ai-aerospace/ams_data_full_2000-2020 huggingface dataset:
Click-bon ds (#10-32 stud s) were to be adhe sively at tache d to the  optical bench. The compl iance in the   mounting pl ate was sized to maintain the minimum resonant freq uency requi rement and al so to ensure   that thermall y induce d stresse s would remai n well below the all owable st ress calcul ated for the Cli ckbond  to com posite  structural adhe sive bond. The  titanium  did prese nt a th ermal p roblem since maxim um  thermal  gradi ent re quirements had been e stablished for the BA M due  to th e close p roximity of the  mech anism to the lase r assemblie s. The actu ators were mounte d to Aluminum 6061 -T6 m otor hou sings  that we re thermally coupl ed directly to t he al uminum  cover. The  housings and the cover act as radi ators   to the interio r of the GLA S instrum ent and re prese nt the prima ry transfe r path to shed h eat from the  actuato rs. The motor ho usings also in corpo rated additional m ass to increa se the the rmal ti me con stant  and minimi ze the maximum temperatu re experie nced unde r ope rating co nditio ns. 

Drive Assem bly Descri ption   A cro ss-secti on and p hotograph of the BAM prism d rive assembly are sho wn in Figure 6. It consi sts of the  followin g com pone nts: step per moto r actuator an d mo unt, GSE step verification sensor, beari ng hou sing,   angul ar cont act ball beari ngs and pr eload spring, spur gear, shaft, ri sley pri sm assembly  and prel oad  spring, an d gear cover. T he ri sley pri sm is prelo aded within th e rotating shaft, which i s driven by the   actuator thru a spur gear attached to the outboar d end of the shaft. The sha ft is supported by angular  conta ct ball b earin gs which are outbo ard of the risley  prism outer di ameter a nd are preloa ded i n a back to  back co nfigu ration into the drive assem bly housin g via a custom designed wa ve sprin g. Labyrinth se als  were utilize d on either side of the bearing pair a nd betwe en the spur gear and gear cove r to minimize  potential cont amination. Photograph s of selected co mpone nts of the BAM du ring the asse mbly of the   mech anism a re shown in F igure-7     - Labyrinth S ealsBearing PairDrive Pin Back PrismPinion to Spur  Gear Engagem entTherm al Connection To Cover Actuator and Mount 

 Figure 6.  Driv e Assembly  Cross -section and Photograph s 

Drive Train   The d rive trai n con sisted of a ste pper motor actuator and sp ur gear. The  actuators we re obtaine d from  CDA Interco rp. They were 30° step per motors wi th a two stage , 50:1 planetary gear reduction. They   incorporated a custom  mounting flan ge and a n integral output pinion. The  spur gears we re obtained from  PIC and where modified t o allow alignment and attachm ent to the drive assembly shaft. The gear also  acted to com press the bearing prel oad spring when bolted to the s haft. The 303 SS gear provided the  addition al 26 4:38 ge ar reduction. Com bining the risley rate of 3 5 microra dians/degree of counter-rotatio n,  51
Bearing  (OD of Bear ings  BAM Structure and  Assem bly Mounting Plate  Com pliant   Shaft Inserte d into  Bore w/ Preload W ave Modified Spur Drive Assembl y  GSE Step  Sensor Actuator in  Motor Bearing Preload  Assem bly  Pinion to Spur  Engagement Prelim inary BAM Assem bly  (No Cover or Mountin g Housin g w/ One  Assem bly  Figure 7.  BAM As sembl y  53
BAME Desig n and Op eration  Given the lim itations in  time, part s, and  comm and capability, the BAME desig n had to be as simpl e and  reliabl e as possible. A latching, me chanical relay (part# JL-D2A-1 08) was included to completely isolate   the BAM fro m the sp acecraft bus when not being operated. Using one HLDC t o close and one HLDC to  open  the l atching relay, th e BAME was co nnected or isolated  from the spare laser heater power li ne  providin g a p eak power capability of 2 8 watts. Oth er HLDCs were used to d rive optically couple d solid  state relays (part# HSSR-7111 ) in an H-bri dge co nfiguration  to driv e current in e ach phase of each moto r  (Figu re 8). Each separate HLDC was a uniqu e commutation st ate for one o f the two BAM motors. By  sequencing t he HL DCs, each motor could be command ed on e step at a t ime clo ckwise or counterclockwise with a 50-m sec pulse at a maximum rate of 10-Hz. Since  only one HL DC could b e activated at  a time, the  H-brid ge could n ot ina dvertently be  sho rted by com manding a n ode high an d low  simultan eously. The inputs diode s of the relay s only requi red b etwee n 5 mA and 20 mA e ach, so the   HLDC limit of  280 mA p rovided am ple m argin to  drive  two at a time . Each moto r coil was about 60 ohm s,  so using the +30-V supply, a maximum curre nt of 0.5 A was dra wn at any time. The soli d stat e relays a nd  spacecraft h eater po wer lines were rated to supply 1 A, so the BAME had 50% deratin g for those  comp onents. All of the BAME parts, in cluding a ssociated re sistors and ca pacitors, were spare flight parts   prima rily left over from th e TRIANA project. Even t he en closure was a flight qualified lefto ver from the  Shuttle small payload s program. T he bi ggest limitati ons we re no telemetry wa s available an d no memory  resided on th e BAME. Th e boresight a lignment i nformation h ad to be determi ned by th e GLAS pick-off  and ea ch m otor's comm utation state had to be tracked by the  spacecraft g round ops. Howeve r, the  proposed operation  of the  BAM co uld easily a ccomm odate th ese i ssues and provide a nd excelle nt mea ns  of fine tuning the GLAS boresig ht alignm ent at any point during the missi on.  Figure 8.  Simplified Schematic of H-Bridge Driver – One Moto r Phase  54
Verification and Test 

Although t he developm ent sched ule fo r the BAM was extrem ely tig ht, all ph ases of the  revie w/verificatio n  process were  completed. The design of the BAM was subjected to prelimi nary and critical d esign revie ws  and a po st environm ental  testing delivery review. A nalysis of all critical comp onent s wa s perfo rmed to   evaluate ma rgins fo r performance and strength. Th e BAM  was subjecte d to ba seline performance testin g  (beam devia tion and  torque m argin), protoflig ht vi bration te sting, thermal  vacuu m cycling and  perfo rman ce testing, EMI/EMC, magn etic leakage, and po st en vironme ntal perfo rman ce testing. Th e  BAM successfully passed all v erification tests. During the  testin g program the BAM exe cuted >80,0 00  motor ste ps which is a pproximately 1.6 times the expecte d on orb it lifetime.  

Vibration Te sting   The Bam an d the BAME were b oth su bjected to p rotoflight ran dom and si ne burst vibration levels at the  GSFC. Both  comp onents passe d the  tests. Post -test inspection of the BAM  did raise some concerns. A  chip o n the edge of on e of the risley  prism s was “discove red”. It was dete rmine d that the chi p was  probably present prio r to vibration. Fu rther discu ssion of this event is given in the  lesson s learned sectio n.  

Therm al Vacuum Te sting   The BAM a nd BAME we re teste d simultaneo usly i n the same chambe r using two separate cold/h eater  plates. The perform ance of the system  was verified over the operating a nd survival tempera ture ranges.  Figure 9 sho ws a schemat ic of the t hermal vacuu m test set-up. Th e units we re subjected to a minimum of  ten hot-to -cold ope rating cycle s and two hot-to-col d su rvival cy cles. Thi s test laste d approximately 2  weeks. Torque margins were measured for the worst -case condi tions utilizing the GSE  step counter  sensor. Beh avior of the  system with the  BAME hot a nd the  BAM cold and vise  versa was chara cterized.  The BAM  and BAME were in strum ented with  thermal couples on critical component s to characterize th e  thermal behavior of the u nits. A typical therm al profile  durin g contin uous operatio n of the BAM is sho wn in  Figure 10  along with  a photo of the BA M in the th ermal vacuum cham ber. In addition, o utgassing data  was coll ected by witness sample s and a  cold fi nger to verify the cleanline ss of the units. 

Chamber 281 With   Temperture Controlled a Shroud  BAME Te mperaure t Controlled Plate   BAME Te mperaure  t Controlled Plate   BAM Wedges  Chamber Wi ndow  BAM Wedge  Theodilite  Refere nce Mirror   TheodiliteTheodilite  Light  Paths Refere nce MirrorBAM/BAME   Computer/GSE   

BAME With    Ther mal Blanket   

 Figure 9.  Thermal Vacuu m Testing Setup 

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BAM  StructureMotor B Hou sing Time12:00 10:48 9:36 8:24 7:12 6:0040 35 30 2025De gre es  C  15 10 5 0Motor B Tempe rature Senso r @ Cold Oper ational (5  C)  

 Figure 10.  Charac terization of Therm al Behav ior During BAM/ BAME Th ermal Vacuum Testing  

Perform ance   The p erformance te st to chara cterize the be am steering capabili ty of the BAM wa s accomplish ed using  an auto-colli mator in a d ouble p ass configuration. A large flat mi rror was pl aced behi nd the BAM. The   autocollimato r was aligned to th e reference mi rror. As the risley prisms rotate  the beam  deviation was  observed by t he autocollim ator. Since the BAM i s only required t o correct for poss ible boresight alignment  errors o n a seasonal basis no dyna mic me asure ments where required. Th e sam e technique was also  utilized during the the rmal vacuu m testi ng ho weve r and ad ditional autocollimato r was used to  monitor t he  referen ce mirror through the thermal vacu um cha mber win dow. Perform ance tests we re co nducted after  assembly (to  obtain  a ba seline) du ring the enviro nmental te sting, an d post enviro nmental testin g. No  degradatio n in perfo rman ce wa s observ ed. Figure 11  shows a ph otograp h of the basi c test set-up an d the  actual  beam  steering capability as  compared to the theoretical curve. It dem onstrat es that the  performance of the BAM was withi n the measurement  capabilit ies of the set-up. 

Figure 11.  Performa nce Verification Setup and El evation Axis Results   (similar results for az imuth)  

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Conclu sions  

The BAM and BAME compone nts were succe ssfully dev elope d from dra wing boa rd co ncept to flight  qualified  unit s in approxi mately 4 m onths. T est results indi cate that all objectives were met and  all  requi rements verified. The  comp onents coul d have been utili zed had they be en required. The BAM is a   precise, rob ust mechani sm with high  torque ma rgins. Th e be am stee ring capability of the BAM is   excelle nt. It demon strated  a null stabi lity of ±2 ar csec, a repeatable beam stee ring rate of 0.62   arcse c/step, and a line ar range of ±30 0 arcsec. An y future mission with si milar requireme nts to GLAS  should seri ously co nsider incorporating a device such as thi s into the desi gn up front. Th e risley p air  concept provi des precise steeri ng capability while not imposi ng precis e alignment requirements of the  mech anism to the re st of the system.  The ri sk of addin g an a dditional m echani sm mu st be wei ghed  again st the compl exity and cost of maintainin g extremely tig ht alignm ent and stability  requirements  betwe en the tran smit and receive p aths of a laser alti meter in strum ent.    Ackn owledgements 

Scien ce Sup port:   Jim Abshire  Optical:      Luis Ramo s-Izqui erdo  Designers:   Bryan Gram mer, David Pf ennin g  Fabri cation/Assembly:   Sid John son, Paul Han ey, Cindy Rosm arynowski  Bearin g/Co ntamination:   Mark McClen don  Analysis:    Jeff Bologne se, Anthony Melak 

Referen ces 

1. IceSAT ’s Laser Mea surements of Polar Ice, Atm osphere, O cean, and La nd, H.J. Zwally, B. Schutz, W.  Abdalati, J. Abshi re, C. Be ntley, A. Brun ner, J, Bufton , J. Dezio, D. Hancock, D. Harding, T. Herri ng, B.  Minste r, K. Quinn, S. Palm, J. Spinhirne,  and R. Thom as, Jo urnal of Geodynami cs 34, 3-4, 200 2 

2. Preci sion Optical Stru ctures & Mech anism s for th e Geo science Lase r Altim etry System  & the Solar  Ozone Lim b Soundin g Experim ent II , J. Budinoff, SPIE 5176-24, Society of  Photo-Opti cal Inst rumentation  Enginee rs (S PIE) August 2003 

3. IceSAT – Ice, Clo ud an d land Elevation Satellite , FS-200 2-9-D47-GSF C 

4. Introdu ction to Infrared System  Design, W. Wolfe,  SPIE Optical Enginee ring Press, Volum e TT24  ©1996 p ages 85-86 

5. GLAS 106 4 nm  Boresig ht Alignm ent Analysis , J.Abshire, E.Ketc hum, H. Riris , X. Sun, B.  Mins ter,  NASA GSFC Laser Remot e Sensing Branc h, 10/04/2002 

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+Z+X +Y LV Interf aceStanchionsXY Mech anism XYZ Mechanism Honeycomb BaseDeployerSpringPinPullerYZ Mec hanismFlyaway Spacecraf t Bus

 Figure 1.  Initial Design of the ST5 Deploy er Structure show ing Spin  Direction and Semi-Kinematic Restraint 

 Figure 2.  Initial Design of ST5 YZ Mechanism  60
Thrust  Bearin gPushe r Pusher  Housi ng  Force  Adjust ing  Rod  TufLite® BushingsSprin g

Figure 3.  Cutaw ay View  of Initial Design of ST5 Pusher Assembly  

Butterfly  Key  Sprin g    XYZ  Stanchion  Interface  TufLite® Bushings  

Cup for  Z Cone for Z  Restrain tKey to Drive  XYZ Shaft  Rotating  XYZ Shaf t  Slot driven by   Key in XYZ  Hard point Thrust   Bearin g  TufLite® Bushing 

Figure 4.  Initial Design of XYZ Hardpoint (right) and Cutaw ay View  of XYZ Mechanism (left)  61
Prior to every test, low-vapor-pressure lubricant Bray cote 601 was applied to the pin. Prior to each hot or  cold test, the test fixture was held for one hour at a soak temperature of 70 oC or -70 oC, res pectively, to  simulate the possibility that the spacecraft would be at a more extreme environment before moderating to  one of the allowable temperatures for deployment. Duri ng each test, the Pinpuller was fired with a custom  circuit that limited pulse voltage, current, and durati on to prevent damage to the shape-memory-alloy wire  inside the Pinpuller. After each test, the chamber had to be brought back to ambient temperature in order  to res et the Pinpuller. 

A typical actuation time versus temperature profile for the Pinpuller at the two current levels appears in  Figure 5. In all 44 tests at the various temperatur e and current levels, the Pinpuller fully retracted under  goal shear loading of 1000 N (225 lbf). This demons trated a Pinpuller release margin of 1.5.   0.010.020.030.040.050.060.070.080.090.0 -40 -20 0 20 406 08 0 Temperature (C)Actuation Ti me (msec)6.8 - 7.4 Amp Draw 4.6 - 5.0 Amp Draw   

Figure 5.  Actuation Time Versus Temperature for Tw o Pinpuller Testing Current Cases 

Structural Analy sis 

The ST5 deployer was structurally analyzed using a NASTRAN finite element analysis (FEA) model to  predict its natural frequencies and structural response to the worst case LV loads [3]. The minimum  predicted natural frequency for the final deployer design was 51.1 Hz, which exceeds the minimum  natural frequency requirement of 35 Hz. A thorough stress analysis of the deployer  structure was also  conducted using classical analytical techniques. The results appear in Table 1. The minimum fastener  margin of safety was +0.52 including 50% preload and occurred in the XYZ mechanism. The minimum  component margin of safety was +0.55 and occurr ed in the bushings in the XYZ and XY mechanism.  Based on this analysis, the deployer structure was deem ed in compliance with the requirements set out in  the ST5 Mechanical System Specification document [4].   

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Table 1.  ST5 Deploy er Structure Margin of Safety  Table 

Deploy ment Analy sis 

Deployment analysis was conducted using several different methods to determine how much the  deployer spring should be compressed to deploy the spac ecraft at the desired spin rate. The first analysis  method used was the energy balance method. To co mpare with testing, analysis was conducted  assuming that the spacecraft was deployed in a gr avity field. The energy balance equation for the  deployment is:  (1)  where  kpusher = the spring constant of the pusher spring   x 1 = the amount the spring is compressed   m sc = the mass of the spacecraft  Ig.sc = the mass moment of inertia of the spacecraft in the direction of spin   v release = the speed of the spacecraft at the instant of release   ωrelease = the spin rate of the spacec raft at the instant of release  hcg = the initial height of the spacecraft center of gravity (CG) above its height at release  g1= the acceleration of gravity  

From dynamics, it can be shown that:       vrelease = L hinge.cg  . ωrelease               (2) 

where L hinge.cg  is the distance from the hinge line to the s pacecraft CG. Substituting (2) into (1) gives: 

ωrelease1 2kpusher⋅ x12⋅ mscg1⋅hcg⋅ + 1 2mscLhinge.cg()2⋅ Ig.sc+⋅:=   (3)  

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By solving (3), the theoretical spin rate for t he spacecraft can be determined. This equation can also  easily be solved for x 1 to determine how far the spring should be compressed to produce a certain spin  rate. In addition to energy methods, a Newton force balance approach was used to determine a  differential equation for the angular acceleration. Nume rically solving this equation and then integrating to  find angular velocity gave a release spin rate that  was within 0.5% of the energy balance result. 

Figure 6.  ETU Deploy er with ETU Spacecraft Du ring Initial Qualification Vibration Testing 

Two 3-D simulation software packages were also us ed to visualize the deployment and further confirm  the computational results. ADAMS,  the industry-standard 3-D simulation software, was used to develop a  spacecraft and mechanisms model. The resulting depl oyed spin rate again agreed with both the energy  and differential equation methods to within 0.5% . Working Model Software was also used and again  results were consistent with previous  analytical methods to within 0.5%.  

Initial Qualification Vibration Testing  

A flight-like engineering test unit (ETU) spacecraft and ETU deployer were subjected to qualification  vibration testing in March 2003 as shown in Figure 6.  In all three axes it received 17.1-Grms random  vibration, 20-G sine burst, 250-G shock, 1-G sine sweep, and low level sine signature testing. The  requirements for the structure to be successf ully qualified were that in each axis:  • The pre- and post-vibration sine signatur es must be substantially identical  • The FEA model must be correlated  • The natural frequency must be above the 35 Hz requirement  • The Pinpuller must fire and deploy the spacecraft successfully  • The structure must be pass visual inspection and show no significant wear 

During this initial qualification vibration test the first four requirements were satisfied. The testing  produced substantial wear at three of the mechanism  interfaces, requiring redesign in those areas.  

Discussion of Wear and Mechanism Redesign 

After vibration, significant wear, shown in Figure 7, was seen in the PTFE-wrapped journal bushings in  both the XY and XYZ hardpoints. This wear was likely caused by Z-axis motion of the two stanchions,  which caused the XY and XYZ shafts to slide in and out of the XY and XYZ hardpoints. Because the keydriven slot was a through feature with fairly sharp edges in the end of each shaft, this motion ultimately  chewed up the bushings.  

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The shafts were redesigned to incorporate a pocket by  closing off the ends, which allowed full material  contact between shaft and bushing, espec ially in the Y-axis. The tang was shortened to fit in the shaft  pocket with 0.013-cm (0.005-in) clearance on all sides to prevent load transfer through it. In addition, the  Tuf-Lite bushing was replaced by custom-machined Vespel. A brace was added to the XY stanchion to  reduce the Z translation of the shafts.  

Significant wear also occurred at end of the alum inum pusher. There was major dimpling of the pusher,  shown in Figure 8, where it met the stainless st eel roller ball on the deployer. The pusher was redesigned  by cutting it back 0.508 cm (0.2 in) and installing a 17- 7-stainless-steel tip to contact the roller ball. 

Finally, there was significant wear on the cup and cone interface of the XYZ mechanism. Brinelling  occurred where the stainless-steel-410 cone rubbed on the aluminum-7075 cup and was likely due to the  previously discussed Z translation of the spacecra ft. This interface was redesigned as an aluminum-7075  tongue in a Vespel bushing as shown in Figure 9. The Ve spel slot aligns itself in the housing and contains  a diagonal cut-back that assures that  the tongue is clear of the Ves pel slot when the spacecraft has  deployed 6 deg, before the pusher stroke ends at 7.25 deg.  

The deployer was analyzed structura lly again. All of these design c hanges had positive margins despite  using conservative assumptions. It was also shown by  test (see next section) that none of these changes  would affect the deployment of the spacecraft.   

Figure 7.  Wear of XY and XYZ Journal Bushings (right) and Tang in Pocket Redesign (left)          Wear

Figure 8.  Brinelling of the Cup on the XYZ Hardpoint (left) and Dimpling of Pusher (right) 

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4 Front Ambient 6.6  5 Front Hot 6.9  6 Front Cold 6.2  Table 4.  Natural Frequency  Response Results Ob tained During Requalification Vibration Testing 

Axis Pre-Sweep Low Random High Random Post-SweepAnalysis   (Hz) (Hz) (Hz) (Hz) (Hz)         X 49 44 50 49 55  Y 125 100 124 114 68  Z 41 42 50 41 52 

Figure 12.  Comparison of X Force Sum in X-Axis  Vibration Pre (red) and Post (blue) Sine  Signatures During the Requalification Vibration Test  

Requalification Vibration Testing 

The final step that remained to qualify the deployer was to show that the redesigned mechanisms would  not wear during qualification vibration levels. Just as  in the initial qualification vibration test, the ETU  deployer and ETU spacecraft were subjected to 17. 1-Grms random vibration, 20-G sine burst, 250-G  shock, 1-G sine sweep, and low level sine signature testi ng in all three axes. After each axis of testing,  the Pinpuller fired and the S/C was successfully deployed.  

In each axis, initial sine signature sweeps and those r un post testing remained very similar to each other.  These frequencies are shown in Table 4 and a repr esentative frequency plot comparing pre- and postsignatures in the X-axis is shown in Figure 12. The ETU successfully with stood structural loads and  demonstrated substantially identical pre- and post-  sine signatures and its fundamental frequencies of 49  Hz in the X-axis, 114 Hz in the Y-axis, and 41 Hz  in the Z-axis all exceeded the minimum frequency  requirement of 35 Hz. After each axis of testing, the Pinpuller fired and the S/C was successfully  deployed.  

It should be noted that due to clearances in the me chanisms, the response of the ETU spacecraft and  deployer is very nonlinear. In all the three axes , the fundamental frequencies differed significantly  between the high (full) level random and the low (-18 dB ) level random runs. The requirement to correlate  71