Test and Evaluation Aspects of the Nimbus II Program Useful to Other Long Life Space Programs

The Space Congress® Proceedings 1968 (5th) The Challenge of the 1970's

Apr 1st, 8:00 AM

S. Charp Consulting Scientist, General Electric Company, Missile and Space Division

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Scholarly Commons Citation Charp, S., "Test and Evaluation Aspects of the Nimbus II Program Useful to Other Long Life Space Programs" (1968). The Space Congress® Proceedings. 4. https://commons.erau.edu/space-congress-proceedings/proceedings-1968-5th/session-17/4

This Event is brought to you for free and open access by the Conferences at Scholarly Commons. It has been accepted for inclusion in The Space Congress® Proceedings by an authorized administrator of Scholarly Commons. For more information, please contact [email protected]. TEST AND EVALUATION ASPECTS OF THE NIMBUS II PROGRAM USEFUL TO OTHER LONG LIFE SPACE PROGRAMS*

S. Charp Consulting Scientist . General Electric Company Missile and Space Division Philadelphia, Pennsylvania

the spacecraft's development history to the on-orbit performance? The Nimbus II satellite has been in orbit 20 months (to January 15, 1968), operating well with The planned life of Nimbus was in design specifications as a six months at carrier spacecraft the time of the start of the for meteorological instruments. program, long by any then-existing demonstrated performance. The payloads were new and at the edge of This spacecraft includes the state-of-art . a three axis active Man's knowledge of the space environment attitude control system which was quite itself is more com meager. As a space carrier for plex than many other satellites experimental-type and their included meteorological instruments, the life payloads. An essential element required of of the total the spacecraft should be at least development of Nimbus that of the most is the intensive and exten significant instruments being sive test and evaluated. Con evaluation programs at the General versely, to be most effective, Electric Company, a spacecraft capable to which the prototype/qualifi of a six-month life should carry cation and flight instruments like spacecraft were exposed, coupled wise designed to live for six with selected calendar months. design features which contribute to Thus the life requirements and demonstrated long life. Although per a rigid causal relation can formance for the spacecraft and its housekeeping not be established between the programs' adminis electronics equipment can be treated separately trative and technical activities associated with from that of the "passengers". the design of the spacecraft, integration of payloads and the test and evaluation program, and What constitutes the the orbital success of Nimbus long-life on-orbit performance experienced, II? In-orbit success they must be measured only in are considered important contributing factors. terms of the performance This of the bus. Did the bus paper reviews: the essential elements of the function in orbit General in a way to allow the payload Electric Company's program, long-life instruments to perform aspects their intended functions of the spacecraft f s basic design, the over a time period fundamental of at least six months? From philosophies which guide the test and the point of view evaluation of the research and development program, the multiple nature of the mission, the entire Nimbus test and evaluation II orbital program was program, review and analysis an unqualified success. To a somewhat of failures encountered in lesser the systems test pro degree Nimbus I was almost as successful grams from fabrication to alunch in spite and of anomalies of its near-one-month life, because in its encountered in the orbital performance, time and some environment the total data taking mission representative lessons learned was more from the program successful than had actually been anticipated, which can be applied directly to other long life prior to flight, by those individuals close to the space programs. program. I. Introduction To what factors can the a posteriori success of a single long life spacecraft The Nimbus II be attributed? spacecraft was launched into a Causal relations are hardly near polar orbit in the likely to be devel early morning of May 15, oped between the flight performance 1966, and has been performing of a single its planned func spacecraft and planning methods, tions well within its procedures, tests design specifications for and evaluations. A more meaningful over 1.6 years at the answer to this time of this writing. The question can be had by outlining Nimbus I spacecraft was the activities at launched into a similar the General Electric Company which orbit in the early morning preceded the of August 28, 1964, successful flight; and relating these and lasted for 26 days, to a. review at which time the solar and analysis of the failures which array drive ceased occurred during rotating and power was flight. If similar analyses were depleted. Both spacecraft made with were unqualified respect to other programs of like complexity, successes in 'terms of their useful operating life which, resulted in long .life (greater than and in terms of the vast one amounts of meteorological year) orbital performance, and the data acquired by the payload respective data f s sensors and were examined for commonalities, then collected by the data center. perhaps These spacecraft collectively causal relations could be derived. were launched without benefit of prior engineering flight tests to assess potential orbital weak Constrained by data accumulated from only a nesses. To what degree is it meaningful to relate single example of a long life complex spacecraft, it is clear a causal relationship which defines a - differential sensitivity between 'cThe National Aeronautics and 'What and low Space Administration* things were done and in-orbit Goddard Space Flight performance cannot Center, Nimbus Project Office,. be. established on the Greenbelt, Maryland, basis of scientific, techni has Systems Management respon cal, and logical considerations,, sibility for the series of Neither is it: Nimbus spacecraft* Under meaningful to dogmatically state contract to this group, the and assign, General Electric requirements for activities Company, Spacecraft- Department, which.,, if conducted, designed and devel would insure successful orbital performance, oped the three axis, active with stabilization and atti a p reassigned, numerical assessment tude control system; designed, developed! and of reliability* fabricated the spacecraft structure, thermal con trol system, and various components; integrated experiments and meteorological sensors; and tested the complete spacecraft. 17.1-1 II. J^. JjnbugL P rogram o f Ac t ivi t ie s 3. Mechanical and electrical interfaces were very loosely defined for Nimbus I, but some The GE responsibilities as a major contractor what more tightly for Nimbus II. to the NASA/GSFC Systems Manager were: 4. Sensor/pay loads and spacecraft design 1* Conduct detailed electrical and mech represented advanced state-of-art , with their anical design studies. understanding and appreciation maturing as the efforts progressed. 2. Design, build, test, and evaluate the stabilization and attitude control system to ' 5. A comparatively short development time broadly defined goals/specifications. existed,, considering the absence of a technical baseline design from which to grow, 3. Design, build, test, and evaluate the passive and active thermal control system, The complexity of the control system, with its own composition of basic elements provided by 4. Design,, build, test, and evaluate the many internal design, groups and external vendors, antenna systems, based upon a feasibility study resulted in problems of program management similar provided by the University of New Mexico, to those encountered by the multiplicity of CO' -con tractors. Supporting the GE in-house efforts 5* Design, build, test, and evaluate the was the entire complex of NASA/GSFC, including, spacecraft structure to the general outline and but not limited to personnel from groups concerned systems design provided by NASA/GSFC, with: equipment., and component design; materials; test and evaluation; reliability; tracking and. data 6. Design, test, and evaluate all models* systems ; and most important , the NASA/GSFC Nimbus Including: full-scale mock-up (pre-prototype) ; Project group itself* 'The devoted effort of the prototype ; flight; structural dynamics; antenna; NASA/ Indus try team jointly attacking problems as thermal; separation. they arose in these, sample areas may be considered am essential factor in the resulting long life. 7* Integrate all equipments/components into the several models spacecraft (engineerings, prototype/qualification, flight)* conduct inte Ill , Ba s ic De s i.gn Cons idera t ions grated systems level tests and evaluate the per formance under a wide range of environmental A. Pes ign Ph i losop fay conditions. Except for the comparatively simple spin stabilized spacecraft, which have demonstrated 8. Provide the adapter structure to the extremely long; life records, Nimbus has demonstrated launch vehicle. the longest life in orbit of the complex spacecraft. 16 a not inconsiderable extent, this performance nay 9. Design and fabricate all special test be attributed to the depth to which this spacecraft and ground handling-equipment for the spacecraft* and its related ground systems vere conceived, analyzed, planned t and conceptually integrated. 10'* Integrate the ground stations required • before extensive detailed design effort occurred. to coronand» teat, evaluate the spacecraft at In essence, .Nimbus is well "system engineered", the plant in Alaska• with the system design obviously being the result of thorough consideration having been given to lit Design and provide computer programs mission objectives and test, qualification and for engineering evaluation at the plant and for evaluation requirements » operational evaluation during flight* The basic hardware design was planned to 11* Provide operational systems engineer provide means for extensive evaluation of the space ing to evaluate in*flight performance craft through its environmental test program, and lift*off to completion of in»orbit system checks* s isolated ., opera t iona lly-controlled , orbital flights In both ambient and vacuum/ thermal environments* II* Provide launch support services* Conservatively high estimates of the environmental stresses expected to be encountered by the space 14* Bstablish, staff» and operate the craft were selected, these being embodied In both Technical Control Center it M,SA,/iSFG in the environmental specifications and designs » suppo11 § f orbi ta1 oper*tions. overall system design recognised the For a number of reasons* teat and evaluation dearth of quantitative design data and the prepon (concept if "test/fix11 ) occupies mm essential derance of qualitative design factors* A priori portion of the activity; confidence of success fill orbital performance provided to the engineering design with a **test» I* Minibus Is a complex spacecraft* having evaluate* re-design » retest** or **test - fiat - test* i, cwpltx control system and i, variety of individ* philosophy :s such that iterative application of 'lit! senior loads» these elements would result In a control system in m spacecraft which would have but twill tlfcell* 1* Many co*contree tors, each an indepen hood of failure after launch. dent contractor to NASA/GSFC» provided equipment services to tint spacecraft tt * but iaii no tittrit eontrtetuel * I, formil redundancy policy did wist, responsibility to the integration end teet ecu- nevertheless* fu&etio&fcl end block is it, 1 provided as failure protection at the system, level 4, __ in several ways. These can be categorized as Failure mode and effects analyses were conducted being related to the overall aerospace/ground on much of the control •system* especially during system, the spacecraft and its "housekeeping" the early phases of design where its impact was systems3 and the sensor instrumentation payloads. greatest. This technique was just being developed Figure 1 shows the redundancy built into the as a design tool in I960 so its application here stabilization and attitude control system. was somewhat of a "first". Later this analysis was if plied to some of the integration and test The large number of spacecraft oper effort including the operational programs * ations which are controllable from the ground by command provide a failure protection mode* By 5. jfcdglg_ior_De^t^!i - light model space* ground command all but essential subsystems can craft or major sections (pre -prototype s antenna be set to their off or standby modes of operat ion, And adapter* ' thermal, structural dynamics, separa resulting in a standby mode or minimum spacecraft. tion,. prototype/qualification, electrical systems , This capability has proved to be quite important flight) were fabricated, tested, and evaluated as during all phases of system testing as well as farts of the total design cycle. Each is con during orbit, when it was important to "sit tight" sider® J to have contributed significantly to the and reflect upon performance preparatory to final design* hence their inferred contribution issuing commands. The ground command capability, to a long*life design. for example, allows selection of different payloads, modes of operation^ voltage levels for 6. Flece Parts- - The Nimbus spacecraft motors, and redundant elements in the command contains approximately 40300 electronic -type piece receiver. It is the key element which allows farts* The engineering and pro to type/qua lifica - energy to be managed during the orbit, thus insur tion of the control system were built with ing batteries are not discharged at rates higher coflMercial quality parts (5000) . No testing than or to levels lower than allowed by their or evaluation of parts was done except for brief design. Most pre-programmed on-orbit operations test ing of functional performance. can be over-ridden from the ground through real time and stored commands. The flight model spacecrafts were built vith the commercial quality parts. However, 2. Con f igura t ion ... Mana gement - Forma 1 by the time farts were to be ordered and accumu configuration management was not an engineering lated for producing three flight model spacecraft, design requirement. Following release of engi required power aging (burn -in) of piece neering data, design change control formally part a 9 especially transistors and diodes, restricted alterations to designs and drawings. A Design Change Control Board met almost daily Occasional circuit failures caused by to review and act upon proposed changes required piece parts occurred* but elements were to continue activity, normally general improve replaced during the almost continuous testing to ments in designs were not acceptable reasons for wnich the control system, was exposed. Tine: "make a change - rather changes were authorized pri and test" philosophy which was central to the marily because of design errors, manufacturing development of the control system, served effec or materials problems. tively as paver conditioning and -parts selection controls. The degree to which activity 3. Interface Controls - General Electric "cost effective" Is problematical* and was never Company practices were followed for defining assessed* electrical, mechanical, and thermal Interfaces for all components and subsystems provided B, De s i gn Be a t tires internally. These were controlled, by the docu mented engineering data multistage release .1, " Modular layout - 'Hie design system in effect then. of the spacecraft provided fieri: 'two structural by a struc Interface information necessary for ture, the upper unit being stabilization design of mating components and elements being attitude controls 9 developed by'independent internal groups flowed pay loads. long-life of Is readily from one to the other* allowing changes essentially that of - teat- to be made informally during the formative design ing which could he period.. .An extensive and most valuable set of informational drawings were developed, each set Itself Is of devoted to a particular part of design* e*g«; design, of IS a) signal flow; b) command flow; c) grounding to be in -or •with and d) electric power and distribution* Bach' out: «o;f cover plates* showed interfaces in detail* with brief verbal by of either m mx: m torn and technical descriptions of the input/output is a •matching stages (voltages, impedances, mi constants, etc.). t>f flexibility tut ximg*

'^• j * '" 'he 'Spacecraft Jiiiitsox Ting, with itss titts'talled equip- tm be mm is mounted to the sensor ring framework, with "space Hoisting the spacecraft and locating it approxi grease" at interfaces, providing good thermal mately 15 feet from a reflecting surface minimized conduction, and in a sense leading the heat to this type of interference. be dissipated from the spacecraft to those regions of the ring having either flat panels coated with Had electromagnetic susceptibility paint having a specified ratio of absorptivity requirements been placed on the engineering design to emmissivity or to panels having Venetian blind and performance specifications for each subsystem, type louvres. Their slats opened in response to most of the noise sensitive problems encountered the mechanical motion of a bellows filled with during the electrical systems test would never freon which has a large coefficient of thermal- have occurred nor would the flight data exhibit the expansion. In the event the primary bellows degrees and kinds of interference which they do. failed, a secondary bellows acted to move the slats to a fixed 30-degree open position. Failure 4. Maintainability - Two features of the has never occurred. conceptual design characterize the consideration given to maintenance. The first of these has A truly benign environment was achieved already been noted, namely, the modular design of in the sensor ring on both Nimbus I and Nimbus II. the sensor ring, the standardized sizes of enclo Whereas the system design called for a temperature sures and the placement of connectors on hardware of 25° + 10 C, the actual temperature range mea allowed easy removal and replacement of components sured during orbit has been close to + 2°C. The during the integration test and evaluation activi thermal design for the control system was also ties. The high degree of selective redundancy extremely conservative. Again passive (paint) allowed alternative operating modes thus minimiz and active (louvres) systems were used. The ing down-time during tests and what is, of course, actual temperature range measured during orbit has most important, failure of the mission in orbit. been close to + 5 C. An important feature of the Nimbus 3 . Electromagnetic Compatibility Control - design not to be overlooked is the inherent capa Potential electromagnetic interference among the bility which exists for remote "maintenance and subsystems through conduction and radiation was repair" while the spacecraft is in orbit. The recognized in the very early phases of system capacity of the command system, directed from the design. However, specifications for the design ground, to accept many alternative directions and and performance of subsystems did not include the orders, coupled with the data relayed through the subjects of limiting the levels of electromagnetic telemetry system, provided a total maintenance interfering signals within the subsystems nor pro capability which has proven most effective during tection against incoming interfering signals the life of Nimbus II. This concept of in orbit (susceptibility). Only mild requirements were maintenance and repair through cooperative ground included in the power subsystem limiting the gener activity should be expanded and applied as part of ation of sharp pulses which could be placed on the the engineering systems design of new spacecraft. regulated bus. As a consequence, many electro magnetic interference problems were generated 5. Command - The large number of commands immediately as subsystems were interconnected and capable of ordering functions on the spacecraft during the first series of electrical systems provided the flexibility necessary for alternative integration tests. selection of operating payloads and orbital manage ment of energy. The system allowed 128 coded and The control system was particularly 4 unencoded (emergency) commands. Sixteen of the susceptible to conducted interference pulses prop coded commands could be stored in a memory, to be agated on the power bus, generated by switching actuated at specified times during a given orbit. transients. Most of this susceptance was elimi These commands could be revised once each orbit. nated by revisions in grounding and by filtering. This large number allowed equipments to be turned Camera shutters introduced substantial transients on or off individually and to select among the until the input circuits to the camera subsystems redundant alternative modes of operation, as noted were revised by substituting inductive for capaci- earlier. It is clear, had this capability not tative input circuits. Circuits in the clock and been available to the Nimbus II spacecraft, little telemetry subsystems were also susceptible to likelihood exists it would have remained as useful a power line transients, requiring changes to be spacecraft for the long period it has. made to them. 6. Telemetry - Several references already In addition to determining and elimi have been made to the role of the telemetry system nating the sources and effects of conducted inter and its alternative modes of operation as con ference, radiated interference existed. Leads tributing factors to the long life of Nimbus II. and cables had to be re-routed; shielded leads In spite of the failure of the telemetry system's replaced unshielded leads; unbalanced a.c. line tape recorder several months ago, almost the full pairs which were twisted were untwisted; balanced status of all subsystems can be monitored nearly a.c. line pairs which were not twisted were continuously through the real-time telemetry twisted; many cables were wrapped with metal foil, transmissions. This mode allows reception of status some grounded; and fine mesh screening was installed information by only those properly equipped receiv around cameras. ing stations in the tracking network. Without them real-time data energy management would have been An electromagnetic anechoic chamber not impossible and the entire spacecraft would have being available during systems tests, radiated failed at the time the telemetry system's tape interference entered the antennas, energy being recorder failed. reflected from the laboratory's floor and walls.

17.1-4 An, essential feature of the Nimbus space 1. Very conservatively define the total craft is the total dependence on the telemetry environment and the stress levels expected to be system for all systems testing, countdown, launch experienced by the subsystem and/or system (the and orbital phases. Although provision was avail" spacecraft). Include as environments: manu able for transmitting telemetry data through hard facturing, ambient, laboratory, handling, shipping, wire connection to a telemetry interfacing con storage, launch, space, nector, except during the earliest periods of testing this method was not used. To provide a 2. Separately qualify each prototype sub 'more realistic situation, all telemetry data were system and sensor/instrument payload before in transmitted through radiation from the telemetry stalling it into the spacecraft. Study of avail system's quadra loop antenna system, the data being able specifications indicate for some subsystems picked up by a receiving antenna (s) connected to loosely defined demonstrations of performance were the ground station's telemetry receivers. This acceptable evidences of qualification. Further, allowed assessment of the full telemetry system individual subsystems were not required, at the under all conditions including the time when the time of their initial development, to adhere to spacecraft was undergoing environmental tests in the same single set of evaluative tests to demon the vacuum./ thermal chamber, strate qualification of design for space use. Some subsystems required engineering development IV. gli^ beyond their anticipated scheduled completion dates, resulting in fully qualified hardware not being available for installation into the space A. Gene ra 1 Phi lo goghy craft at the times planned. The GE-operated bench acceptance test activities screened some mal Any effective test and evaluation program is functioning equipment, but it also allowed other the fulfillment of a logically planned effort, marginally-performing equipment to be installed the conduct of which is traceable to some under into the spacecraft. lying philosophy. The Nimbus test and evaluation effort, as actually conducted, is the eventual 3. A prototype/qualification spacecraft, realization of a set of guidelines established by complete with identical equipment, and payload NASA/GSFC prior to 1962. . It is understood these sensors as will be incorporated into the final have not changed materially in the intervening flight spacecraft, and in full operating condition, years. including electro-explosive devices, as required, should be tested in a series of environments simu Designing and manufacturing a complex space lating those anticipated, augmented by appropriate craft is more of an art than a science, and safety factors as noted in 1. above. The time subject to a large number of alterations and duration of testing in a vacuum/thermal environ constraints representative of many unknown, factors ment should be two weeks. The conservative assump and inter-relations which become apparent during tion is made that a spacecraft whose design is to the development period. A complex spacecraft is be evaluated which could not survive these tests a good example of a system which usually is not is poorly designed and is not suitable for a space fully operable upon "rolling off the end of the program. Obviously all components and subsystems, production line" - the whole is not the sum of its even those whose orbital lifetimes are short, must parts. The performance of a spacecraft, be it an survive the test period and not constrain the tests. engineering, prototype/qualification or flight model, must be evaluated empirically. Performance 4. Separately expose to vibration and of each model must be demonstrated under simulated vacuum/thermal acceptance tests each flight sub conditions representative of the environmental system and sensor/instrument payload before in stresses expected. In essence, tests must supple stalling it into the flight spacecraft. For the ment analyses. most part, equipments destined for flight use were exposed to the environments noted and to full accep This point of view is particularly applicable • tance tests. The problems and failures encountered to a space program which is essentially research during the integrated systems level tests indicated and development in .nature, and especially so in a even these tests were not, in many cases, effective program which, features a spacecraft of unusual means for screening potential failures. It is pre design and equipment which pushes state-of-art , sumed, where opportunity existed, equipment intended Conventional approaches to establishing reliability for Nimbus II was tested to a better degree than and confidence levels, based upon measurements made similar equipment which was incorporated into Nimbus on a relatively large population, simply do not I. This trend is continuing for successive Nimbus pertain. Evaluating the performance of the com spacecraft. pleted spacecraft, however, must not be an isolated activity; it must be viewed in the context of the 5. Similarly, each flight spacecraft should activities of a total program, i.e., test and eval be tested to a set of environmental stresses which uation activities should not be isolated from other define an. acceptance test. Obviously, to conduct activities. Accepting this premise leads to con such program requires the spacecraft: to be amply sidering test and evaluation requirements very early provided with command capability to allow it to be in the planning stage of the total program. thoroughly exercised during the environmental tests; to be amply instrumented with telemetry to allow A basic philosophy to guide test and evalua rap id asses sment o f per fG nuance of s ub systerns; and tion. activities, expecially at systems level, can to communicate between the ground control reception be expressed in seven brief statements. point and the spacecraft through only a radio link

17.1-5 (no hardwire), The Nimbus spacecraft is designed It was therefore necessary to institute a to allow testing to be so accomplished. vigorous and carefully documented acceptance test program for all equipment provided by co-contractors, 6. A strong policy should be followed mini with product assurance controls patterned somewhat mizing the replacement or removal of any equipment after those required by the total spacecraft* from the spacecraft between the start of flight This program only allowed either acceptance or acceptance tests and launch. Access to, adjust rejection of the elements of or a complete sub ment, or disassembly of the spacecraft or its sub system, based on functional performance tests in systems should be discouraged. The system should an ambient laboratory environment. Based on accum remain intact and last-minute changes avoided. ulated experience, it is postulated 'many functional Make no hardware changes at the launch site. performance problems which were encountered during the systems test program would never have occurred 7. Prior to start of the actual test and had each of the subsystems been operated under evaluation program, the participating active appropriate environmental stress conditions for a ground equipment, facilities, handling mechanisms, time duration sufficient to establish little further detailed test procedures, and test conductors and likelihood of additional failures. supporting personnel should, in a sense, be "qualified", V, Test and Evaluation Programs

B. Functions of the GE Test and Evaluation Program A. Ba s ic C qns i d er a. t i on s

The primary functions of the GE Test and Eval To a great extent high emphasis was placed on uation Program can be summed as: the sp a c ecra ft' s sy s t em, 1 eve 1 te s t and e va lua t ion p ro g ram ( " - - - - -de s i gn. -1 es t - fix -1 es t - fix- - - - - '*') Estimate orbital environments because it served as an overall screening function, to detect and indicate corrective actions which, Establish and prepare general, specific, and had not previously been indicated in the individual detailed test specifications and procedures subsystems because of the limited testing; to which they had. been exposed. System, level testing.,, how Conduct subsystem tests for equipment pro ever, should n.ot be considered as an isolated. vided by GE activity; its role must: be part of a comprehensive test and evaluation program,.. Conduct systems test and evaluation using antenna, structural, thermal, electrical, A stru.ctu.ral concept of the test and evalu prototype/qualification, and flight models ation program, fox Nimbus (and one which is also applicable to other programs) is shown in Figure 2, Redesign or recommend redesign., as required, Each level of testing is designed to discover weak- based on results of the test and evaluation ne s s e s , a noma 1 o u s b e h a v i Q' r a nd. fa i. lur e s win i, c h program passed through a lower level testing,, and. those similar characteristics of performance which, may be Provide certified personnel, procedures, and generated by interactions created by the .functional/ facilities for all tests except the accelera physical forms associated with that level. tion test Theoretica1ly, i f a suffieiently comprehensive Subsystems were intended to have been, evalu and, effective test and evaluation program bad 'been ated prior to their integration into the spacecraft consummated at Level i, failures in testing; at for system level tests, to the same regime as Level (i+1) should not be anticipated as cowing planned for the system,, i.e ( , -engineering 3 proto from Level i, but rather as coming from unantici*- type/qualification or flight acceptance test levels pated causes peculiar to' the total environment to their own specific set of specifications,. Detec "associated with Level (i+1) . Corrective action. tion of marginal performance conditions, early could therefore be localized and minimized. wear-out and failures were the primary results to be achieved from a strong subsystems test program, Levels 1 to 4 pertain, to the design, ties t, and, evaluation of subsystems. T'hey are designed to Notwithstanding the detailed overall program demonstrate: achievement of design,, goals, perfor planning to have available for systems level test mance, structural and functional integrity, and ing fully qualified, ready to use subsystems, assurance particular pieces of hardware are in fact schedule changes in various parts of the program of sufficiently high quality to be incorporated resulted in some co*contractor's equipment being into a., flight spacecraft* Presumably, if the flight delivered to the General Electric Company in a acceptance test program at subsystem level does not-fully performance-evaluated condition* Each realistica1ly take into account anticipated environ piece of co-eontrietor 1 s equipment ©r sensor pay* mental stresses, and if the subsystem has been load was to have passed its prototype/qualifica tested for a time period to have eliminated infant tion and/or flight acceptance tests before deliv* mortality failures, testing at systems level should ery for incorporation into the spacecraft, encounter pxoblems princIpa1ly attributable to undetected failure or out-of"specification con integrating subsysterns, dition which passed these co-contractors 1 screens theoretically would be picked up in the acceptanct Level 5 ;|l , Systems Testing, encompasses a vide tests for government-furnished equipment or in flit spectrum of testing activities on the Nimbus pro systems level tests * gram since it includes evaluation of * large number

11,14 of special-purpose spacecraft models. Final veri between the test personnel and the fication of the adequacy of design of the space C0*c0fityact0r0 f ''representative** Existence of craft (as a bus) and its complete integrated pay- formal procedural documents forced consideration load was demonstrated by the performance of the to be given to the exact environments to be prototype/qualification spacecraft under extreme simulated and to the specification of criteria environmental stresses. Exposing each flight for evaluating performance of the individual sub spacecraft to a flight acceptance test program systems total spacecraft as an integrated wherein it must perform while subjected to 0 y s tern * environmental stresses similar to those expected during and after launch is the ultimate means for c* instilling confidence it is ready for launch. The Test Program for tbe Mimflw* 11 satellite Level 6, Flight Readiness Testing, demon was based upon tbe wealth of experience strates no degradation to have occurred during •by its predecessor program, Ninbus I* A full flight transportation from the factory to the launch site, spacecraft acceptance test program at flight envi no short-term degradation to have developed during ronmental levels was conducted as were environmental the conduct of the launch-site activities, con qualification tests of new subsystem designs* A cluding with the final checking of the spacecraft f s full spacecraft qualification program was carried critical parameters as part of the launch count out for Nimbus 1, but not for Nimbus II, down prior to lift-off. The complete Nimbus II system test program In spite of differences which may in reality constittjted four stages; have existed between what each level of test should have accomplished and what it did in fact A prototype test program - intended pri accomplish, including both depth and time duration marily to complete the design development and of testing, the Nimbus 1 and II spacecraft were system integration, exposed to a more comprehensive and longer series of tests than any other spacecraft developed under A sensory ring test program - intended the aegis or at the Goddard Space Flight. Center. to integrate all flight quality subsystems within As noted earlier, although causal relations cannot the sensor ring and essentially subject this be established, the fact the Nimbus II spacecraft identifiable, major part of the spacecraft to has demonstrated exemplary in-orbit performance is flight acceptance testing, sufficient for other spacecraft programs to con sider Nimbus 1 test and evaluation program as a A spacecraft test program - intended to model for planning purposes. integrate all elements of the complete spacecraft and subject it to flight acceptance testing. B. Test Procedures and Plans A launch test program - intended to pre The complex nature of the"Nimbus spacecraft pare the flight spacecraft for launch, and the many alternative modes of operation pos sible with the housekeeping and sensor/payload 1. Prototype Test Program Plan - The proto subsystems preclude leaving to chance its operation type/qualification spacecraft from, the Nimbus I and control. This is especially so during the program was updated to the Nimbus II design con initial system level integration tests when test figuration. It served also as the initial conductors and supporting personnel are becoming .for integrating subsystems, the composite resulting acquainted with the expensive spacecraft and in the Electrical Systems Model for limb us II. supporting test facilities. Careful pre-planning This spacecraft was used for electrical., electro- of all factors of the test and evaluation program nagnetic compatibility control, mechanical and is essential. Simultaneous with intensive reviews Interference testing. Following its use within of the first drafts of documents within the the test - program, the prototype spacecraft was General Electric Company, representatives of exposed to the launch test program in -house and at NASA/GSFC, and to some extent co-contractors, the Western Test Range. reviewed plans. These were comprehensive and in step-by-step-detail to insure little likelihood 2. Sensor Ring Test. Program Flan - Follow of errors of omission and/or commission during ing physical Installation of flight quality hard the test program. The reviews were not infallible; ware into the sensor ring, the subsystems were corrections in test plans still had to be made energized to form an operating system. Electrical during the conduct of the tests. Following review te s t Ing e 1 Imlna ted technica 1 d 1 sc repanc ie s , int e r- and approval by NASA/GSFC representatives,,, the ferences and Incompatibilities not previously procedures and test plans were the guiding docu encountered in the prototype test program. Follow ments for the actual conduct of systems level ing the electrical testing, the sensor ring as a tests, major subasserafoly was subjected to an acceptance test program designed as a prelude to a similar Several auxiliary advantages result from the program for the entire spacecraft.., preparation of these plans* Their value as educa tional and training aids is unquestioned* Each 3» Flight., Spacecraft Test. Program Plan - participant having a defined responsibility for The flight spacecraft was subjected to a compre part of the test program is forced to pre-plan, his hensive acceptance test program, including vibra activities and to consider the impacts and impli tion testing; at levels for the Thrust Augmented cations of his activities on those of other sub Thor»Ageaa launch vehicle QfLmbus j wa§ placed in systems* Close working relations developed orbit by the, Thor-Agena vehicle) tbetaial

17.1-7 vacuum conditions. Briefly, the attitude control programs with the exception of a small number of subsystem (the upper section of the spacecraft) components. Therefore, the design being qualified was mated to the sensor ring, and the spacecraft "-s by the integrated systems test program was essen electrical and mechanical integrity was established. tially an evaluation of the spacecraft as a Significant performance characteristics are mea bus/carrier and of the mechanical and electrical sured before and after each major environmental interfaces among any subsystem and among the sub exposure. Tuneable antennas were adjusted prior systems and the spacecraft, created by the instal to shipment to the Western Test Range. lation of the subsystems into the spacecraft.

4. Launch Test Program Plan - The launch The purpose of the program specifically test program included a series of pre-shipment was not to evaluate the design of the subsystems, tests, tests after receipt at the Satellite although their performances in the system test pro Assembly Building (SAB/WTR), and tests on the gram did result in recommendations for the redesign launch pad after the spacecraft was mated to the of some of them. The test program neither ascer launch vehicle. Pre- and post-shipment tests tains the true levels of stresses which result from were preceded by a low- level (workmanship) vibra the exposure nor are the tests made to establish tion test to insure the absence of extraneous or stress levels to failure since only one spacecraft loose parts. is available and it cannot be tested to destruction.

D. Aspects of the Qualification and Flight Realistic simulation of performance of Acceptance Test Programs the spacecraft can only be achieved if the space craft is in fact energized and operating as it 1. Prototype/Qualification Test Program - would normally operate in the expected environment. The concepts developed for evaluating the perfor Thus, during the vibration test there were ener mance of the complete Nimbus spacecraft repre gized the command receiver, the power system on sented a substantial departure from those applied batteries, the telemetry/beacon transmitter, and to other spacecraft developed in the same time the drive motor on the High Resolution Infra-red period. Even at the start of the Nimbus program Radiometer. During the vacuum/thermal tests, all in 1960 there already had been outlined the basic subsystems which would normally operate in orbit elements of the qualification test program. These were exercised as part of simulated in-orbit oper were specifically defined in a set of test speci ations. Commands were received and acted upon, fications which described the five environments subsystems were turned on and off, measurements and the stress levels to which the qualification were made by sensors activated by appropriate spacecraft was to be exposed as part of the systems stimulators, television cameras recorded projected integration and test program. Four elements per test patterns, sensor and telemetry data were taining to the specification and the test program stored on tape recorders, and data were trans should be noted: mitted through radio in both the real-time and stored data modes of information transfer. a) The purpose of the qualification test program was to discover basic and/or inherent Exercising the spacecraft effectively weaknesses in the design of the entire spacecraft implies the existence of two additional necessary packaged in a form which represented the flight factors - instructions and personnel. The quali spacecraft. fication test program was conducted in response to sets of essentially step-by-step instructions b) The performance of the entire space embodied in written detailed test plans which were craft was to be evaluated while it was in the continually reviewed and revised. Personnel oper operating condition expected during that portion ating the spacecraft were being trained to be of its flight history being simulated by the called upon later to operate the flight spacecraft environmental exposure. in its acceptance test program and in orbit. Operational systems planning was completed and c) Stress levels for vibration tests written in detail. Computer programs were devised were to be set at approximately 150% of those for use with the telemetry data to allow its expected by the flight spacecraft (95% level) reformatting, special read-outs, energy balance, during launch and placement into orbit. The conversion from binary digital counting to practi extremes of temperature were set by specifying cal engineering units of measure, and other factors an increase of + 10°C of those expected during as desired. These were all essentially completed orbit. and ready at the start of the simulated flight operations tests in the vacuum/thermal test chamber. d) The order of exposure to the five Similar preparations, suitably up-dated,, were made environments were: as part of the flight acceptance test program..

Humidity The simulated flight tests started with Vibration the spacecraft on the launch pad, with appropriate Acceleration equipment in full operating condition as they would Pre-vacuum/Therma1 be in truth. At lift-off, the pressure in the Va cuum/ The rma 1 vacuum/thermal chamber was reduced,, simulating the expected change in pressure/altitude. Subsystems As noted earlier, the systems test pro were energized (solar paddles were in the unfolded grams were concerned with evaluating performance position)., monitored, and exercised on a real time of the integrated spacecraft. The subsystems 'basis as they would be in the early and later which were provided as government-furnished equip orbits after several days following launch., all ment 'had completed their own qualification test such activity in accordance with the guidelines 17.1-8 and instructions embodied in the operational of stress for these tests comfortable margins of systems documents. Commands to the spacecraft safety were chosen. While inside the "space were generated by the Test Director, written chamber" for the vacuum/thermal tests, the space and forwarded to the ground command equipment, craft was exposed to radiant heating to achieve and then transmitted via radio to the space the specified profile. All temperatures were craft's command receiver. Telemetry and sensor measured on the spacecraft ! s most massive non-heat derived data were transmitted from the space producing section, part of the sensor ring. The craft via radio to the ground receiving equip structure had mounted thereon a number of tempera ment. Antennas were mounted inside the vacuum/ ture detectors. The lowest temperatures read on thermal chamber to permit this transmission. the structure was taken to be the level specified Hardwire communications to and from the space as the high temperature, and the highest tempera craft was not permitted. There was no access ture on the structure was taken as the level to the normal test points used in the develop specified as the lowest temperature. Temperature ment or subsystems test programs. readings were made when the pressure was less than or equal to 10" mm Hg; the rate of pressure During the several days operation, change was to be no greater than the pressure/time thermal stress cycling was provided with real profile expected in flight. The spacecraft was time and/or near real-time evaluation of all interrogated bi-hourly, looking for any changes or data transmitted from the spacecraft. All anomalous performance which would be interpreted decisions regarding the performance of the as a forerunner of an equipment failure or harmful, spacecraft and its subsystems were made on the but slow degradation or out-of-specification con basis of data emitted by the spacecraft. It dition. was necessary to rapidly interpret data, and assess whether the test program should cease As data have been accumulated over the because of the exhibition of malfunctions, past several years, and the skills of spacecraft anomalous behavior, or failure. These require designers have improved, the needs for these ments provided on-the-job training for the test extremes have diminished and tighter tolerances conductor, the entire test crew, the personnel will no doubt be specified. operating the ground equipment, and all NASA personnel who were later to be faced with true While the qualification test program situations. pertained specifically to the Nimbus I program, a partial qualification test program was in effect When the environmental test speci for Nimbus II. The sensor ring with its subsystems/ fication was written, a comparatively limited experiments were exposed to a qualification test amount of data were available regarding the program, including exposures to humidity, gas leak, vibration environment during the launch phase. vibration, acceleration, and vacuum/ thermal. A central problem attendant to an environmental test program is the establishment of appropriate Two additional factors were necessary stress levels. The vibration specification was before initiating the qualification test program: based upon the knowledge available at the time for the Atlas/Agena launch vehicle, updated as 1) Criteria and/or guidelines to judge more data became available between the time of whether the spacecraft and its full complement of the initial specification and the time of the installed equipment "passed" the qualification test. tests. The stress level for the vibration test was set to be a 99% probability level, i.e., 2) An operating plan to define what during flight, there would be only one chance should be done in the event of failure of an element in a hundred the flight spacecraft would experi during the test program. ence a vibration environment more severe than that experienced by the prototype/qualification Formal and rigid acceptance criteria spacecraft in its test program. Assuming a were not established during the qualification test Gaussian distribution of the vibration stress program. Presumably, the absence of formal tech levels, the 99% level is set at the mean level nical criteria provided the NASA Project and Space (i.e., the mean vibration level expected based craft Managers with more freedom to exercise tech upon actual measurements) plus 1.5 times the nical judgments related to program flight schedule, difference between the mean level and the 95% and expenditure commitments. Collective technical level. In essence, the 150% vibration test judgment is a most important factor and quite cor level is based on the 95% level on a Gaussian rectly should take precedence over detailed formal distribution of expected stresses not on the planning, which should be subject to amendment as 150% of the mean value. Thus, all test levels the test program proceeds. In a sense, this was are far more severe than the average stress acceptable since in many cases quantitative descrip level which may be encountered in a flight. tions of performance would not have had firm bases The duration of the vibration test was selected for their specification. For example, although the to give assurance the flight units would survive resolution of the television camera system had been the environments imposed both in the acceptance quoted at 800 lines (in one diameter inch tube), tests and in the actual launch activity. this actually pertained » only to the central section of the tube. If 700 lines resolution were found in The temperatures to which the space a tube, this would hardly be cause for rejecting craft would be exposed in orbit, the quality the spacecraft. As the test program developed, of the thermal design and the temperature regu specific criteria were formulated for some sub lating system were not known with high confidence. systems. The NASA Project Manager relied heavily Consequently, in establishing the extreme levels on the judgment of the members of the Nimbus

17.1-9 Environmental Test Committee, whose membership The stress levels used in the flight included NASA/GSFC and GE representatives. This acceptance tests were the best estimates of these group monitored, reviewed, and recommended changes anticipated in the launch/flight environment. The in the rationale for testing and evaluating test test levels were to be no greater than those data, and it reviewed all test data and made expected in these environments. In the case of recommendations to the NASA Project Manager as to the vibration tests, the stress level was set by whether the spacecraft and its installed equipment first assuming a Gaussian distribution of vibration met design objectives and was qualified. levels (based on accumulated experience). The test level was the 957o probability level, i.e., one The rejection retest plan was formu chance in twenty this level would be exceeded dur lated before the start, of the program. Its essen- ing the actual launch. The 95% level is the 1.650* t ia 1 features were: point.

a) If the spacecraft is rejected before, An innovation introduced into the vibra during, or after an exposure, discontinue the tion test program for Nimbus II was the use of exposure . notching techniques. The combination of the space craft and the drive power/frequency characteristics b) Determine the cause of failure, and of the vibration table were such that if the orig co rrec t i t a , inc lud ing any d e s i gn defects. inal amplitude/frequency specification were adhered to, the spacecraft structure could be overs tressed. c) Starting at the event, timing, tem Analysis and vibration survey data indicated fre perature where the failure occurred, repeat the quency bands wherein the drive power of the vibra complete test program until successful. tion table should be reduced with no resulting deleterious effect on the purposiveness of the 2. Flight Acceptance Test Program - Little vibration test program. These specific notch differentiated the flight acceptance program for points were then incorporated as part of the final the Nimbus I and II spacecraft. Basically, the specification for the vibration tests. same set of environmental specifications pertained. Whereas the qualification test program searched The same rejection and test plan was for inherent design weaknesses, the flight accep- stipulated for the flight acceptance test program program could be characterized by the following: as for the qualification test program, and just as easily interpreted on the basis of program a) 'The purpose of this test program was needs as in the earlier tests. to determine the flight, worthiness of the flight spacecraft and to discover workmanship defects In addition to the systems test program and/or errors in use of materials. for the entire Nimbus II spacecraft, the sensor ring with all equipment installed was exposed to b) The performance of the entire space a flight acceptance test program in all environ craft, which, was to be an exact replica of the ments, and with its own vacuum/thermal test cycle. p ro t o^ t yp e / qua 1 i, f i ca t ion. spa cec r a f t., wa s to b e evaluated, in the' operating, condition expected The rejection and retest plan was not during, that portion of its flight history being followed, in spite of its good intentions. Pro simulated by the environmental exposure. gram management considerations, trading off per formance, cost, and schedule factors resulted in c) Stress levels were to be set at continued testing in accordance with the planned approximately 100% of those expected by the flight profile, following repairs or replacement of failed spacecraft during launch and orbit... equipments, rather than recycling the tests.

d) 'The order of exposure to the three 3. Launch Pad Test Program - The test pro environments were: gram at the Western Test Range was conducted to ascertain that the flight spacecraft had not Vibration suffered any detrimental damage in being shipped Pre - va cuum/Therma 1 from GE/VFSTC to WTR. Va cuum/Therma 1 Since flight acceptance testing; follows VI. Analys is of Failure Records the qualification acceptance test of a single - System Tests and Flight spacecraft, no prior knowledge has been developed regarding, the expected variations among supposedly A. The Raw Data identical spacecraft.. The assumption is made no variation exists, which, of course,, is statisti Quantitative analyses were made to determine cally invalid, but economically wholesome* On. causes of anomalous behavior, out of specification the presumption that all design, defects have been performance, and/or failures of components and eliminated as a, result of the qualification test subsystems which form the total spacecraft, encoun program, the flight acceptance test program is tered during the flight acceptance tests for Nimbus presumed to screen only for workmanship and /or I and II and their performances in orbit. All material errors and defects* That this is not so available reports within the General Electric Is ably demons tra t ed by analysis of the failures Company pertaining to these system tests were ing the flight acceptance test pro- individually examined, including: System Failure I, Nimbus II, and many other Reports; Malfunction Reports; Final Test Reports; ______for which comparable data, and Flight Reports* These data were supplemented hav by records of in orbit malfunctions as reported in GE-originated reports. Failure data pertaining 17.1-10 to the Prototype/Qualification Test Program were Design a thermally benign environment for all not reviewed at this time, equipment, to the greatest extent possible,

B. Method of Data Analysis - Provide a structure and heat control ling system having high thermal inertia,. The principle method of analysis of the raw data is categorization into factors of interest, - Provide a thermally conservative design showing for each classification the absolute whereby the heat dissipating capability is much numbers of occurrences. The general classifica greater than the heat producing sources, tions are by: - Provide for thermal load sharing in 1. Importance of the failure to the sue- . the event of failure or removal of planned equip cessful operation of the mission or of the sub ment. system. Long life in orbit may be greatly enhanced by: 2. Attributed cause of failure. - Condition/power-age piece parts before 3. Environment and test activity existing installation, into components. at the time of failure. - Specify conservative stress conditions 4. Level of testing, other than system and environments for systems level tests. level, at which the failure could have been detected. - Apply experienced engineers to design, test and evaluate subsystems and the spacecraft. 5. Subsystem. - Incorporate functional redundancy 6. Type and complexity of subsystem. an appropriate command system to allow ground control. 7. Source of hardware. - Provide a benign thermal environment C. Significant Failures Encountered to minimiz e in -orbi t the rma1 st re s se s,

The most significant failures encountered B - Test"Operations during the Nimbus I Prototype/Qualification, Nimbus I Flight and Nimbus II Flight Systems Acceptance Qualification tests were not made to destruc Test Programs are shown in Tables 1, 2, and 3. tion levels y hence mo real measure exists of the degree of conservative design in the spacecraft. VII. Lessons Learned and Applied There is no way to evaluate the applicability of all tests to which the spacecraft and/or subsystems The experiences encountered during the develop were exposed. ment of the Nimbus I and II spacecraft may have application to other space programs. Operate each subsystem under normal conditions (life-test) after passing its flight acceptance test A. Design Practices to discover latent failures. This would reduce failures and anomalies encountered during the more Provide extensive alternative modes of func expensive integrated systems test program. tional redundancy as a means of failure protection, thus enhancing long life potential. Expose the spacecraft to an orderly planned series of tests, each successively encompassing a Provide an extensive command system with larger number of subsystems which require integrated ground command override for programmed functions performance.» to effectively use alternative modes of operation. Flan the test and evaluation activities early Fabricate all cables and wire groupings on a in the program, preferably concurrently with engi full-scale, exact replica three dimensional mock-up neering design* Potential problems in the conduct model spacecraft. This minimizes stresses and of system tests and simulated orbital operations - strains in wiring and connectors, will be exposed early,

Give most attention in design and test to non The 'best efforts expended in: conservative electronic and complex equipments, since most fail design, engineering, management and test practices.;,, ures and anomalies occur here* selection, and training of personnel are still not sufficient to absolutely screen, and eliminate all Be wary of all specifications for components latent failures from, individual subsystems or and subsystems which contain idealized, "text book" the complete spacecraft* A strong test and evalu conditions for design* test and performance; e,g«, ation program helps to isolate these types o£ a transmitter designed for a 50 ohm resistive load failures prior to launch; e.g»> which actually is to be connected to a transmission line, impedance matcher and antenna, - The control system on I, after more than 750 hours of operation, including - Evaluate all equipment under conditions weeks of vacuum/thermal testing. of the total expected environment.

VIVM1 - The (2N768) transistor in the command Provide long term personnel assignments con clock failed after more than 1000 hours of oper sistent with separable phases/elements of the ation, in spite of conservative circuit design and engineering, test and evaluation program and the power conditioning of piece parts. professional interests of the personnel.

C. Failure Records and Analyses - Better use is made of accumulated experience than if frequent training and retrain Report and record all failures and anomalous, ing is required. out-of-specification performance of components 9 subsystems and the spacecraft system as soon as - Methods are developed for the expe observed. ditious conduct of routine tasks.

- Provide expeditious diagnosis and - Develops healthy inter-personal determination of the specific reason for the fail relations among members of the project team. ure. Rehearse all new test activities and phases - Provide closed-loop record of the of the test program using the previously published action taken. detailed test plan containing step-by-step instruc tions . - Analyze and record failures and final actions taken on a near real time basis. Failure F. Technical Reviews records are intended to be used for management control, not for purposes of historical record. Prepare information flow schematics early in the program. D. Documentation - Show individual schematics for: signal, Write all plans, specifications, requirements grounding, power, command, telemetry. and procedures necessary for a test and evaluation program in detail, step by step, before the start - Very useful in system trouble shooting of the functional activity, and planning operational activities.

Focus attention on specific problems as they Require design reviews for all hardware developed. arise through rigorous use of the detailed test requirements and plans documents. - The periodic reviews cover activities from the initial concept through qualification test - Confidence is provided to the Test ing. Conductor and to the Program Manager by isolating activities and problems. - Personnel responsible for hardware must be made to understand the review is an essen E. Personnel tial part of the design cycle, intended to help them strengthen the design, and is not a "police" Define fully all responsibilities and author action. ities vested in personnel directly handling, work ing on or operating the spacecraft. - Technically strong, competent, per ceptive personnel should lead design review activ - Do not permit random assumptions to ities and recommend action. replace such assignment, - All action items resulting from a - Avoid misunderstandings which can review are accountable to be accomplished as part lead to strained inter-personal relations and con of the design effort. sequent loss of operating efficiency. Permanently staff the program with Senior Certify spacecraft Test Conductors as to their Consulting Engineers. Their need and contributing ability to operate the spacecraft, to supervise value were well recognized by the Program Manage the test program, and to take effective measures ment . should an emergency arise. - The Nimbus program had two,, one in - Provide a unique certificate. the area of mechanical engineering, another in the area of electrical/electronic engineering and Encourage and train all personnel, through sys terns, direct, conscious management effort, to be per formance oriented, technically honest, and to - Each had full responsibility and author demonstrate integrity, ity for overview of the overall technical perfor mance of the electrical and mechanical systems and - Personnel should feel free to expose to enter into any technical area for analysis/review errors and faux pas without fear of management and to recommend changes to all levels of partici discipline either to themselves or to others. pating personnel and management,

Establish a highly disciplined environment The GSFC 'Nimbus Environmental Test Committee for the test and evaluation program wherein each provided independent guidance directly to the participant knows his job and the personal impact NASA/GSFC Project Manager regarding: Test Program; he has on the total task. Test Levels; Acceptance of Test Results; le-fest Requirements, 17J-I2 - GE had two members on this committee, the two Consulting Engineers, responsible for the areas of "Electronic Systems" and "Design and Reliability".

A Launch Readiness Review preceded certifica tion of flight readiness of the spacecraft.

TABLE I MAJOR FAILURES DISCOVERED DURING INTEGRATED SYSTEMS TEST PROGRAM - NIMBUS I PROTOTYPE SPACECRAFT

Potential Effect Attributed Cau.se Subject Action Taken System Test on of Failure Related? Remarks Mission

HRIR Catastrophic to Mechanical redes igp of Functional design No Loss of synchronizing the HRIR subsystem the mount for the mag signal netic pulse pickup which moved when vibrated

AVCS TAPE RECORDER Catastrophic to stored Component redesigned Functional design No Failed as a result Shaft of recorder broke AVCS data of extreme tests; in vibration test failure occurred during a second vibration test at the prototype/qualifica tion test levels.

COMMAND CLOCK Catastrophic to total Bo'Ught-off, Parts/material No Circuit was very One transistor (2N768) mission Replaced the piece conservatively de exhibited excessive part with another of signed with respect leakage current after the same type. to expected variations more than 1000 hours in the transistor's operation parameters. This explains the long time period of operating without failure.

ATTITU DE C ON TRO L Po t en t i a 1 ly ca ta s t rop h i c Replaced overheated Test procedure No j Solenoids had not Solenoid valves overheated to total mission solenoids with others failed in performance; of the same type • insulation appeared damaged; replaced to enhance confidence.

17.1-13 CORONA

ATTITUDE

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ELECTROMAGNETIC

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Other

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Cause

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Thermal

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MOMENTUM MOMENTUM

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LEVEL LEVEL

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ACCEPTANCE PRE-LAUNCH