ai-aerospace/ams_data_full_2000-2020 huggingface dataset:
Introduc tion 

The PRA CA system has been used since 1997 t o track, d ocument, an d approve the investigatio n  analysi s, root  cause dete rmination, corrective a ction, and re currence cont rol for all hardware probl ems on  ISS. The S&M SPRT man ages the re solution of all structural and  mech anical system p roblems. Sin ce its  inception the  S&M SPRT h as worked on over 350 problems rangin g in seve rity from galled  fastene rs on   the gro und to  failed deploy ment syste ms on-orbit. The PRACA dat abase can b e searched fo r key wo rds  or othe r field identifiers and  is use d for trendin g of system probl ems on the vehicl e.  

Problem Reportabilit y Criteria 

Most ha rdware non -conformances are not co nsidered to be PRA CA reporta ble. Hardware discre pancies  that occu r during th e ma nufacturi ng proce ss are usually dispo sitioned th rough a lo cal M aterial Revie w  Board s with stand ard repair proced ures. A PRAC A re porta ble proble m is an omal ous hardwa re  perfo rman ce that occurs d uring Qualification or Acce ptance testin g, or after del ivery of the hard ware from   the cont ractor to NASA while the sy stems are being readi ed for l aunch, or  occur during or after activation  on-orbit. PRACA re portable problems that o ccur on the ground are inve stigated a nd worked until the   hardware is either b rought back to p rint or to a waiverable conditio n where it can be a ssured  to meet all  system  requireme nts. O n-orbit th e problems a re worked  until t he h ardware can be repaired or an  accepta ble condition can  be found so the hard ware can co ntinue to funct ion and mee t most of its   perfo rman ce requi rements until a future time whe n a repair can b e impleme nted. 

PRA CA Rev iew 

A detailed re view of all S&M PRACA’s was perfo rmed. The root cau se of e ach PRACA wa s noted and  categ orized. It was found that most PRACA’s co uld be gro uped into one of 9 types. Figure 3 gives a   percenta ge breakdown for these nin e categori es. As can be seem toleranci ng errors and design errors  accou nt for almost 40%  of the PR ACA repo rtable problems that occu rred in the structu res and   mech anical system s. M anufa cturing type pro blems m ade up 15%.  Faste ner probl ems and  environ ment/ procedure problem s acco unted fo r 9% and 8% resp ectfully. Te st and STE  related p roblems  each accou nted for 4% of the PR ACA’s. Panel Ret ention ma kes up 2% with the other miscella neou s  probl ems making up 23%. While th e miscell aneous category makes up a sig nificant numb er of  anomali es, there did n ot appea r to be a ny recu rring theme s althou gh ea ch on e rep resents a probl em that  had to be worked with the impleme ntation of a corre ctive action an d recu rrence control. 

PRACA Brea kdow n by Type Test 4% Fasteners 9%STE 4%Panel Retention 2% Environment/Proced ure 8%Manufacturing 15%Design 17%Tolerancing 18%Misc. 23%

Figure 3.  Problem Repor ting Bre akdo wn  293
repai r and rework. Fo r future ma nned vehicle d esigns, care should be taken to provide ve ry robu st and   easily re placeable faste ners for ha rdware like pa nels that will be fre quently utilize d by the cre w.   

Other Mi scellaneo us   Failure that  could not b e placed into one of th e above catego ries, and the total numbe r of si milar events  was less then 5. More categories were used in t he initial review of the PRACA’s  like corros ion,  config uratio n control, and unexplain ed anomali es, but none of them resulted i n more than 4 occurren ces  and thu s were grou ped int o the miscella neou s catego ry. 

PRACA’s with Larges t Program Impact 

The o riginal p urpose of thi s revie w was to dete rmine the type s of problems that occurre d most freque ntly  in a large p rogram li ke ISS . It became clear d uring thi s revie w that it would al so be be neficial to discu ss  some of the failure s that have had the  large st impa ct or potentia l impact to missi on obje ctives, co st or  sched ule.   

Figure 4.  Photo of BGA and 4-Bar As sembly  

BGA Hinge L ock Failu re and Excessive  Deployme nt Force (P RACA 2389 & 243 5)   This fail ure occurred during the d eploy ment of the f irst solar arrays on th e P6  eleme nt of ISS. Durin g  activation of the element a four bar m echanism is us ed to rotate the sola r array  mast canisters and  blanket boxe s from a lau nch position to an on -orbit position. Lo cking featu res at the ba se of four ba rs  enga ge whe n full rotatio n is achieve d. The deplo yment force on-orbit was significantly higher tha n  expecte d an d not all locks could be  prop erly en gage d. An exhaustive inv estigatio n revealed that   insuffici ent control of the  manufa cturing toleran ces allowed binding to o ccur in the m echanism which  increa sed the required fo rce to d eploy a nd caused the hinge lo ck to malfun ction. Also, a  design feature   that relied o n friction to m aintain alig nment of critic al part s failed  when subjected to the bin ding loa ds. A  redesign was unde rtaken to correct thi s on the remaining six solar array four ba r mechanisms along with  tighter co ntrol of the asse mbly process a nd functio nal checks to veri fy deploymen t force s.   

297
Figure 5.  Photo of comm on hatc h.  Incorrec t handle position in red. 

Conclu sion 

Every failure should be l ooked at for le sson s lea rned to preve nt the same p roblem from recurrin g. This  report attemp ts to loo k at the failures on ISS not on a case-by-ca se ba sis, but rather from a more global   view of the  identifying the  most frequ ent types of fa ilures and th eir root causes. A p rogram as large as  ISS affords t he uni que opportu nity to do this because of the l arge number of mechani sms all b eing  manag ed by one team. It was found th at what appe ar to be the simplest of pro blems li ke toleran cing and  fastene r design are still the most com mon problem s in the dev elopme nt and operation of spa cecraft  mech anisms. A large number of de sign problems were  not identified until qua lification and  accepta nce  tests we re perform ed, rei nforcing the validity of these  test s ev en in  the fa ce of cost and sched ule  pressures. Care mu st be  given in de sign revi ews of even the simple st mech anisms t hat adeq uate  attention is paid to detai ls especially  whe n they can directly affect syste m perfo rmance. Man datory  inspectio n points should be called  out to assu re all critical dime nsions are check and functional  tests  should always try to re plicate the enviro nmental conditions that the  hard ware must op erate i n. It is hoped  that these results will be used to focus designers of f uture spacecraft mechani cal system to be sensitive   to the most common failu res that have been exp erienced on ISS. 

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Space Stat ion Berthing Mechanisms,  Attaching Large Structures On-Orbit th at were Never Ma ted on the Ground 

Robert M. Fo ster∗, John G. Coo k*, Paul R. Smudde∗∗ and Ma rk A. Hen ry+ 

Abstract 

A significa nt benefit re sulting from the Intern ational Space Station (ISS) prog ram h as be en the  developm ent of commo n mechani sms to attach  the various modula r compone nts to gether in  orbit.  Several different mechani sms were develope d to ma ke assembly and g rowth of the ISS possible. Thi s  pape r add resses the d riving desig n requi rements and ve rification activities a ssociate d with these   mech anisms that ensu re that all ISS module s, buil t in different parts of the world and no t coming in   conta ct prio r to on-o rbit assembly, would fit together wi thout probl ems. 

Introduc tion 

Two basic t ypes of berthing m echan isms are u sed on ISS. The first type, pre ssurized be rthing  mech anisms, req uire the e stablishment of a pressu rized pa ssag eway fo r transfer of crew and equip ment  betwe en inhabited volum es. Incl uded in this type  are the Boein g-designed Com mon Be rthing  Mechani sm  and a number of Ru ssian-designed mech anisms that were used in t heir very succe ssful MIR S pace  Station Prog ram. The second type, un pressurized berth ing me chani sms, i s used to attach the vari ous  module s on the out side of the ISS. Included i n this type are the Truss elem ent assem bly system s and  their de rivatives u sed to attach external payload s and other compo nents to the ISS. 

All major ISS element s are positio ned u sing a roboti c arm, either th e Shuttle Rem ote Manip ulator System   (SRMS) o r the Space Station Rem ote Manipulato r System (SSRMS), re sulting in similar initial mating  conditions for all berthin g mecha nisms. Howeve r, the interfa ce geomet ry and the resultin g post-m ate  config uratio ns are very di fferent. Therefore, ea ch of the berthi ng me chanisms is uniquely design ed to  meet the spe cific re quirements an d nee ds of the interface s that they join. 

This pa per focu ses on f our be rthing mech anisms used to a ssembl e ISS comp onents. They are a s  follows : 

Comm on Berthing Mechan ism (CBM)  Segment-to-Segment Attac hment Sys tem (SSAS)  Rocketdyne Tru ss Attach ment System (RTAS )  Common Attac h System (CAS) 

An explode d diagram of th e Intern ational Space Station ide ntifying the majo r assembly elem ents i s shown  in Figure 1. T his figure sho ws which ele ments a re already assem bled in orbit an d the addition al element s,  built by the United State s and ou r Internatio nal Pa rtners, that are to be inst alled as the Internation al  Space Statio n is compl eted. Each of th e mechani sms discusse d in this p aper has already been used at  least on ce during a ssembl y of the ISS. 

                                                  ∗   The Boeing Comp any, Hou ston, TX  ∗∗ The Boeing Comp any, Huntin gton Beach, CA  +   The Boeing Company, Hunt sville, AL 

Procee dings of the 37th Aeros pace Mec hanisms Sy mpos ium, Joh nson S pace Ce nter, May 19-2 1, 200 4  301
Figure 1.  Internatio nal Space Sta tion Assembly1 

Generic Ber thing Require ments 

There a re two distin ct met hods, be rthing and docking, for atta ching ele ments together in space. Both of  these  metho ds are used on ISS. Bert hing i s defin ed as assisted atta chment usi ng either th e Orbiter  (SRMS) or the Space Station Remote M anipul ator Sy stem s (SSRMS). Do cking is una ssiste d attachm ent  of an autono mous vehi cle or element. Examples of doc king are the Orbite r to ISS and Russian Soyu z or  Progress veh icles to ISS. All of the ISS elements that are deli vered in the Orbite r Paylo ad Bay are   berthe d to the Space Stati on. 

Many of the requireme nts that dr ove the  berthin g mechani sm de signs apply to all berthing m echanisms.2   Each mechanism must accommo date uncertainties in Remote  Manip ulator System (RMS) control,  dictating a relatively large captu re envelope for t he transitio n from RMS ope ration to berthi ng mechani sm  captu re device ope ration. Final ali gnme nt, on the other hand, m ust be  preci sely cont rolled in orde r to  orient th e elements correctly, and the  berthing  mechanism  design must be able to p ull the element s into   their final alig nments, p reload the inte rface, a nd the n provid e a structu ral atta chme nt sufficient to rea ct  the interface loads. For the CB M, SSAS and CAS, the large capture envelope and remote operation of   the mechani sm req uires feedba ck to ind icate when th e eleme nts are in ra nge of the ca pture mech anism,  hence th e incorpo ration  of ready-to-lat ch indicators and  telemet ry. The RTAS has a similar captu re  envelop e; ho wever, the re ady-to-l atch indication is visual o nly due  to the RTAS being ma nually operated.  

                                                  1 NASA Illustration: http://www.h q.nasa.gov/osf/ISS_Core_P rogram.pdf  2 SSP 41000, “System Specification for the In ternational Space Stat ion”, 01 Oct 2003  302
fine alignin g pins o n the A ctive half eng age sockets of the Passive half to begi n the final ali gnment of th e  two h alves. Standoff plu nger a ssemb lies (sprin g loade d sho ck ab sorbers) of the Pa ssive half rea ct  again st Strike  Plates of the  Active half t o ensure relative motion of  the halve s is damp ened and prevent  uncontroll ed seal conta ct and se al da mage befo re captu re is complete. Once  capture and  alignment is  compl ete, the  two halve s are ready to b e structurally prelo aded together.  This is accomplished by sixtee n  electro-m echanical a ctuator driven  powered  bolt assemblie s, equally sp aced around the perimeter of the   Active half, extending and  engagin g floating nut assemblie s on th e Passive h alf. The powered bolts a re  loade d in a serie s of step s to evenly distribute t he lo ads, en suring prope r seal conta ct and compressio n  along with re quire d preload for p ressurization and on orbit loa ding of the be rthed ISS module s. Shea r and  torsio n loads are  reacted by the fully e ngag ed fine  alignme nt pins and sockets. The  electro-me chanical  actuato rs on the powered bolts and ca pture latch assemblie s are po wered and controlled  by individu al  motor controllers group ed into four sets of five c ontrol lers on removable controller p anel assemblie s. An  RS-4 85 lo cal bus n etwork within the A ctive CBM in terlinks the controlle r panel assem blies. The Active  CBM is, in turn, linked to th e ISS internal  compute r through a MIL -STD-155 3 Bus network. 

After berthing  and pre ssurization i s com plete, the co ntrolle r panel  assemblie s can b e remo ved, thereb y  enlarging th e available p assage way b etween th e ISS module s to the maxim um allo wed by the hatch   envelop e (an approximatel y 127 cm (5 0 inch ) square openi ng). 

Nut  (16)  Powered Bolt   (16) Capture Fitti ng (4)Passive CBM Ring (1) Standof f  (16)  Strike plate (16) Ready  to Latch Indicator (4)Capture latch (4)Alignment  Guide  (16)  Contro l Elec. Panel (4)Active  CBM Ring  (1) Socket  (4)  Alignment Pin  (4)  Figure 3.  CBM major co mponents 

CBM Verification   The ve rificati on of  the CBM sy stem was accomplished using a combinatio n of toleran ce analysis, FEM  and T hermal  analysi s, capture dynami cs an alysis and testing, comp onent level testing,  stan dalon e  system level t esting a nd flight article te sting.  

Comp onent Level Testing   Captu re Latch, Powe red Bolt and Flo ating Nut, Ready to Lat ch, Coa rse Al ignment G uides and Fi ne  Alignment Pins and So ckets verificatio n testing we re performed at compo nent levels. Alignment Guide s  and Pins/Sockets we re subje cted to static load testing  only. Captu re Latch testin g established th e  comp onent mech anism op eration, rea ch ability and capture operation. Powe red Bolt/Floating Nut te sting  establi shed the comp onent mechani sm ope ration, rea ch and pre-loading capability. Re ady to L atch  testing e stablished the compon ent m echanism op eration. Ve rification of op eration  for t hese three   304
RTAS Design Feature s and Operation   ifficulty in getting po wer to  the outboar d truss elem ents on ISS   As the truss segm y-to-Latch (RTL )  hen the fine alignment mech anism fully se ats in its recepta cles, the capture bar is approximately 6 mm  The RTAS d esign was driven by the d durin g assembly. It was d ecided early-on to u se EV A cre wmembers to op erate the RTAS mech anisms  instea d of providing com puter cont rols and po wered  mechani sms. Rea dy-to-l atch i ndication  is provide d  by the EVA c rew.  ent b eing berthed i s brought into physi cal p roximity, the vi sual ReadCapture Latch Fine Alignment Bolt MechanismCoar se Alignment Receptacle

Coarse Ali gnment Z  X Y Free in Y-Z  Plane Note : in a ddition to the  in-plane   freedoms identified, all four  corners h ave limited swivel  freedom ( O y ,  O z ) Fixed in Y-Z  Plane Free in Y-Z Plane  Free  in  +  Y  Capture Bar  Figure 11.  Passiv e RTAS Componen ts (P6 Trus s Inter face Plane)  Figure 10.  Activ e RTAS Compon ents (Z1 Inter face Plane) 

indicator (a painted tip on the coarse alignme nt cones) di sappears from  view as the coarse alignmen t  cones pen etrate their re ceptacle s. The captu re procedure requi res that three of four cones line up fo r  “RTL crite ria” to be met. This place s the seg ment within the ca pture envelop e of the RTAS.  Crewmembers  then us e the electric EVA power tool to ac tuate th e capture lat ch from the EVA works ite  location. As th e latch jaws close on the capture bar, the segm ents are p ulled together.  At a bulkhea d-tobulkhea d dist ance of appro ximately 13 mm (1/2”), the fine alignm ent mech anism begin s to engag e  .  W (1/4”) from ful ly seated.  Continuing to  close, th e capture latch deflects the capture bar an additional 6  mm  (1/4”), which produces approximately 8 .2 kN (1850 l b) of pre-load between th e structu res. At this point ,  the EVA crewmem bers engage the prim ary structural attachment bolts  to fasten the two segm ents to  each other. 

310
Final On-Or bit Assembly   Verification 

Asse mbly of  the Intern ational  Space  Statio n, acco mplish ed fa r  from the m anufacturi ng flo or onorbit, introdu ces some unique  probl ems. Each stage of  assembly, a s ha rdware and   increa sed capability is ad ded to  the vehicle,  must re sult in a  viable, life-suppo rting ve hicle,   with no l oose end s or safety  issue s for the vehicle or the   crew inhabiting it. Failu re to   berth a ne w eleme nt of the  Station incu rs immense co sts a s  that elemen t must the n be  returned to  earth an d relaunched or, possibl y,  discarded. Comp ounding the  probl em - ti me, distan ce and   budg et con straints in almost   every ca se pre clude trial   assembly an d fit on the groun d  prior to la unch. In many cases, the new element was not even  built  when the  element it is to b e berthed  to  was laun ched.   PlatformGuide Vane Asse mbly #3Guide Vane Assemb ly #1 Capture Latch Assembly Umbilical Mechanism Asse mblyMCAS/Payload Coordinate System* *X-Y Plane located on  GVA well bottom  cen terlines. X-Z Plane located at C LA cen terline.Guide Vane Assemb ly #2Z YX Figure 12.  Activ e CAS Platform 

A pro cess was d evelop ed to minimize the ri sk inhe rent in a ssembling the elements fo r the first time o norbit. Thi s process, titl ed “A ssembl y Analysis”,  uses a combinatio n of CAD models, as-buil t  measurement s, and comp uter-based an alysis tool s to verify the berthing me chanism will fit prope rly, that  pressure and  thermally ind uced defle ctions and di stortions can be accomm odated, and that no ha rdware  external  to th e berthing mechanism will cause i nterference proble ms.  Use of this process h as resulte d in  a numbe r of hard ware modificatio ns prior to lau nch and a 10 0% succe ss rate for on-orbit berthing   operation s.      Asse mbly An alysis uses ki nematic CAD software  to process multipl e interference checks in an automate d  fashio n. The  CAD software move s the incoming  element throug h a set  of lateral and ang ular  misalig nment s that bound the nomin al ap proach corridor. During  move ment, at specific interva ls,  interferen ce che cks are c ondu cted. T he first i nterference fo r each path i s identified, and  the in comin g  element mov es to the next start point. The interferen ces can the n be plotted as a mesh (Fi gure 13). Thi s  mesh  is a method to  visu alize the hardwa re constraint corrido r. Use of this method makes it easy to  visuali ze intrusio ns into t he n ominal  approach corridor. By m oving the  incoming el ement t hrough thi s  regim en of interferen ce che cks, we can show that  there are no inte rferen ces bet ween the geo metry. 

As-built mea sureme nts are use d to update the CA D model s. This is a ccomplished by measurin g the   hardware in a comm on coordi nate sy stem. This i s accompli shed by using a digital phot ogrammetry   system. This system allows  for the precision me asurement of the hardware relative to the  CAD mo del of  the hard ware. The measu rement data for ea ch pa rt is saved a s a clou d of point s. The CAD g eometry for  each part is t hen be st fit to  the respectiv e clou d of poi nts.  312
On-o rbit effe cts a re add ed by a simila r  process. Fin ite element  model s are   created of the interfa ce hardware and  they are subje cted to  thermal a nd  pressure lo ads to d etermine  the   deflectio ns. These d eflection s a re  captu red a s a cloud of p oints and th e  CAD geom etry is be st fit to  this cloud of  points.  Figure 13.  Assembly  Analysis Map of Repr esentative  Berthing Co rridor 

   Figure 14.   Photo of Common Captu re Cla w  Com pared with CA D Mod el   A critical a spect of this process i s  control of h ardware an d CAD mo del  config uratio n. In order to  obtain vali d  results, the CAD mod el must mat ch the  hardware. Config uratio n control i s  further complicate d as the on-o rbit  hardware changes from  the as-la unched  “Cargo Elem ent” i nto an integrate d ISS  module th rough a series of on -orbit  modificatio ns and depl oyments. An   example of the level of model fidelity is  shown in Figure 14, where the  hardware  photo a nd CAD imag e are presented  side-by-side.   

Lesso ns Lea rned 

1) Qualification of mech anical syst ems should, a s much as possible, establi sh the full ca pabili ties of  the system. Minimum test ing that ju st meets the re quire ment s, while initially less expensiv e, leaves  no margin tha t can be expl oited later to resolve u nforeseen problems or op eration s changes. 

As ea ch of these me chanism designs mature d, and ou r understanding of the overall inte gratio n  task for ISS gre w, additio nal req uirements su rfaced. A prime example of this is with the modified  RTAS mec hanism us ed to assemble the outboard truss eleme nts. Initially the sh ort sp acer between  the two solar array truss es was  planned to be in stalled by the EVA c rew and the mec hanisms  designed with that in  mind. Only recently, ope rational concern s have  led to a requireme nt to use  the  SSRMS to perform the in itial installatio n of the short spacers (P 5 and S5, See Figu re 1).  This  change h as resulte d in a tremend ous amount of ana lysis an d testi ng to determi ne the comp atibility  313
Figure 1.  Representativ e ISS Pressurized Elements and Trusses. P

The pressurized modules provide a habitable environm ent at Earth-normal atmospheric temperature and  pressure. This environment allows research to be conducted in a shirtsleeve laboratory environment for  the astronauts. The trusses are an open framework stru cture similar to a bridge truss. The trusses  support the array of systems that are needed to make the ISS a viable orbital vehicle, such as the  electrical power system; the ther mal control system; the guidance, navigation, and control system; and  the mobile servicing system.  ressurized ModuleTruss Segment

The US-built elements of the ISS are assembled using berthing mechanisms, the types and uses of  which are given in Table 1. 

Table 1.  Assembly  Mechanism Ty pes and Usage Summary  

Assembly  Mechanism Qty Type First-tim e use Final Attachment  Mode  Common Berthing  Mechanism 23 Pressurized 16 Automated  Truss Attachment System 6 Non-pressurized 4 Automated  Truss Attachment System 8 Non-pressurized 8 EVA Attached 

The berthing mechanism used to attach pressuriz ed modules together provides a pressurized path  between the ISS habitable elements and maintains stru ctural integrity between them. Another type of  berthing mechanism is used to join the truss struct ures that are the backbone of the ISS. These truss  attachment mechanisms provide structural attachment  only, since truss segments are not pressurized.  There are 48 first-time uses of berthing mechani sms. The Common Berthing Mechanism (CBM) is used  to berth pressurized US elements and International Partners, except for the Russians.  The Russians  provide their own docking mechanisms that serve the same function as  the US assembly mechanisms.  The truss attachment systems are used only to join US-built truss elements11.  Each berthing task is accomplished using a robotic mani pulator system to maneuver the element from the  US Space Shuttle cargo bay to the assembly locati on on the ISS. The element is positioned at the final  pre-assembly location with an accuracy of  inches during berthing operat ions. The final phases of  316