa ilu res of the light occurred. The problem w a s

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Intensity: 0.5 ft-cat CM mold lineEV A floodlightpot1 gh t0

-2. ,0 Boresiahted o X a x i s at 300 R

ColorYellowRedGreen

AYntenna- ---Runninq

Number of lights Minimum intensi ty, cp Location4 0 . 2 3 Bottom2 . 1 5 Left side2 . 1 5 Right side

Green

Figure 1 2 . - Lunar module externallights, forward view.brightness0 Visible to ey e at 60 insufficient protection at the l amp termi-

nals, which resulted in corona and burn-0 1 flashlsec ing out the terminals. This problem wassolved by providing better environmentalprotection to the lamp terminals and re-ducing the input voltage.

miles or to telescopeat 160 mile s

Figure 1 1 . - Command and se rv ic emodule external lightingsystem. Fo r orientation and altitude aline-ment at a distance of 1 mile to 500 feet,running lights were used on both the LM and CM (figs. 11 and 1 2 ) . The color codingw a s the same as standard ai rc ra ft coding. Ther e were eight running lights mounted onthe CM/SM syst em. The four front lights wer e mounted on the CM adapter section,aft of the CM adapter/SM separa tion point. The four rear lights wer e located near therear bulkhead of the SM. The followingwere the locations, co lo rs , and intensit ies ofthe eight SM lights.

The light fixture s consisted of five grain-of-wheat lamps enclosed within a lensed- housing.The sa me type of lights and color coding fo r orientation and altitude alinementwere used on the LM, except that the re were no lights at the bottom. The LM al so hadrunning lights fore and aft. The following we re the locations, co lo rs , and intensit ies.

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Color Number of lights Minimum intensity, cp LocationWhite 1 0.23 AftWhite 1 .23 ForwardYellow 1 .23 Forward-

There w a s concern about the color discrimination range of the running lights andabout whether the crewmen could actually make positive discriminations between there d and yellow color s at 1000 feet. Therefore, tests were conducted at the FederalAviation Administrat ion fog chamber located in Oakland, California, to confirm that therunning lights could be detected. The resu lt s of the test showed that the lights were de-tected at 2000 feet and colors could be discriminated at 1000 feet.During the stationkeeping phase of the rendezvous and docking maneuver (500 to50 feet), the crew must be presented a three-dimensional view of the passive vehicle.A spacecraft docking floodlight (fig. 11)was added to the CM to provide illumination of

dlepower. The light w a s located on the EVA compartment door on the SM behind thecommand module pilot. The light illuminated the LM docking interface with 0.03 ft -cat 500 feet. The usab le distance fo r stationkeeping w a s 500 to 50 feet.

the passive LM during CM-active rendezvous. The light intensity w a s 8000-beam can-

The docking floodlight w a s originally the basic light used on the Gemini space -craft. It w a s discovered ear ly in the qualifications testing that the lamp and tr ans forme rmounting would not tol erate the high vibrations that were cha racte risti c of an Apollolaunch. Therefore, the light had to be mounted on a special isolation syste m to atten-uate the vibration.

The docked CM/LM and the crewmen's line of sight were not coincident. There-fo re , optical sight s and suitable tar get s were provided to give cues fo r the docking ma-neuver fr om clo se range (75 to 50 feet) to soft dock (figs. 13 to 15). The crewmanoptical alinement sight (COAS) w a s a columniated reticl e sim ilar to a gunsight used onaircraft and was attached to the rendezvous window before the final docking maneuver.The target in the CM w a s attached to the r ight rendezvous window before the dock-ing maneuver. The LM target w a s fix-mounted outside the spacecraft . The tar get sprovide a cue fo r the COAS. In thi s manner, proper orienta tion and spacecra ft aline-ment were provided for docking. The CM target consisted of a base 8 inches in diam-

et er and lighted by a green EL lamp, with a red incandescent cross placed 4 inches infront of the base to provide a three-dimensional effect. The cr os s w a s resolvable at75 feet . The brightness of the target w a s 28 ft-L on high intensi ty and 17 ft- L on lowintensity. .The LM target was twice the si ze of the CM tar get and was illuminated by RLdisks . The disks wer e 5/8 inch in diameter with a 1/2-inch-diameter circular area ofillumination. The intensity of the disks was 0.8 ft-L at time of launch. The half lifeof the RL disk w a s 18 months.

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LM activedocking target

Crew m an opticala l i n e m e n t s i g h t

Figure 1 3 . - Command module sightingaids.

Figure 15. - Command module dockingtarget.

Figure 14. - Lunar module dockingtarget.

The prim ary docking procedurewas per forme d using sunlight reflections .All of th e previously mentioned arti fic ia lillumination sourc es, excluding targ et,we re used to provide visual aid to the crew-men during darkside contingency docking.The basi c rendezvous and docking maneuvertechniques we re developed and verifiedduring the Gemini Pro gra m. These tech-niques wer e used f or docking the commandand se rv ic e module (CSM)/LM. One maj orchange to the CSM/LM vehicle was the ex-ter nal ther mal control coating and the in-cr ea se d geomet rica l complexity of thevehicle shapes . Both the CM and LMthe rma l control coatings wer e highly spec-~ul ar ; fo r example, the CM w a s covered with aluminized Mylar and th e LM with ano-dized aluminum. Coupling thes e vehicle reflection char act eri sti cs with the operationalvisua l environment, where the incident light fro m the s ola r di s k is collimated, the vis-ually perceived details of the two vehicles change marked ly as the relative positions Ofthe viewed vehicle change in relation to the sun and the obs erv er. These char acte ris-ti cs required definition for docking as a function of target vehicle and illumination en-vironment relationships to determine limiting conditions fo r astronaut visual capabilities.

The extre me complexity of the photometri c phenomena involved in the docking maneuversuggested that simulation using vehicle models and a simulated solar sou rce would benecessa ry to define the visual environment. Therefore, a cont ract was negotiated toconduct an Apollo illumination environment simula tion. The study was divided intothre e major phases.

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Phase I:1. CSM/Saturn IVB (S-IVB) separation and turnaround, and CSM/LM dockingand LM withdrawal from the S-IVB2. CSM/LM docking at various solar incidence angles and viewing angles3 . Operational requirements f o r photographing the LM on the lunar surfacePhase 11:1. Apollo Lunar Surface Experiments Package (ALSEP) deployment2. ALSEP visual "near-field" work ar eaPhase III:1. Visual detection of the sunlit CSM and LM2. Lunar horizon visibilityThese stud ies provided the following support to the Apollo mission operation .1. Removed the original sun constraint f o r LM withdrawal2. Established a set of sun const raints w h e r e high refl ect ive illumination wouldbe a visual problem during docking3 . Provided the bas is to establish sun/vehicle reflection conditions to demon-s tr at e and explain the COAS washout problem that occur red on Apollo 9, resulting ina modification to the COAS (The Apollo 9 COAS problem is discussed in detail in later

paragraphs. )4 . Provided information for the logical selection of the lunar surf ace modelingmat eri al used fo r the LM landing and ascent simulator5 . Established the exposure required for LM lunar surface photography, whichreduced the br iginal number of photographs of each area of the LM from four to twoexposures6. Provided the Apollo 11and 12 crews with a set of training photographs7. Resulted in two minor hardware changes to the ALSEP to establ ish maximummeter bezel heights for shadowing constraints and to establish reflectivity requirementsfo r the modular equipment stowage assembly and ALSEP decals8. Resulted i n the relocation of the bubble on the passive seismograph9. Provided maximum distance fr om which a sunlit vehicle may be detected,which provided backup navigational guidance data by optical tracking of the LM usingsunlit refle ctio ns off the therm al control coatings during LM descent, ascent, andrendezvous

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10. Established the range of LM specular reflectivi ty from the lunar su rface s othat specific acc ess ti mes could be determined fo r orbiting CSM sightingsThe Apollo 9 rendezvous and docking profi le was performe d with spacecr aft at -titudes and sun angles that produced ve ry bright spe cular reflec tions off the CM. Thisref lec ted gl ar e impinged di rec tly on the LM window and into the LM COAS, washingout the illuminated reticle. Two conditions that attr ibuted to the washout of the r et ic le

we re the following.1. The intensity of the COAS had been reduced by 90 percent with the addition ofan internal neutral density filter.2 . Specular reflec tions off the CM caused excessive g la re that impinged on theLM COAS optics with a flare ratio in exce ss of the design limits .The intensi ty of the COAS wa s originally 1000 ft-L o r more. This brightnesswould guarantee visibility against a background brightness of 10 000 ft- L. The Apollo9

COAS full intensity was between 50 to 90 ft- L, which was too low to provide prop erbrightness compatibility between the COAS and the specular gla re f ro m the passi ve ve-hicle. This situation resul ted in the crew man not being able to see the reticl e againstthe brighter gla re off the passive vehicle. Subsequent to the Apollo 9 experience, twochanges were made, one in the design of the COAS and one in procedure.

1. The COAS was modified by removing the internal filter and remounting it out-si de the COAS and by providing a means of removing the filt er i n the event a brighterreticle is neces sary. This change inc rea sed the full bright ness of the retic le to1000 ft-L.2. Procedurally, i f the docking profile di ctat es such a condition as occurr ed onApollo 9, a passi ve o r active vehicle ro ll attitude change will be initiated to precludespecular glare i n the d irection of the docking window of the active vehicle.Extravehicular lighting. - Scheduledand contingency EVA operat ions requ iredspecia l extravehicular lighting. A singlefloodlight used t o ill umina te the CM/LMexternal area was mounted on the CM(fig. 16). The light was positioned be-tween the righ t rendezvous window andsi de window by a 24-inch pole, mountedto the SM, that was extended automatica llyafter the boost protective cover w as jct-tisoned. The light fixture was the sam eas that used for the running lights exceptthat the color fil te rs had been removed.The light was oriented to illuminate theCM hatch, right-side EVA handrails, andLM EVA trans fer handrails. The illumi-

nation of t h i s light was 0. 2 to 0. 5 ft-c

20

Figure 16. - Extravehicular activitylight and handrai l identificationlight.


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on the CM mold line. This light als o served as a docking floodlight to illuminate theCSM for LM active docking.To aid the EVA ast ronau t in locomotion about the exter io r of the CM/LM w a s aser ies of EVA handrai ls. Radioluminescent disks were used to aid the crew in locating

the handrails as well as the environmental control system exter ior dump valve andhatch-opening mechanism and handle. These disks were the sa me as those used onthe LM ta rget . The location of the disk on the EVA handrail s is shown in figure 16.Two disks are mounted at each end on the handrail base.

,

CO NCLUDI NG REMARKSThe following are the four types of lighting used on the Apollo lunar landingmission.1. Fluorescent (with dimming capability achieved by varying voltage, frequency,and wave shape)2 . Incandescent (with and without dimming capabil ity)3 . Electroluminescent (with dimming capability, ac voltage control)4. Radioluminescent (such as promethium-1 47)A conservative approach has been used in developing spacecraft lighting to satisfyApollo illumination requ irements. Although other ideas wer e considered, the pr im ar ydesign cr it er ia for the Apollo lighting systems were reliability, crew safety, and mini-mization of spacecraft weight. Proven methods of illumination consistent with estab-lished aviation standards have been implemented when possible.

Manned Spacecraf t CenterNational Aeronautics and Space AdministrationHouston, Texas, November 9, 1972924-23-D7- 82-72

NASA-Langley, 1973 1 S-360 21


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apollo experience report crew station integration. volume v lighting considerations

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