ai-aerospace/ams_data_full_2000-2020 huggingface dataset:
drive ele ctronics for te sting. Table 4 summ arizes the types of t esting gene rally performe d on ste pper  motors and m echanisms a nd ho w these  shoul d follow from spe cification and a nalysis st rategie s. 

Table 4.  Bas ic Stepper / Mecha nism Testing for Correla tion with Specific ations an d Analy ses 

Specification Paramete r   Analysis Co rrelation      Motor Compo nent Te st   Mech anism T est 

Electri cal  Characterization     Estimate mot or Kt, R ph,  and L ph   Verify electri cal  parameters   (Usually only minimal  electri cal ve rification ) 

Output Re solution   Define Step Angle     Step Accuracy Test  if  appli cable     Output Re solution 

Unpowere d Holding  Torque   Estimate Mini mum  Detent an d Bearin g  Drag     Motor – mea sure dete nt  Bearin gs – m easure  drag ove r temperature   Measure unp owered  holdin g torqu e over  temperature 

Powe red Hol ding  Torque   Motor g ross o utput  torque le ss internal  losse s must p rovide  margi n over l oads;  simulate with margin on  losse s and lo ads     Measure net output  torque a nd int ernal  losse s; adju st analyse s  and sim ulation model s  for any loss variation s   Measure po wered  output over te mperature  to confirm P HT spec is  met (no ma rgin applied ) 

Specific T orque and  Rate Op eration (u sing  driver p ulse width,  current limit, etc.)     Optimum gear reduction  with motor rate and  torque, simul ate over  wors t cas es with flight  driver limits     Dyno test mo tor and  include ma rgin on  losse s over worst case  thermal con cerns; use  flight-like driv er   Test outp ut rate and  torque at worse case w/  flight driver. Can te st  above rate a nd torque  spec for reference. 

Dynami c Perf orma nce  (New stepper motor  characterization sh ould  go beyon d mech anism  requi rements)     Predi ct & analyze pull-i n  and pull -out torqu e  perfo rman ce and if  electroni c rate ramp i s  need ed for o peratio n 

   Espe cially for new  designs, perf orm pull -in  and pull -out torqu e  characterization;  evaluate rate ramp   Verify mechanism  perfo rman ce over  dynamic ran ges  requi red and all worst  case conditio ns 

The ma rgin approa ch used in Tabl e 4 is the sam e as u sed in Table 3, whi ch does not comp ound the  margi n facto rs.  Ho weve r, if the margin must be demonstrated d uring test bey ond that act ually needed t o  meet the spe cificatio n, the n the additio nal margin must be b uilt into the analy ses and motor testin g as  well. Some te st anom alies are the result of not followi ng the same margi n strate gy throug h from sp ec to   test.  

Lesso ns Lea rned in Esta blishing Ade quate Perfor mance Margi n 

Prope r sizing with th orough a nalysi s and te st will usually a ddress mo st st eppe r motor appli cations  concerns. However, the re are some problem scena rios that sh ould be id entified in a le ssons l earne d  format. The se are largely  from histori cal lesson s and sho uld be considered relevant whe n establi shing  127
and position al sta bility. Testing  at the  mechani sm level loo ks at torque  outp ut and  margin, rate and  resolution, and step stability. 

Inadeq uate o r inapprop riate uses of m argin eithe r by the torq ue or volta ge method i s sure to cause  perfo rman ce deficie ncies. It should al so be unde rstoo d that the inclusio n of such  margin facto rs does not  guarantee th at perfo rman ce will be satisfacto ry und er every condition, be cause other influ ential condition s  may exist. In a broad er sense, to cov er the full perspective of perfo rman ce margi n, it is necessa ry to   evaluate not just the  torqu e requi rements and asso ciated m argin factors but also the  dynami c stability of   the motor in its system. Ta ble 3 provide s guideli nes that the syste m or mech anism level desi gner can u se  to spe cify margin for space stepp er mo tors, so t hat whe n the de sign i s create d and an alyzed suffici ent  design ma rgin will be p resent to help a ssure actual torqu es and speed s are met over all th e wo rst case  conditions at test. It is important that t he spe cifications be cl ear as to wheth er the ma rgin applie s for  design an alysis purpo ses only or if fact ors sh all be ap plied at test a s well. 

Stepper mot ors have  been an impo rtant motor type fo r use in  space applications offerin g adva ntages  over b rushless and brushed moto rs in many ap plications. It re mains relevant that those  who design  stepp er m otor based me chani sms an d are i nvolve d in d evelop ing spe cifications fo r stepper m otors  unde rstand their a pplication con cerns.  This i s truer  with ste ppers than brushl ess an d brushed moto rs,  becau se steppers tend to chang e thei r individual pe rformance characteristics fro m one syste m to anothe r.  Establishing adeq uate p erforman ce margin i s therefore a critical process and should b e caref ully explore d  as ne w ap plications em erge for step per motors in sp ace flight.   130
and can be manually rota ted out of th e way after latch re lease to aid in rem oval of the SA3. This latch  provide s 1 degree of con straint. Latch  4, shown in t he lowe r left corn er of Fi gure 1, is a v ise-type lat ch  that has a n internal me chanism simila r to that of Latch 2. A 7/16” (11-mm ) hex  drive lo cated on the Lat ch  3 housing, coupled to a dri ve extens ion rod, transmit s torque to the input shaft. Semi-spheri cal pads on  this lat ch clamp down on attachme nt fittings on th e corne r of th e SA3 to p rovide 1  degree of co nstraint.  Latch 5, sh own at the top of Figure 1, con sists of a CSS mounted towe r with  a lockin g bolt and clevis,   which pick an attachm ent fitting on th e stowed sol ar array ma st. The uniqu e degree s of con straint  provide d by each latch al low the overall system to  accommo date misalign ments an d me chanical and   thermal di stortions. 

Latch  3 Latch1Latch5 Latch4 Latch2 Figure 1.  Latch Moun ting Locatio ns with SA3  

The pall et and Center S uppo rt Struct ure a re an integral asse mbly provid ed by Orbital Scien ces  Corpo ration a nd is de signated the Rigid Array Ca rrier (RAC). Figu re 2 sho ws an end view of the RAC a nd  SA3s in  the launch configuration  prior to shipm ent to Kenn edy Sp ace Cente r. The gold colo red structu re  in the center of the p hoto is the CSS, and the  SA3s are mounted to the left an d right side s and are  silver  colored with gold h andrails on th e edges. Th e SA3 mast s are not  shown in this photo be cause th ey we re  shipped separately to Kennedy Space Cente r and in stalled p rior to laun ch.     

Figure 2.  Launch Co nfiguration o f the SA3  132
Overall Sy stem Require ments 

• High load carrying capabilit y during launch (m aximum  of 20Gs supporting a 318-kg (700-lb) Solar  Array  Assembly).  • Very high pre load capa bility.  • Ability to accommodate large misali gnments due to on-orbit therm al gradi ents.  • Acce ssibility for Extravehicular Activity – All latches a re astro naut a ctuated.  • Compatible with exis ting EVA power tools .  • Operational t orque less th an 34 N-m (2 5 ft-lb) fo r on-orbit diseng agement an d engag ement (power  tool limits).  • Low turn co unt for on-o rbit disen gagement and en gagement (mi nimized crew workl oad).  • High repe atability in torque and turn count.  • Allow for rel atively blind re-i nstallatio n of the SA3 if unable to install onto HST (S A3 would b e returne d  to earth).  • Independent EVA override.  • Indepe ndent laun ch lo ck at  each lat ch drive. 

Latch Desig n Details 

Photos of the  actual flight latche s 1 thro ugh 5 in t he laun ch co nfiguration a re shown in Figures 3 thro ugh  7, resp ectivel y. Basic de sign feature s of the SA3/RAC latch es are detailed bel ow: 

 Drive Mechanism    • All latc hes are rotationally driven through 7/16”  (11-mm) EVA tool c ompatible hex nuts .    ¾ Latch 1 – Sp ring load ed directly driven A CME screw with alignment guide ri des on oppo sing  deep g roove ball bea rings.  ¾ Latch es 2 a nd 4 – Input sh aft drives 2:1 integral 9 0° bevel gear set whi ch outp uts to ACME  screw and lin ear sta ge.  ¾ Latch 3 – Directly driven A CME screw d rives cl ampi ng pad.  ¾ Latch 5  - EV A-rem ovable bolt with align ment guide o n end.  • All latch drive s incorp orate indep ende nt positive lo cking feature s for laun ch.  • Latch es 1, 2, and 4 drive trains in corporate matche d deep g roove ball bea ring p airs fo r maximum  thrust capa city and minimu m torque.  Nomin al Latch Preloa d  • Latch 1 = 158 8 kg (3 500 lb ); Latch 2 = 13 61 kg (300 0 lb); Latch 3 = 272 kg (60 0 lb);   • Latch 4 = 318  kg (70 0 lb); L atch 5 = 2 72 kg (6 00 lb ).  • Belleville wa shers incorporated into load  path to allow preloa d compliance and accurate settin g    Torque an d Turn s Required to Dise ngage Latch   • Latch 1: T < 31 N-m (23 ft -lb) (20 turns); Latch 2: T < 20 N-m (15 ft-lb) (12 turns);   • Latch 3: T < 20 N-m (15 ft -lb) (8 turn s) Latch 4: T < 20 N-m (15 ft -lb) (12 turns);   • Latch 5: T < 20 N-m (15 ft -lb) (8 turn s)  Dimen sions and Wei ght  • Latch 1: Dim ensions: 298 mm (11.75 ”)L X 146mm (5. 75”)W X 89m m (3.5”)H   M ass = 6.3 kg  • Latch 2: Dim ensions: 381 mm (15 ”)L X 254mm (10 ”)W X 197mm (7.75”)H       M ass = 17 kg           • Latch 3: Dim ensions: 432 mm (17 ”)L X 152mm (6”)W X 135mm (5. 3”)H           Mass = 6.3 kg  • Latch 4: Dim ensions: 635 mm (25 ”)L X 152mm (6”)W X 203mm (8 ”)H              Mass = 23 kg  • Latch 5: Dim ensions: 102 mm (4”)L X 1 02mm (4 ”)W X 191mm (7. 5”)H             Mass = 1.4 kg   Material s for Major Comp onents    • ACME Sc rew: CRES 17-4PH  • Barden 107H Ball B earings: CRES 4 40C  • Latch Housing: Al 7075-T7 3  

133
Figure 3.  Latch 1 (Array  Stowed) Figure 4.  Latch 2 (Array  Stowed)  Figure 5.  Latch 3 (Array  Stowed)  Figure 6.  Latch 4 (Array  Stowed)  Figure 7.  Latch 5 (M ast Sto wed)  Latch 5 consists of a to wer mount ed to the top of the CSS with a  clevis and captive bolt which pi ck up a fitting on the  Solar Arra y  Mast.  The mast fitting contai ns a floating spherical bearing  which accommo dates thermal an d mechanical misalignments.   The basic mechanism of Latch 4 is similar to that o f Latch 2.   The vise action of the latch c auses semi-spherical spring pads  to clamp do wn on two separat e fittings at the lower  left corner  of the Solar Ar ray.  A floati ng spa cer provides separation  between the  two Solar Arra y fitting s. Latch 3 consists of a sw ing a rm with integral directly driven  spring pad that clamps on a fitting at the uppe r left corner of  the Solar Arra y. Latch 2 incorpor ates a moving lin ear stage with a spring pad  that clamps down on a semi-c ylindrical fitting on the low er right  corner of th e Solar Arra y structu re.  The fitting is constrained b y  a “V” guide o n the latch housing. Latch 1 consists of an acti ve rectangular shaped h ousing  mounted to the  SA3 structure with internal 7/16” (11-mm)  hex d riven ACME screw .  A sphe rical fitting of the end of  this housing mates to a passive cone on the CSS that  contains aninternal swivelnut. 134
Environmental Test Progr am 

Enginee ring Units   Prior to  build of the flig ht units, E ngineerin g Test Units (ET Us) of all l atches  were built to qualify the  design for space-flight u se. Strength tests, qu alificat ion level a coustics test s, and thermal vacu um test s  were su ccessfully performe d on all latch es. 

Flight Units  

Random Vibration Test  An acce ptance level rando m vibration te st wa s performed on  each of the f light l atches. Te st fixtures were  designed to realistically simulate the l aunch configu ration. During this te sting , several de sign flaws in  Latch 3 were uncovere d.  1.) Excessive we ar an d gallin g occurre d between the se mi-sphe rical spring pad on the latch, an d the   fitting on the  Solar Array.  Both part s were fab ricated from Al uminum Alloy 7075 coated with a  Teflon impreg nated ha rdcoat. Changi ng the material o f the latch spring pad to un coate d phosphor  bronze solve d this pro blem.  2.) Wobbling of the drive screw and spring pad during lateral  vibratio n caused the scre w to “wal k out”  of the launch  lock ja ws an d unscrew. T his problem was solve d by tightening up cle arances to  eliminate wobble, and redesigning th e lock jaws to provide a more positive hold on th e drive  screw.  

Acoustic Te st  An acou stic test at the payload level wa s perfo rmed, and the latch es operate d nominally. 

Therm al Vacuum  Test  Each flight la tch wa s subjected to 8 thermal cycle s in a vacuu m at the worst-ca se predi cted missio n  thermal  environment plus margin (-40°C to  +60°C). The  latche s were o perated after so ak at ea ch  temperature  extreme, a nd during transitions. T orque versu s turns wa s measured during functional tests  using a digital  torque tra nsducer. Typica l data for Latch 2 is sh own in Figure 8.   

02468101214161820 0.00.30.50.81.04.08.0 12.016.0 Number of TurnsTorque (N-m) 

 Figure 8.  T ypical Therm al Vacuum Torque v s Turn Data  

During testin g of Latch 3 , a design d eficien cy wa s un covered.  Excessive wea r and g alling occu rred  betwe en th e laun ch lock jaws an d the cam on which the jaws sli de. T his problem was solve d by  specifying a much smoother surfa ce finish o n the cam, and treating it with  a dry lubri cant coatin g  (Dicronite). T his pro blem was present on flight unit, and not o n the ETU du e to the fa ct that t he ja w spring  force was increa sed afte r the pro blems encounte red durin g vibrati on testing. 

135
Lesso ns Lea rned Summa ry 

Throughout the devel opment of the SA3/RAC L atches, m any le sson s we re learn ed. Ran dom vibratio n  and the rmal vacuum te sts on the flight units un cove red several d esign deficie ncies that we re not  uncovere d during ETU te sting.  1.) Perform  as thorough a test program as possible on the En gineering T est Units prior to building  an testing flig ht hard ware  2.) Random Vibration test ma y uncover d esign an d wo rkman ship flaws u ndete cted by an acou stics  test.  3.) Hard coating s in high contact stress areas can b e problemati c. Softer no n-coated m aterial s  should mate with hard-co ated surfa ces. Softer par t should be ma de repla ceable in the event that  excessive wear occurs.  4.) If possi ble, a combi nation of desi gn approaches should be utilized to minimi ze friction and wear  on highly stre ssed sli ding p arts. The se include:  • Prope r materi al sele ction  • Smooth su rface finish  • Surface co atings (Tufram, Tiodize, Dicro nite, etc.)  • Additional Lu brication (B raycote 601, 60 2)  5.) Mech anisms shoul d avoid slop an d exce ssive pl ay wherever possible in  order to elim inate  unfore seen m otion duri ng vibration. 

Conclu sions  

The SA3/RA C Latch development a nd space flig ht qualificatio n has been fully  com pleted culminating i n  the su ccessful laun ch and co mpletion  of the Se rvicing Mission  3B mission.  The SA3/RAC Lat ches  operated fla wlessly du ring the missio n, and were instrume ntal in th e tran sport and in stallation  of the Sol ar  Array 3 Assemblies.    

Data obtaine d from  thermal vacuum,  stre ngth, vi bration, and stiffness te sts perfo rmed during the  developm ent process as well as de sign improvem ents along th e way we re instrumental i n the success of  the flight hardwa re. The e nvironm ental test progra m uncove red design fla ws prior to laun ch that we re  corre cted as descri bed in previou s sections. This is definite proof that a robu st test plan is essential i n the  developm ent of any mecha nism de signed for harsh e nvironm ents, wheth er for spac e-flight u se or merely  for use on th e gro und. To rque versu s turn data m easured at variou s tempe ratures du ring thermal vacu um  testing provid ed esse ntial data for on -orbit verificatio n of prope r latch ope ration.  This data was also used  for pro gramming the HST p ower tool whi ch can be set to shut-off at a pred etermi ned torq ue or turn co unt.  

136
Lessons Learned on Cry ogenic Rocket Engine’s   Gimbal Bearing Lubrication Selection 

Christian Neugebauer * and Manfred Falkner * 

Abstract 

The bearings for a gimbal of a cryogenic upper  stage rocket engine are highly loaded, exposed to  corrosive environment and have to cope with a wide temperature range down to cryogenic temperatures.  “Cronidur 30” steel was chosen for the bearing materi al. To satisfy the demand of low friction moment  under these conditions a tribological concept should be found. This paper describes the selection process  for the tribological concept suitable for this specific application that resulted in the selection of sputtered  lead as lubricant for the gimbal bearings. 

Introduction 

Austrian Aerospace has the privilege to develop the gimbal for the next generation upper stage engine of  the ARIANE 5 Launcher family, the VI NCI engine. The challenges for beari ngs of a gimbal for a cryogenic  upper stage rocket engine are:  • high loads due to the transmission of the thrust of the engine,  • high reliability requirements for both environment s: atmospheric conditi ons and space vacuum,  • wide temperature ranges down to cryogenic temperatures,  • considerable number of only small oscillatory cycles,  • low friction and wear requirements. 

Previous gimbals use common bearing materials simila r to the bearing materials used in today’s satellite  mechanisms such as SAE 52100 or AISI 440C steele. FAG (FAG Kugelfischer AG, Germany) recently  developed a novel bearing steel “Cronidur 30” which combines the advantage of high load capability with  the advantage of high corrosion resist ance, better than AISI 440C. Moreov er “Cronidur 30” demonstrates  excellent mechanical properties at cryogenic temperatures [1]. For this reason we selected “Cronidur 30”  as the material for the gimbal bearings. 

We expect “Cronidur 30” to increasingly become the substitute for today’s commonly used bearing steels  for space mechanisms. Currently little information is available and little experience exists concerning  lubrication of “Cronidur 30” under space environment and specifically the behaviour of solid lubricant  coated bearings made of Cronidur 30. 

Keeping in mind the needs of the rocket engine gim bal bearing, Austrian Aerospace had the task to  investigate possible lubricants; to make a trade betw een those candidates by aid of development tests;to  select a suitable lubricant; and to qualify the chosen tribological system for the use on the ARIANE 5  launcher. 

This paper details the development wo rk performed to date (December 2003). 

Description of the Mechanism 

The main functions of the gimbal are to fix the engine in its defined position, to transfer the thrust and all  inertial forces from the engine to the stage and to allow the gimbal operation.  The gimbal consists mainly of an engine-side bracket, the cardan cross and the stage-side bracket,  connected by two pairs of bearings.                                                    * Austrian Aerospace GmbH, Vienna, Austria  

Proceedings of the 37th Aerospace Mechanisms Symposium, Johnson Space Center, May 19-21, 2004  137
Needle roller types were selected due to the high load capability and the low friction moment under the  specified environmental conditions, combined with a small envelope. 

Figure 1.  Gimbal design  The bearings are kept as small as possible due to  the stringent mass and envel ope restrictions. This  leads to high Hertzian contact pressures between the needle rollers and the bearing rings. The Hertzian  pressure will be more than 4000 MPa. 

Demands on the Tribological Sy stem 

The main demand on the tribological system formed by  the bearing’s inner and outer rings, needle rollers  and the lubricant is to attain a low coefficient of friction during the whole lifetime of the bearings and after  exposition to environmental attacks. This low coeffi cient of friction shall be attained despite the high  Hertzian pressure and for operation in air as well as in vacuum.  The tribological system shall be comp liant to the following requirements:  • compatibility to corrosive media, LOx and LH 2, and the operating temperature range that includes  cryogenic temperatures.  • material compatibility to the bearing material Cronidur 30.  • no impact on the bearing precision  • the application process shall be suitable for the needle roller geometry selected. 

Selection of the Lubricant Candidates 

The decision either to operate without a lubricant or to use a lubricant was driven by the risk of cold  welding under vacuum environment using a full “Cronidur 30” design comprised of rollers as well as races  made of “Cronidur 30” without a lubricant. To avoid this risk the use of a ceramic material for the rollers  was discussed. However utilizing ceramic needle rollers  is not feasible for the gimbal bearings as  bending stresses in the needle rollers would possibly cause destruction of ceramic needle rollers. 

Fluid lubricants are not usable for the gimbal bearing lubrication at cryogenic te mperatures and heating of  the bearings is not possible. The operating temper ature range goes down far below -75°C, which is  known to be the lowest temperature limit for fluid lubricants. For these reasons a dry lubricant was  select ed.  138
Dry lubrication by sputtered silver was first discussi on due to the vast heritage of the use of silver as  lubricant in space mechanisms on conventional bear ing steels. However a systematic approach was  taken.  For a pre-selection, 27 dry lubricants were assessed in terms of friction coefficient, wear rate in air and  vacuum, applicability for high prec ision bearings, heritage, temperature range, corrosion and electrogalvanic aspects and development ri sk for the application process.  Without going into more details concerning this pre-se lection, the remaining promising candidates of this  pre-selection were:  • Sputtered MoS 2 (Molybdenum Disulphide)  • Sputtered Ag (Silver)  • Sputtered Pb (Lead)  • Air-blasted WS 2 (Tungsten Disulphide).  Trial depositions with these candidates were perfo rmed on representative rollers of “Cronidur 30”. 

Performance Testing of the Pre-Selected Lubricant Candidates 

No information could be found in the literature regar ding the behavior of these candidates together with  “Cronidur 30” under corrosive environment. Therefore env ironmental tests were carried out as a first step.  To investigate the behavior of the coatings applied on the “Cronidur 30” base material under corrosive  environment, a salt spray test was performed as an a ccelerated test simulating environmental conditions  that may occur during storage, integration, transport and test of the gimbal it self and of the gimbal  integrated on the engine. 

A transparent test rig was manufactured that held the sample rollers on both ends. The samples were  exposed to a salt fog for two cycles comprising of  24 h with and 24 h without salt fog (pH=6.85±0.35)  each at 35°C±1°C. 

 MoS 2 Ag            Pb WS 2 Uncoated                  Figure 2.  Salt Fog Test Results  In Figure 2, the results of the environmental tests are shown. The pict ures show magnifications of two  roller surfaces of each candidate of un-tested rollers (left roller surface in each picture) and tested rollers  after exposure to the environmental test (right roller surface in each picture):  139