Systems Engineering - Manned Space Flight

8/7/2019 Systems Engineering - Manned Space Flight

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SYSTEMS ENGINEERING, MANNED SPACE FLIGHTCharles W. MathewsNational Aeronautics and Space Administration

IntroductionThe systems engineering activities of the United States

in manned space flight have involved the design, development,manufacture, test, and operation of three flight systems--Mercury, Gemini, and Apollo. The overall arrangement ofthese configurations is illustrated in Figure 1. The systemsengineering effort has not only encompassed these spacevehicles, but also encompasses many other aspects of theprograms such as worldwide instrumentation and communicationsnetworks, control centers, recovery support and equipment forthe conduct of various scientific and technical experiments,As experience has evolved from development efforts and flightoperations, certain factors have become apparent related tothe achievement of well balanced and functional systems.Some of these factors will be discussed herein.

Program ObjectivesThe point of departure for systems engineering work

involves the definition of objectives for any given spaceflight program. In the programs just mentioned, primaryobjectives have been stated in fairly simple and direct terms,


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as indicated in Figure 2. The Mercury Program was aimed atthe demonstration of manned orbital flight and safe return.The Apollo Program is aimed at manned lunar landing and safereturn. The Gemini Program involved a somewhat broader rangeof objectives, but in the conceptual design, long durationflight and rendezvous received the main emphasis. A l l ofthese programs carry with them a substantial number ofsecondary objectives; however, the primary objectives tendto dominate the design and it is quite important not toconfuse early efforts with a large multiplicity of objectives.At the same time, sufficient consideration of secondaryobjectives must be applied in order to avoid basic constraintsthat might make their achievement impossible. In Gemini, forexample, such considerations involved the hatch design toafford extravehicular possibilities, the docking adapterdesign to afford maneuvering in orbit with a large propulsivestage, and design control of center of gravity offset toafford the possibilities of maneuvering reentry.

Mission ModesOnce the objectives of the flight program have been

clearly established, it is necessary to analyze variousmission modes by which these objectives can be achieved.This mission mode analysis ultimately establishes the design

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missions and environmental conditions upon which th econf igura t ion of th e f l i g h t hardware i s based. The broadestand most basic example of t h i s systems engineering task i sa s s o c i a t e d w i t h th e Apol lo objec t ive and i s t h e work whichl e d t o t h e d e c i s i o n t o u s e th e lunar orbit rendezvous mode.

T h i s mode se l ec t io n qu es tio n involv ed many of th ec l a s s i c a l e le me nt s of a complicated systems engineeringproblem. I t involved money, t i m e , th e s t a t e of c u r r e n ttechnology, t h e s t a t u s of space hardware development i n t h eUnited Sta te s , the f a l l b a c k p o s i t i o n s a v a i l a b l e i n c a s e oftechnical problems, t h e probabi l i ty of miss ion success and,l a s t but not l eas t , t h e ques t ion of a s t r o n a u t safe ty . Eachof th ree mission modes were eva lua ted with respect t o thesekinds of f a c t o r s and t h e f i n a l s e l e c t i o n made wi t h t h e b es tp o s s i b l e intercomparison among a l l of them.

The three modes which were se r i ous ly cons ide red( t o g e t h e r wi t h v a r i a t i o n s ) are i l l u s t r a t e d i n Fig ure 3 andb r i e f l y de sc ri be d i n th e fol lowing paragraphs.

Direct Launch Mode - The b a si c p r o f i l e of t h i s modes t a r t s w i t h a simple launch from t h e ear th d i r e c t l y i n t oearth-moon t r a n s f e r much l i k e s e v e r a l o f t h e Surveyormissions. No ear th o r b i t p hase i s included. A t t h e moon,t h e descent i s d i r e c t l y t o t h e lunar sur face wi thout f i r s te s t a b l i s h i n g a c i r c u l a r l u n a r o r b i t , m u c h l i k e Surveyor.


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Two hardware c o n f i g u r a t i o n s were considered for t h ed i r e c t mode. One co ns is te d of a l aunch vehic le of th eSa turn V class w i t h a somewhat cramped two-man s p ac e c ra f t .The o t h e r p l an u t i l i z e d a l a rge r (NOVA) des ign toge therwith a three-man spacecraft of more comfortable design.

E a r t h O r b i t Rendezvous - The ear th o rb i t r endezvousmode involves two or more ear th launches ( two was consideredi n th e Apollo s tudy ) and assembly of a s p a c e c r a f t / i n j e c t i o nstage i n o rb i t . This mode provided th e p o t e n t i a l f o r u s i n ga three-man sp ac ecr af t without going t o a b o o s t e r of theNOVA c la ss ( abou t 75 t ons t o escape ve lo c i ty ) . The lu na rphase o f t h e miss ion fo r the case cons idered was the same asf o r th e d i r e c t a s ce n t mode with a three-man spacecraft .Obviously, there are p o s s i b i l i t i e s f o r combining ear threndezvous and lunar orbit rendezvous modes. These werenot cons idered. ' s ince t he y of fe r ed no advantage i n t h e Apollomode trade-off study.

Lunar O r b i t Rendezvous Mode - For t h e d i r e c t a s c e n t a n dear th orbit rendezvous modes already discussed, a r e l a t i v e l yheavy s p a c e c r a f t had t o be landed on the moon and subsequentlylaunched back toward earth. All o f t h e f u e l for t h e t r a n s -ear th journey and the heat s h i e l d for r e e n t ry i n t o the ea r th ' satmosphere are par t of t h e payload landed on the moon. With

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the lu na r o r b i t rendezvous mode, both of these items, i na d d i t i o n t o s t r u c t u r e an d e ng in es , remain i n l u n a r parkingo r b i t . T h i s avoids both t h e need t o slow down a l l of th eweight f o r l a nd i ng a nd t h en a c c e l e r a t e i t during launch andt r a ns ea r th in j ec t i on maneuvers. The spacec ra f t f o r the luna rorbit rendezvous mode i s thus composed of a d i f f e r e n tcombination of stages than would be a p pr o p ri a te f o r d i r e c ta s c e n t t o the sur f ace . S ince i t i s l i g h t e r f o r th e samepayload t o t h e moon, th e ear th launch payload requirementsare lower.

The var ious modes s tudied requi red di f fe rent hardwaredevelopments , and design studies were accompl i shed to c l a r i fyboth hardware and mission con figu rati ons . From these s t u d i e s ,estimates were developed for cos ts , t i m e , s u c c e s s p r o b a b i l i t i e sand safe ty . Some of these r a t h e r e l u s i v e f a c t o r s were eas ie rt o compare th an t o estimate on an absolute basis . The r e s u l t sof some of t h e s t u d i e s are shown i n Fig ure 4 . I t can b e seent ha t t h e lunar orbi t rendezvous mode was e qua l o r bes t i n a l lc h a r a c t e r i s t i c s l i s t e d . T h i s i s c o n s i s t e n t with t h e choicebeing unanimous among NASA management a t th e t i m e i t was made.

Operat ional Features of Lunar O r b i t Rendezvous - Anotherimpor tant fe a t ur e of a mission i s t h e crew a c t i v i t i e s t i m e l i n e .I n t h i s r e spec t , t oo , the lunar orbi t rendezvous mode has somea t t r a c t i v e f e a t ur e s. I t i s g e n e r a l l y desirable t o h av e t h e

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mission arranged so t h a t t r a n s i t i o n p e ri od s of r e l a t i v e l yhigh a c t i v i t y and s t ress are r e l i e v e d by p e r i od s o f r e l a t i v es t a b i l i t y and freedom from imminent pressures, In missionplanning language, these pe r iods a re sometimes called s t epsand ' ' p l at eaus , r e spec t ive ly . The lunar orbi t r endezvous11mode provides t h e maximum number of plateaus among t h evarious modes considered,

To lea ve one pla teau and t ra n sf e r t o another normallyinvolves a powered f l i g h t maneuver. After t h i s event , therei s an oppor tuni ty (on t h e p l a t e a u ) t o assess t h e s t a t u s o fth e s p a c e c r a f t i t s e l f , assess t h e maneuver just completed,and t o review t h e a v a i l a b l e o p ti o n s f o r t h e nex t even t o rs t e p . The opt io ns usua l ly inc lude th e nominal nex t s tep ,pe rhaps cor r ec ted fo r t h e dispers ions encountered previous ly ,p lus one o r more a l t er n a t e s . One of t h e v ar io u s a l t e r n a t e si s chosen only i f the re i s a reason t o di scon t inue thenominal mission, The di f f e r ences be tween a l t e rna te s can beca tegor ized by t h e r e l a t i v e accomplishment i n terms ofcomple ting miss ion ob j ec t i ves o r ga in ing f l i g h t exper ience,i n terms of t h e t i m e o r p ro p ul si on r e qu i re d , o r i n terms oft h e le ss en in g of burdens on var ious subsystems. An a l t e r n a t emission of unusual urgency i s commonly called an abort.

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The existence of convenient plateaus, coupled with theplanning of mission alternatives which can be selected inreal time, provides considerable flexibility in conduct ofthe mission. Before launch the mission plan is the nominalone including accomplishment of all objectives. As time goeson, modifications can be made as necessary to accommodatereal time events while achieving as much applicable flightexperience as possible. For a mission as long as the lunarmission, this approach is expected to provide real advantages.It allows making as much progress as conditions will permit,i.e., capitalizing on success while being able to respond toadversity.which may be contrasted t o the more common approach ofarranging a long series of successive missions, each slightlymore complicated than the previous one.

This approach is part of an "all-up" concept

The plateaus which naturally occur in the lunar orbitrendezvous mission are listed in Figure 5. The end pointsof these plateaus representing major 11commit" points in thelunar landing mission are characterized by propulsivemaneuvers resulting in major changes in the spacecraftenergy. These commit points and mission plateaus can bothbe represented schematically on a single chart as shown inFigure 6. This figure illustrates the major maneuvers during

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the lunar landing mission in terms of both delta V (on theleft) and pounds of propellant (on the right). These maneuversrepresent therepresents the plateaus,

I?commit" points, and the space in9between

A similar approach was taken in establishing the missionmode for the initial Gemini rendezvous, although later in theflight program many different techniques were explored.Three different r.endezvous techniques were analyzed. Thesetechniques are illustrated in Figure 7. Rendezvous planningbegan over three years before the first rendezvous missionwas performed. *This planning began with the selection ofthe basic mission design criteria which would be used forall of the eventual studies that were performed and conducted.

The basic criteria were really very simple. Considerationwas given almost.exclusively to providing the highest proba-bility of mission success. That is, the intention was todesign a mission which could routinely depart from thenominal in response to trajectory dispersions and/or space-craft systems degradation, while s t i l l providing minimumdispersion of the conditions going into the terminal phase.More specifically, it was to provide flexibility withoutintroducing undue complexity, thus giving the astronauts anaflight controllers a capability of choosing alternate

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maneuver sequences not dissimilar to the basic maneuversequence, but which would provide a capability of reactingto the aforementioned dispersions and problems, (I

Returning to Figure 7, the three plans that wereprepared and documented were:

Plan 1 - The so-called tangential mission plan, whichprovided rendezvous after approximately three and a halforbits;

Plan 2 - The coelliptic, which utilized the same basicmidcourse phase maneuver sequence but which had a signifi-cantly different terminal phase; and

Plan 3 - A rendezvous at first apogee.Based on analysis of these plans, a recommendation was

made to adopt the second. The basic desired feature of thecoelliptic plan was that the relative terminal phasetrajectory of the spacecraft with respect to the target wasnot particularly affected by reasonable dispersions in themidcourse phase maneuvers. More simply stated, the coellipticapproach afforded a "standardized" terminal phase trajectorywhich yielded obvious benefits with regard to establishmentof flight crew procedures and training.

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Systems Selection and ConfigurationOnce the selections of mission modes and design missions

have been made, a further clarification of the approach tothe hardware design is possible. This approach involvesconsideration of the state-of-the-art of potential systems,the developmental status of systems that could be applicable,and the requirements for new systems development. In theselection of the systems and types of operations to bedemonstrated, a strong effort is made to consider therequirements of future programs. It is not anticipated thatsuch systems necessarily will be directly used in otherprograms; however, their operating principles should besufficiently close that the concepts f o r their use can bevalidated. Where possible and to minimize development time,.systems that already have some development status areselected. The Gemini spacecraft guidance system typicallyrepresents this approach. A simplified block diagram ofthis system is shown in Figure 8.carrying out navigation, guidance, and the precise spacemaneuvers needed for such activities as rendezvous, maneuvering,reentry, and launch guidance, At the same time, such majorelements of the system as the inertial platform, the digitalcomputer, the radar, and the flight-director display drewheavily on previous developments.

The system is capable of

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There are cases, however, where the state-of-the-artmust be extended. A fact of extreme importance is thenecessity to recognize the need f o r extending capabilitieswhere the requirements of present or future programs lead tothat conclusion. The Gemini fuel cell development isbelieved to be a good example of the exercise and recognitionof such a requirement. These electric power generatingdevices, along with the long term cryogenic storage ofhydrogen and oxygen, were an entirely new development notpreviously utilized even in ground applications. The Apollosystem also uses fuel cells and can now draw on the Geminiexperience. This development has also stimulated groundapplications of similar devices.

Another factor requiring due consideration involves thechoice of systems configurations achieving desirable opera-tional characteristics. Some of the factors involved in thisconsideration are listed in Figure 9. Considerable operationaldifficulty was experienced with the Mercury spacecraft confi-guration which was very weight critical and could not affordmany desirable operational characteristics. These difficultiesresulted in launch delays, lengthy retest periods aftermodifications or component replacements, and a lower level ofreliability than desired. As a result, heavy emphasis wasplaced upon achieving truly operational configurations forGemini and Apollo. The Gemini vehicle, for example, was

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assembled from a number of discrete modules which generallyaccomplished a s i n g l e p a r t i c u l a r f un ct io n . Interfacesbetween th e modules were as simple as po ss ib le . Each modulewas capab le o f be ing bu i l t and t e s t ed independently of t h eo th er modules. The p a r a l l e l approach that t h i s makesp o s s i b l e has gr ea t ly f a c i l i t a t e d Gemini a ssembly, t e s t i ng ,and l a t e replacement of malfunctioning modules.

I n a d d it i o n t o t h e ph ys ica l modul arization , Geminia t tempted t o make each subsystem as independent as p o s s i b l eof o th er subsystems. Th is permits more thorough testing a ta lower assembly level and avoids lengthy re te s t s of manysubsystems when a single component f a i l s . I n most cases,i t was necessary t o b r e a k e l e c t r i c a l o r plumbing connectionst o Mercury equipments for pre - f l igh t t e s t in g and checkout.This r e q u i re d v a l i d a t i o n of the reconnec t ion a f t e r t e s tcompletion. During system des ign of t h e Gemini and Apollo,a dete rmined e f for t was made t o make a l l necessary t e s tp o i n t s , r o u t i n e as w e l l as d i a g n o s t i c , a v a i l a b l e i n th e formo f s p e c i a l t e s t c o n n e c t o r s . D i f f i c u l t i e s w i t h the " layered"Mercury equipment insta l la t ion l e d d e s i g n er s t o adopt theground rule t h a t each uni t must b e i n d i v i d u a l l y a c c e s s i b l eand in d iv idu a l ly removable w i thou t d i s tu r b in g o t he r un i t s .No u n i t c o u l d be i n s t a l l e d i n t h e pressure cab in un le ss i twas s p e c i f i c a l l y r e q u i r e d by f u n c t i o n t o b e there . The

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philosophy o f i n d i v i d ua l a c c e s s i b i l i t y was a p p li e d a l s o t owire bundles and plumbing t o avoid any need for threadingthese through s t r u c t u r e o r d i s t o r t i n g them duririg i n s t a l l a t i o n .Mercury experience taught us that a i r c r a f t processes were n o tne ce ss ar i ly adequate f o r manned sp acec raf t . The manned spaceve hi cl e programs pioneered i n a number of areas such asa l l brazed propulsio n system plumbing, crimped el e c t r i c a lconnect ions, s a l t f ree c o l d p l a t e b r az i ng , e t c .

The modular concept was a n o t h e r f a c t o r i n v o l v e d i n these le c t io n of . t he lu na r o rb i t r endezvous mode f o r the Apol losystem. T h i s mode afforas the use o f a sp ec ia l purposev e h i c le f o r c a r r y i n g o u t t h e descent t o and landing on t h emoon and t h e subsequent launch and re-rendezvous wi t h th emother spacecraf t . T h i s s p e c i a l v e h i c le i s c a l l e d t h e l u n a rmodule and i s shown In Figure 10, The e f f i c a c y of t h eapproach taken can b e en vi si on ed when one co ns id er s t h e manyope ra t ing regimes t h a t must b e t r ave r sed dur ing th e l u n a rmiss ion. Provis ions m u s t be made i n th e sys tem for t r ave r s ingt h e launch environment, f o r a b o r t i n g t h e launch, f o r longdura t ion ope ra t ions $0 and from t h e moon, f o r lunar landing,and f o r s u p e r - c r i t i c a l r e e n t r y on the r e t u rn t o ear th ,


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A single vehicle or module encompassing all thesecapabilities would be severely compromised with respect tothe lunar landing operation. For example, a very high landedweight would be required, flexibility of the system layoutand visibility would be compromised because of reentry heatprotection considerations, and the design of the propulsion,guidance and control systems for ascent and descent would becompromised by the maneuvering requirements associated withother parts of the mission. In the design of the lunarmodule, many of these compromises can be avoided and,therefore, this module becomes a truly special purpose space-craft whose structure, external and internal configuration,and system operating characteristics reflect safely itsunique function in the mission.

Redundancy and Crew IntegrationAnother important facet of systems engineering for

manned space flight involves the use of redundant or backupsystems. This technique must be applied judiciously and, asa result, in the spacecraft systems utilized to date a completerange of combinations exists. For systems directly affectingcrew safety where failures are of a time-critical nature, on-line parallel redundancy is often employed. In the spacecraftpyrotechnics system, a complete parallel redundancy is oftencarried to the extent of running separate wire bundles on

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opposite sides of the spacecraft. In a few time-criticalcases, off-line redundancy with automatic failure sensing isrequired. The flight control system of a launch vehicle isan example where this technique is frequently employed, Inmost crew-safety cases which are not time critical, crew-controlled off-line redundancy o r backup is utilized with thecrew exercising management of the systems configuration. Inthe spacecraft propulsion system, the reentry attitudecontrol system is utilized solely for that phase of themission. This reentry propulsion in turn involves parallelredundancy because of the critical nature of this missionphase. Many systems not required f o r essential missionphases are basically single systems with internal redundancyfeatures commensurate with the requirements f o r overallmission success. Certain systems have sufficient inherentreliability once their operation has been demonstrated andno special redundant features are required. The spacecraftheat protection system is one of this type.

For systems which are very critical from the flight-safetystandpoint, much attention is paid t o obtaining flexibility insystem operation and safety against successive failures ofindividual components. The Gemini electrical power anddistribution system will be used here to illustrate atypical approach. A simplified schematic of this system is


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p r e s en t e d i n F i g u r e 11. The p ri ma ry s o u r ce o f e l e c t r i c a lpower i n o r b i t dur ing long dura t ion ope ra t ions i s the f u e lc e l l . Two f u e l c e l l s e ct i on s are c a r r i e d i n t h e equipmentadapter module t o provide t h e b a s i c system redundancy. Eachs e c t i o n c o n t a i n s three s t a c k s wired i n pa ra l l e l and eachstack can be opera ted independently a t the d i s c r e t i o n of t h ec r e w . N o ~ ~ ~ ~ a l l y ja l l s t a c k s are placed on t h e l i n e becauselow l o a d o p e r a t i o n has been shown by t e s t t o be extremelyb e ne f ic i a l t o f u e l c e l l l i f e . The crew, by monitoring t h ev o l t ag e ou t pu t o f e ach s t a c k i n r e l a t i o n t o t h e load, cane v a l u a t e t h e load shar ing of th e s t a c k s a nd determinewhether a l l are ope ra t in g wi th in to l e r anc e . I n t h e eventof a s t a c k f a i l u r e , and t h i s d i d occur on a number ofoccasions, th e crew may s h u t i t down and, i f necessary,e x e r c i s e management ov er power consuming sy stem s t o reducet h e o v e r a l l l o a d ,

The f u e l c e l l system i s fur ther backed up by a b a t t e r ysystem contained within t h e re en tr y module. T h i s b a t t e r ysystem has t h e capac i ty to hand le t h e power drains r e q u i r e dfor one and one-half o r b i t s a t f u l l load, o r much longer i ft h e s p a c e c r a f t i s powered down. I n ad d i ti o n t o t h i s c a p a b i l i t y ,i t prov ides r equ i r ed power f o r r e t ro - f i r e and r een t ry , andt h a t r e qu i re d f o r t h i r t y - s i x hours of ope ra t io n o f necessa ryequipment a f t e r landing. The ba t t e r y system i t s e l f has two

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sections--a main bat tery sect ion which provides f o r g e n er a lsupp ly of power and a squib bat tery sect ion which providespower f o r c r i t i c a l o p e r a t i o n s such as t h e f i r i n g o f r e t r o-rocke ts and pyrotechnics and the opera t ion of f l i g h t c o n t r o l s ,Again, i t i s p o s s i b l e f o r t h e crew, through manual switching,t o provid e main bus power t o t h e c r i t i c a l s qu ib c i r c u i t s .P r o p e r i s o l a t i o n i s provided so that f a i l u r e o f o ne systemdoes not jeopardize t h e o t h e r .

A s shown by t h i s example, manned spacecraft systemsdesigns employ manual sequencing and systems management t o alarge e x t e n t . T h i s f e a t u r e a f fo r d s s i m p l i c i t y by u t i l i z i n gman's c a p a b i l i t y t o d ia gn os e f a i l u r e s an d t o t ake c o r r e c t i v ea c t i o n . I t f a c i l i t a t e s f l e x i b i l i t y i n t h e i n co r po r at i on o fnecessary redundancy or backup conf igurat ions of t h e systems.For example, i n t h e spacec ra f t e l ec t r i ca l power sys t em jus ti l l u s t r a t e d , the redundancy involved would make automaticf a i l u r e sensing, in ter lo ck in g, and switchi ng both complexa n d d i f f i c u l t , i f not imposs ible .

Related Systems Englneering FactorsI n s p i t e of t h e u t i l i z a t i o n of t h e t e chn iques jus t

o u t l i n e d , we have ge ne ra l ly found t h a t t h e y i n themselves don o t e l i m i n a t e a l l c a t a s t r o p h i c s i n g l e p o in t f a i l u r e s . It i snecessa ry t o have separate design reviews and ana ly ses o f

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failure modes and their effects to assure the identificationof all single point failures. These activities are under-taken by special reliability engineers who do not have theinvented-here" feeling of the design engineer.11

Another factor related t o systems engineering is therecognition that even the best designs may fall prey t odiscrepancies which occur in the manufacturing process,Routine aircraft quality control methods are inadequate toachieve manned space flight goals. For this reason,inspection efforts are materially increased both in-plantand at suppliers' facilities. Sampling techniques arecompletely eliminated except for items such as nuts andbolts and it is necessary t o carefully record each step ina component's history and t o establish a central recordsystem to assure positive retention and rapid access to theserecords.

The practice of writing off failures as isolated casescannot be tolerated and, by careful analysis and testing,the causes of most malfunctions which occur during ground orflight testing can be identified. Meticulous attention tothis failure analysis and follow-up with corrective actionis an essential feature of any safety program, It is highlyundesirable to fly equipment with a questionable history,

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Still another factor strongly related to systemsengineering is the approach taken to subsystem and systemtesting. The Gemini development philosophy was based on thepremise that confidence could be achieved through heavyemphasis on ground testing and that a very limited number ofunmanned flights could serve to validate the approach. Inthe Apollo Program, the Gemini philosophy of limited unmannedflights is being continued, but adding to it the all-up spacevehicle concept. This concept requires that, to the extentpracticable, all flights will be scheduled as complete spasevehicles, This approach intensifies reliance on a compre-hensive ground test program which involves development,qualification and integrated systems tests. The ability tocapitalize on success offers us the potential for earlymeaningful missions at a significant cost saving.

Emphasis in the ground test program is focused on designverification, qualification and acceptance testing ofcomponents and subsystems, and comprehensive factory andlaunch site checkout of total systems. The ground testprogram, however, not only involves rigorous qualificationand checkout, but also includes many special test articlesfor integrated testing. They include those for propulsiontests, systems compatibility tests, facility checkout,dynamic and structural tests and many others. Twenty-one

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additional spacecraft test articles provide test data ondynamic loads and response, water landing, parachute recovery,flotation, structural integrity, thermal-vacuum characteristicsand abort characteristics to insure operational reliabilityover the entire flight regime. A typical example of thistype of an integrated test article is shown in Figure 12.The photograph shows an Apollo command and service module ina 60 x 120 foot Space Environmental Simulation Chamber.Complete Systems Functional Tests are performed in a mannedoperation while the spacecraft is exposed to liquid nitrogencooled balls and a vacuum of 10-5 torr.radiates heat onto the spacecraft surfaces, programmed aswould occur in an actual space flight. This test facilityin combination with a flight-type test article permits theevaluation of the combined capability o f man and spacecraftsystems to perform to the requirements of present and futurespace flights.

A solar simulator

An Apollo test program phasing chart is presented inFigure 13. A high level of ground test efforts commenced atthe outset of the Apollo Program and will be sustainedthrough the early manned flights. It is planned that allground qualification tests be completed prior to initiationof the first manned Apollo flights.

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Systems Engineer ing of Special ProjectsMany s p e c i a l s ys te ms i n t e g r a t i o n a c t i v i t i e s o c cu r

throughout t h e l i f e of a program. Incl ude d i n these e f f o r t sare such areas as e x tr a ve h ic u la r a c t i v i t i e s and s c i e n t i f i cand t e ch n i c a l experim ents. The Gemini experiment programw i l l b e used t o i l l u s t r a t e t h e f a c t o r s i nv ol ve d.

The experiments in teg ra t ed i n th e spacecraf t rangedf r o m r e l a t i v e l y simple ones such as cameras t o complexexperiments which had t o be s t ru c t ur a l ly mounted, thermal lycont rol led and automat ica l ly deployed for taking measurements.In add it io n, some experiments had ex tens ive data recordingand tra ns mi ss io n require ments, Even the r e l a t i v e l y s i m p l eexperiments, however, presented a s i g n i f i c a n t i n t e g ra t i o nproblem due t o th e s m a l l conf ines of th e spacec ra f t cab in ,Most of the remaining usable space was f i l l e d w i t h experimentand crew equipment. Experiments were mounted o r stowed onthe overhead hatches, t h e cabin walls, and even on t h e f l o o r ,as shown i n F igu re 14.o u t s i d e t h e cabin d i d n o t presen t a s ign i f i can t s towage

Experiments which could b e mounted

problem; however, t he y tended t o b e more complex since theycou ld no t gene ra l ly be manually operated and had t o haveaut oma tic pr ov is io ns . An example of an experiment w i t hcomplex integrat ion requirements i s shown i n F igu re 15.

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Por t ions of t h i s experiment which measured ultraviolet andi n f r a r e d ra di at io n i n space were mounted throughout t h espacec ra f t . The sensors , one a l i q u i d neon cooled spectro-meter, i nvo lve pyro techn ical ly ope ra ted ha tches , and spe c ia ldeployment and alignment provisions.

A s i g n i f i c a n t p o r t io n o f the crew t ra in ing was devotedt o t h e experiments. The t r a i n i n g i n vo lv e d b o th l e a rn i n g t h eexperiment objectives and equipment operation. The crewt r a i n i n g t ur ne d o u t t o be a two-way s t r e e t since manym o d i f ic a t i on s i n t h e experiment design were suggested by th ecrews t o improve t h e e qu ip me nt 's o p e r a t i o n a l c h a r a c t e r i s t i c s .I t was impor tant for the crew t o under s tand th e experimento b j e c t i v e s a nd u n d er l yi n g p r i n c i p l e s t o take f u l l a d va nt ag eof t h e i r di sc r e t iona ry ab i l i t y and modi fy the experiment planas requi red t o adap t t o t h e s i t u a t i o n s .

Mission planning for the experiments was a very complexproblem involv ing many i t e r a t i o n s of the f l i g h t p la n t o f i tth e exper iments i n t o th e proper place. Some experiments hadt o b e performed on the nigh t s ide o f t h e o r b i t , o t h e r s ont h e d a y l i g h t s ide. Some experiments had t o b e performedear ly i n the mission , o th e r s l a t e i n the mission. Someexperiments were designed t o measure the nea r earth r a d i a t i o n ,o t h e r s were damaged by i t and provisions had t o be made t o

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minimize the e f f e c t s of ra d ia ti on . Some experiments workedbes t a t high a l t i t u d e , o t h e r s a t low a l t i t u d e . I n a d d i t io nt o c o n s i d e r a t i o n s su ch as these, there were o t h e r s such aspayload margin and pr op el la nt consumption. Every e f f o r t wasmade t o u t i l i z e the payload c a p a b i l i t y a v a i l a b l e i n o r d e r t oprov ide fo r the greatest p o s s i b l e exper iment capabi l i ty .P rope l l an t t anks were added t o t h e spa cec raf t t o maximizeth e amount of prope l lant ava i lable f o r t h i s p a r t of thef l i g h t program.

There were a t o t a l o f f i f ty- two exper iments i n theGemini Program. In general , each exper iment was flowns e v e r a l times t o take advantage o f varying f l i g h t condi t ions .This re su l t ed i n one hundred and eleven experiment missions,an average of eleven experiments per f l i g h t . The experimentswere d i v id e d i n t o three c a t e g o r i e s ; s c i e n t i f i c , t e c h n o l o g i c a l ,and medical. A wide va r i e ty o f very in t e r e s t in g and use fu l.r e s u l t s were obtained f rom these experiments and have beenrepor te d on i n numerous tec hni ca l and s c ie n t i f i c papers . A san example, t he r e are t h e ul t r av io le t spec t rograph ic pho to-graphs obta ined of t h e s t a r Canopus during standup extra-v e h i c u la r a c t i v i t y ( Fi gu re 16) . The l i n e s i n t h e upper endof the spectrum { l i n e s or magnesium and iron) are n o tt r a n s m i t t e d by th e ea r th s atmosphere and were recorded fort h e f i r s t time i n t h e spectrum of a s tar o t h e r t ha n our sun.

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In-Flight and Post-Flight ConsiderationsAlthough the main emphasis in systems engineering prior

to flight is to achieve a high degree of assurance that amission will proceed as planned, one must also consider thepossibilities that events can occur which will change themission. Therefore, a substantial amount of effort has beenplaced on considerations of alternate missions which willachieve maximum benefits from the flight in the face ofcertain contingencies. This aspect of planning involvessuch factors as re-cycle plans as a result of launch delays,alternate mission plans related to major failures duringlaunch which abort a mission, and contingency flight plansbecause of problems evidenced during the orbital flight.Obviously, it is not possible to do planning for eachdetailed failure that might occur, but rather to considerbasic classes of problems for which other approaches can beestablished. In the flight program to date, a substantialamount of effort has been placed on this aspect of our flightoperations and, in the majority of the flights, this activityhas paid off in some way. Generally, this planning evolvesfrom a series of logic diagrams from which such things asre-cycle plans, alternate mission plans, contingency flightplans and the like are developed.

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A highly simplified version of a logic diagramindicating alternate mission possibilities concerned withone Gemini flight is presented in Figure 17. Also includedon the figure is an indication of how the mission was actuallycarried out in the face of certain contingencies. In theGemini IX mission, the basic plan was to launch a targetvehicle into orbit, followed by the launch of the spacecraftwhich would then rendezvous and dock with the target vehicle.Subsequent to docking an extravehicular program was to beconducted, followed by a series of re-rendezvous operationsto investigate purely optical rendezvous techniques as wellas more difficult approach geometries. The mission thenended with the reentry and landing of the spacecraft. Theattempt to launch the target vehicle was a failure becauseof a control system problem in the launch vehicle. A planhad been previously developed to immediately place a backuplaunch vehicle on the launch pad and to utilize a simplertarget vehicle, also maintained in readiness. The detailsof this re-cycle plan showed that the checkout process ofthe replacement system could be accomplished in two weeks,providing the determination of the cause of failure and theresulting corrective action could be concluded within thattime period.

The re-cycle plan was initiated immediately after theIt wasfailure and in parallel with the failure analysis.

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p o s si b le t o i s o l a t e th e most probable cause of fa i l u r e andsecondary p o s s i b i l i t i e s a nd t o make approp r i a t e modi fi ca t ionst o th e backup launch vehicle within the t i m e frame planned.Thus, t h e new t a rge t was la un ch ed i n t o o r b i t w i t h i n t h e t i m ep e r i o d s p e c i f i e d a n d was f ol lo we d i n t o o r b i t by t h e space-c r a f t . I t may be noted a t t h i s p o i n t t h a t contingency f l i g h tplans had been de t a i l ed for a number of bas ic in- f l ightcont ingencies such as not be ing ab l e t o achieve rendezvousor t h e i n a b i l i t y t o dock. On t h i s p a r t i c u l a r f l i g h t t h erendezvous was achieved wi thout inc ide nt , but because apro tec t ive sh roud on the t a r g e t d i d not deploy proper lydur ing launch, t h e docking mechanism was not exposed and i twas impossible t o dock. A t t h a t p o i n t , a contingency f l i g h tplan was i n i t i a t e d as had been pr ev io us ly developed. Thus,i n t h e ac tu a l m ission , i t was necessary t o per form t h ere - rendezvous .exerc ises prior t o t h e e x t r a v e h i c u l a r a c t i v i t ybecause undocked station keeping of th e t a rg e t would havere s u l t e d i n p ro h ib i t i ve p ro pe l l an t consumption and fo rcede l imina t ion , no t on ly of e x t r a v e h i c u l a r a c t i v i t i e s , b u t manyof t h e programmed experiments as well . In t h i s manner, th ef l i g h t o p e r a t i o n was able t o proceed smoobhly i n t o t h e newplan and achieve a major i ty o f t h e miss ion ob jec t ives des i r edf o r t h i s f l i g h t .

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Even when missions proceed essentially as planned orproceed on a contingency basis, system or operationaldifficulties that occur must be corrected prior to initiationof the next flight. In the Gemini Program, an attempt wasmade to establish an analysis, reporting, and correctiveaction system which avoided this potential delay in theprogress of the program and it was very successful.

In targeting for two-month launch intervals thepublication of the mission evaluation report was set atthirty days In turn, a major part of the data handling,reduction, and analysis activities took place in a period ofapproximately two weeks following each mission. This timescale presented a rather formidable task when one considersthe volume of data accumulated during the mission. TheGemini flight data production rate can be envisioned from theinformation presented in Figure 18.the number of tabulations and plots required for the analysisof a typical long duration mission. The acquisitfon of largequantities of data, combined with the need to evaluate andquickly resolve anomalies, resulted in the utilization of amethodology for selective reduction of data which has provedeffective. This selection process involves real time missionmonitoring by evaluation engineers which, in turn, resultsin a judicious selection of flight segments for which data

The figure also indicates

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are needed. Th is monito ring, combined wi t h t h e a p p l i c a t i o nof compression methods f o r t h e p r e s e n t a t i o n of data, hasmade i t poss ible to comple te eva lua t ions on a t ime ly basis .

A 1 1 problems were n o t n e c e s s a r i l y s o l v e d a t t h e end ofth e thi r ty-d ay repo r t in g per iod, bu t problem is ol a t io n, impac te v al u at i on , a nd c o r re c t i v e a c t i o n i n i t i a t i o n were p o s s i b l e i nt h i s t i m e pe r iod . I n ca r ry ing ou t these a c t i v i t i e s , a formaltask group i s s e t up with personnel assigned who have beena c t i v e l y w orking i n s p e c i f i c areas b e f o r e t h e f l i g h t anddur ing the f l i g h t . T h i s approach provides personnel al readyknowledgeable wi th t h e background of t h e p a r t i c u l a r f l i g h t ,C o r r ec t i v e a c t i o n i s i n i t i a t e d as soon as t h e problem i si s o l a t ed and aef' ined,

Swn-naryA s implied rrom t h e di scuss ions i n t h i s paper, systems

eng inee rin g encompasses a l l phases of a manned space f l i g h tprogram, An attempt has been made t o d escr ibe some of t h emajor c on ce pt s u t i l f z e d i n t h e U ni te d S t a t e s t o p r ov id e f o ra logical development of t h e f l i g h t hardware, for t h esuccessful conduct of f l i g h t opera t ions , and f o r achievementof maximum benefits from each mission. We a r e n a t u r al l y i nth e process of continu.ing these concepts i n t h e on-goingApollo Program i n o rd er t o meet i t s l u n a r l a nd i n g o b j e c t i v e .

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The greatest unde rly ing in f luence t o our approach i s t h eassurance of f l i g h t sa f e ty , bu t w i th in tha t c o n s t r a i n t , as t rong secondary inf luence i s th e achievement of missionsuc cess . Undoubtedly, i n the f u t u re , new and more extend edmiss ion ob jec t ives w i l l be def ined and some a l te ra t ion to thepresent concepts can b e expected because of the p e c u l i a r i t i e sof th es e miss ion s. However, th e major f ac to r s of t h e presen tapproach appear sound and worthy of continued consideration.

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Systems Engineering - Manned Space Flight

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Transcript of Systems Engineering - Manned Space Flight