ai-aerospace/ams_data_full_2000-2020 huggingface dataset:
Obviou sly, meeting the se requi rements wa s a challenge. Further compli catin g the task wa s the fact that  no laun ch ve hicle ha d been sel ected, thus re quiring the worst case lo ad co nditions envel oping Aria ne,  Delta, and Atlas be u sed. 

The Initial Concept 

From th e start, it made sense to use a comp osite  boom for th e magnetom eter to minimize t he ma gneti c  contami natio n. The  boom needs to be fold ed around  the periphe ry to  get the le ngth requi red and  minimize volume and de ployer interfe rence. This boom wrap -arou nd de sign requi res three hi nges,  anthro pomorphically refe rred to a s the shoulder, elbow and wrist joints. Integ ral folding “ca rpenter tape ”  hinge s with n o slidi ng or rotating pa rts were chosen as the simplest way to ensure positive d eployment  torque -ratios at the least mass whil e avo iding the is sues of de signing for frictio n and dam ping.  This boo m  design reli es on the strain energy in the buckle d hinge to straighte n the hinge a nd depl oy the boom. 

Figure 2.  Bo om La yout 

Mounting 

The sides of the sp acecraft are covered with solar panel s and hard point s are avail able only at two   vertice s of the octagon. These coincide with  the cast aluminum  card cage th at provide s the primary loa d  path of the satellite. T he shoulder joi nt of the boom i s attached at  one of these reinforc ed vert ices and the  pin p uller that  hold s the ma gnetomete r uses the othe r. Thu s the ends holdi ng th e boom a re conne cted to  the ca rd cage rathe r than t he sheet met al sides that sup port the solar a rrays. T he intermedi ate elbo w an d  wrist joints a re sup porte d at the intermed iate vertice s by snub bers mounted to th e she et metal.  

Boom Relea se 

The meth od of boom release was driven by the l ow-power l ow-mass characteristi cs of the  ST-5 mi crosatellite. Low-power actuat ors from bot h Starsy s and TiNi Aerospace were cons idered. P yro actuators  had b een ruled out early on be cause of co ncern for shock in a  very co mpact structure and the current  requi rements on a 5-volt sp acecraft bu s. In the e nd TiNi produced a mo dified ve rsion of their P5-40 4-6SC  pin-p uller th at was sele cted. The mag netometer i s held to the spa cecraft throu gh a tang and clevis design  74
that utilize s the pin -puller pin in d ouble shear.  This hold s the bo om to  the spa cecraft in the  stowed con figuration. The mag netometer  mount is a se mi-ki nemati c desi gn co nsisting of  a cup-co ne and spring arran geme nt. The  relatively low minimum l ateral load capability o f  the pin pulle r, only 67 N (15 lb) shea r in the  worst ca se condition over the entire o peratin g  range, further drives the stowa ge sch eme. To  establi sh this numbe r rather exten sive testing  was done wit h a test ri g using parts that were  as flight li ke as possible. To maintain a release  force margin of two to one , 33 N (7.5 l b) was  used as the maximum all owable shea r load o n  the pin pulle r. This value l imits the ma ximum  tensio n that  can be applied to  the  boom   stowed. 

 Figure 3.  Bo om mounted  on Spacecr aft   Design Tr ades 

Initially it was h oped th at an all -composite  “mon olithic” de sign co uld b e made with  section s of the boo m cut  away to cre ate the  hinge s. Whil e gra phite fiber i s not n ormally   considered a fle xible material, when used in  this de sign, it allows th e boom to b end over 9 0°  and has many feature s that ma ke it optimal for  its intende d use. The graphite pro vides an  electri cally condu ctive pa th to bleed  off static  charge – a great b enefit o n the hig hly cha rged  missi on o rbit. It resist s creep a nd remain s  stable under large te mperatu re flu ctuation s,  allowin g for pre cise sci ence mea surements  even und er direct sol ar radiation. Th e thermal  conductance is optimum fo r controlling heat  loss from t he s pacecraft. Furthermore, the  comp osite m aterial exhibit s nearly non e of the   thermal ma gnetic effects th at cu rrently pl ague the met al hinge u nits.  Figure 4.  Boom Release 

Considerable effort was e xpende d dev elopin g this concept. Vari ous shapes were t ried fo r the cut out s  and diffe rent material s were tried, i ncluding Kevlar®,  E-Gla ss, S -Glass, an d various we aves and type s of  carbon  fiber.  Differe nt lay-ups of the ma terials were also tri ed. On e notable result from this study was the   fact that the  desired radiu s of a tub ular boom  sectio n was not usually the optimal  radiu s of curvature o f  the leaf or bl ade eleme nt of the hinge. An oval section boom  with cutout s for the hinge section  was  considered, but reje cted in  favor of a sa ddle pie ce to adapt the dif ferent ra dii. On ce it wa s realized that  the hinge de sign req uirement wa s necessarily differe nt than that of the basic tu be, the con cept of cutting  windows in  a tube to create a mo nolith ic boom was aban doned. It was also f ound th at ad ding a  fourth  facet in attempt to move the magnetometer further from the  spacecraft re sulted in the boom hitting th e  solar array on  deployment.  

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Monolithi c Boom With Inte gral Hi nge  Cut Out  

Oval Attempt to Have Larger Radius  for  Hinge Area  

Saddle Ada pter to Provide  Larger Radiu s for Hi nge  

Figure 5.  Hinge 

Design Dev elopment 

Titanium adapter shoes were  bon ded to the boom with the  hinge blades bolted to th em. A variety of  compos ite and metallic  blades were tried. The problem was  to make a hinge that was stiff enough when  deploye d, but that would  buckle at a l ow enough strain energy n ot to de stroy itself. The  boom joint s are  made up of t wo diffe rent l ength carpenter tap e hinges (longer at the shoul der joint and  short er at the  wrist   and elbow). The le ngth of the hin ges is driven by the a ngle th rough which the hing e has to be nd (120  degree s, sho ulder and 45  degrees, wri st and elbo w). Each hinge is ma de up of fo ur "carpente r-tape"  blade s sta cked two thick o n each sid e.  Analysis a nd developm ent testing sho wed that the bending strain   in a single blade th at was as thick as the two stacked bl ades was too  high. Every time the  thick hi nges  were bent th ey perma nently deformed.  Splitting the hinge into two sep arate ta pes cut the b endin g strai n  in half fo r each individual  blade. T his allows t he boo m to buckle and deploy a s designed  and still have   almost all the  stiffness of the ea rlier single hing e design. While th ese exp eriments and t rades continue d,  anothe r app roach usi ng ei ther Elgiloy or Bery llium-Co pper hi nges was pursu ed as a ba ck up.  

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Figure 6.  Hinge Tape Stacks 

The An alysis Program  

Hinge devel opment involv ed both  analysis and te sting. Th is boom design relies on the  strain en ergy in  the buckle d hinge to straig hten the hing e and de ploy the boom. Th e deploye d boom re sults in the hinge s  being i n a state that req uires considerably more fo rce to bu ckle again. Deploy ment test s of the co mplete  boom syste m have sho wn that the st ored strain energy is suffici ent to ca use th e shoul der hi nge to  snap  over. Thi s snap ove r is th e bu ckling of  a hing e in the di rection of deploym ent (opposite its originally  stowed state). However, af ter seve ral cycles the syste m damp s out with the b oom deploye d. The bu ckled   state requi red a n on-linear approa ch for a nalysi s. There was signifi cant effort in dev elopin g and  correl ating st owed hing e behavior m odels. Throug h strain -gauged mech anical bendi ng test s and photo - stress meth ods a hig h degree of correlation b etwee n the analysi s models and the p hysical model s was  achieved. This correl ation betwe en analysis and test has given  confiden ce to the mod eling tech nique  employed. T hrough thi s analytical tool material, lay -ups, hi nge l ength an d sub tended angle were tra ded  off against st rain an d pea k snap-over moment. Thu s a few hig hly likely candi dates were selecte d for  further te sting. The stre ss/strain results indicat ed th at a succe ssful comp osite  blade wo uld have to be  very thin. Subse quent testing sh owed that the very thinne st co mposite hinges did wo rk, but the re sulting   boom s had v ery low de ployed natural freque ncy an d long snap -over dam ping ti mes. 

The stiffest 1. 9-cm (3/4" ) diameter boom  with co mposite hinge s tested had a f requen cy of ab out 2.7 Hz.   The diamete r of th e boom was in creased above the nomi nal 3/4" in an effort to g et the de ploye d  freque ncy ab ove the req uired 5 Hz. It wa s determin ed that even  significant in creases in tu be diamete r  resulted in o nly small in creases in th e deploye d boom’ s natu ral freque ncy. In addition, the proje ct  77
objected to the in creased shadowing of the sola r arrays. In the  summ er of 2002 th e design effort wa s  concentrated on the metalli c BeCu hi nge.  

Figure 7.  An alytical snap-thr u and sto w sequence for joint with   0.15-mm (0.0 06-in ) thick tape and 8.2 6-cm (3.25 -in) window 

Figure 8.  Co rrelation o f Analy sis to Experiment for the 72 -degree ca se  78
Figure 9.  Co rrelation o f Analy sis to Experiment for the 22 -degree corr elation  

Moment /Life test fixture 

Hinge testi ng falls in to two  categories: characterization and  qualification.  Sometimes, of  course, these distin ctions are   artificial. In o rder to determine the  deploye d frequen cy, the deploye d  stiffness in  the foldi ng plane and  norm al to it n eed to be me asured.  Since the  deployment force is the   buckle d restoring force, this force  need s to be mea sured as a  function  of angle of bu ckle. The  cycle life, how many bucklin g  operation s can the  hing e survive,  also ne eded to be known. A  Moment /Life  test fixture wa s built  to make th ese measurements.  The torq ue resi stance of the   harness was also me asured with  this rig. T he fixture was d esigned  to operate in thermal chambers so  that measurements co uld be taken  over a ra nge of environm ents.    Figure 10.  Test Fix ture 

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Figure 11.  Torque v s. Angle 

The ratio  of pea k sna p-back moment  is less than half of pea k sna p-thru mo ment. This is the dampi ng  action that m akes the deployment stabl e. 

Deplo yment Testing o f a Complete Boom . 

Once it is de ployed, the b oom is a ble to sup port itself under  gravity; however it need s a "g " neg ation system  while it  deploy s. Duri ng depl oyme nt, the collap sed hi nges do not have  enou gh latera l rigidity to su pport the b oom. A low-fri ction trolley  was develop ed to support the boom  payloa d mass du ring  deployme nt. The trolley consi sts sim ply of a spring-loaded  support fo r the mag netometer o n a trolley ba se. This trolley  moves along a flat pl ate of alumin um. A sim ulated spacecraft  structure hol ding the  boom is also mounted  to this plate.   Becau se of the boom' s deployment characteristics, the pl ate is  1.5 m x 1.8 m (5 ft x 6 ft). This trolley h as been designed to be  comp atible with the requi rements of the thermal -vacuum te st  cham bers. T he bo om de ploys fro m the outb oard segm ents  inward (e.g. the wrist joint locks in first, followe d by the elbo w  and then the shoulder joint ). Extensive testing of the full boom  has shown that the sho ulder joint snap s over (t he hin ge  collapses in the oppo site dire ction) on all depl oyments. Th e  numbe r of times that the  hinge snap s over i s a function of the depl oying  energy. During ambie nt,  horizontal te sting the boo m sna ps over an average of just on e time. The a ngle that the pla te is mou nted  with re spect to the ho rizont al allo ws gravity to simula te the rotatio nal "g" field. Du ring life testin g, when th e  centri petal a cceleration of the spin ning spacecraft is take n into acc ount the boo m snap s over an averag e  of seven time s at the lowe st energy level  and nine tim es at the high est ene rgy le vel.  Figure 12. Spe cial Test Equipment  

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Figure 13.  Trolle y on Tilt  table 

Vibration Te sts 

Another du mmy spacecraf t was built as a vibration fi xture adapte r to allow the st owed co nfiguratio n to be   tested. The v ibration survival of the stowe d boom t urns out to b e a real concern. The m agnetom eter  mount i s held securely by the pin puller, but not much  tension force i s available to  retain the boom  section s in the snu bbers. Consequently  during vibrat ion the flexible colla psed h inges allow consi derable  motion of the  boom segm ents. It is a rattling fit in th e snubbers. This in turn result s in the boom system   behavin g no n-line arly un der vibration . A finite el ement mod el exists for the boom in  its stowe d  config uratio n, but becau se of the non-li near n ature of the stowe d boom exte nsive testin g wa s need ed.  The sto wed boom “natura l frequen cy” is very depe ndent on the vibration envi ronment. Extensive sine  sweep  testin g has determined th at at l ow levels, l ess than  1G, the b oom resonate s around 50 Hz. A s  input levels i ncrease towa rds 8 G’s, the  boom “natu ral freque ncy” drop s to aro und 20 Hz. Naturally this  mean s qualifying the sto wed configu ration by test. S uch extensiv e testing, o bviously, is going to re sult i n  a few failures. 

In Figu res 14 and 15, note the Y response of the  center segment of the boom  moving t ransversely h as  the mo st no n-line ar characteri stics, fo llowe d by th e X re sponse moving o ut of the snu bbers, a s the  collapsed hin ges a re softest in these di rection s.  81
Figure 15.  Accelerome ter Resp onse s (X-axis an d Y-axis) 

Failures 

During an ETU spacecraft qualification t est, the  stowed boom was incl uded to gather informa tion ab out it s  capabilities. I n gen eral, the bo om b ehaved a s expecte d. Ho wever, the Ultem tang th at holds the   magneto meter mou nt to the sp acecraft bro ke during a 17 G RMS random vib ration ru n. The t ang, it turne d  out, had  been manufactured thin ner th an the  drawing called for. Material s investigatio n of the part  showed there were two fail ures. There was an initial  fatigue failu re at the jun ction betwe en the  top hat and  the tang a nd this led to a subsequ ent fatigue failu re of the tang itsel f. The tang was rema de to  the co rrect  dimen sions and the testin g pro ceeded. It was later decided to re place the Ultem tang with  a titanium  piece, whi ch was non -magnetic an d stro nger. 

During a subsequent 8.5G , 15-Hz si ne burst test  in the last axi s (durin g the  same time fram e) the  oute r  elbow tape fa iled. A materials inve stigati on of this hi nge sh owed th at it had failed due to low cycle fatigu e.  Furthe r stu dy of this probl em reveal ed that t he hinge sh oes that hold the hinge s to the boom were  causing a sh arp be nd in the hinge ju st before it atta ched to the bo om. The fatigue failure of the hinge h ad  occurre d right  where the sh arp be nd was identified. The sho es have been modifie d to cure this.  

The o uter el bow ta pe a gain failed during a 17 G, 14-Hz si ne b urst te st. Thi s failu re was due to an  oversi ght by the mechani cal team. Foll owing the ET U sine burst failure it wa s reali zed that  the natural   freque ncy of the sto wed boom was so close to the  sine burst fre quen cy that a burst te st on the stowe d  boom was out of the question. Durin g planni ng fo r this testin g, this detail was ove rlooked. Material s  83
investigatio n of the faile d hinge  showe d that it h ad failed as a result of a  single event o verload. No  changes to the desi gn were need ed to fix the problem . 

Figure 16.  Before a nd After Sine Bur st 

Production Problems 

The ET U 1.9-cm (3/4") T 300 boom  section s were f abricated by a local contractor,  and the initial te st  pieces worked fine. The flight units, ho wever, were made by an “improved ” techni que with the mand rel  contin uously wrapped with the lay-u p rather th an b y using  cut layers. The resultin g bo oms had resinstarved sections. Anoth er delay to ord er new materi als and do it ri ght. Since hin ge blad es were not goi ng  to be made from the materi al, it was ch anged to M55 J, higher mo dulus but le ss to ugh.  

Thermal Ma gnetic Consi deratio ns 

Now some thermal  results beg an comi ng in  and th e scienti st became concern ed with th ermal g radients  producing Pe ltier voltage s whi ch in tu rn wo uld p roduce ed dy currents and a resulting magneti c field  higher than the extremely l ow fields he was setting out to measure. All metallic pieces became suspect. In  particula r the titanium tang  with its low h eat cond uctivity gave ri se to big gradie nt. The tan g was change d  to polish ed al uminum a nd the su rface was coated with SiO. The hinge leave s were al so gol d plated to cu t  down o n the thermal  gradients across them. Th e composite  boom had already been  blanketed, a s had  the cabli ng. The de sire o n the part of  scie nce to apply therm al blankets to  the workin g hinge s wa s  resisted i n the favor of t heir working. The stainle ss screws were replaced  by cu stom-built titanium   shoulder screws. 

A ne w Launch Vehicle 

At this point the idea  of g etting a ri de as a second ary paylo ad on a l arge Expenda ble L aunch Vehi cle  faded as th e prim aries raised obje ctions. Th e option of a  ride to orbit as th e pri mary on a Peg asus  became a re al possibility. A new su pport stru cture was required  beca use the  honeycomb base of the  baseline depl oyer system wa s designed as an adapter to a n ELV Payload Attach Fitting . Most of the   loads are we ll under th ose envelopi ng Ariane, Delt a, and Atlas.  Becau se the ST5 missio n con sists of  three spacecraft, the dimensional limit ations of the P egasus fairi ng drov e the support st ructure to a new  cantileve red design. Thi s has l ed to concerns about the couplin g betwe en the Pegasus environments and   the prim ary mode s of the  support stru cture, th e ST5 spacecraft bus and the stowed b oom. The d riving  techni cal question i s the response of the b oom to the Pe gasus drop transient and  captive ca rry random  vibration. Aga in the non -linear a spect of the boom p recludes a di rect analytical a pproach. Orbi tal Scien ce  Corpo ration has been given a finite ele ment mod el of the payload with the b oom represented by a sprung   84
Latch-up Anomaly  Study  for Reflector Deploy ment of INSAT-2E Satellite 

G. Nage sh*, S. Ravindran* and N. C. Bh at*  

Abstract 

INSAT–2E Spacecraft (S/C) h as two d eployable antenna reflectors of which o ne of the refl ectors showed  a non-nomin al latchi ng behavior during the ground deployment te st with a couple of boun ce backs befo re  a final p ositive latch -up. Similar p henomenon  was not ob served duri ng the  test on  the ot her reflector.  Detaile d analysis reveal ed the inh erent limitation in t he de sign for the form er reflector. T his limitation was  confirmed by  subsequ ent tests. This p aper de scrib es in detail the analysi s carried out, the variou s test s  conducted an d the optional  desig n modif ication worke d out to corre ct the sam e. However, in view of the   capability of the pre sent design to latch po sitively , it was deci ded to fly this mecha nism without any  changes in INSAT-2E. The observed o n-orbit pe rform ance is al so reporte d here.       

Introduc tion 

INSAT–2E i s a multip urpose spa cecraft intende d for communi catio n and meteorologi cal purposes. It h as  two ante nna reflecto rs on e on th e west side and th e othe r on the east side henceforth called as West  reflec tor and Eas t reflec tor respec tively as  shown in Figure 1 . These reflectors a re in the stowed   config uratio n duri ng la unch a nd a re deploye d in orbit. As th e config uratio n of the two  refle ctors a re  different a nd due to th e constrai nts on the spa cecraft stru cture with re spect to t he reflecto r interfaces the   reflecto r deployment me chanism s are di fferent for bot h the refle ctors. Durin g the grou nd depl oyment tests  on the reflect ors, it was observed that the depl oyment of the Ea st reflector was nominal, whe reas the  West refle ctor after the first latch -up was fou nd to rebou nd an d the forward - reverse o scillatory motion   repeated a co uple of times before the fin al latch up of the reflecto r.  

     S/C  East ReflectorWest Refle ctorØ 1600Ø 200 0EV 

Figure 1.  INSAT –2E S/C with Both Reflec tors Deployed 

                                                            * Space craft Mech anisms Grou p, ISAC, Bangalo re, India 

Procee dings of the 37th Aeros pace Mec hanisms Sy mpos ium, Joh nson S pace Ce nter, May 19-2 1, 200 4  87
Anomaly  description 

The two reflectors we re received at an advan ced sta ge of the  INSAT-2E p rogram. Both the reflectors and  the deployme nt mecha nism were take n up dire ctly for assem bly and tests at S/C level.  

The East reflector after in tegration was succe ssfully deployed a nd the perfo rmance of all  mech anism  element s were nomi nal. Howeve r, the West refle ctor was fo und t o rebound a couple of tim es before the   final latch -up. A thorough  analytical study of the tw o hin ge a ssembli es was carried o ut in orde r to  unde rstand the above be havior. Thi s was followe d up by  tests on hinge a ssembly of both the reflecto rs.  The latching characteristics of both refl ectors we re studied by a l atching  test under simulated  zero ’g’. As  this pro blem occurre d late in the progra m, there wa s a need to verify the required functio ning even with  this an omaly by subj ectin g the refle ctor to the expe cted on – orbit disturban ces. A desi gn mo dification of  West refle ctor to corre ct the above an omaly was al so wo rked out as a ba ckup /  or for use in future S/C. 

Description of Reflector deplo yment mechanism 

The Ea st and West reflect ors in th e deployed configu ration are interfaced to  the spa cecraft through a pair  of hinge s each. A typical hinge a ssembl y of the East re flecto r and We st refle ctor is a s shown in Figure 2a  and 2b. Ea ch hin ge asse mbly consi sts of two  brackets - an Inboard bracket conne cted to S/C an d an  Outboa rd bracket conn ected to the refl ector. Th ese two bra ckets are pivoted by a mono ball sphe rical  beari ng mo unted on a shaft supp orted  by the inboa rd bracket. The dry lu bricated sp herical beari ng ta kes  care of misal ignment s caused du e to assembly or therm al environment s. A locking lin kage is provide d  with one e nd hinge d to the inboa rd bracket. The out-b oard b racket has a circula r cam se ctor over whi ch a  roller mounted on th e oth er end of the  locking li nkag e rolls during the depl oyme nt and l atches into th e slot  provide d in the se ctor at the end of de ployment. Two torsi on spri ngs m ounted  on the sh aft provide th e  necessa ry energy for deployment. Th e torsi on springs  on the l ocking lin kage shaft en sures positive   latchin g at the end of depl oyment. The locking lin kage also p rovides for min or adjustment of the openin g  angle of the reflecto r. A micro-swit ch mounted o n the inboa rd bracket co nfirm s the hinge l ockup at the  end of depl oyment.  

Figure 2 a.  Hinge assembly of Eas t reflector         Figur e 2b.  Hin ge assembly  of Wes t reflector          – INSAT -2E     – INSAT -2E  

88