Apollo Experience Report Development Flight Instrumentation

8/8/2019 Apollo Experience Report Development Flight Instrumentation

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N A S A T E C H N I C A L N O T E NASATN D-7598

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BRIENCE REPORT -T FLIGHT INSTRUMENTATION

Lyndon B. Johnson Space CenterHouston, Texas 77058

NATI ONA L AERONAUTICS AND SPACE ADMINISTRATIO N WASHINGTON, D.C. MARCH 1974


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CONTENTS

Section

SUMMARY . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .INTRODUCTION . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .INITIAL PROJECT CONSIDERATIONS . . . . . . . . . . . . . . . . . . . .

DESIGN CONSIDERATIONS . . . . . . . . . . . . . . . . . . . . . . . . . .FLIGHT SYSTEM BREADBOARDING . . . . . . . . . . . . . . . . . . . . .

ACCEPTANCE AND QUALIFICATION TESTING . . . . . . . . . . . . . . .

COMPONENT FAILURES . . . . . . . . . . . . . . . . . . . . . . . . . . .SPECIAL SYSTEMS . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .SCHEDULES AND MANPOWER LEVELS . . . . . . . . . . . . . . . . . . .PROBLEM AREAS . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Tape Recorders . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .Commutators . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .Vehicle System Cabling . . . . . . . . . . . . . . . . . . . . . . . . . .

Handling Damage . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

CONCLUDING REMARKS . . . . . . . . . . . . . . . . . . . . . . . . . . .APPENDIX A- BOILERPLATE PHASE . . . . . . . . . . . . . . . . . . .

APPENDIX B - OMMAND AND SERVICE MODULE PHASE . . . . . . .APPENDIX C - UNAR MODULE PHASE . . . . . . . . . . . . . . . . . .

APPENDIX D - ANNED FLIGHT VEHICLE PHASE . . . . . . . . . . . .

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TABLES

Table Page

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A-I

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A-III

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D-III

DFI EQUIPMENT FAILURES AFTER DELIVERY TO

PRIME CONTRACTOR . . . . . . . . . . . . . . . . . . . . .DF I FAILURES IN FLIGHT . . . . . . . . . . . . . . . . . . . . .(a) Measuremen t fai lu res . . . . . . . . . . . . . . . . . . . .(b) Causes . . . . . . . . . . . . . . . . . . . . . . . . . . . . .APOLLO DFI PROGRAM MILESTONES . . . . . . . . . . . . .BP-6 FLIGHT EQUIPMENT . . . . . . . . . . . . . . . . . . . .BP- 12 FLIGHT EQUIPMENT . . . . . . . . . . . . . . . . . . .

BP- 13 FLIGHT EQUIPMENT . . . . . . . . . . . . . . . . . . .

BP-15 FLIGHT EQUIPMENT . . . . . . . . . . . . . . . . . . .BP-22 FLIGHT EQUIPMENT . . . . . . . . . . . . . . . . . . .BP-23 FLIGHT EQUIPMENT . . . . . . . . . . . . . . . . . . .

SC-002 FLIGHT EQUIPMENT . . . . . . . . . . . . . . . . . .SC-009 AND SERVICE MODULE FLIGHT EQUIPMENT

SC-011 FLIGHT EQUIPMENT . . . . . . . . . . . . . . . . . . .SC-017 FLIGHT EQUIPMENT . . . . . . . . . . . . . . . . . .

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SC-020 PLIGHT EQUIPMENT . . . . . . . . . . . . . . . . . .LTA-10 FLIGHT EQUIPMENT . . . . . . . . . . . . . . . . . .LM- 1 FLIGHT EQUIPMENT . . . . . . . . . . . . . . . . . . .CSM- 10 1 FLIGHT EQUIPMENT . . . . . . . . . . . . . . . . . .CSM- 103 FLIGHT EQUIPMENT . . . . . . . . . . . . . . . . .LM-3 FLIGHT EQUIPMENT . . . . . . . . . . . . . . . . . . .

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Figure

FIGURES

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The DFI system block diagram . . . . . . . . . . . . . . . . . . . . 9A typical breadboard system . . . . . . . . . . . . . . . . . . . . . 10

The DFI tes t console . . . . . . . . . . . . . . . . . . . . . . . . . 10The DFI system installed in LM-2 . . . . . . . . . . . . . . . . . . 10A photograph of White Sands, New Mexico; one of 370 Earth-

mapping pictu res taken du ring the Apollo 6 mission . . . . . . . 14

The fi rs t photograph of the E ar th taken fr om an altitude of10 000 nautical miles . . . . . . . . . . . . . . . . . . . . . . . . 14

Variation of DFI manpower lev els with ti me . . . . . . . . . . . . . 15

The BP-6 vehicle during launch escape syste m te st firi ng . . . . . 22The launch of BP-12 . . . . . . . . . . . . . . . . . . . . . . . . . 22The instru mentation and communications subs ystem s used on

BP-15 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 24The CSM-101 flight-qualification instrumentation diagram . . . . . 42

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APOLLO EXPERI NCE REPORT

DEVELOPMENT FLIGHT INSTRUMENTATION

By Norma n B. Fa rmerLyndon B.Johnson Space C e n t e r

SUMMARY

In this report, the development flight instrumentation s yst ems that were suppliedas Government-furnished equipment for 25 Apollo vehicles a r e discussed. These

syste ms, which were used to flight-test and qualify the earlier Apollo vehicles, wer edesigned, integ rated, and teste d at the NASA Lyndon B. Johnson Space Center (for-merly the Manned Spacecraft Center) and shipped to the pr im e co ntr act ors o r launchsites fo r vehicle installation. The 8-ye ar activity supported se ve ra l types of missionsvarying from the first pad-abort launch (boilerplate 6 ) from White Sands Missile Rangeto the Apollo 13 lunar flight from the NASA John F. Kennedy Space Cen ter . Duringthis time, techniques were developed to provide for flexible and relia ble sys tem s.Thes e techniques involved p rim ari ly the us e of rel ati vel y low-cost off-the-shelf equip-ment and an accompanying qualification program that ensured the nec essar y degree ofreliability for the flight test program.

Component qualification and acceptance tes ting was devised to place emph asis onthe environmental tes ting of all flight components. This testing was considered amajor f act or in obtaining the ove rall 98 percent successful data re tri eva l for the Apollodevelopment flight instrumentation sys tem s. Components were standardized betweenvehicles because the development flight instrumentation used a "building block" ap-proach that easil y accommodated any variation in quantity o r type s of meas ureme nts.A maximum deg ree of s tandardi zati on was also used for component input/output ch ar -acteristics, common component test levels, component range selection lists and testprocedures, and so forth.

The significant problem s encountered in the component. areas involved electro-mechanical devices. Tape re co rd er s were of particul ar concern and required con-piderable attention to achieve improved reliability. Some of the mos t serio us problemsoccurred with failures in the vehicle wiring harn ess es, which affected the integrity ofthe development flight instrumentatio n syste ms. One lesson that was learned was theimportance of performing a prop er sys tem installation design within the vehicle. Poorcomponent accessibility and high-density wiring ha rnes s design were major facto rs insystem fai lures.


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INTRODUCTION

In late 1961, the NASA Manned Spacecraft Center (MSC) (now the Lyndon B. JohnsonSpace Center) began an activity to provide t he development flight in strumentation (DFI)fo r the Apollo vehicles. The instrumentat ion involved in this activity was prim ari lyused to provide da ta during the flight test ing phase s of the ope rational subsy stems andconsisted of tra nsdu cers, amplifiers, modulation packages, tape rec ord ers , tra nsm itt ers ,and so forth. The sys te ms were flexible (variable in capacity) and could accommodatehigh-bandwidth analytical para meter s; fu rther more, the sys tem s were designed to bephased out of he Apollo Pro gra m as the vehicles became operational. The DFI equip-ment w a s generally disp ersed over al l par ts of the vehicle and frequent ly occupiedlocations that were destined la te r to be taken by the operational systems. Tran sduc erswere located in suc h positions as the inside of rocket thru st cha mber s, the outside skinof s er vi ce modules, t he pipes on hydraulic sys tem s, and the surf ace of s tr uc tu ra lmemb ers. The la rg er elements of the sys tem (modulation packages, tape re co rd er s,etc.) were generally centrally sited on shelves (if available) fo r accessibility. Incont rast to the DFI, the "operational" instrumenta tion syst em w a s designed as a perma-nent entity and w a s lat er phased into the Apollo P rogr am to monitor the vehicle opera-tional "housekeeping" par ame ter s. The operationa l instrumentati on w a s essentially alow-bandwidth system and w a s not concerned with analytical-type m easu remen ts thatrequired high frequency response rates. Although the original intention w a s to providethe DFI syste ms for the developmental par t of the Apollo P rog ram , the equipment w a sused throughout the operationa l phases, including the early lunar missions. All DFIsyst ems were supplied as Government-furnished equipment (GFE) fr om MSC. A DFIgroup managed the MSC in-house activities for the fabrication and delivery of he s y s -tems. These activ ities involved syst em design, component procurement , componentqualification, syste m integration, and so forth.

In all, 25 flight sys tem s were designed and constructed. Eighteen of the se wereactually flown on miss ion s; the remaining seven, because of p rog ram changes, wer e

used in ground tes t vehicles or w ere reassigned f o r use as sp ar es . Additional equipment(not full systems) was flown on se ve ra l othe r vehicles. The first flight syste m w a s usedon the boilerp late (BP) 6 pad-abort miss ion a t White Sands Missi le Range (WSMR) nNovember 1963. The DFI w a s supplied to the pr ime cont racto rs fo r installation in theflight vehicles. A total of 8000 replaceabl e units (tran sdu cer s, ampli fiers , t ra ns mi tt er s,etc.) w a s qualified and supplied as DFI to supp ort the Apollo Pr og ra m. Some of thehardware w a s designed and fabr icate d at MSC; the majori ty, however, w a s obtainedunder contract with 45 equipment vendors.

The D F I perfo rmed well throughout the flight pro gra m: no maj or equipment fai lur esoccurred and the overall average data ret rie val was 98 percent. A secondary benefitderived from the DFI activity w a s t h e engineering train ing of pe rson nel in spacecr aft

equipment. Engineers received a prac tica l indoctrination in the design and operatio n ofvarious equipment and in the development of gen era l testing req uir eme nts and specifi-cations. The gained knowledge, especially the genera l testing requ ire men ts andspecifications, became an as se t to the Apollo Pro gr am and helped in the prep arat ionof the operational equipment. La ter , DFI specification s we re reflected in theoperational specifications and procedures.

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The first DFI requirement originated from aerodynamic and str uctur al me asure -ments on Apollo/Saturn miss ions 101 and 102, which car rie d both ser vic e modules andboilerpla te command modules. The initial syste m design w as predicated on largequantities of mea sur eme nts using radio-frequency (13) tr ansmis sion as the means ofdata ret rie val (nonrecoverable mission). Development of the instrumentation design w a sbased on sys te ms with approximately 1200 data channels and 7 trans mitter s; however,the initial requirement fo r DFI on thes e vehicles w a s postponed because of the BP-6mission at WSMR. The objectives of tha t te st were to demonstrate the launch escapesystem and to tes t various par ts of the boilerplate; 94 measurements were made withthe DFI sys tem on the BP-6 ehicle. The system performed satisfactoril y and providedconfidence to continue with the ba si c block design fo r futu re Apollo D F I systems.Flight perfo rmanc e information f or each mission is given in appendixes A to D.

IN IT IA L PROJECT CONSIDERATIONS

The decision to supply the DFI as GFE was made for two pri me reasons: to permitNASA o maintain schedule and system flexibility during the flight tes t prog ram and torelieve the prime contractor workload to allow greater concentration on the operationalsystems. During the initial sys tem design, th e fact became apparent that meetingschedule deadlines would preclude excessive component development and require thatsystems be designed to use reliable aff -the-shelf equipment.

done with the NASA George C. Marshall Space Flight Center , the NASA Langley Rese ar chCenter, and other institutions. The resulting information revealed that mor e re se ar chand component evaluation would be necessary in the general areas of low-level commu-tation, signal condi&ming, and frequency modulation (FM). Because of the extent ofthe projected activity, the decision w as made to begin on a long-term design with abuilding-block approach. In the beginning, neither weight nor s iz e was a problem becausethe early boilerplate vehicles were not equipped with operational equipment. However,as the activity continued, weight and s i ze beca me increasingly impor tant as operationalequipment was ntroduced and less room was left fo r the DFI. The later requirementfo r DFI in the lunar module (LM) necessitated a weight and size reduction of theindividual components. A comparison of the resultant reduction in DFI componentcharacteristics between the command and service module (CSM) and the LM is indicatedby the following statistics.

To determine the best select ion of t eleme try hardware, considerable research w a s

1. Average LM component volume = 20 pe rc en t of CSM counterpa rt

2. Average LM component weight = 32 percent of CSM counterpart

3. Average LM component power = 27 pe rc en t of CSM counterpart

The DFI was not subject, as w a s the operational equipment, to the mo re str ingentreliability and quality assurance (R&QA) ontrol meas ur es a t vendor plants. To removemost reliability ri sk fa ct or s associated with off -the-shelf equipment, applicable R&QAsta ndar ds were adopted and a qualification and test program w a s performed at MSC onal l components. Considerable emphasis w a s placed on acceptance testing to reveal

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fabrication and quality control weakness es in prospectiv e flight hardwa re. Most of theDFI ground-support equipment used by the pri me cont racto rs w a s supplied by MSC.This was the most economical approach because it en sured a maximum standardizationof te st equipment and test pr oced ures f or the project.

One significant fa ct or in en suring the inte grity of the DFI s ys te ms was th e adoptionof the breadboard integration tes ts. In thes e predelivery te sts , the actual flight compo-nents of a particular system were strapped to a large, metal-covered breadboard. Theent ire system was energized, checked, and subjected to electromagnetic-in terferenceconditions. The cabling on the brea dboa rd w as si mi la r to that used on the flight vehicle.The test method ensured, as far as possible, that the system would work aft er it w a sdelivered and that i t als o would function within the elec tromagne tic-i nter ference envi-ronment inside the vehicle. Both the sys tem and the breadboard wer e shipped to theprime contractor s o that the system could be checked and accepted before installation.Furt herm ore, if pro blems occur red in the vehicle, the contr acto r could us e the brea d-board for additional system checking.

Project engineers were assigned the responsibility fo r specific vehicle syste ms.For instance, a project engineer at MSC would be responsible for a syste m and thetes tin g of that sys tem . He would then monitor and approve the insta llati on and tes tin gperfo rmed at the pri me contractor facility, ultimately following through to the NASAJohn F. Kennedy Space Center or WSMR and the launch. For backup, resi dent te am sof DFI -trained cont ractor personne l were ass igned by MSC to provide daily lia iso nwith Houston, to implement modifications, and gene rally to expedite the solut ions tosystem problems.

DESIGN C ONSIDERATIONS

In the general system design, certain sa lient ite ms helped to stan dardize the

approach and to provide a sound bas is for the sys tem s. One such item and one of thefi rs t to be decided on w a s the operating signal voltage for sub car rie r oscillators,ampl ifier s, commutators, and tr'ansducers. A cr os s secti on of the vari ous tra nsd uce rtypes that would have had to be used could have produced output voltages va rying fr ommillivolts to volts in combination with variou s pol ari tie s and bia s offsets. Althoughsub car rie r oscillators that accommodated these varying sensitivities and conditions ofsign als could be acquired, the decision w as made to standardize all subcarrier osci l latorsat a signal input of 0 to 5 volts. Therefore, the amplifier or trans duce r would have tosupply 0 to 5 volts re ga rd le ss of th e polarity o r input swing of the par ame ter . Appar-ently, in many cases, tailoring the subc arr ier osc illat ors to suit the varying par ame terconditions would have been easie r. However, considerin g that each su bc ar ri er oscillat orwithin a given package w a s of a different frequency, the fac t became readily apparent

that an additional variable, the sub ca rr ie r osc ill ato r input voltage, would compound theproblems of spares and interchangeability.

Consideration then w a s given to the low-level pa rt of the syst em in which commu-tation apparently would be most economically advantageous beca use of the req uir ed lowfrequency response and the profusion of measu rem ent s (th erm al surv eys, st ra inmeasurements, et c. ). The original sys tem s were designed in accordance with a

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10-millivolt signal input standard. No range fr om any of the low-level devices w a s ex-pected to give less than thi s value, considering any possibl e dynamic range of th e forcingfunctions. Prev ious experie nce with noise problems at these levels indicated that thesy st em s should be designed with balanced signal-conditioning networks having outputimpe danc es of 100 ohms. The fac t that th e required low-level data were generally lessthan 10 hert z permitted filtering in the frequency spe ctr um above which most electrical

interference occurred. This approach permitted f a i r l y long line lengths (approximatelyof f u l l scale. Low-level tr an sdu ce rs requiring high frequency res pons e were fed directlyinto high-gain dir ec t-c urr en t ampli fiers and converted to a high level (0 to 5 volts) fo rinsertion into the subcarrier osci lla tors . The output and input impedances of amp li fi er sand conversion devices were designed to be approximately 1000 and 50 000 ohms, respec-tively. High-level tr an sd uc er output impedances were generally designed to be a maxi-mum of 1000 ohms. By establishing f a i r l y rigid rules on syst em impedances and voltages,component engineers (who had to deal with 45 vendors) w er e provided with fai rly specif icsta ndar ds f o r procurement and interface specifications.

40 feet) f o r remote se ns or locations and kept noise lev els to approximately 1 percent

Before the individual des ign of many components could be established, the matter

of a utomatic calibration had to be decided. Automatic calibration w a s considered adesirable system feature, providing it did not com promise the s yst em in cost andreliab ility. One of the benefits to be derived fro m this approach is that it provides afast and easy way to check ;i system before launch and in flight. However, many validarguments exist for not using such a method. In many cases, calibration functions donot e xer cis e tel eme try channels completely from end to end; therefore, the sys tem isonly part ly checked. Many high-level transducers were not equipped to accommoda teautomatic calibrations and would have had to be specially modified. This modificationwould have des troyed the integri ty already built into off -the-shelf components. Certaintra nsdu cer s (resistance thermometers, thermocouples, stra in gages, etc. ) can bechecked effectively and easily by minor cir cui t additions into signal-conditioning net-works. Others, such as variable-reluctance pre ss ure trans duce rs and crysta l accel-

erometers, cannot be checked as eas i ly and are scarcely worthy of th e attempt becauseof the complexity involved. Also, when an ambient condition (temperature, pr es su re ,etc.) falls within the dynamic range of a sensor, a good channel-verification point isprovided. Furthermore, certain units a r e so reliable that to check them is probablysuperfluous. It also must be remembered that when a system is commanded into acalibration mode, the syst em must necessarily switch from the true data position. There-fore, the r isk of inflight malfunction ex is ts ; the system may either shift or remain inthe cal ibra tion position and re tu rn only calibration information instead of data. Althoughthere were many re ason s f o r not using such a feature, the convenience of a fast built-intest arrangement overrode the objections and an automatic calibration sys tem wa sdesigned into the DFI. This system included the transducer elements, w h e r e practical,and all signal ampl ifie rs except vibration-charge amplifiers. The subcarri er osci llators

also were switched into either five- o r two-step calibration levels, and the commutatorshad fixe d-re fer enc e voltages impre ss ed on cer tai n channels. This genera l philosophyof syst em calibration w a s followed throughout the DFI activity because it provided aquick check f o r faultfinding and reduced un necess ary probing into alre ady functioningsystem areas.

The DFI s ystem design was not founded on a redundant concept as were th e lateroperational CSM and LM systems. At the time, redundancy, in the t rue sense ofduplication, see med to be extremely complex and expensive; furthermore, the DFI

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equipment was not sub ject to the stringent miss ion-essent ial requ irem ents that governedthe later operational equipment. Certain as pec ts of the design compensated somewhat fo rthe lack of redundancy. Where pra ct ica l, mos t units had interna lly regu late d powersup pli es , which, in effect , made them independent of any commonly regulated supply(excluding the pr im e battery) and thus fr e e fr om mos t single-point power fai lure s. Unitsdesigned with this cha ract eris tic consisted of high-level trans duce rs, signal amplifiers/converters, s ubcar rier oscillators, mixer amplifiers, transmitters, and, to some

power-to-signal line s fo r most of the front-end instrumentation, thus ass is tin g thextent, low-level conditioning units. Thi s concept al so made poss ible the us e of i sola ted

overall design to achieve a single-point grounding system with a minimum of potentialloops.

Another f eat ure that helped to c ompensate for no redundancy w a s the cross loadingof batches of measurement s to different modulation packages and trans mitte rs. In otherwords, if a given a rea of th e space craft w a s subjected to a specific survey of measure-ments, where practical, the measurements would be distributed ac ro ss various r f links.Thus, in the ca se of the part ial o r entire loss of a link, a perc enta ge of the pertinentmeasurement s would sti ll be preserve d.

The cen tra l pa rt of t he sys te m was the signal-conditioning unit, which channeledmos t of the signals. Thi s unit se rv ed as a central checking point and w a s capable ofquick acc es s fo r channel and range changes when the system w a s already installed inthe vehicle. Furthermore, this unit permitted programed inflight time sha ring for thepurpos e of maximum us e of high frequency channels . Th is unit a n d the c entral powerunit contained al l the syst em fusing, which w a s used in both single and grouped compo-nent powerlines . Fusing arrays we re optimally designed into the DFI sys tem because ahigh possibili ty of syst em damage existed during a mission. The equipment w a s to beused in the initial proving sta ges of the Apollo Pro gra m, during which high st r e s s e sand violent maneuvers we re imposed on the vehicles with som e possible c atas troph iceffects. One such case w a s BP-12, which w a s launched f ro m WSMR. P a r t s of theser vic e module ramme d the bottom of the CSM boilerplate and seve re d som e DFI cab lesfeeding several pre ssu re transduc ers. Some fu se s were blown in the affected sys temarea, but the short-circuiting effect on the powerlines w a s not tra nsfer red to theremai nde r of the sys tem , which continued to function.

Components of the s yste m wer e designed and spec ified to include pol ari ty- re ver sa lci rc ui tr y with one exception: the power and control unit, through which all pri me powerwas fed and switched. . This overall precautionary design concept saved schedules andequipment because many of the sy st em s we re subjec ted to inadvertent power-polarityrev ers al during testing.

The overall selection of the DFI system, regarding modulation transm issi on andtracking, w a s controlled by the existing tele met ry ground stati ons and radars in theManned Space Flight Network (MSFN) and at WSMR. The requirements for instrumen-tation called f o r wide-band multichannel sy st em s capable of c ater ing to data r a t e s f rom2 kilohertz to steady stat es. The choice w a s a pulse amplitude modulation (PAM)/pulse

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duration modulation (PDM)/FM/FM sys tem; the PDM pa rt w as used exclusively formultiplex tape recording. Briefly, this pa rt of the sys tem conformed to the followingdesign param eters .

1. Commutation

a. 90 channels X 10 sampl es/s ec (PAM/PDM)

b. 90 channels X 1-1/4 samples/sec (PAM/PDM)

c. 3 channels x 0.8 sampl e/se c (PAM) special high frequency data commutator(2 kilohertz)

2. Modulation

a. Proportional bandwidth (PBW) sy st em s interra nge instrumentation group(IFUG) channels 2 to 16 and channel E

b. Constant bandwidth (CBW) sys te ms Aerospace Indus tri es Associationchannels IC to 1O c

3. Transmission: Very high frequency (vhf) 228.2 to 257.3 megahertz ( F M )

4. Tracking

a. C-band transponders

b. 5690 to 5765 megahertz

5. Recording

a. 14 racks on 1-inch tape at 15 in/sec

b. Commutated data - 0 x 10 PDM and 90 X 1-1/4 PDM

c. Continuous da ta - PBW channels 2 to 16

d. Wide-band dat a - M record 14.5 kilohertz (2-kilohertz data)

The low-speed commutation ra te of 90 channels X 1-1/4 samples/sec w as previ-ously used in Proj ect Merc ury and, although suitable for many quasi-static pa ra me te rs ,w a s not considered fa st enough to resolve di sc re te events adequately for the Apollo

Prog ram ; hence, a rate of 90 channels X 10 sam ple s/s ec was additionally adopted. Thecommuta tors that had a ra te of 3 channels X 0 .8 sampl e/se c were specially developed totime-share high frequency data (as much as 2 kilohertz) effectively on single continuouschannels. Various makes and si ze s of PBW modulation packages, the large r onescontaining 16 channels, were used throughout the DF'I project. Channel E w a s generallyused for the loading of high-rate commutators. The CBW packages, which we re usedin the L M phase of the project, were introduced to accommodate the la rg e amount ofhigh frequency dat a and t o reduce the number of rf links. These packages could abso rb

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as ma ny a s 10 channels of 2-kilohertz information and were, in fact , the first CBWsy st em s ever flown. The sys tem s, although having relatively lar ge capacities forhigh frequency data, had certain drawbacks. First, the i nformation loading on the sub-

ca rr ie rs resulted in a modulation index (MI) of 1 (2-kilohertz data). Second, beca useof the re la tively high number of channels occu rr ing at the high end of t he sp ec trum(compared to PBW IRIG), t he tra ns mi tt er MI pe r channel w a s down to 0.3, the overallper mis sib le vhf t ra ns mi tt er deviation being 128 kilohe rtz (ground-receiver bandwidth

limitations). Thi s loading, with its re su ltan t low values of modula tion index, gavereduced F M improvement with consequently lower signal-to-noise ratios than the P Bwsystems. Altogether, the sys tem s were very useful as convenient large-c apacity toolsto obviate the need for additional r f links but had to be used within the sense of theirlimitations.

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2

During the boilerplate stages of the DFI activity, 12-watt tube-type t ra ns mi tt er swere used. Lat er, thes e tr an sm it te rs were superseded by 5-watt and then by 10-wattsolid-state units as the DFI equipment w a s phased into the CSM and LM vehic les . Ea rlymissions generally conformed to short-range ballistic or 100-nautical-mile-altitudeEarth-orbital pat hs; in these ca ses, the radiated power w a s mor e than sufficient to give

high positive circuit margins to MSFN and WSMR.appear to be conservative ly designed. With the inception of the LM Earth-orbital vehi-cles, which used as many as five 10-watt DFI trans mitte rs, the sys tem s were designedto meet a worst -cas e 3-decibel margin in conjunction with the M S F N helic al vhf antennasat a requi red slant range of 1200 nautical miles . Data-channel noise under thes e condi-tions generally conformed to le ss than 2 perc ent of ful l sca le fo r the PBW sys tem s.However, the CBW sys tem s, fo r the rea son s previously disc usse d, gave bette r pe r-forma nce near the stations and assumed channel-noise leve ls of approximately 3. 5 percentfull scal e for slant ranges of 600 nautical miles.

From this viewpoint, the rf links

FLIGHT SYSTEM BREA DBOARDING

The block diagram of the basi c DFI sy st em s (fig. 1) illustrates a design approach.The only difference s in the various sy st em s ar e in the number and types of measurementsmade and in the number of pie ces of tel eme try equipment (commutators, modulationpackages, and trans mitte rs) used to accommodate these measurements . The photo-graphic cameras and onboard tape recorders that were used on the CSM boilerplate andproduction vehicles, however, are not shown in fi gur e 1.

'It is preferable to u s e an MI value of 2 (or greater) in this part of the system,ca rr ie r deviationsince lower values produce poorer signal-to-noise rat ios ; MI = modulation frequency *

6l

A higher value of M I (say 1) would be pre fera ble a t thi s par t of the system topermit a better signal-to-noise ra tio. (See al so footnote 1 .

8


14/47

I

Modu lation package DKBW)

I I I IVibration Vibrationcommutator commutator

I

Battery

I

Modu lation packageC Modula tion package A I Modul ation packaqe B 1

5-V ref- 5-V ref- High-levelinputs.erence transducers as requiredpower

to measure acceleration,ower

supply supply

Amp1i ie rPower Power tocontrol

-fitem

Strainox

high-levelcommutator

temperature low-levelsignal

t an d u c e r

Iianal conditioner

IPBW)

1 1 .90 lohigh-levelcommutator

Vibration

-

1

I w-leve1

temperaturesignal

Temperature

Ground-support-equipmentcontrol and monitor

Figure 1. - The DFI system block diagram.

Afte r acceptance testing, each component w as mounted on a breadboard system(fig. 2). The breadboard consisted of a movable A-frame-shaped st ru ct ur e on which theequipment could be eas ily mounted and elec tri cal ly interconnected. The breadboardproved to be a convenient, useful, and necessary tool for verifying an entire systemoperation . The breadboard also served as a handy shipping container.

Afte r breadboard drawings had been prepared , the flight-qualified components wereinstalled on the breadboard. The components were interconnected according to theschematics and prepared for syst em test s. The DFI tes t console (fig. 3) w a s developedto verify the operation of the sys tem and components and to isolate system wiring mal-

functions and other operating problems. A typical DFI subsystem installation i n anL M is seen in figure 4. (This particular example is from LM-2.)

9


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Figure 2. - A typical breadboard syste m. Figure 3. - The DFI tes t console.

Figure 4. - The DFI system installedin LM-2 .

ACCEPTANCE AND QUAL1F IC AT I O NTEST1NG

A s early as 1961, a gen era l plan ofcomponent qualification w a s outlined fo r theDFI . Although at that ti me the main con-cern w a s selecting and evaluating certaincomponents, the ground rule w a s establishedthat the overall project would include envi-ronmental testing for acceptance as w e l l asfo r qualification. The DFI w a s based onoff-the-shelf equipment, and, to detectmanufactur ing defects and provi de confidencefo r flightworthiness, the decision was madeto subject each flight unit t o the most crit icalflight environment s. The selected environ-ments were vibration, temperature, vacuum,and accelerati on. Each unit als o was benchtested to elec tric al specifications. This modeof acceptance testing w a s pursued throughoutthe DFI activity and w a s considered to bethe majo r contributor in ensuring reliableflight component operation.

The qualification testing f or each com-ponent type that w a s evolved contained 12separate environmen tal tes ts. However, thenumber of te st s vari ed slightly throughoutth e activity. The peak test levels werefounded on values above the maximum designlimits; that is , any level expected in any

10


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vehicle area that might contain DFI equipment. This was a maj or difference betweenthe DFI and operational qualification philosophy. Whereas most DFI components werequalified a t one level, the opera tiona l equipment w as tailored for specific environmentalzones within the vehicles. The standardized concept was used on the DFI to en su re thatmost equipment could be used and installed in any par t of a vehicle without incurringdifferent or additional qualification. U s e of this concept not only permitted generalmounting flexibility but als o simplified procedures , procureme nt, and paperwork.

environments. Examples of this exposure occurred in the se rvi ce modules, wherecomponents were mounted on thin shelving and were subjected to high vibrationalg-for ces. Other components were mounted through the service module skin w h e r e bothhigh acoustic noise and vibration levels were imposed. To cater to these types of condi-tions, a pa rt of the gen era l qualification document w a s written to include a section for"space-exposed and/or ext re me conditions. '' This section allowed the imposition ofespecially high levels where applicable. The use o r partial us e of the lat ter proce durew a s more th e exception than the rule.

In a f e w cases, components (generally transducers) were exposed to high-intensity

An example of the difference of leve ls between the general and extreme qualifi-cations is shown by a compar iso n of the vibration specifications. The forme r required12.3g; the latter required 45g. In general, the extre me qualification section w a s a"catchall" that satisfied all high environmental requirements and, at the sa me time,standardized the qudification testing.

COMPONENTFAILURES

The failure rates and problem areas associated with equipment after delivery tothe prime contractor are shown in table I. The survey involves the preinstal lation andpostihstallation component failures on seven LM-type vehicles. The following fact swere derived from the i.-+ormation in table I.

TABLE . - DFI EQUIPMENT FAILURES AFTER DELIVERY TO PRIME CONTRACTOR^

1 F a i l u r e s reported for LT A - l , b LTA - 2 R , LTA - 8 , LTA - l O R , L M - 1, L M - 2 , and LM-31

aAn additional 43 f a i l u r e s reported were determined to be incorrect (i. e . , no failures occurred).

b~~ test article (LTA) 1.

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Most of the fai lu re s we re induced by som e external means into the DFI. Of he106 induced failures, most (71) were caused by personnel action; that is, handling,installation, and so forth. It is als o intere sting to note in the induced-failure area thatpostinstallation fail ure s outweigh preinstallation f ail ure s by approximately 2 o 1; thisrat io indicates that most component damage occurr ed a fter vehicle installation.

Of a total of 45 equipment fai lur es, 16 were major. The remaining 29 were minorand, although attributable to vario us cau ses, wer e primar ily out-of -tole rance calibrationconditions. Fa il ur es of t his kind would warrant r eplac emen t only if repl aceme nt wasconvenient. The 43 ncorrectly reported DFI failu res (table I footnote (a)) ar e the inev-itable resu lts of pr ema tur e diagnosis and preli mina ry test investigations.

In summary, the information in table I indicates that the highest failure rate on theequipment was caused by handling. Ther e a r e approximately 2 . 4 induced failu res f o revery 1 equipment fai lur e. Some of the pro ble ms were prompted by inaccessibility, byvehicle cabling design, and by the fact that some equipment w a s within the person nel"traffic" flow through the vehicle. Also, the fact that only four maj or fail ure s occurr edbecause of component pa rt s def ect s (r es is to rs , capacitors, etc.) is significant. The DFIused mature hardware; hence, most of the poo rer component pa rt s were alrea dy scree ned

fr om the equipment. This factor, together with the additional scree ning effected by theacceptance testing a t MSC, reduced the pa rt s probl ems to minor proportions. Althoughinvolving a different prime contractor, the tre nds displayed were sim ila r fo r the DFIsys tem s supplied for the boilerplate vehic les and the command and se rv ic e modules.

Some actual flight failu res a r e listed in table 11, which contains information froma survey including 10 DFI flight sy st em s on vehicles BP-6 o spacecraft 011 (SC-011).Thes e vehicles afforded a bette r opportunity to investigate flight failu res tha n theL M vehicles because most of the s ys te ms were r.ecovered and postflight an alys is w a spossible. From th e information in table 11, i t is apparent that the caus es and typesof fai lu re s covered a r e varied. The number of stati c surface -pres sure failu res (8)se em s relatively high; but this can be explained on the b asis that thi s area sometimes

included mor e than 50 per cen t of all syste m measu remen ts. However, vibration andacoustic failures did prompt a revision of qualification level s to accommodate extr emeenvironments as mentioned previously. The rel ati vel y high number of wiring fa il ur es(10) accounts fo r nearly 50 percent of all fai lur es and st re ss es the criticality of thisarea.

12


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TABLE 11. - DFI FAILURES IN FLIGHT^

(a) Measurement failures

Measurement type I Number of fa il ur esPr es su re (motor combustion)

Pressure (static surface)

P res su re (fluc hat ing su rfac e)

Calorimeter (surface)

Thermocouple (surface)

Vibration (surface)

Strain (beam)

Acoustic (interior)

Total

1

8

3

3

1

3

3

1

23-

Probable cause of failure

(b) Causes

Number of fa il ur es_ _ ~

Vibration

Constricted pre ssu re line s

Acoustics

Wiring (short or open circuit)

Unknown

.

Total

4

3

3

1 0

3

23

aA to ta l of 1104 measurements w a s made; 9 complete and 14 partial measurementfailures occurred. Failures were reported fo r BP-6, BP-12, BP-13, BP-15, BP-22,BP-23, BP-23A, SC-002, SC-009, and SC-011.

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SPECIALSYSTEMS

One interesting par t of the pro gram was the supplying of some special GFE instru-mentation systems. These sys tem s were sma ll compared with the average DFI vehiclesys tem s, but each one had to be supplied on a comp ress ed schedule. Some were identifiedas being require d 2 or 3 months before the vehicle launch time. In fact, the vehicle was

frequently on the launch pad when the initial r equiremen t ar ose.

One such cam er a system, which was installed on CSM-020, generated stereographicmapping photographs ac ro ss the width of the United States; it additionally providedphotographs of the plasma phenomenon during reentr y. The entire sy ste m was self-contained because i t w a s not to int er fer e with, o r impact in any way, the operationaland DFI equipment alr eady on board the spac ecra ft. Th e equipment consiste d of twocameras, a timing and dela y unit, a battery, and an accelerometer start mechanismthat was triggered by the ascent g-force of the vehicle. The syst em w a s designed,fabricated, and qualified within a 3-month period. The syst em was used successfu llyand provided, fo r the first time, a se t of overlapping ster eogr aphi c photographs ac ro ssthe United States. One of the photographs is shown in fi gur e 5.

Another cam er a system, supplied on a previous spacecraft (CSM-017), produceda sequence of color pictures of t he Ea rth fr om a 10 000-nautical-mile altitude. The sepictures, which were th e first taken of the Ea rt h fr om that distance, were publishedworldwide. . An example is shown in fi gure 6.

Figure 5. - A photograph of White Sands, Figure 6. - Th e first photograph of theNew Mexico; one of 370 Earth-mappingpic ture s taken during the Apollo 6mission,

Earth taken from an altitude of10 000 nautical miles,

1 4


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MSC personnel-ontractor personnel

-----40 -120 -

SCHEDULESANDMANPOWERLEVELS

The manpower peak for the DFI activ-i t y was reached between December 1965and June 1966. During that time, the NASA

and contractor support w a s approximately40 and 125 men, respective ly. This peaksupport w a s prompted by the requirementto deliver seven LM sys tem s within an8-month period. Curves representing themanpower applied to the DF I progr am areshown in fi gu re 7. All DFI sys tems weredelivered to the contra ctors on schedule.A list of the delivery date s is included intable 111.

loo-

Time, yr

Figure 7. - Variation of DFImanpower le vel s with time.

?

Delivery date t oprime contractor Launch dat eehicle Apollo mis sio n

TABLE IU. - APOLLO DFI PROGRAM MILESTONES

PA- 1

A-00 1AS- 10 1AS- 102

--A-002

PA - 2

A-003

December 1962

March 1963April 1963

July 1963

February 1965

May 1963

January 1965

July 1964

B P - 6

B P - 1 2B P - 1 3

B P - 1 5

aBP- 14

B P - 2 3

BP-23A

BP-22

sc-002 A-004 December 1964 January 20, 1966

bSC-O1O -- January 1965 ---

November 7, 1963

May 13, 1964May 28, 1964

September 18, 1964

- -December 8, 1964

June 29, 1965

May 19, 1965

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TABLE III. - APOLLO DFI PROGRAM MILESTONES - Concluded

Vehicle Launch dat eelivery date t o

prime contractorpollo missio n

Command module - ConcludedaSC -006

a SC-008

sc-009

sc-011

SC-017

sc-020

sc -101

SC- 103

HS- 1

a LTA-8

LTA-2

LTA- 10

LM- 1

aLM-2

LM- 3

LTA-B

- -

--

AS-201

AS-202

4

6

7

8

--6

4

5

--

9

8

December 1964

February 1965

December 1964

March 1965

August 19 6 5

November 19 6 5

June 1966

August 1966

Lunar module

Febr uary 1966

May 1966

October 1966

June 1966

June 1966

August 19 66

October 1966

August 1968 (to KennedySpace Center )

--

Feb ru ary 26, 1966

August 25, 1966

November 9, 1967

Apri l 4, 1968

October 11, 1968

Dec emb er 21, 1968

- -Apri l 4, 1968

November 9, 1967

Jan ua ry 22, 1968

- -

March 3, 1969

Dece mber 21, 1968

Ground te st vehicle.

House spac ecra ft (HS) 1.

a

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PROBLEMAREAS

The major component problem a r ea s were in electromechanic al devi ces and in-volved attitude gyros, cine camer as, and tape rec ord ers . Although all the se componentswere troublesome in regard to reliability and delivery schedules, tape rec ord ers werethe major concern.

Tape Recorders

Tape re corde rs were key components in the DFI s yst ems because they were apr ime sour ce of data. Most DFI w a s selected on a proven off-the-shelf basis. Thef i rs t recorders were selected similarly because they were involved in production runstotaling more than 100 units and were proved in high-linear-g environmen ts on sled tests.Despite the ir apparent ruggedness, they failed in the req uir ed 12.5g vibration environ-ment and had to undergo major redesign modifications on the main housing castings tomeet this requirement.

Various troubles continued to plague the r ec or de rs during the DFI activity: motorsburned out, tape jammed or would not st ar t, tape curled on re el edges, and, periodical ly,certain tracks would not record because of improper tape-to-head tension. After variousmodifications, however, they we re made operable fo r flight usage. One of the contrib-uting fa ct or s to the successful flight operation of these recorders w a s the exclusivehandling and tape loading by a few trained personnel who stayed with them until launch.Later in the program, a higher capacity recorder w a s introduced; this recorder w a s acomplete depart ure fro m the fo rm er type and used a co-belt syst em with flangelesstape reels. In effect, this design was used to compensate fo r the previous p robl ems offlutt er and tape packing. However, a new set of prob lems ar os e rega rdin g belt tensionand belt ma teri als. Various modifications to the capstans and belts had to be made tomake the recor der operational fo r the later boilerplate and ai r f rame ser ie s of launches.

The problem of rec or der qualification w a s in the vibration areas. It w a s difficultto make a recor der with a rela tively high tape capacity conform to ."wow and flu tte r"specifications under a 12.5g random environment. One fac tor that alleviated the situ-ation and perm itte d ea si er qualification was the introduction of shock mounts. Theprob lems that involved bearings, motors, and tape a r e standar d prob lems that continuein present-day flight tape record ers. The reason for this situation is probably twofold.First, the fundamental principl e of pulling tape is stil l the same. Furthermore, thefact that flight re co rd er s ar e frequently redesigned or greatly modified f or v ario us pro-gr ams tends to negate the maturity o r reliability that might be established w i t h a giventype.

Commutators

Another area of concern that can be categorized as electromechanical is commu-tators. In 1963, low-level commutators were not available in a reliable solid-stateconfiguration. At that time, the D F I used mechanical devic es for multiplexing thelow-level measurements. Initially, these device s proved very reliable, having lifetimes

17


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(800 hours) approaching those of the electroni c units. However, the reliabili ty w a scompromised when the dev ices wer e modified to pro gram channel and r at e selection.Thi s modification incre ase d the number of switching poles and caused heav ier motorloading. This w a s a clas sical ca se of improving an item that was highly relia ble t o thepoint that it became complicated and unreliable. The resultant lifetime was reduced toapproximately 250 hours, and the units were phased out of the LM sy st ems and repla cedby soli d-st ate dev ices , which by then had become available. All high-level-type commu-tators used solid-state techniques throughout the project and proved to be reliablecomponents.

Vehic le Sys tem Cab l ing

The ar ea of concern that seri ousl y affects the integrity of the en ti re syste m is thevehicle cabling. Neglect in the pr ope r design, accessibili ty, and installat ion of cablescan offset all the reliabi lity engineered into the components of the sys tem. The factthat the DFI har nes s was completely rep laced twice on B P - 6 and once in LM-1 isevidence of a problem a re a. In the B P - 6 harness, the problems were design and fab ri -cation err or s; fo r instance, al l the signal shields were inadvertently short-ci rcuitedbecause uninsulated instead of insulated shields were used. This type of cable problemw a s not as detrimental to the syst em as the LM-1 problem.

cer tai n DFI components led to th e design of a cen tra l signal-conditioning unit t hat hada den sit y of 1600 connector pins over a 45-square-inch faceplate. This unit was demon-stra ted on the sy stem breadboard at MSC as being operational but requirin g ca re ininstallation to provide for cable accessibility. The unit was inst alled initially in LM-1in a relatively inaccessible position, and the mating cable harness consisted primarilyof no. 26 AWG wire. After a se ri es of requ irem ents changes and troubleshootingprocedures that involved moving and opening the signal-cond itioning unit, so me of t hewire s in the h arne ss became fatigued and broken. This problem w a s also manifestedin the harness in other a re as where cable movement w a s excessive. The situationdeteriorated to the point at which attempts to rectify cert ain cable breakages precipi-tated fur the r breakages in adjacent are as. The Apollo Pro gra m w a s forced into themajo r decision of reinstall ing t he signal-conditioning unit with a new swing-out shelf(for accessibility) and refa bric atin g and installin g a new DFI ha rn es s with heav ier gagew i r e (24 AWG).

In LM-1, th e scar ci ty of available spa ce and the consequent miniatur izat ion Of

H a n d l i n g D am a ge

Handling fail ur es w ere induced by s uch things as excess power voltage, reversepower polarity, over pres sure, riveting, and physical damage during and afte r instal-lation. It is impossible to avoid handling fa il ur es at f ield si t es despite al l the Precau-tions and procedure s. However, any equipment installat ion should be delayed to the las tpossible moment to avoid vehicle manufacturin g oper atio ns as much as possible. Theover all provision of sp ar es for components must nece ssari ly take into account theattr itio n resulting from handling damage; thi s point is amplified by the commentsalr ead y made in the "Component Fai lur es" sect ion, which shows the rela tive ly highincidence of handling failures.

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CONCLUDINGREMARKS

The building-block approach used on the DFI project permitted a standardizedsystem that w a s basically common to the Apollo vehicles. This feature, together withthe progr ama ble signal-conditioning units, allowed flexibility in progr am changes and

new requir emen ts. Measurement changes were implemented on flight vehicles some timeswithin a mat ter of hour s; some sma ll syste ms were designed, qualified, and installedin per iods of 3 months.

Certain features of the activity helped to per mit this versatili ty. The project hadaccess to fast purchas ing through the NASA Manned Spacec raft Cent er pro cure mentchain, and the ha rdw are that was purchased w a s invariably off-the-shelf type andreadily available. Moreover, the DFI w a s not subject to the "mission critical" standardsas were t he operational sy st em s and consequently had gr eat er latitude in reliability andquality control. This fea tur e eliminated the necessit y fo r the inclusion of "high reli-ability" p ar t s and generally excluded the equipment from any developmental prog ram s.The fact that al l flight hardwar e was acquired on ' 'fixed cost" contracts (ra ther than"cost plus" type contracts generally incurred by developmental contracts ) is an indicationof the sma ll amount of development work involved throughout t he project.

As n most flight test activities, th e DFI requirements were generally defined late.In some cases, th"ts late definition only permitted a period of approximately 12 monthsbetween the go-ahead and the actual syst em delivery to the pr ime contractor. Certainmeasures that we re taken helped allev iate these scheduling probl ems ; one of t he seinvolved the standardization of measurement ranges. Tran sduc ers comprised thelargest area of the syste m, contained the grea test variety of component types, andwere subject to the greatest number of changes (because of varyi ng measur eme ntrequirements). To simplify and expedite procure ment, tra nsd uce r ranges were stand-ardized to cut down the variety of inst ruments. In effect, the pro ject offered a stand-ardized list of transducer measurement ranges instead of responding to specific rangerequests. The standardized leve ls that were established for component qualificationensured that DFI equipment generally could be installed in any part of a vehicle withoutinc ur ri ng additional testing. This not only per mit ted mounting flexib ility but alsosimplified pro ced ur es, procuremen t, and paperwork. Similarly, standardization ofel ec tr ic al specifications fo r components, regarding input/output voltages, impedances,and s o forth, ensu red simplified syst em interfaces and set specific controls and stand-ard s for procurement with the component vendors. This la tt er point w a s significantbecause it provided clear guidelines to component engineers who had to interface withapproximately 45 equipment vendors.

The DFI project obtained an over all 98-percent successful data retrieval. Twofactors were major contributors in achieving t h i s status. Fi rs t, off -the-shelf ma tu rehardware w a s selected; this hardware had already been screened fo r poor pa rt s by usage.(Only minor pa rt s problems were experienced throughout th e activity). Second, allflight hardware was subjected to acceptance procedures that emphasized environmentaltesting and effectively weeded out th e weaker components bef or e commitment f or flightpurposes. Certain precautionary design features were instrumental in minimizingsystem proble ms during the testing and operational p hases of th e DFI activity. Theadoption of internal ly regulated power supplies (where practical ) fo r individual unit s

19


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made most parts of the sy ste m independent of single-point power supply fa il ur es andal so permitted signal-to-powerline isolation, which prevented ground loop probl emsand gave gre ate r immunity fr om ele ctr ica l interfe rence. The DFI was not designedwith tr ue redundancy, and such fe atu re s as independent power reg ul ato rs precludedthe possibility of losing lar ge sec tion s of the sy ste m because of single power unitfail ures . The technique of c ro ss loading measu reme nts fr om a given ar ea onto differentradio-f requency tr an sm it te rs and modulation packages provided som e functional redun-dancy and ensured that a percenta ge of pertine nt pa ra me te rs could be pr ese rve d if amajo r sys tem link were lost. Most units of the syst em were designed to include polarity -re ve rs al c ircu itr y to prevent the gro ss destruction of components during inadvertentsyst em power re ver sal s. The re ve rs al of power polarity on complete sy st em s (andcomponents) actually occur red many ti me s during the project.

Various proble ms are disc usse d in the repo rt, but one of the most significantar ea s fr om which some very important lesso ns were learned concerned system cabling.From the cabling problems cited, th re e conclusions can be drawn. Fi rs t, high-densitywiring configuration should be avoided. Second, sign al conditioning should be dec en tra l-ized or made remote so that low-density connector configuration can be achieved toper mit easy ac ce ss and re pai r and not res ul t in inflexible bundles of ca bles. Third, theDFI system involved frequent equipment changes; ther efor e, it should use a heaviergage w i r e than t h e mo re permanently situated, operational-type equipment. Fo r gene ralsignal wiring, no. 22 AWG wire is recommended, although no. 24 AWG wire is suitableif weight becomes a constraining fact or. Finally, t he installat ion design with re spe ctto accessib ilit y should be given pr op er and time ly attention. Neglect of th is ar e a w i l lpromote needless damage and gener ally w i l l lead to a more unreliable system. In thecomponent area, tape re co rd er s proved t o be the most problematical because theyoccupied key positions and wer e relatively unreliab le. Fail ure in a rec or der wouldhave meant a significant or total lo ss of d ata ; consequently, thes e devices wer e givenexclusive personnel handling treatm ent throughout th eir te st and flight phases.

The benefits gained from keeping the DFI independent a r e worth mentioning. Obvi-ously, the DFI had many interfa ces with oth er syst ems , which were, in fact, the so ur ce sof the measu rement pa ra me te rs . The point to be made, however, is that the control,power, wiring, and calibrat ion functions generally w ere independent of othe r onboardsystems.DFI beca use of measu rement changes could be implemented with litt le or no impact onthe vehicle operational systems. It also meant that the DFI could be checked out with-out disturbing other systems. The DFI w a s used fr equently to take advantage of un-scheduled vehicle downtime fo r additional tes ting because of its over all independenceof operation. It could be quickly ener gize d and checked with its own support equipment.The net resul t was that testing of the s yst em w a s easily dovetailed into the vehiclemas ter tes t plans and provided a convenient mea ns f o r schedule optimization during thevehicle tes t operations at the pr ime cont ract ors' plants.

This fa ct meant that modifications (par ticu larl y lat e ones) requ ired of the

Lyndon B. Johnson Space CenterNational Aeronautics and Space Administration

Houston, Texas, November 16, 19739 4-50- 0-06- 2

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APPENDIX A

BOILERPLATE PHASE

INTRODUCTION

The boilerplate (BP) vehicles were const ructed to simulat e the weight, shape, andce nter of gravity of the man-rated Apollo spa cec raf t. The flight te st s e r i e s consi sted offive launches fro m W h i t e Sands Mis si le Range and two from the NASA John F. KennedySpace Cen ter (KSC). In addition to the rocket launches, boile rpla te vehic les wer edropped from aircraft, impacted into land and water, and used during parachute-evaluation tes ts. The boilerplate vehicle, originally quite simpl e, became a complex,

this phase gathere d data about basi c spa cecraft design, operation of subsy stems , andcompatibility of spa cec raf t and launch vehicles .

functioning unmanned spac ec raf t as th e program progressed. The test program during

APOLLOMISS lO N PA-1, BP-6 VEHICLE

A total of 94 measurements was required on BP- 6 during Apollo mission PA- 1.The development flight i nstrumentation (DFI) light equipment used is listed in table A-I.The BP-6 during launch escape sys tem test firing is shown in figur e A- 1.

The sy ste m operated satisfactor ily except fo r two pressur e-me asure ment channels.One incurr ed a 4-second inter ruption and the other a 1-psi calibrat ion displacement.

TABLE A-I. - BP-e FLIGHT EQUIPMENT

Component

Te leme t ry modu lat i on package

Te leme t ry r ad io - fr equency (rf) package

T i m e r

Tape r eco rde r

Signal condi t ioner

Power ontrol

Junction box

Main battery

Pyrotechnic bat tery

Ra te gy roscope packageAtt i tude gyroscope

L i n e a r a c c e l e r o m e t e r

Vibrat ion system

Ampl i f i e r r ack

Pres su re t r ansduce r

Tempera tu re sys t em

Ampl i f i e r

Res i s t ance t he rmomete r

Quantity

1

1

1

1

1

1

1

1

2

1

3

6

6

3

42

4

5

4

2 1


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Figure A-1. - The BP-6 vehicle duringlaunch escape syste m test firi ng.

i !.-* .

Figure A-2. - The launch Of BP- 12.

APOLLOMISSIONA-001, BP-12VEHICLE

Apollo missio n A-001, the second abor ttes t, used BP-12 to test the launch escapesystem under high dynamic pre ss ur e andtranson ic speed conditions. The tes t vehicle

consist ed of the spac ecra ft (launch esca pesystem, BP-12 command module, andboiler plate s er vic e module) and the launchvehicle (Little Jo e II (LJ-II) booster)(fig. A-2).

-

A total of 138 measure ments w a s re -qui red on BP-12 during miss ion A-001. TheDFI equipment complement w a s similar totha t flown on BP-6. The equipment used islisted in table A-11.

The telemetry syst em performed satis-factorily except for a brief disturbance at28.47 seconds into the flight and the fai lur eof thr ee measure ment s to provide data. Thethre e 16-millimeter ca mer as provided flightdata, but the ser vic e module cam er a evident-ly w a s damaged by the explosive fo rc e fr omthe thrust termination of the LJ-I1 booster.

TABLE A-II.- w - 1 2 FLIGHT EQUIPMENT

Component

Telemetry modulation package

Telemetry rI package

C-band transponder

Tape recorder

Ti m e r

Signal conditioner box

Junction bo x

Main battery

Pyrotechnic battery

Power control box

Rate gyroscope package

Attitude gyroscope

Linear accelerometer

Relay box

Pres s u re t r an sduce r

Beacon line filter

Amplifier rack

Amplifier

Resis tance thermomete r

Temperature signal condltioner box

Temperature s imulator box

Commutator

Camera

-antity

1

1

2

1

1

1

1

1I

1

1

36

1

26

2

9

15

3

1

1

1

3

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APOLLOMISSIONAS-101, BP-13VEHICLE

Apollo mission AS-101 w a s the f i r s t flight of an Apollo s pacec raft configurationwith a Saturn boost er vehicle. The te st vehicle consisted of the booster (Block 11Satur n I and an S-IVB) and the spa cecra ft (launch escape sys te m, BP-13 command

module, and boile rpla te ser vi ce module). The DFI equipment i s listed in table A-111.

Per for man ce of the instrume ntation and communications sys tem s throughoutthe mission w a s sat isf act ory . Analysis revealed that 106 of the 112 mea sur eme ntsprovided reliabl e data.

TABLE A-III. - BP- 13 FLIGHT EQUIPMENT

Component

Tele metry modulation package

Telemetry rf package

C-band beacon

Power control box

Junction box

Battery


Amplifier rackAmplifier, dir ect curre nt

Pressu re t ransducer

Accelerometer

Vibration system

Strai n gage

Commutator

Resistance thermometer

CalorimeterZone box

Tempe ratur e s ignal conditioner box

Beacon line filter

Quantity

3

3

2

1

1

6

1

1522

25

7

3

48

1

16

2020

1

2

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APOLLOMISSIONAS-102, BP-1 5 VEHICLE

Apollo miss ion AS-102, the second flight of an Apollo spacecraft configuration witha Saturn launch vehicle , was launched fr om KSC on Septembe r 18, 1964. The BP-15spacecraft w a s used for this test. Figure A-3 is a block diagr am of the inst rumentati onand communications subsy st ems . A total of 133 mea su remen ts w a s required on BP-15during mission AS- 102. The DFI equipment used to obtain the required measurementsis listed in table A-IV.

The flight per formance of the BP- 15 instrumentation syst em w a s satisfactory. Ofa total of 133 mea su rem ent s, 131 provided continuous data . The two C-band beaconsprovided good tr acki ng information throughout the flight,

vhf omnidirectional antenna

C-band C-bandantenna

Power divider

transponder

antenna

Power divider

transponder

90 1-114

mnditionin g box

Zone baxes.temDeratu re

I

rf transmitter

I r - - - - - - - -

;Launch vehicle IMu1 icoupler linstrumen t un it;

ltelemetry 1L,----,-JI I '

B Crf transmitter rf transmitter247.3 MHz 257.3 MHz

Emitter fdlower

Mixer amplifiers Mixer amplifiers

Subcarrier S ubca rrie roscilIators oscillators

commutator

sensorensors andamplifier

sensorsI I -1Main signal-conditioningbax

junctions

Accelerometers

Calorimeters and

I thermocouples

Figu re A-3. - The instrumentation and communications subsystems used on BP-15.

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TABLE A-IV. - BP-15 FLIGHT EQUIPMENT

Comp.onent



90 x 10 commutator90 X 1-1/4 commutator

Main s igna l conditioner box

Temperature signal conditioner box

Amplifier rack

Amplifier

Vibration syste mPre ssu re t ransducer

Accelerometer

Strain gage


Thermocouple

Calorimeter

Zone box

MicrophoneBall assembly

Quantity

33

1

1

1

1

1 1

21

624

7

48

20

9

20

20

11

APOLLOMIS S l O NA-003, BP-22 VEHICLE

Th e BP-22 spacecraft, which was launched on an LJ-I1 booster, contained230 DFI measurements. The flight equipment is listed in table A-V. The flightper for man ce of the B P - 2 2 instrumentation system was proved satisfactory by thepro per function of al l 230 measurements during Apollo mission A-003. However,significant information w a s not obtained from some of the instrumen tation because ofan early abort of the flight.

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TABLE A-V. - BP-22 FLIGHT EQUIPMENT

Component


Tape modulation package

Commutator


C-band beacon

Beacon line filter

Tape record er transport

Tape recorder electronics

Tape recorder line filt er

Camera battery

Logic battery

Coded timer


Signal distribution box

Power control box

Main batter y

Junction box


Attitude gyroscope

Junction box

Inverter

Pres sure t ransducer

Accelerometer

Amplifier

Accelerometer

Pressure transducer (acoustic)

Amplifier rac k

Br eakwir e measurement adapter

Frequency detecto r

Quantity

2

10

1

2

2

2

2

2

2

3

2

1

1

1

1

2

1

1

2

1

1

44

6

49

349

2

1

1

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TABLE A-V. - BP-22 FLIGHT EQUIPMENT - Concluded

Component

Voltage monitor, alterna ting curr ent

Temperature signal conditioner box

Zone box

Thermocouple

Calorimeter


Camera

Camera mount

Came ra control box

Pulse generator

Strain gage

Displacement t ransduce r

Quantity

1

1

19

3

19

6

3

3

3

3

8

1

APOLLOMISSIONA-002, BP-23VEHICLE

The BP-23 pacecraft, which was launched on an W-11 ooster, contained 190 D F Imeas ure ment s. The flight equipment is listed in table A-VI.

Of th e 132 tape-recorded channels, 128 provided s atis fac tory data d uring Apollomission A-002. The C-band beacons were interrogated by three PPS-16 ada rs duringflig ht. The command module w a s tracked throughout the flight, and satisfactory returnsignals were receiv ed fr om both transponders. The tower and command module ca me ra sperf orme d satisfa ctorily. The service module ca mer a provided only 4 seconds of usablefilm before it jammed because of mechan ical malfunction.

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TABLE A-VI. - BP-23 FLIGHT EQUIPMENT

Component

AmplifierResistance thermometer

Pressure transducer


Junction box

Transponder

Beacon line fil ter

Relay box

Pres sure t ransducerAmplifier



Onboard ti mer

Tape recorder


Rate gyroscope

Attitude gyroscopePower control box

Main battery

Pyrotechnic battery


Strain gage

Displacement transd ucer

Camera

Ca mer a mountCamera control box

Camera pulse generator

Quantity

33

53

2

1

2

2

1

1 1

16

1

1

1

1

1

1

3

1

1

7

4

8

1

3

3

3

3

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APOLLOMIS S l O NPA-2, BP-23AVEHICLE

The vehicle DFI used on BP-23A (pad-abort mission PA-2) w a s the refurbishedsys tem that was used on BP-23 for Apollo mission A-002. The instrumentationsubsystem provided good quality data f rom all measurem ents on the vehicle. Complete

telemetr y data coverage w a s obtained fr om the ground stations .

The two onboard ca me ra s, one mounted in the launch esc ape sy ste m tower andone in the command module, oper ated as programed by the subsystem control box. Thetower c am er a showed launch esca pe sy st em motor plume impingement on the commandmodule until the lens w a s sooted during tail-off of the launch escape moto r. The com-mand module came ra , sta rte d at 5 seconds after launch, gave excellent covera ge ofapex-cover jettison, tower jettison , and all parachut e deployments.

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APPENDIXB

COMMAND ANDSERVICE MODULE PHASE

INTRODUCTION

The command and service module (CSM) hase of the te st prog ram used vehicles

These vehicles were used to confirm the structural integrity and thethat were co nstr ucte d to the production model configuration of the man- ra ted Apollospacecraft.operational capabilities of the Apollo subsystems. A ll Apollo vehicles li ste d in th isappendix, with the exception of spacecraft 002 (SC-002), were launched from the NASAJohn F. Kennedy Space Center on Saturn boosters.

APOLLOMISSION A-004, SC-002VEHICLE

A tot al of 242 onboard mea sur eme nts was req uir ed on Apollo mission A-004.Satisfactory data were obtained from 237 of the 242 onboard mea sur eme nts . Fai lu reof the single operational sc im it ar antenna (because of damage by protec tive-cov erjet tiso n) following abort caused a loss of both radio-frequency telemetry links at 74.7seconds after lift-off. However, data tr ans mit ted by these links were recorded on theonboard tape re cor der and recovered. The spacecraft w a s launched on a Little Joe IJboost er . The SC-002 flight equipment is listed in table B-I.

TABLE B-I. - SC-002 FLIGHT EQUIPMENT

I Component I QuantityTelemetry package

Tape modulation package

90 X 10 high-level commutator

Main i nstrument battery

Tape transpor tTape electronics

C-band transponder

Line fi lt er C-band transponder

Signal distribution box


Powe r control box

2

10

2

2

22

2

2

1

1

1

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TABLE B-I. - SC-002 FLIGHT EQUIPMENT - Concluded

Component

5-volt r ef ere nc e supply25-volt reference supply

Time reference generator

Logic battery

Amplifier, direct current

Microphone assembly

Amplifier, alternating curr ent


Attitude gyroscope systemGyroscope package

Inverter

Junction box

Pres sure transducer

Amplifier, al ternating cu rre nt

Pres sure transducer

Angular dis placeme nt

Resistance thermomete rThermocouple

Calorimeter

Zone box

Accelerometer


Strain gage

Camera

Camera mountCamera control box

Camera pulse generat or

Camera batterv

Quantity

11

1

2

98

1

1

1

12

1

1

51

31

2

2

133

6

9

29

6

164

1

1

1.

1

1

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APOLLOMISSIONAS-201, SC-009 VEHICLE

A n instrumenta tion package in the C S M nterfac ed with the communicationssubsystem and processe d 37 measuremen ts from the str uctur al subsystem, the in-strumentation syst em, the serv ice propulsion syst em, and the reaction control syste mduring Apollo mi ss ion AS-201. An instrumentation package in the s er vic e moduleinterfac ed with the communications subsystem and proc ess ed 93 meas urem ents f romthe str uct ura l subsystem. The SC-009 and service module flight equipment is listedin tabl e B-11. The perfo rma nce of al l development flight inst rumentat ion (DFI) equip-ment w a s sat isfactory.

TABLE B-II. - SC-009 AND SERVICE MODULE FLIGHT EQUPMENT

Component

Tele metr y package

Modulation assembly

Transmitter

90 x 10 high-level commutator

5 -point calibrator

90 x 10 low-level commutator

Time-code generator

Command receiver

Auxiliary decoder

Command receiver voltage regulator

Modulation kit pulse amplitude modulation(PAM)/frequency modulation (FM)/FMsystem

5 -point calibrator

Quantity

1

2

2

2

3

3

1

1

1

1

1

1

SPACECRAFT010

A DFI system for SC-010 was designed, procure d, checked, ca librated, and in-stalled on a breadboard, The integrated sys tems checkout of the breadboard was com-pleted succ essfully , and the breadboar d was shipped to the contract or facility. However,because of the success of he SC-002 mission, SC-010 wa s not flown. Components we rereturned to bonded storage areas at the NASA Manned Spacecraft Center to be used inothe r instrumentation systems.

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APOLLOMISSIONAS-202, SC-011 VEHICLF

The D F I system for SC-011 was composed of the ite ms liste d in table B-111. Aninstrumentati on package interface d with the communications subsyst em and pro ce sse d60 measurements from the stru ctur al subsystem, the operational instrumentationsyst em, the s erv ice propulsion syste m, and the reaction control system. A secondpackage processed 9 3 additional measurements from the structu ral subsystem. The fourca me ra sys tem s were used to rec or d panel and apex-cover deployment. All DFI equip-ment opera ted sat isfa ctor ily during Apollo mis sio n AS-202.

TABLE B-111. - SC-011 FLIGHT EQUIPMENT

Component

Tel eme try package

Modulation assembly

FM transmitter

9 0 x 10 high-level commutator

5 -point calibrator

Tape re cor der support equipment


Time-code generator

Modification kit PAM/FM/FM syst emCamera

Camera power distribution

Timing pulse generator

Battery

Baroswitch

Cont rol box

Quantity

12

2

2

3

3

3

1

14

1

1

2

1

APOLLO3 M IS S ION, SC-012 VEHlCLE

Instrumentation equipment was supplied for the Apollo 3 mi ss io n (SC -012) butwas nev er require d because of the acc ident and fire within the space craf t.

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APOLLO4MISSION, SC-017 VEHICLE

The DF'I equipment used on the Apollo 4 mission is listed in table B- N. A self-contained ca mera s yst em was included for purposes of photographing the Ea rt h froma 10 000-nautical-mile apogee. Ove ral l performance of the SC-017 tel eme try sy st em

and camera system w a s satisfactory.

TABLE B-IV. - SC-017 FLIGHT EQUIPMENT

Component

Modulation package

90 x 10 high-level commutator90 x 10 low-level commutator

Time-code generator

Camer a system

Camera

Lens

Cam era control box

"Gf' switch

Battery

Quantity

1

22

11

1

1

1

1

1

APOLLO6 MISSION, SC-020VEHICLE

The SC -020 vehicle contained instrumentation and came ra sys tem s (table B-V)similar to those used on the Apollo 4 mission. Equipment was provided to acquireEarth-mapping and CSM plume-impingement photographs. Per for mance of thes e sy st em swas sati sfac tory, and useful data were obtained throughout the Apollo 6 mission.

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TA B L E B - V. - SC-020 F L I G H T E Q U I P M E N T

Component

Modulation package90 x 10 high-level commutator


Time -code gene rato r

Boost/entry camera

Boost/entry-camera control box

Earth-mapping camera

Earth-mapping-camera control box

I ' G" switchBattery

36

~~~

Quantity

12

2

1

1

1

1

1

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A P P E N D I XC

LUNAR MODULE PHASE

INTRODUCTION

In addition to the flight vehicles, sev era l ground test vehi cles were const ructedfo r the luna r module (LM) phase. The LM test art icl e (LTA) vehicles contained onlydevelopment flight instrumenta tion (DFI) syste ms and were designed to gather informationon vibration and st ra in and to obtain acoustical data. Later configurations of the LMvehicle contained additional sub syste ms and mor e sophisticat ed instrumentation.instrumentation measured additional par ame ter s such as temperature, pressure, and

The

acceleration . A C-band ra da r w a s al so included.

HOUSESPACECRAFT1

A total of 241 DFI measurements w a s initially installed on house space craf t 1(HS-1). The equipment was similar to that used on LM-1.structed for ground test purp oses and for in-house training.

The HS-1 vehicle was con-

LUNAR MODULE TEST ARTICLE1

The DFI syst em installed in HS-1 was later r eins tall ed in LTA-1. The LTA-1vehicle was constructed fo r ground test and training purposes.

APOLLO4 MISSION, LTA-10VEHICLE

A total of 38 DFI measur ement s was made on the'LTA-10 vehicle for the Apollo 4mission (table C -I). The instrumentation system performed satisfactorily except forthr ee ac cel ero met ers that exhibited temporary bias offsets during maximum dynamicpr es su re conditions. Thi s vehicle also car rie d the first constant bandwidth (CBW)modulation package to be launched into space.

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TABLE C-I . - LTA- 10 FLIGHT EQUIPMENT '

I Component

Modulation assembly, CBW

Modulation assembly, propor tio nal bandwidth (PBW)

90 x 10 high-level commutato r

25 -volt power supply


Power cont rol box

Temperature resis tance thermomete r

Vibration system

Linear acc eleromet er

10-watt transm itte r

Acoustic pre s sure transducer

Amplifier

5-volt r ef ere nc e supply

Strain gage

Quantity

1

1

1

1

11

6

9

52

2

22

1

64

APOLLO6 MISSION, LTA-2 VEHICLE

The DFI syste m for the LTA-2 vehicle was very si mi la r to the one installed onLTA- 10 and performe d s atis fac tori ly throughout t he Apollo 6 mission.

APOLLO5 M I S S I O N ,LM-1VEHICLE

A total of 195 DFI measure ments was installed on the LM-1 vehicle fo r the Apollo 5mission. The flight equipment used is list ed in tabl e C-11. The instrume ntation sys te mperformed s atisfac torily, with the exception of seve n me as ur em en ts that did not re-

spond as planned. Four thermocouple me as ur em en ts malfunctioned (probably becameunbonded), and three low-pressure measurements failed (probably because of con-stri cted ports from engine firing).

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TABLE C-II. - LM-1 FLIGHT EQUIPMENT

Component

10-watt transmitter

Tele metr y modulation assembly, PBW

Telemetry modulation assembly, CB W

Signal conditioner

90 X 10 high-level commutator

90 X 10 low-level commutator

Vibration commutator

90 x 1-1/4 low-level commutator

Power control box

5-volt reference supply

C-band beacon

C-band filter

25-volt re fer enc e supply

Amplifier

Strain gage

Pressure transducer

Vibration system

Temperature resistance thermometer

Temperature thermocouple

Zone box

St r ain/t empe r atur e system

Quantity

5

3

1

1

2

1

3

1

1

1

2

2

1

38

80

45

18

36

25

41

2

LUNAR MODULE2

Th e DFI system on LM-2 w a s near ly identica l to that installed on LM-1. Thesystem was completely checked and accepted at the contractor facility before thedecision w a s made t o place LM-2 n a nonflight status.

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APPENDIXD

MANNED FLIGHT VEHl CLE PHASE

INTRODUCTION

The launch of the Apollo 7 spacec raft on October 11 , 1968, marked the beginningof the manned Apollo vehicle flights . Development flight ins trum enta tion (DFI) wasprovided by NASAfor thr ee manned missions: Apollo 7, Apollo 8, and,Apollo 9.

APOLLO7 M I S S I O N , C O M M A N D A N DSE RV l CE MO DULE101VE Hl CLE

A total of 167 DFI mea sure ment s was made on command and ser vic e module 101

(CSM-101) during the Apollo 7 mission. The flight equipment is diagramed in figure D -1and list ed in tabl e D-I. The telemetry syst eri performed satisfactorily except for onehigh-level commutator that became er ra ti c during the r eent ry phase.

APOLLO8 MISSION, CSM-l tY3 VEHICLE

A total of 36 DFI measure ments w a s made on CSM-103 during the Apollo 8 mission.Equipment sim ila r to that used fo r CSM-101 (table D-I) wa s used fo r CSM-103 (table D-II).The telemet ry sys tem performed satisfactorily during the mission.

APOLLO8MISSION, LTA-B VEHICLE

A tot al of six DFI measurements w a s used on LTA-B fo r the Apollo 8 mission.The instrumentation performed satisfactorily during the mission.

A P O L L O 9 M I S S I O N , L M - 3 V E H I C L E

A total of 248 DFI measurements w a s made on LM-3 during the Apollo 9 mission.

The.fli ght equipment used is list ed in table D-III.248 meas urem ents indicated an average return of 98.7 percent of the data over the10-hour period of sy ste m operation.

The over all evaluation of the to tal

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~ i i q h tualification tape recorder

ParametersCommand module

StructureTemperature (251Pressure (91

Temperature (481Flux rate 1141

Vibration (21

Guidance and navlqatiResolver 161Vibration 131Events (21Voltage (101Temperature (11

Environmental controTemperature (61 Operational pulse codemodulated telemetry

Service moduleEnvironmental contro

Temperature (2 )

Service propulsionTemperature (5 1

StructureVibration 18)Temperature (61

NoteThe thick black lines indicateOF1 equipment.

CB W* constant bandwidthDF I-development flight instrumentationlR lC-interrange instrumentation groupPDM-pulse duration modulationre f -reference

Key.

Figure D-1, - The CSM- 01 flight-qualification instrumentation diagram.

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TABLE D-I. - CSM-101 FLIGHT EQUIPMENT

90 X 10 high-level commuta tor


Modulation package, constant bandwidth (CBW)

Time-code generator

ComponentI2

1

21

Modulation package, pro por tion al bandwidth (PBW)

Time-code generator

TABLE D-II.- CSM-103 FLIGHT EQUIPMENT

1

1

1 Component 1 Quantity

TABLE D-111. - LN1-3 FLIGHT EQUlPMENT

Component

10-watt transmitter

Telem etry modulation assemb ly, PBW

Telemetry modulation assembly, CBW. Signal conditioner

90 x 10 high-level commutator

90 x 10 low-level commutatorVibration commutator

90 x 1-1/4 ow-level commutator

Power control box

5-volt r efer ence supply

C-band beacon

C-band filte r

25-volt reference supply

Strain gage

Press ure transducer

Temperature resistance thermometer

Tem pera ture thermocoupleZone box

Strain/temperature system

Vibration system

Press ure transducer

Amplifier

Power control timer asse

Apollo Experience Report Development Flight Instrumentation

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