apollo experience report spacecraft pyrotechnic systems

8/8/2019 Apollo Experience Report Spacecraft Pyrotechnic Systems

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S A T E C H N IC A L N O T E

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APOLLO EXPERIENCE REPORT -SPACECRAFT PYROTECHNIC SYSTEMSby Mario J . Falbo and Robert L . RobinsonManned Spacecrafi CenterHoustoiz, Texas 77058N A T I O N A L A E R O N A U T I C S A N D S P AC E A D M I N I S T R A T I O N W A S H I N G T O N , D. C. M A R C H 1 9 7 3


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1. Report No.NASA TN D-7141

APOLLO EXPERIENCE REPORTSPACECRAFT PYROTECHNTC SYSTEMS

2. Government Accession No. 3. Recipient's Catalog No.

March 19736. Performing Organization Code

4. Title and Subtitle

7. Author(s)Mario J. Falbo and Rober t L. Robinson, MSC

5. Report Date

8. Performing Organization Report No.

I MSC S-3409. Performing Organization Name and AddressManned Spacecraft CenterHouston, Texas 77058

10. Work Unit No.914 50-17-08-7211. Contract or Grant No.

13 . Type of Report and Period CoveredTechnical Note. Sponsoring Agency Name and Address

19 . Security Classif. (of this report)None

National Aeronautics and Space AdministrationWashington, D.C. 20546

20. Security Classif. (of this page) 21 . NO. of Pages 22. PriceNone 46 $3.00

14 . Sponsoring Agency Code1

15. Supplementary NotesThe MSC Directo r waived the use of the International System of Units (SI) for th is ApolloExperience Report, because, in hi s judgment, the use of SI Units would impair the usefulnessof the repor t or resul t in excessive cost.

16. Abstractmr ot ec hn ic devices were used successful ly in many systems of the Apollo spacecra ft.physical and functional ch ara cte ris tic s of each device ar e desc ribed. The development,.qualification, and performance tests of the devices and the ground-support equipment a r ediscussed briefly. Recommendations for pyrotechnic devices on future space vehicles a r egiven.

The

17. Key Words (Suggested by Author(s) )Initiator * Disconnects Linear -Shaped ChargeCar tri dges Mild Detonating Fuse ValvesThrusters ' High Explosive ' Parachute MortarsActuators ' Circuit InterruptersGuillotines .Separation System

18. Distribution Statement


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CONTENTS

Section Page

SUMMARY . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .INTRODUCTION . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .DESCRIPTION OF SYSTEMS . . . . . . . . . . . . . . . . . . . . . . . . . . .

Launch-Escape System . . . . . . . . . . . . . . . . . . . . . . . . . . . . .Command and Service Module Systems . . . . . . . . . . . . . . . . . . . .Lunar Module Syst em s. . . . . . . . . . . . . . . . . . . . . . . . . . . . .

COMPONENT DESIGN AND DEVELOPMENT . . . . . . . . . . . . . . . . . .Single Bridgewire Apollo Standard Initiator . . . . . . . . . . . . . . . . . .Cartridge Assemblies . . . . . . . . . . . . . . . . . . . . . . . . . . . . .Detonators . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .Core Charges . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .Line Cutters . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

GROUND-SUPPORTEQUIPMENT . . . . . . . . . . . . . . . . . . . . . . . .Pyrotechnic Simulators . . . . . . . . . . . . . . . . . . . . . . . . . . . .Spacecraft Verification Equipment . . . . . . . . . . . . . . . . . . . . . .

CONCLUSIONS . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .RECOMMENDATIONS . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Single Bridgewire Apollo Standard InitiatorExplosive Tr ains and Interfaces . . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . .Separation Systems. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .Reefing Line Cutter s . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .Firing Circuitry . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

81417182021232526263236373838393939

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Table

TABLES

PageI APOLLO PYROTECHNIC CARTRIDGES

11

(a) Pr es su re and igniter cartridges . . . . . . . . . . . . . . . . . .(b) Detonator cartridges . . . . . . . . . . . . . . . . . . . . . . . 55APOLLO CORE-CHARGE ASSEMBLIES . . . . . . . . . . . . . . . . 6

FIGURES

Figure1 Locations of Apollo pyrotechnic devices . . . . . . . . . . . . . . . .2 Single bridgewire Apollo standard initiator . . . . . . . . . . . . . .3 Apollo pyrotechnics configuration . . . . . . . . . . . . . . . . . . .4 Launch-escape-tower jettison . . . . . . . . . . . . . . . . . . . . .5 Tower -separation system . . . . . . . . . . . . . . . . . . . . . . .6 Lunar module separation system . . . . . . . . . . . . . . . . . . . .7 Adapter panel explosive train system . . . . . . . . . . . . . . . . .8 Separation of SLA panels

(a) Operating detail s . . . . . . . . . . . . . . . . . . . . . . . . . .(b) Panel jettison . . . . . . . . . . . . . . . . . . . . . . . . . . .

9 Electrical circuit interrupter s(a ) Command module . . . . . . . . . . . . . . . . . . . . . . . . .(b) Service module . . . . . . . . . . . . . . . . . . . . . . . . . . .

10 Pyrotechnically operated valve(a) Closed position . . . . . . . . . . . . . . . . . . . . . . . . . . .(b) Open position . . . . . . . . . . . . . . . . . . . . . . . . . . . .

11 Command module and se rv ice module s tr uc tu ra l separationsystem . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .12 Command module/service module guillotine . . . . . . . . . . . . . .13 Command module forward compartment . . . . . . . . . . . . . . . .

Page4778899

1010

1111

1111

121213

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Figure14 Apex-cover thruster locations . . . . . . . . . . . . . . . . . . . . . .

Reefing line cutter . . . . . . . . . . . . . . . . . . . . . . . . . . . .16 Parachute disconnect system . . . . . . . . . . . . . . . . . . . . . .17 Earth landing sequence. . . . . . . . . . . . . . . . . . . . . . . . . .18 Lunar module pyrotechnics . . . . . . . . . . . . . . . . . . . . . . .19 Lunar module landing-gear uplock and cutte r assembly . . . . . . . .

. 15

2021

22

232425262728

293031

Lunar module electrical circuit interrupter. . . . . . . . . . . . . . .Lunar module interstage nut and bolt assembly

Page13131314141415

(a) Before firing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 16(b) After firing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 16Lunar module interstage umbilical guillotine(a) Cross section . . . . . . . . . . . . . . . . . . . . . . . . . . . . 16(b) Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 16Initiator indexing 19Electrically initiated cartridge . . . . . . . . . . . . . . . . . . . . . 20. . . . . . . . . . . . . . . . . . . . . . . . . . . . .Spacecraft cartridges . . . . . . . . . . . . . . . . . . . . . . . . . . 20Spacecraft/lunar module adapter thruster cart ridge 21. . . . . . . . .

21nitiator with weld washer . . . . . . . . . . . . . . . . . . . . . . . .Apollo s tandard detonator

. 21212222

(a) Detonator car tridge assembly . . . . . . . . . . . . . . . . . . . .(b) Cro ss section . . . . . . . . . . . . . . . . . . . . . . . . . . . .End-detonating cartridge . . . . . . . . . . . . . . . . . . . . . . . .Long-reach detonator . . . . . . . . . . . . . . . . . . . . . . . . . .Core charges . . . . . . . . . . . . . . . . . . . . . . . . . 23a) Mild detonating fuse 23. . . . . . . . . . . . . . . . . . . . . . . . 23c) Linear-shaped charge(b) Confined detonating cord . . . . . . . . . . . . . . . . . . . . . .

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Figure Page32 Command and service module initiator simulator

(a) Frontview . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 27(b) Rearview . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2733 Schematic ofCSM initiator simulator . . . . . . . . . . . . . . . . . . 27

34 Functional block diagram of LM initiator simulator . . . . . . . . . . . 2935 Environmental initiator simulator . . . . . . . . . . . . . . . . . . . . 3036 Static EM 1 device . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3037 Stray electri cal energy indicator . . . . . . . . . . . . . . . . . . . . . 3138 . High-energy -initiator simulator . . . . . . . . . . . . . . . . . . . . . 3139 Pyrotechnic checkout test set

(a) Dummy ini tiator . . . . . . . . . . . . . . . . . . . . . . . . . . . 32(b) Portable measuring device . . . . . . . . . . . . . . . . . . . . . . 3240 Pyrotechnic bridge checkout unit

(a) Remote control assembly . . . . . . . . . . . . . . . . . . . . . . 33(b) B r i d g e u n i t . , . . . . . . . . . . . . . . . . . . . . . . . . . . . . 33

41 Explosive device test set . . . . . . . . . . . . . . . . . . . . . . . . . 3442 Current regulator . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3543 Functional block diagram of the cur ren t regulator . . . . . . . . . . . 3544 Pyroharness shorting plugs and cable set . . . . . . . . . . . . . . . . 35

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APOLLO EXPER I NCE REPORTSPACECRAFT PYROTECHNIC SYSTEMS

By M a r i o J. Fa lbo an d Robert L. R o b i n s o nM a n n e d S p a c e cr af t C e n t e r

S U M M A R YPyrotechnic devices were used in the Apollo spacecraft sy st em s to perform thefollowing functions: launch-escape-tower separation, separation-rocket ignition, sepa-ration of the booster stage from the lunar module, forward-heat-shield jettison,spacecraf t/lunar module adapter panel separat ion, lunar module landing -gear deploy -ment, pressur ization and activation of the lunar module propulsion systems, deploymentand release of parachutes, opening and closing of elec tr ical circu its, execution oftiming and delayed-time functions, and cutting of lines and cables .Requirements for high reliability and maximum safety were met with devices ofminimum weight and volume.low-energy charges for puncturing gas bottles to high-energy charge s for cutting0.153-inch-thick steel.

The capabilit ies of the devices ranged from relatively

Conventional elect ric al and mechanical components were used, when possible, tominimize potential design problems, Selection of proper explosive materia ls a lso w a svery important. Te st and evaluation prog rams were continued after the flight qualifica-tion of the devices to unders tand bett er the reliability, safety, vulnerability, and outputfac tors . Redundancy and common usage enhanced confidence in the overa ll pyrotechnicsystems. To protect against spurious form s of electrical energy, standard safetypracti ces wer e followed in the design of the el ectric al sys tems.

NO failures of pyrotechnic devices have been detected during any of the Apollomissions. This reliability probably is attributable, at least in larg e measure, to theconservative design, closely controlled manufacturing pro cesses and testing techniques,and thorough acceptance procedures.

INTRODUCTIONThe total number of pyrotechnic devices in the Apollo spacecraf t sys te ms variedfo r different spacecraf t. More than 210 pyrotechnic devices were used per flight toperform a myria d of onboard, inflight, timed, and controlled tasks automatically o r oncommand in the Apollo spacecraft sy stems. All devices required high reliability ands a f e t y . Most w ere classified as either crew-safety critic al or mission critica l, because


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improper operation o r failure to operate could have resu lted in los s of the crew, infailure to meet a mission objective, o r in an aborted mission.The high speci fic energy and other unique pr operti es of explosive and pyrotechnic

ma te ri al s afforded the method for providing a large energy s ourc e in a smal l package.By using explosives, numerous functions were accomplished re liably and safely withminimum weight and space limitations. These proper tie s, coupled with the capabilityof the pyrotechnic devices to release energy at a high ra te in a shor t time, made wideacceptance in the Apollo Program a natural result.

Confidence in the subsystems was further enhanced with maximum use of redun-dancy. When complete system o r device redundancy was not possible because of spaceor weight limitations, redundant cartridges or single cartr idges with dual initiatorswere used. Two separat e and electrically independent sys tems operated in parallel andprovided complete redundancy in the firing circuitry.

Early in the Apollo Program, the concept of modular cartr idge assembl ies basedon a standard ized hot-wire initiator w a s adopted to avoid the expense and time involvedin extensive testing associa ted with the development of different init iators for variousapplications. A higher confidence in reliability was achieved by the us e of standard ized,high-volume items. Components, subassemblies, and ass embli es were qualified s e r i -a l l y during development of complete systems . Thus, confidence in the reliability of theinitiator was enhanced through increased testing with its common use in components.

When possible, the principle of commonality was extended to other assemblies.That is , where a new application required an assembly that used a device (or devices)in a manner almost identical to the use for which existing devices were originally de-signed and tested, the new assembly would be closely re lat ed in design and functionalcharacteristics. Additional confidence w a s attained through extensive performance datathat we re compiled for all applications, because the systems were largely dependent oncomponent interactions. The policy of standardization of components, which wasachieved i n the Apollo Program to a grea ter extent than in any other space program,w a s not easy to implement, primarily because of natural tendencies of various primecontrac tors and subcontractors to diverge on the basi s of unique technical requirements,both real and unreal,

Where use of the s am e hardware in different applications was not feasible, theFor example, the opposing-blade guillotine ,ame or s i m i l a r techniques were used.which severed the umbilical between the command module (CM) and the serv ice module(SM), w a s used as the basis for the designs of the following: (1) the lunar module (LM)inte rstage guillotine, (2) he two guillotines for umbilicals between the LM and thespacecraft/lunar module adapter (SLA), and (3 ) the landing-gear uplock cutter on theLM.

The quality of explosive materials was very important in the reliability of eachsys tem. Only newly manufactured, specification-controlled cyclotrimethylenetrinitra-mine (RDX)nd hexanitrostilbene (HNS) were used to ensure consistent quality andtraceabili ty of the high-explosive mater ia ls .HNS 11, differed in particle size and purity. The two types of HNS used, HN S I and

2


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vhf antenna deployment

Pitch-control motor

separation thrust e

LM ascent propulsion system(hel ium and propellant) valvesM separation system

tiedown lin eguillotineNotes R C S - reaqtion control systemGSE - ground-support equipmentS B A S I - single bridgewire Apollo standard initiatorvhf = very-high frequency

Figure 1. - Locations of Apollo pyrotechnic devices .

main

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TABLE I. - AF'OLLO PYROTECHNIC CARTFUDGES(a) Pre ssur e and igniter cartr idges

Cartridge1 Cartr idee character is t ics 1 Nominal I Romh I

Use~ perform-ance. Volumein. in. psi TypeDiameter, Threads/ Type in

Type n 3 4 16

Pressur e car tr iQes

Right hand 2 100 10 Closed Tower-Jettison motor 2 (e)

CanardType 1

Type nType N

Type NType N

Type VIType 100

Type 100Type 200Drogue disk

Main disk

LM valveElectrical

circuitinterrupterExplosive nu l

Explosivebolt

SBASI

SLA thrust er

Nominal performance capability

0.045-in. dent in aluminum

,045-i n, dent in aluminum

,018-in. dent in steel

,022-in. dent in steel

1-1/2718

11/16

15/16

15/1615/16

1-1/1611/16

11/163 413/16

1

3/81/16

9/16

1 1/16

3/8

1-1/16

Use

Various locations onCSM

SLA separation

LM guillotine andlanding-gear uplock.

Docking ring separatio

1214

12

16

1616

1824

241620

16

2424

24

18

24

18

h a m e t e r ,in.

9/16

9/16

9/16

5 8

Right handRight hand

Right hand

Right hand

Rtght handLeft hand

Right handRighl hand

Left handLeft handRight hand

Right hand

Left handRight hand

Right hand

Right hand

Right hand

Right hand

TypeThreads/in.18 Right hand

18 Left hand

18 Right hand

18 Right hand

13 50014 50 0

11 200

2 250

2 2502 250

14 5009 000

9 00012 9005 800

10 500

1 600I 000

6 800

23 000

65 0

4 200

2

10

-0

4 . 8

. 9

1 . 9

4-(1)

2000

-52

8. 9

8 .8

8 . 88.8

. 5

. 5

I. 0

1010

2. 7

2 . 5

10

Igniter cartridges

-losed

Venied

Vented

Closed

ClosedClosed

VentedClosed

ClosedClosedClosed

Closed

ClosedClosed

Closed

Closed

Closed

Closed

Cmar d thrusterDrogue parachutemortarPilot and drag par a-

chute mortarCM RCS propell ant

valveSM circuit interrupterCM RCS propell ant

valveApex-cover th ruste rCM and LM RCS

helium valvesCM RCS helium valveCM circuit interrupterDropue parachutedisconneciMain parachute

disconnectLM propulsion systemLM circuit interrupler

LM interstageseparation

LM tnterstage system

Vent valve and dockingprobe retraction

SLA-panel deployment

Number ofcar tr idgeson eachspacecraft

24

8

5

41

412

24

C

C

224

4

4

4

8

Approximatenumber fire da f t e rqualificalion

5001000

2000

(a)

(a )(a )

1000(b)

(b)200300

400

300200

200

200

(b)

600

I I (e)Type I I 5/ 8 I 18 1 Right hand I 2 100 I I 10 I Closed I Launch-escape/pitch-control motors(b) Detonator cartridges

Cartridgetype

ApollostandarddetonatorApollostandard

detonatorEnd-type

detonatorLong-reach

detonator

Number ofcar tr idgeson eachSpacecraft

26

Approximatenumber fire da f t e rqualification

I 300a ~ t a ~umber of type N i r ings is 100.bTotal number of SBASI firin gs i s approximately 1000.c ~ oBAWS per c a r t r i a e .

dNonelectric c artri dge initiated by confined detonating cord.e ~ ~ t a lumber of igniter cartr!dge fir ing s i s 1500.'Total number of standard detonator firings i s 5000.

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TABLE II. - APOLLO CORE-CHARGE ASSEMBLIESConcentrat ion Num ber ofan d type ofexplosive

L ine a r - hagA sse mbly na me

!d charge100-grain/ft RDX I P b I 2 (6) I T e ns ion- t ie c u t t e r2-grain/f t RDXZ-grain/ft RDX2-grain/ft RDX3-grain/f t HNS IITwo 5-grain/ft RDX

T w o ZO-grain/ft RDXB-grain/ft RDXTwo B-grain/ft RDXTwo 5-gra in/ft RDXTwo 5-grain/ft RDXT w o 5-grainlf t RDXTwo 5-grain/ft RDXTwo 5-grain/ft RDXTwo B-grain /ft RDXTwo 7-grain/ft RDX1B-grain/ft RDXTwo O-grain/ft RDXTwo lO-grain/ft RDX28 . L-grain/ft HNS II

Confined detonating c or dP bP bP bAgP b

P bP bP bP bP bP bP bP bP bP bP bP bP bAg

Lower umb i l ica l gui llo t ineGround- sup port equipment (GSE) um bilica l guil lot ineSM/SLA umbilical disconnectGSE umb ilical guil lot ineLM gui l lo t ineSLA pane l thrus te r

Mild detonating fus eCSM umbilical guil lot ineCSM umbilical guil lot ineSL A - orwa rd ins ide longi tudinalSL A - orwa rd outs ide longi tudinalSL A - f t c i r c umf e r e n t i a lSL A - ower ins ide longitudina lSL A - ower outside longitudinalCen te r outs ide longi tudinalSL A - ente r ins ide longitudina lSL A- or w a r d c i r c umf e r e n t i a lSM/SLA umb i l tca l d iskLower umb i l ica l gui llo t ineGSE umbilical guil lot ineLM guil lot ine

% umbe rs i n pa r e n the se s r e p r e s e n t t o ta l numbe r of i t e ms on a spa c e c r a f t .

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Spark gap detail

Pressure cartridge(type UIForward-heat-shield augmented m ortarPilot parachute mortar

Ignit er cartridge(types I and 11)- Launch-escape-motor igniter- Pitch-control-motor igniterTower-jettison-motor igniteri nd-detonating cartridge 1

LM landing-gear uplockLM interstage guillotine

Pressure cartridge l t p ElSM circuit interrupterCM RCS propellant Valve

Pressure cartridge ItypeV11

L C h a r g e w a s h e rhole L eld washer

- Canard thrus ter pressure cartridge- LM circuit interru pter pressure cartridge- Drogue parachute mortar pressure cartridge- CM circ uit in terrup ter pressure cartridge- LM RCS helium valve cartridge- LM propulsion system valve cartridge- LM interstage separation nut cartridge- LM inte rstage separation bolt cartridge- Long-reach detonator for docking ri ng- Docking probe he liu m bottle valve- Docking tunnel pressurization valve- Main parach ute disconnect cartridgeL Drogue parachute disconnect cartridge

,

separation charge

Figure 2. - Single bridgewire Apollostandard initiator.

the explosive tr ain syste ms were designedto be dependent on the proper interactionof the standard initiator and on exactlymatched input and output characteristics.The selection of a standard initiator also(1)played an important role in the over-all reliability of the pyrotechnic systems,(2) resulted in sho rt er development timesfor higher assemblie s, and (3 ) alloweddemonstration of high reliability at a highconfidence level fo r each device.

The Apollo pyrotechnics configurationis shown in figure 3. In som e applications,the initiator w a s used alone as a pressurecartridge. However, for most applications,additional explosive elements or explosivetrains were assembled to the initiator toproduce a desi red effect. For example, anintermediate charge w a s used in a detonatorto augment and transfer a detonation waveto the main charge of high explosive. Otherexamples ar e certain types of cartridges fo rpres sure generation systems in which an in -termediate and an output charge are required.

Apollo initiatorLApollo standard detonator

Linear-shaped charge for CMlSMtension-tie cutterCMlSM umbilical gu illotineTower-separation frangible nutLMlSLA lower umbilical guillotineLMlSLA frangible link

Thruster cartridgeJoints and interfaces

LMlSLA umbilical guillotineSMlSLA umbilical guillotine

Figure 3. - Apollo pyrotechnics configuration.7


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Launch-Escape SystemThe LES included the launch-escape motor, the pitch-control motor, the canardactuating mechanism, and the tower-jettison motor. The LES w a s designed to pull theCM from the launch vehicle if a pad abort o r an abort during first-stag e ascent w a s nec-

essary , To the present, it hasnot been necessary to use the LES for its intended function on any Apollo mission.Frangible nuts were used to s ec ure the LES to the CM.

In a nominal flight, the LES is jettisoned after the ignition of the second-stagebooster (fig. 4). The separation is accomplished by simultaneously igniting the tower-jettison motor and the frangible nuts in the base of each tower leg (fig. 5). However, ifan emergency should occur, the CM would be separated from the launchvehicle immedi-ately, and the pitch-control motor would ignite simultaneously with the launch-escapemotor to provide late ral translation and to as sur e safety of separation. Eleven secondsaf ter LES abort initiation, the canards would deploy to stabil ize the system. Three sec -onds later, the docking ring would separate from the CM to be jettisoned with the LES.The jettison of the apex cover and the deployment of landing parachutes would completethe abort sequence.

Identical igniter cartridges were used in the launch-escape and pitch-controlrocket -motor igni ters; only the th read and connector indexing differed for the tower -jettison-motor igniter cartridge. A power cartridge was used to pres sur ize a thrusterlinkage, which deployed the canards . Detonators were used for fracturing the frangi-ble nuts that secured the tower to the CM;

Launch-escape motor

Figure 4. - Launch-escape-towerjettison. Figure 5. - Tower-separation system.

Command and Service Module SystemsThe CM and SM pyrotechnic devices performed a multitude of ta sk s in each flight.On a typical lunar mission, these tasks began approximately 4 hours into the missionwith CSM. separation from the third stage (S-WB) of the launch vehicle and were

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completed a few moments after CM splashdown. The pyrotechnic events accomplishedduring a mission are described in the following paragraphs.Separa tion of the CS M and the SLA. - The SLA connected the CS M to the S-IVBuntil separation. Until then, the LM w a s contained in the SLA and w a s secu red to thelower section by means of four tiedown straps attached to the apex of the four outriggers

(fig. 6). An explosive tr ain system formed an integral par t of the SLA system and wasused to disengage the four SLA panels on command. The explosive train system con-sist ed of a number of M D F lines, panel thrus ters, an umbilical disconnect, and an um-bil ical guillotine. The components were interconnected by means of CDC and transfercharges. Details of the explosive tr ain sy stem ar e shown in figure 7.

A

LFlexiblepolyurethanetlock

systemDetail A

L o e t o n a t o rDetail B

Figure 6. - Lunar module separationsystem.

shedSection A-A Section 6-B Section C-CAsxxiated devices

-SLApanel

. . .

Umbilical guillotine

Detail Chousing- / (adapter)

C h a r g e h o w Detailretainer

SMlSLA electricalumbilical disconnect

Detail 0Figure 7, - Adapter panel explosivetrain system.

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Detonators were fired to initiate the explosive tra in system, which caused thefour SLA panels to be jettisoned and permitt ed separa tion of the CSM from the S-N B(fig, 8). The M D F w a s used to seve r the forward and aft circumferential and the innerand outer longitudinal splice pla tes of the SLA. The four panels we re folded back andwere je ttisoned by means of the cartridge-actuated th rus ter s. When the splice platewas severed, an explosive-operated guillotine severed the umbilical between the LMand the SLA. A spring reel then retracted the umbilical arm, which w a s jettisoned withthe panel. An explosive charge separated the SM/SLA umbilical disconnect.

/-Support

After panel deployment,dart of jettison

(a) Operating details. (b) Panel jettison.Figure 8. - Separation of SLA panels.

Separation of the SLA and the LM. - After the CSM separated from the S-IVB, theCSM w a s docked to the LM , and an elec tric al umbilical was connected to the LM/SLAseparation firing circuits through the docking tunnel. Detonators in frangib le links,attached to the four tiedown straps, were fired to permi t separa tion of the LM fromthe SLA. Thirty milliseconds after firing the frangible-link detonators, an explosiveguillotine was fired to sev er the umbilicals fo r the frangible-link firing circuit. De-tails of the LM separation system a r e shown in figure 6.

Docking probe retract ion. - During docking of the CM to the LM, the dockingprobe (which provided for CM/LM coupling) was re tr ac te d by means of SBASI units inthe probe retraction system. The SBASI, when fired, punctured a gas bottle to re leas egas that drove a piston to re tr ac t the probe.Docking ring separation. - The LM docking ring separation sys tem was p ar t of theCM. The system contained M D F charges that were used to cut the ring and to permitfinal separation of the LM from the CM. During a norm al mission, the docking ringvas separated from the CM and remained with the LM. In the event of a launch emer-gency requiring L E S abort, the docking ring would have been jettisoned with the launch-escape tower.

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Separation of the CM and the SM. - On return to earth and before the CSM enteredthe atmosphere, the propellant tanks in the CM reaction control system (RCS) werepressurized by opening pyrotechnically operated isolation valves. At separation, sev -era1 pyrotechnic events occurred. The circuit interrupt ers in the CM and SM (fig. 9)were actuated by means of electrica lly in-itiated pr es su re cartr idges to dead-facethe CM and SM elec tr ical circuits . Nu-merous fuel and oxidizer dump and purgingfunctions were performed by pyrotechnicallyoperated valves. The genera l configura-tion of the valves is shown in figure 10.

,n,t,ator

(a) Command module. (b) Service module.Figure 9. - Electrical circuit interrupters.

m - nitiator -InitiatorBody Body

(a) Closed posit ion. (b) Open position.Figure 10. - Pyrotechnically operated valve.

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The CM RCS contained 16 pyrotechnically opera ted valves, which were used to controlthe distribution of helium and propellants. Each valve (normally closed) was actuatedby an electr ically initiated car tridge. Upon firing, the valves remained openpermanently.

Tension ties connecting the SM and the CM (fig. 11)were cut by means of linear-shaped charges. The CM/SM guillotine (fig. 12) cut the umbilical between the CM andthe SM to allow the hinged umbilical boom to swing c lear of the CM and permitt ed s ep-arat ion of the C M from the SM. Similar pyrotechnic events would have occurred i m-mediately upon initiation of an LES abort .

Era(

- Tension ties

Tension-tie strap

Figure 11. - Command module andservice module structura l sepa ra-tion system.

Manifold chargeooster chargeBlade-

Booster c h a r g e lFigure 12. - Command module/servicemodule guillotine.

Ear th landing system. - The installa-tion of ear th landing syst em (ELS) equip-ment in the for ward compartmen t of he C Mis shown in figures 13 and 14. During nor -mal entry, the apex cover was jettisonedwhen the spacecraft had descended to ap-proximately 24 000 feet. As the cover sep -ar at ed from the CM, a lanyard-operatedswitch fired a drag parachute mor ta r at-tached to the cover, When deployed, thedrag parachute prevented the apex coverfrom recontacting the CM or from int er -fering w i t h drogue parachute deployment,which followed 2 seconds after cover jetti-son. The drogue parachutes wer e deployedin a reefed condition. At line -st retch, the

time-delay line cutters (fig. 15)were actuated; the line cu tters disreefed the drogues10 seconds later. At approximately 10 000 feet, the drogue parachutes were releas edby severing the r i s e r s with cartridge-actuated guillotines (fig. 16). The main par a-chutes were deployed in the reefed condition at the sam e time by means of mortar -ejected pilot parachutes. At line-s tretch, 6- nd 10-second-time-delay line cutter swere actuated to effect disreef in two stages. The deployment of the main parachuteactuated line cut ter s that, 8 seconds later, automatically deployed tw o very-high fre-quency (vhf) recovery antennas and a recovery beacon from the forward compartment.

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R C S engineprotector-

LES towerMra ch u te \ electricalSea recovery Droguecable7\mortar \receptacle 7

L arachute \\ r =FMain land ing and Swimmerparachute umbi I ca l

r iserprotectorL prighting bags L-High-frequencyunder main parachutes recovery

mortar 13 places) antenna

Forward-heat-shieldthrust p ressurecartridges(2 required) -l

-Y

Figure 13, - Command module forward Figure 14. - Apex-cover thrustercompartment. locations.

7RetainerFi r ing p in

Time-delav mix-R N dhesiveJ

Figure 15. - Reefing line cutter.Immediately afte r splashdown, the mainparachutes were rel ease d by car tridge-actuated blades in the parachute disconnectassembly . After touchdown and main pa ra -chute release, all pyrotechnic functionswe re completed for the CSM syste ms. TheELS sequence is illustrated in figure 17.

Note A l l risers shown are bundles of steel cables

Section A -A

Figure 16. - Parachu te disconnectsystem.13


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Lunar Module SystemsThe LM pyrotechnic devices andsy st em s were used for deployment of the

landing gear; fo r opening of valves f o r pr es -surizat ion of the descent, ascent , and RCSpropellant tanks; for venting of descent pr o-pellant tanks; for electrical circuit inter -ruption; for interstage umbilical severance;and for separation of the ascent and descentstages. The'general locations of the LMpyrotechnic devices are shown in f igure 18.

1. Apex cover jettisoned at24ooO ft + 4 sec2. Drogue parachutes deployedreefed at 24 OOO fl + 2 sec3. Drogue parachute single-stagedisreef, 10 sec4. Ma in para chute deployed reefed by way of pi lot Fand drogu e parachutes released at 10ooO fl5. Mai n parachute ini t ial inf lat ion

I

6. Ma in parachute first-stage disreef, 6 sec7 . vhf recovery antennas and flash ing beacondeployed. 8 sec8. Ma in parach ute second-stage d isreef, 10 sec9. Main parachutes released

Figure 17. - Earth landing sequence.

RCS hel ium AscentpropulsionED contro l panel isolation valves compatibil ity valves

Descent propulsion Landi ng gear ED relay box (2)hel ium isolation valve uplock (descentlascent

14 places) stagelNote: ED explosive devices

Figure 18 . - Lunar module pyrotechnics.

Landing-gear operation. - The LMlanding gea r was retracted during tr an s-lunar flight. Before separation of the LMfrom the CSM, detonators in the uplockdevices (fig. 19) were fired to drive a bladethat severed the s tr ap and permitted springsin the deployment mechanism to extend thelanding gear . The design of the landing-ge ar uplock device included two opposingblades to sever the holding st rap. Before aproblem that was uncovered during qualifi-cation testing, both blades were drivensimultaneously by firing both detonators atthe sa me time. The problem w a s that thisprocedure sometimes resulted in the str apbeing "captured" between the two blades and,

Cutter blade-

Strap-

Landing-gearpr imary strutmounting inter face

L t

-Detonatorcartridgepin

Strap housingattachment

71Descent-staymountinginterfaceFigure 1 9 . - Lunar module landing-gearuplock and cutte r assembly.

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because of simultaneous fir ing, caused the blades to meet at the center of travel andreliability of a single blade severing the st ra p already had been demonst rated. Ther e-fo re , the sequence of firing the detonator for each blade was staggered to eliminate theprobability of capture of the st ra p. The second blade acts as a backup if the fi rs t blademalfunctions. Otherwise, the second system is fired only to eliminate the livedetonator.

deflec t one another. Capture of the s tr ap could prevent landing-gear release. The

Pressur izat ion valvAs. - When opened, pyrotechnically actuated valves installedin the LM propulsion sys tem s allowed descent propellant tank pressur ization and vent-ing, ascent propellant tank pressur ization, and R C S propellant tank pressurization.The valves operated instantaneously upon command b y firing self-contained explosivecharges, which provided the nec ess ary impulse for valve functioning. The valves nor -mally were closed to provide complete shutoff of flow, and, after actuation, valveswere opened permanently. The configuration of the valves wa s simil ar to the valveshown in figure 10. Helium-isolation valves, fuel and oxidizer valves, and vent valveswer e the pyrotechnically operated valves in the descent propulsion system. The ascentpropulsion syste m contained helium-isolation valves and fuel and oxidizer compatibilityvalves. Helium-pressurization valves were used in the RCS. Electrically initiatedexplosive cart ridges were used for actuation.

Separation of ascent stage. - Four explosive nuts and bolts wer e used in the inter-stage s tructur al connections, and a n interstage umbilical cut ter was used to s eve r theinters tage elec tri cal umbilical. The separation of the ascent and descent stages in-volved the following steps: (1)operation of the cir cuit inte rrupter (fig. 20) to breakthe interstage electrical circuits, (2) separat ion of the in terstage nuts and bolts(fig. 21), and ( 3 ) severance of the umbilical by the in terstage umbil ical guillotine(fig. 22) .

Ascentconnector

lock detent

disconnectedin 4 (springloaded)IFigure 20. - Lunar module electrical

circuit interrupter.15


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,nitiator r B o l t artridge

Nut cartridge.-(a) Before firing.

Figure 21. - Lunar module

1nitiator

Shea r blade

Detonator

U(a) Cross section.

interstage nut(b) After firing.

and bolt assembly.

(b) Installation.Figure 22 . - Lunar module interstage umbilical guillotine.

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COMPONENT DES I G N AND DEVELOPMENTThe pyrotechnic devices were of pr ime importance fo r flight safety. The safetydesign reliability goal was established to be 0.9999 at the 95 percent confidence level.

The demonstration of such reliability was impractical by direc t testing methods (becauseit would have requir ed demonstra tion of no more than one fai lure in approximately45 000 firings); however, statistical test methods were used to obtain data upon whichacceptable est ima tes could be made. Under conditions simulating the complete missionprofile (including.launch, space flight, and recovery), abbreviated te st progr ams wereconducted to determine that each device tested was functionally safe and reliable.Te st s of the init iato rs were performed in all applicable types of environment andstor age conditions to obtain information on the safety aspec ts or the no-fire capabilitiesof a part icular device. Sensitivity and output tests were conducted to determine thecapabili ties of the initiator s (1) as separate components and as integral pa rt s of otherdevices and (2) as parts of complete systems. Test fixtures and conditions simulatedthe intended applications that incorporated the electrical input sou rc es and char acte ris -ti cs and the output characte ristics. Physical and environmental surroundings weresimulated to assess functional capabilities and requirements.The performance evaluation of each device consis ted mainly of obtaining informa-tion on input and output cha racteris tics. In general, testing of input characteristicsconsisted of sensitivity measurements to determine energy requirements f or sat isfa c-tory fir ing of the device. Testing of output cha racte risti cs consisted of obtaining dataon the physical phenomena that resulted from the firing of the device.during the performance evaluation of the SBASI, sufficient data were obtained to predictthe true all-fire and no-fire current levels for par ticu lar functioning times. Another

example is the reefing line cut ter, f o r which tests were conducted to determine , at vari-ous pull angles, the pull force r equired to trigger the ignition system for satisfacto ryperformance. Also, the reefing line cutter was tested at different temperatures to ob-tain cur ves of the shif t in functioning times.

For example,

Depending on the circums tances, te st s to determine the haza rds associated withthe pyrotechnic devices required special test procedures and equipment. Because mostpyrotechnic devices were initiated electrically and normally would not discriminate be -tween sour ce s of ignition energy, te st s were conducted under various conditions with arange of electrical sources of energy. Test hardware also was subjected to miscella-neous tests that cannot be considered input, output, o r electrical hazards te st s, but wereperf ormed to observe and resolve the effects of different conditions on the devices.Surveillance tests, sealing and moisture-proofing te st s, and vibration and shock tests'wer e included in the development program.

The types of inspection and nondestructive tests common to all units included vi s-ual and X-ray examination of the product, bridgewire and insulation res ist ance tes ts,leakage tests, and neutron radiographic (N-ray) inspections. However, N-ray inspec-tions wer e not performed on the initiator s,Neutron radiography is a relatively new technique. In a number of instances,such as examining the explosive core in an MDF or discontinuities, the N-ray technique

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w a s superior to X-ray. The opacity of the lead sheath and of the explosive core to ther-mal neutrons is the re ve rse of that with X-rays; however, the advantage was lost whenthe M D F was bonded into a chargeholder with a hydrogenous materi al such as epoxy.Therefore, the N-ray technique was applied only selectively to Apollo pyrotechnics tosupplement X-ray examination.Over the past 6 years, all Apollo pyrotechnic devices and syst em s have beentes ted extensively on the ground, in unmanned flights , and in manned flights. Thefollowing is a summary of the evolution of the components and accumulated history fr omall test programs conducted before the Apollo 1 3 mission.

Single Bridgewire Apol lo Standard InitiatorAs stated earlier, a double bridgewire initia tor, the ASI, originally was developedfor use i n the Apollo Program but was replaced by the single bridgewire initiator, theSBASI. More than 20 000 AS1 units were tes ted or used satisfactorily during the space-craft development program. Approximately 7000 SBASZ units have been tested and usedsince. Approximately 140 init iato rs were installed for each Apollo mission. The AS1units were used in all Block I Apollo spacecraft, except fo r the vehicles used fo r theApollo 4 and 6 missions , in which both AS1 and SBASI units were used. All Block 11Apollo spacecraft were equipped with SBASI unit s.Except for having only one bridgewire, the SBASI w a s identical physically to thedual-bridge ASI. The modification to a single bridgewire unit was made primarily be-cause of electri cal sensitivity problems associated with the bridgewire -to-bridgewire

mode,firings were attributed to buildup and discharge of accumulated electrostatic charges.Other recorded data showed that a difference of approximately a 50-volt potential be-tween the bridgewires would cause degradation of the primary charge. The built-inspar k gap w a s not intended to protect against ignition from spurious forms of strayelectri cal energy in the bridgewire-to-bridgewire mode.

In addition to low interbridge electri cal resi stan ce, occasional, inadvertent

The design of the SBASI retained the performance and desirable e lec tric al ch ar -acteristics of the original AS1 and incorporated the following changes.1. To improve the impact resist ance a t tempera ture s below -65" F, the bodymaterial was changed from type 17-4 P H st ee l to Inconel 718.2 . To increa se internal pr es su re capability, the w a l l thickness surrounding thecharge cup w a s increased, the shape of the header was modified, and the header w a swelded to the body.3 . To ensure that the pins would remain secu rely fixed under high internal pr es -sures, the header material w a s changed from ce rami c to Inconel 718, and the pins wereglassed to the header instead of being brazed to a plating materi al coating the ceramic.4. To protect against environmental contamination (i . e. , humidity, air densitychanges, and dust particles), the spar k gap location was changed to the interior of theunit. The sp ar k gap w a s required to prevent inadvertent firing fr om extraneous high-voltage discharges by diverting discharge to ground.

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A design deficiency that remained with the SBASI involved the built-in sp ar k gapand w a s associated with the breakdown of insulation res istance . Contamination couldbe unintentionally introduced in the spark gap during manufacturing operat ions, andsome units were rejected because of insulation resistance failu re. However, unitsexhibiting resistance failure were rejected on a n individual basis.

During development of the SBASI, the body-header assembly was tested hydro-statically aft er r epeated thermal shocks, ranging from - 3 2 0 " to 500" F, to over100 000 ps i without failure. During production, all units were tested to 40 000 psi.All production units also were tested for electrostatic survival capability to withstand2 5 000 volts (from pin to case) and were leak tested with helium to ensu re proper he r-.metic sealing.extensively to ensure complete interchangeability. Sectioned and exploded views of theSBASI a r e shown in figure 2.

The AS1 and SBASI units produced by the two manufacturers were tested

Indexing the connector end of the SBASI after manufacture permit ted manufactureand stocking of a standard unit. Indexing of the units was accomplished as needed bystaking the ba rr el s to meet a specific keyway combination (fig. 23). Nine specia l key-way combinations were used to meet special requi rements.

Keyway

7numbersMasterXX1 XX2 XX3 XX4 XX5 XX6 XX7 XXB XX9 (XXOl

Basic part Dashnumber numberA -SEB26100001 - 1XXSEB26100001 - ZXXSEB26100001 - BXXSEB261MXX)l - X1XSEB26100001- X5XSEBZ6100001- X X lSEBZ6100001- XXZSEB26100001- XX3SEBZ6100001- XX4SEB26100001- XX5SEB26100001- XX6SEB26100001 - XX7SEB26100001 - XX8SEB26100001 - XX9SEB26100001 - XXO

Designation

Prototype, developmental experimentalFlight configuration, qualified orIn er t deviceWeld washer attachedNo weld washerKeyways 1and 6 closed (nonfligh t)Keyways 2 and 6 closed (flight)Keyways 3 and 6 closed (flig ht)Keyways 4 and 6 closed (fli ght)Keyways 5 and 6 closed (flight)Keyways 1and 2 closed (flight)Keyways 1and 3 closed (flight)Keyways 1and 4 closed (flight)Keyways 1and 5 closed Iflight)Al l keyways open (required specific

MSC authori ty for use)

qualifiable

Staking w a s accomplished by blockingkeyways. The keyways were blocked bycrimping the outer lip of the keyway inwardto a dimension tha t prevented the matingconnector key from entering the slot. A f t e rqualification of the staking procedure, it w a sfound that improper cr im ps could resu ltfrom the use of worn crimping tools. Theresultant improper staking in the electricalconnector portion of the SBASI allowed amating electrical plug of different key con-figuration to be connected.

To prevent the use of crimping toolsworn beyond dimensional tolerances , man-datory dimensional checks of the crimpingtool were established. Dimensional checksalso were established on the initiator itselfand were ver ified by using a comparator fo racceptance inspection.

Figure 23. - Initiator indexing.

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C a r t r i d g e A s s e m b l ie sCartridges of various siz es and configuration wer e used in the following applica-

tions: (1) actuation of electr ical circuit int erru pter s and disconnects, ( 2 ) operation ofthrusters, ( 3 ) deployment of parachutes, (4) peration of valves, and (5) as componentpa rt s of separation systems. Most cart ridg es were simil ar in construction, but dif-fered in thread size and type and in amount of output charge. Figure 24 i s a sketch ofa typical electrically initiated cartridge assembly.various cartridges, including detonator cartridges. Figure 25 is a photograph of the

Figure 24. - Electrically initiatedcartridge. -Figure 25 . - Spacecraft cartridges.All b u t one type of car tr idge were initiated elect rically by an SBASI. The onlynonelectric cartridge (fig. 26) was initiated by CDC and was used to oper ate the SLA-

panel thruster s. By adding booster modules containing var ious char ges, special pur -pose cartr idge s were obtained. The physical configuration, performance, and us es ofall cartridge and detonator assemblies a re listed in table I. The SBASI is included intable I because the unit was used alone as a pr ess ur e cartridge in the docking proberetract ion system and in the docking tunnel vent valve. The indexing of the SBASI inthe cartridges is shown in figure 23. Car tr idges with different outputs had differ entthreads to prevent improper installation.close to each other in the spacecraft and fir ed at different times, were indexed differ-ently. Thus, the same thread and indexing could be used in various locations on thespacecraft.

Cart ridges with the same output, but located

Each cartridge assembly (except the SLA-panel thruster cartridge, which w a sfi red by CDC) consisted of one o r two SBASI units hermetically sea led to a cart ridgebody by a weld washer.cartridge contained no charge other than that in the SBASI; the cartridge module w a san adapter necessary to install the SBASI in a sma ll explosive valve, Approximately250 units of each type of car tr idge assembly were f ir ed during qualification te stprograms.

The weld washe r is shown in figur e 2 7 . The type 100 pressure

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Output charge containing boronand potassium nitrate to providegas pressurer 7 nitiatina train"1entaerythritol tetranitrate

used to transfer chargethrough bulkhead

Figure 26 . - Spacecraft/lunar moduleadapter thruster cartridge.

Torquing

(welded undertorquing section)

Figure 27 . - Initiator with weld washer.

DetonatorsThree types of detonator car tr idges were used. The Apollo standard detonator(ASD) (fig. 28) w a s used in the following locations: the tower -separat ion frangible nut,the CM/SM guillotine, the tension-tie cutter, and the lower LM guillotine. The ASDw a s used fo r SLA-panel separat ion and for LM/SLA separation, The ASD also w a sused to initiate up to 20 grains/ft of MDF or 100 grains/ft of LSC. The ASD consisted

of a modular assembly of the initiator with a primer, an intermediate charge of leadazide, and an output charge of RDX.

Initiator\

.Initiator

Ceramic header - \

Primary E

SpacerLead azide

(a) Detonator cart ridge assembly. (b) Cross section.Figure 28. - Apollo standard detonator.

!xplosive charge

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The end-detonating cartridge (EDC)was developed for high-temperature appli-cations where a directional shock w a s re-quired for the initiation of high-explosiveelements . The EDC had a heavy wallthickness along the threaded length of theoutput end (fig. 29) . The EDC was used onthe LM guillotine and on the landing-gearuplock. The EDC had an intermediatecharge of lead azide and an output chargeof HNS.

The long-reach detonator (LRD)(fig. 30) was used in the docking ring'as-sembly only. The configuration was nec-es sa ry to extend the output charge to aninterface are a that was inaccessible witheither the ASD o r the EDC. The LRD hadan intermediate charge of lead azide andan output charge of HNS.

A significant problem, which wasrelevant to lack of manufacturing controlrather than to design, ar os e on July 30,1969. Two of four detonators tested in LMguillotine lot acceptance tes ts failed tofire high order. Results of the failure an-alysis showed that the failures werecaused by alcohol contamination and thatthi s contamination apparently was isolatedto only one lot of detonators. This con-tamination was confirmed by both N - r a yinspection and ma ss spect rometric analy-sis. Spectrometric analysis of one suspectunit showed it to have an alcohol content of18. 3 microliters. Further investigation in-dicated that the threshold could be below6 . 7 5 microlite rs and that the N - r a y tech-nique could not be used alone to determinewhether units contain alcohol at thi s level.

In addition to the requirement thatno alcohols or solvents be permitted inloading rooms , corrective action requiredthat two units from each lot be selected atrandom for mass spectrometric analysis.Total volatiles, excluding water, couldnot exceed 0 . 0 4 0 microliter per unit.This value was arrived at b y sampling anormal lot of detonators and determiningthe level of contamination, which turnedout to be 0 . 0 2 to 0.03 microliter.

f = = L

Figure 29 . - End-detonatingcartridge.

r 0 . 8 1 2 i n . ( r e f l y

- ea dazide-H N S

-lead a z i d e

\ c

Figure'30. - Long-reach detonator.22


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Core ChargesThree types of co re charges wer e used (fig. 31) . An LSC w a s used to s eve r thetension ties connecting the SM and the CM. An MDF was used to s ev er the forward

and aft circumferential and the inner and outer longitudinal splice plates of the SLAand for separation of the docking ring from the CM . The MDF also w a s used as theexplosive element in the guillotines for driving the blade that cut the umbilicals . ACDC was used fo r detonation tra nsf er between various points in the SLA separationsystem. Core-charge assem blies and their uses are listed in table 11.

Sheath

Explosive MDFcore

(a) Mild detonating fuse. (b) Confined detonating cord.

(c) Linear-shaped charge.Figure 31. - Core charges.

The principal explosive used as the co re explosive was RDX; however, in theCM-to-LM docking tunnel and in the LM interstage guillotine, where tempera ture re -quirements were more seve re, a heat- and vacuum-resistant explosive (HNS) as used.In most applications, booster charges containing lead azide were attached t o both endsof the linear charg e to ensure detonation tra ns fe r. In the SLA, in addition to theboosters on the ends of the linear charge, separate boost ers were employed for addi-tional assuranc e of reliable detonation transf er.

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Approximately 100 sepa rate explosive tr ai n assemblies incorporating linearcharges were installed on each spac ecra ft. The assembl ies were developed duringqualification of the components containing linear charges. More than 2500 fee t of linearchar ges were tested satisfactorily during component qualification.Problems re lated to the us e of co re cha rges that occur red during the Apollo Pro-gr am include the following.Tension-tie cutter. - The tension-tie-cutter co re char ge consisted of an LSC oflOO-grain/ft RDX in lead sheath. Boosters were attached to the ends to effect reliabledetonation transfe r f ro m detonator to cord. During the the rmal vacuum exposure por-tion of the qualification program, blistering of the booster charges and bulging of theV-angles to out-of-specification values occurred af ter exposure to a tempera ture of120" F fo r 72 hours. This phenomenon had been anticipated during development becauseknowledge existed concerning simil ar problems in other military and space programs ;however, because the expected blistering or bulging did not occur during the develop-ment program, no measure s were taken in the design of the syst em to ensure str uc tu ra lintegrity when exposed to elevated temperatures.After the anomaly occurred, a program w a s conducted to determine whether deg-radation had occu rred as a res ul t of impuri ties in the explosives o r if other s ou rc es ofcontamination were responsible for the bulging and bliste ring. Results of analysesindicated that no degradation of the explosive products had occurred and that no contam-inant w a s introduced into the system to cause the angle change or the blistering. A l lte st s conducted to determine the reason for angle change and blister ing proved incon-clusive. However, t es ts conducted on "worst parts" showed breakage of 120-percentplates.Lack of positive information concerning bulging and blister ing was sufficientjustification for modifying the design, even though it had been demonstrated tha t even

"worst case" units were capable of severing the ti es . The mate ria l for end ca ps , whichoriginally were made of preformed lead soldered to the LSC, was changed to aluminumwith epoxy adhesive for mating to the LSC. This change in design eliminated the proba-bility of blistered end caps. To control V-angle bulging, the boostered LSC's wereinspected and selected only after being subjected to a se ri es of temper ature screeningcycles; the acceptance criter ion specified that all charges selected maintain a 96" f 2 "angle during and after exposure to all temperature cycles.Spacecraft/LM adapter. - The SLA core charges consisted of MDF's of 5- andT-grain/ft RDX in lead sheath. During disassembly of charges fro m the chargeholdersin tests conducted at MSC, se ve re degradation of the lead sheath w a s observed f or thatportion of the MDF in contact with the RTV adhesive used fo r bonding the cha rge to the

holder.

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32/46

anomaly occurred because of primer header assembly blowback during ignition andthat the phenomenon actually had occur red on other units as well, although previousoccurrences were not obse rved visually. Review of X-rays taken of the ent ire lotbefore functioning revealed that the genera l quality of manufacture, assembly, and in-spection were f a r below par.A s a resul t of analyses conducted, the following determinations were made.1. A new lot of M42 pr im er s was used in the assembly of the cutters. Untilthen, a single lot of p r ime rs had been used in the assembly of all cutters.2. Because of the difference in output of the new lot of primers, a change indelay column height was made to accommodate the change in primer output pressure.3. The change in pr ime rs and in delay column height resulted in stackup oftolerances, which w a s cause for a less than satisfactory crimp of cri tical interfaces,Although all requir ements fo r function time, reefing line severance, and requir edpull force for initiation were met, the lot w a s rejected because of the header blowback

anomaly, attributed to the substandard quality of workmanship coupled with the higherpr es su re output of the primer. Better quality control me as ur es for correct ion of theproblem included conducting appropria te t es ts for making determinat ions upon whichcri ter ia would be based for the use of certain p r ime r s in specific applications,

GROUN D-SUPPORT EQUl PMENTGround-support equipment (GSE) w a s used to se rv ic e and check out spacecraf tsy st ems before flight.la to rs and spacecraft verification equipment,The pyrotechnic checkout equipment included pyrotechnic simu-

P y r o te c h n i c S i m u l a t o r sThe CSM nitiator simulator. - The CSM initiator simulator w a s an electrome-chanical device used in lieu of actual hot bridgewire ini tiators during s ys te ms andintegrated tes ts of the C M and the SM. The device (fig. 32)w a s a suitcase-typeenclosure that housed six bridgewire simulator circuit s. Each circuit w a s composed

of relays, diodes, fixed and variable re si st or s, and tes t jacks in a combined networkand w a s joined to external connectors (fig, 33). The unit was portable and w a s designedto be ignition-proof for use in hazardous areas. The case assembly was constructedwith a f i l l valve to pres sur ize the case wi th an iner t ga s for ignition-proofing purposes.The case measured 12 by 10 by 7 inches and weighed 20 pounds. The power re qu ir e-ments were 28 volts direct current (dc) at 0.7 ampere,The unit contained six circuits capable of substituting electrically for six pyro-technic initiator bridgewires and provided a simulated resu lt of initiator bridgewireaction.cuit to simulate bridgewire burnout while protecting spacec raf t circui try from over-cur re nt drain. A l l units were capable of being r es et remotely,

The input resistance to each simulator w a s 1 ohm, which became an open ci r -Signal input originated

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Spacecraftsignal inoutKey:

K, Go and reset relay

Relief valveFill valve

Bolt W(a) Front view.

I PI I 1 1

(b) Rear view.

. , . KZ Transient relayK Time-delay relayK4 Transient holding relay

28 V dc reset

-I

1

28 V dc return

TransienoutputGo output

Figure 33. - Schematic of CSMinitiator simulator.

Figure 32. - Command and servicemodule initiator simulator.

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and was controlled by the spacecraft sequencers. The input w a s detectable by eachsubcircu it in th re e modes: an all -f ir e condition, positive and negative transient condi-tions, and a no-f ire condition. Output signals from the device wer e transmi tted throughan interface junction, the initiator st imul i unit, to the acceptance checkout equipment(ACE) consoles, which provided di sc re te light indications. A transient output signalof 6 volts dc w as generated when a signal input voltage w a s gr eat er than 0 .17 volt(positive or negative polarity) but w a s of insufficient magnitude to initiate a "go" outputsignal. The go output signal voltage w a s also 6 volts dc when an input cu rrent level of5 amperes o r greater w a s applied for 10 milliseconds. The 6-volt output signalseventually controlled the go/no-go lamps on the ACE console.

Many problems occurred dur ing the evolution of the unit. Examples of suchproblems ar e listed here to point out a re as where special attention should be given onany future work.1. The simulator must provide a 1-ohm load to the spacecraft circuitry andmust become an open circuit after a specified curr ent pulse has been received. If thedevice failed to become an open circuit , the spacecraf t ci rcu it ry would be damaged byextended current drain. The circuit ry was damaged on th re e spacecr aft because theoriginal design required facility power to operate the circuit-opening relay. In eachcase, the damage w a s caused by a lo ss of facility power. The present Apollo designincludes fail-saf e requirements and redundant circuit-opening features. The fai l-saferequirement w a s met by a relay circuit design that required facility power to completethe 1-ohm spacecraf t load. Loss of facility power would result in an open-circuit load,which would protect t he vehicle ci rcuit s. Redundant circuit opening w a s accomplishedby using two separate relays with contacts wired in series so that fa ilure of a relaywould still allow the 1-ohm load to become an open circuit in the specified time and,thus, prevent damage to the spacecraft cir cuit.2. The devices, as originally designed, gave fa lse no-go indications becausethe open circuit did not occur as fast as required. Even though the lag did not damagespacecra ft wiring, overlapping of the time-sequenced functions resulted in higher thannormal loads on the spacecraf t circuitry. To co rr ec t thi s problem, a relay and ca-paci tor combination that would open a t the proper time sequence to prevent overlappingof functions w a s added in se r ies with the 1-ohm load circu it.3 . In one early design, cur ren t flow w a s significantly longer than planned. Thissituation occurred because tr ansient actuation did not result in the simulation of anopen circui t when the spacecraft bus voltage was low. With no open circuit, the space -craft circuitry could be damaged. To co rr ec t thi s problem, the transien t detectingrelay was wired s o that either it o r the go relay could open the circuit.4. The time response of the transient detecto r was not fast enough for detectionof short-duration pulses. With very high current levels of very shor t duration, theinitiator actually could be fired without detection of a curr ent pulse by the simulator.The fa st response time could not be obtained by using rel ays . A change to solid-statedetection would not have supported the schedule fo r the CSM; therefore, a "workaround"procedure that consisted of monitoring the 1-ohm load resis tance with an oscilloscopew a s used.

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The LM initiator simulator. - The L M initiator simulator (fig. 34)w a s a portableunit used to check out the function of electrical systems by simulating the electricalcharact eristics of the initia tor. The simulator also was capable of detecting trans ien tso r interference signals at the initiator connector.The equipment consisted of 40 circuit -monitor modules having dimensions of

5 by 3 by 3 inches. Each module weighed approximately 2 pounds and was hermet icallysealed. The modules were stored in a car rying case capable of accommodating45 modules. The car rying case w a s a suitcase-type enclosure with dimensions of30 by 30 by 30 inches and weighed 10 pounds. The power requirements were 28 voltsdc at 9 . 0 amperes.

rr

Approximately 34 individual simulatorsI III

ehicle under testJ'-tI

II-

1!Digital signal Iconditioner and ImultiplexerI I IIII

ACE carry-on equipment

IIACE carry-on equipment I

III

lntelieaverTo AC E station

Figure 34. Functional block diagramof LM initiator simulator.

Each module provided a 1-ohm inputto simulate the bridgewire resistance ofeach initiator. The 1-ohm input became anopen circuit after receiving either an all-fire or a transien t signal; a curren t of7 .5 amperes or grea ter resulted in a goindication; and a 200-milliampere orgreater current, with a duration of 5 micro-seconds or longer, would result in a t ran-sien t indication. The output signals wereprocessed through the ACE to the go andtransient lamps located on ACE consoles.Solid-state electronic devices were usedin each module to obtain the faster responsetimes.The first design w a s a solid-state

device but had th re e major undesirablecharacteristics: (1) The unit did not haveredundant capability to obtain an open cir-cuit after firing and, thus, prevent damageto spacecra ft firing circuit s. The redun-dancy w a s accomplished later by providingse ri es redundant relay contacts actuated byboth the go and transient circuits. (2) Useof a common power supply without isolationbetween modules caused sneak circuits andground loops that resulted in false outputindications. These problems were solvedby providing isolation resistors betweeneach simulator module and the power sup-ply. (3) The unit could be actuated pre-maturely by EMI. This problem wassolved by incorporation of EM1 fil ters in thepower supply input circuit and by the use ofshielded cables on all input and output leads.

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The environmental initiator simulator. - The environmental initiator simulato r(fig. 35) was developed fo r us e during vacuum-chamber testing of the spacecra ft. Thedevice was an iner t SBASI with a normal bridgewire. Each device w a s hermeticallysealed fo r operation in an environment containing hazardous propellant gases. Theinitiator was sealed by welding the unit into a stainles s-steel block; thus, inadvertentinstal lation of an iner t unit in a flight vehicle w a s prevented. Seventy-five of the inertdevices were packaged in a portable carrying case.

The inert SBASI provided an exact e lect rical simulation of the flight initiator be-cause a flight-type bridgewire w a s used. When a sufficient spacecraft firing currentw a s received, the bridgewire was burned and a n open circui t resulted. Because no ex-ternal output indicators were used. each device had to be inspected with an ohmmeterafter Spacecraft testing w a s completed.- W e l d washer7 onnector

Figure 35. - Environmental initiatorsimulator.

7r-pacecraft connector FuseholderFigure 36. - Static EM1 device.

The static EM1 device. - The staticEM1 device w a s developed fo r use duringone of the spacecraft sys te ms verificationtests in which the GSE was not connectedto the spacecraf t (plugs out). The devicewas used for passive monitoring of theordnance electrical system to ensure thatinterf erence from other sy st em s would notaffect the ordnance undesirably. Al/lS-ampere fuse in the device w a s mateddirect ly with spacecraft wiring.

Each device (fig. 36) w a s containedin a cylindrical module 2-3/4 inches longwith a diameter of 1 inch. The device hada spacecraft mating connector on one endand a fuseholder on the othe r end. The de-vices were transported and stored in aportable container having a capacity of43 units. Because the unit had no externalread-out capability, an ohmmeter inspectionhad to be made after the plugs-out testingw a s completed. The device was not her-metical ly sealed and could be used onlyfo r ambient environmental testing.

During use in the Apollo P ro gr am ,two problems were encountered. Some ofthe units were found to have actuated pre-maturely, and others failed to actuate tr.ien exposed to sufficient actuation current . In-vestigation showed that the units were actuated prematurely during checkout by the us eof a non-current-limiting ohmmeter. The output current w a s sufficient to actuate theunit. This problem was correc ted by requiring the use of current-limiting ohmmeters.Investigation of the units that failed to actuate showed that the problem w a s caused byuse of improper fuses. The module required plug-in-type fus es . Some of the fuse sused were of the pigtail type, which had longer leads. These leads were cut to theapproximate length of the plug-in type. Cutting the leads resulted in a flattened pin,

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which caused improper contact with the fuseholder. This problem was cor rec ted byrequiring the use of proper fuses.The str ay elect rical energy indicator. - The stray elect rical energy indicator w a sused to simulate the electr ica l cha rac ter ist ics of the initiator while detecting excessive

level s of RFI on the ordnance electr ica l system of the spacecraft . The st ra y electr ica lenergy indicators were installed in lieu of flight initiators on the totally stacked vehicleat the launch pad. All onboard and ground radio-frequency (rf)emi tte rs were energizedto verify R F I compatibility.The str ay ele ctric al energy indicator (fig. 37) consisted of a specia l SBASI thatwas screwed into an adapter. The unit w a s 3 inches long with a diameter of 1 inch.

The indicator consisted of a bellows that expanded to give visual indication of unit actu-ation. The special SBASI contained a more sensitive bridgewire (higher resistance)and an rf-sensitive ignition explosive. The unit was hermetically sealed , even afterfiring. When the initi ator was fir ed, the plunger in the indicator w a s forced outward toexpand the metal indicator bellows. The unit connector had all keyways left open sothat the unit could mate directly with any spacecraft ordnance connector.

Considerable effort w a s expended in the development of an acceptable RFI detectorbecause of the problem of obtaining a proper impedance simulation ac ro ss the frequencyspectrum. The use of an actual explosive device w a s the only solution.

0. 3 i n . H I- in . I

Figure 37. - Stray electri cal energyindicator.

7C o n n e c t o r pins Photoflash bulb

Figure 38. - High-energy-initiatorsimulator.

The high-energy -initiator simulator. -The high -energy-initiator simulator, also adevice to simulate the initiator duringground testing, presented the spacecra ftwith electrical characteristics sim ila r tothe SBASI. This unit w a s used during check-out of all flight vehicles at KSC, includingboth the CSM and the LM. The device w a sused during performance of simulated mi s-sions, including the altitude-chamber runs.A visual indication of status w a s requiredwithout an interface with the chamber wiring.This requirement, along with the high cost ofthe environmental simulator, prevented useof either the LM init iator simula tor o r theenvironmental simulator, described ea rl ie r,in this application. Each high -energy-initiator simulator (fig. 38) was 3 incheslong with a diameter of 0 . 7 5 inch and weighedle ss than one-quarte r pound. The devicew a s hermetically sealed and could be usedin vacuum-chamber testing and in hazardousgas areas, The carrying case was capableof accommodating 100 units. The device wasmated directly with the spacec raft initiatorconnector and provided a resistance of1 . 0 5 ohms before firing. When the device

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received a firing current, an open circui t re sulted within a time interval that wasdependent upon the curr ent level through it.The device consisted of a fuse in series with a 1-ohm resi sto r; a second resistorand a photoflash bulb were i n para lle l with the series circuit. The second re si st or wasadjusted to a value that caused the bulb to flash when the reqyired firing cu rren t levelwas reached. Temperature-sensitive paper w a s placed around the bulb to provide a,visual indication that either an all-fire o r a no-fire cur ren t level was reached.The only fai lur es with the device involved flashbulbs that did not function. Space-craft system fai lur es were distinguished fr o m bulb failu res by verifying that the simu-lated bridgewire (fuse and res is to r) was in an open-circuit condition after t es ts . Uponconsulting the flashbulb manufacturer, it w a s determined that the incidence of fa ilu rewas well within that to be expected and no reasonable correcti ve action could be taken.

S p a c ec ra ft V e r i f i c a t i o n Equ m e nThe pyrotechnic checkout test set. - The pyrotechnic checkout test set was a port-

able, self-contained unit that was used for the following: (1) to detect and measu restray voltages in the spacecraft pyrotechnic circ uits , (2) o measure resistance and toverify continuity of these circuits, and (3) o measure the resistance of the installedpyrotechnic initiators . The test set (fig. 39), which was approximately 11 by 8 b y 8 in-ches in size, had carrying handles and a removable top cover. Included with the testset were 20 dummy ini tiators that were integrated w i t h the pyrotechnic circuitry in

(a) Dummy initia tor. (b) Portable measuring device.Figure 39. - Pyrotechnic checkout test set.

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place of actua l pyrotechnic initiators. The test se t and dummy initiators weighedapproximately 30 pounds. The test set w a s used to verify the pyrotechnic circuitsbefore, during, and af te r installation in the spacecraft.The bridge circuit consisted of a resistance bridge of the Wheatstone type, a

was dual range from 0 to 2 . 9 9 9 ohms and from 0 to 2 9 . 9 9 ohms and included a four-digit selector -indicator f or balancing the circuit and fo r displaying the value of themeasured resistance. Accuracy of the bridge circuit was *O. 15 percent. Theindicator-amplifier portion of the circu it included a null m et er with a zero-center scale.The alternating -current (ac)/dc/ohm circuit consisted of a meter -amplified circuit anda dual-deck rotary switch. The ac/dc/ohm circuit provided the following functions:(1) selection and measurement of both positive and negative dc voltages, (2) selectionand measurement of a c voltages at frequencies as high as 20 000 hertz, and (3) a meansfo r performing continuity tests. The meter-amplifier portion of the circuit included anindicator with dual linear scales of 0 to 10 millivolts and 0 to 100 millivolts and a 0- to5-ohm scale with midscale reading of 1 ohm. Power fo r the bridge circuit and for theac/dc/ohm circuit w a s provided by internal batteries. The se t included 22 two-pindummy initiators. The dummy initiators consisted of a potentiometer that could beadjusted between 0 .95 and 1.15 ohms. The original unit, using the bas ic Wheatstonebridge concept, could not meet the specifications fo r res ist ance measuring accuracy.The requirements were met by including a dc amplifier between the bridge circuit andthe null meter. This arrangement increased the null indication accuracy and, thereby,the accuracy of the measuring unit.

bridge indicator -amplifier, and a 24 -position rotary switch. The resistance bridge

The pyrotechnic bridge checkout unit. - The pyrotechnic bridge checkout unit(fig. 40) w a s used to perform the sa me t y p e resistance measurements as the pyrotechnic

(a) Remote control assembly. (b) Bridge unit.Figure 40 . - Pyrotechnic bridge checkout unit.

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checkout test s e t and, in addition, could be controlled remote ly at the launch sit e. Theunit w a s used to check the circuit resistance of the launch-escape motor, of the tower-jett ison motor, and of the SLA separation system, because these circuits were consid-er ed to be potentially hazardous to personnel in the a rea when fired. In size, thepyrotechnic initiator bridge unit was 1 2 by 1 2 by 18 inches, the remote control panelassembly was 5.5 by 6 by 19 inches, and the dummy was 1. 5 by 1 .5 by 4 inches. Thesystem weighed approximately 100 pounds and was designed to be portable.

The explosive device tes t set. - The explosive device tes t s et w a s used for no-voltage tests and electrical resistance measurements of the LM pyrotechnic circuitrybefore and after installation of flight ordnance. The test set (fig. 41) was a suitcase-type configuration that contained a resistance measuring unit, a 500-volt megohmmeter,and the associated batteries and switches. A cable set w a s provided to effect an inter-face with the LM vehicle. The unit measured 20 by 10 by 10 inches and weighed51 pounds.

Figure 41. - Explosive device test set.

The resistance measuring portion ofthe test se t w a s used to measure and dis -play the circuit resistance of the LMpyrotechnic circuitry. A removable, panel-mounted megohmmeter w a s used to meas ureand display the isolation res ist ance of theLM pyrotechnic circuitry . The power sup-ply provided the necessary power to operaterelays in the relay box. The power supplyiucluded current-limiting and protectivedevices to prevent high curr en ts f ro m flow-ing in the pyrotechnic circuitry i f a failureoccur red. The monitor and control panelwas capable of the following functions:(1) the selection of subassemblies , specificstimuli, and monitoring points; (2) mechani -cal mounting of the resis tance measur ingunit and the megohmmeter ; and (3) visualdisplays of a c and dc voltages, of pyrotech-nic battery voltage, of in ternal power sup-ply voltage, of battery polarity, and of thest at e of gaseous purge when pressurized fo ruse in hazardous gaseous a rea s.

The current regulator. - The current regulator was used in conjunction with theexplosive device tes t s et to limit the spacecraft power bus cur ren t to a safe leve l duringpyrotechnic circu it testing afte r the flight pyrotechnics were installed in the vehicle.The unit (fig. 42) was portable, had dimensions of 18 by 12 by 12 inches, and weighed15 pounds. Mating connectors were supplied both for the LM elect rical power bus andfor the ground power. The cur ren t regulator controlled the output power of a 28-voltdc, 250-ampere power supply. In the bypass mode, the regulator w a s capable Ofcarrying and breaking a current of 150 am pe re s at 28 volts dc. In the current-limitingmode, the minimum output current w a s 100 milliamperes. The limiting curr en t levelw a s adjustable continuously fr om 300 to 500 milliamperes, When the limiting cu rr en tlevel w a s reached, the circu its were opened. The circui ts remained open until the34


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CO NCLUSI O NSThe pyrotechnic device offered a convenient means to the spacecraf t engineer fo raccomplishing numerous functions in the space cra ft systems. The la rg e number andvariety of critical functions performed by Apollo spacecraft pyrotechnic devices im -

posed an obligation to provide the most reliab le and the sa fes t sy stem devised f o r aero-space applications to that point in time.each pyrotechnic component as an integral and critic al pa rt of an overall system and bythen conducting appropriate design reviews and evaluations to ensure that every con-ceivable variable in the sys tem was controlled. The magnitude and complexity of prob-le ms were minimized by using conventional electrical and mechanical componentswherever possible. New and unfamiliar requirements necessitat ed experimental andanalytical studies that, in some cases, uncovered a number of potentially seriousproblems, which, if not uncovered at an opportune time, would have result ed in costlyrework and a se ver e st ra in on space program objectives. An outstanding example isthe incompatibility problem associated with the use of RTV 30-121 adhesive to bondlead-sheath explosive -core charges in the spacecraft/lunar module adapte r charge -holders . Had the problem not been uncovered during almos t routine in-house te st s, itis possible t ha t the condition could have resulted in failure to separate during a mission.Other examples include the docking tunnel separation system redundancy question andthe Apollo standard initia tor sensitivity problem. Costly rework or redesign (or both)of the docking ring separation system w a s avoided because res ul ts of analytical testsand experiments proved the system to be reliable and redundant with only relativelyminor corrections. As a res ult of exhaustive evaluation programs to sea rc h out hiddenweaknesses in the Apollo standard initiator design, the possibilities of both a prematurefire and a misfir e occurr ing during flight were uncovered.

This obligation was met by initially considering

Although there is no question as to the suitability, quality, and reliability ofApollo pyrotechnic devices for applications in Apollo sys tems, and the des irabili ty ofusing devices having high reliability is not open to question, the methods employed toattain and to demonstrate confidence in reliability for a number of Apollo devices werenot the most efficient o r the most practical . As shown in the examples, most Apollopyrotechnic devices did not provide for reliability control o r monitoring before theevaluation phase, and design information was genera ted largely through intuition rat he rthan through a combination of scientific approach, stat is tic al experimentat ion, andsubsequent observation. As a result, in a number of cases, a considerable amountof hardware that w a s orde red o r acqu ired could not be used when the design was laterfound to be deficient, usually at a time which was too late to make modifications eco-nomically o r to redesign for reliability improvement. Often, the funds already spentwere too great to be written off, and it w a s considered to be mor e expedient to addfunds in an attempt to make a product improvement. Another reason fo r the hardwareproblems was that development time w a s not available fo r redesign.

Although the reliability probability of each device w a s established at design con-ception, the evaluation, qualification, and production acceptance tests and the flightuse of the devices contributed to confidence in the rel iabi lity of the design. This con-fidence w a s enhanced further by following the policy of standardization of componentswherever possible.

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Within the pyrotechnics area, the most outstanding example of the potentialitiesin following the policy of standardization of components is the exclusive use by theManned Spacecraft Center of the standard initia tor in all Apollo spacecraft applicationsrequiring elect ric al initiation. Maintaining standardization requi red considerableeffort on a continuing basis but proved worthwhile in enhanced system reliability andreduction of tota l program cost, as a result of eliminating multiple initiator develop-ment programs.

By st r ic t adherence to the philosophy of concentrated attention to all det ail s ofdesign, including iner t and explosive mat er ia ls used in the construction of all compo-nents and systems , the Apollo spacecraft pyrotechnic sy stems played a major and veryimportant role in the resultant highly successf ul Apollo Program.

RECOMMENDATIONSFuture manned-space -flight technology will place a greater demand on extremelysophisticated, rel iab le, compact, and lightweight pyrotechnic devices,purpose of a reliability program for future systems should be to ensure that reliabilityis treated as a design par ameter of equal importance with performance pa ramete rsduring the initial design-conception phase as w e l l a s during the evaluation of a proto-type design. The approach to design and development must be more fundamental thantha t taken during development of Apollo pyrotechnic devices, where the "cut and try"method was used in most instances. Adequate information is available and should beused to assist the design engineer in the selection of p roper m ater ial s in'any givendevelopment program. Pr oper knowledge of the candidate materials and components,of the ir behavior and charac ter ist ics , and of the specific requirements that the end

item must meet should be acquired before design can begin. J u s t as important are themanufacturing, handling, and testing of these components and mat er ia ls , because var i-abl es can have a drama tic effect on the performance and functioning of the end device.

The primary

Qualification of safe, reliable pyrotechnic systems for future programs can comeabout more efficiently and effectively i f , during initial design and development, detail edand in-depth knowledge is obtained of all aspects of the design encompassing a fullcr os s section of the capabi lities as wel l as the limitations of the sys tem. This "keepingabreas t" can be accomplished most effectively by maintaining close control of proc-esses, equipment, and facili ties used to manufacture and qualify any specific item.Unfortunately, it w a s not until the la tt er par t of the Apollo device development programsthat mea s

apollo experience report spacecraft pyrotechnic systems

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