ai-aerospace/ams_data_full_2000-2020 huggingface dataset:
MDD Test Description   The fundamental requirement for the Mast Deployment Drive is to mo ve the mast from the horizontal  position to vertical on Mars. A safety factor of two was imposed upon this requirement to assure adequate  margin over uncertainties. In addition, it was desir ed to be able to deploy the mast on Earth for testing  without counter-balances or other support structure.  And if deployment on Earth was to be feasible, a  factor of 20% was determined to be needed.  

Designing a variable mass to simulate these differ ent conditions on the Pancam Mast Assembly for  testing was considered for a while but found to be very  difficult to implement. Because the MDD primarily  reacts torque, and torque is simply force times dist ance, it was decided to keep the force the same and  change the distance. 

The design of the variable load test  fixture is shown in Figure 10. The  main structure consists of a thinwalled Aluminum tube with flanges  at each end. The threaded rod that  runs the full length of the fixture is  supported at each end by Vespel  bushings. The sliding mass has a  threaded feature in the middle to  engage the rod and a slot on the  outside to prevent it from rotating.  Rotation of the rod causes the  weight to translate, thus changing  the torque on the MDD. The load  test fixture was calibrated using a  torque wrench to verify the torque at  the various weight positions. Sliding We ightSupport Tube Hex F eatureThreaded RodMDD Guide RailMounting Flange

Figure 10.  MDD Load Test Fixture 

For the thermal functional test, the MDD was first bol ted to a surrogate Rover deck. The load test fixture  was then bolted to the MDD. The MDD was then depl oyed and stowed several times to obtain ambient  deployment data. The assembly was then placed in side a large thermal chamber. A socket wrench was  inserted through a small opening opposite the drive nut to rotate the shaft as needed. The weight and  torque change due to the rod rotation is shown in the following table: 

Gravity  Condition Weight of Mast  (Deployed section) Torque on MDD  0.46 M (18.1 in.) Turns  1 X Mars 33 N (7.4 lb) 15 N-m (134 in-lb) 0  2 X Mars 66 N (14.8 lb) 30 N-m (268 in-lb) 163  1 X Earth 87 N (19.6 lb) 40 N-m (355 in-lb) 271  1.2 X Earth 104 N (23.5 lb) 48 N-m (426 in-lb) 357 

Mast Deployment Summary   • Cable is well managed, not subject to extreme bending  • Big central hole for light path  • Meets Envelope and Interface Requirements  • Time to Deploy at -70 C is 2.5 Minutes  • Hard Stops can be adjusted to .012 mm.  • Position against stop repeatable to 0.0025 mm  • Tested under 1 X Mars, 2 X Mars, 1 X Earth, 1.2 X Earth  • Tested at warm, ambient and cold conditions  • Works well with launch restraint mechanisms  269
Azimuth Driv e 

The third mechanism is the Azimuth Drive, which rota tes the upper part of the PMA. This drive allows the  Rover to look left and right. The optical path for the Mini-TES instrument must pass down the middle of  the drive. The principle requirements are as follows: 

Property  Requirement  Angular range +/- 180 degree  Angular accuracy +/- 1 mrad.  Angular speed 0.5 revolutions per minute  Optical clearance 100 mm bore  Electrical interface Upper flange: 4 x 37 pin connectors  Lower flange: 4 x 37 pin connectors  Cable management +/- 180 degree  Mechanical Interface Upper flange to Mast  Lower flange to MDD  The Azimuth drive was originally designed by American Technology Consortium (ATC) for the APEX  program. ATC was then acquired by Starsys Resear ch Corp. who redesigned the drive to reduce the  outer diameter and add a fourth cable path. 

The Azimuth drive is mounted to the top of the MDD to form the Lower  Drive Assembly (Figure 11). By  mounting these two drives at the bas e of the PMA, a significant porti on of the mass is located near the  RED. This has the effect of  lowering the overall mass  center of the assembly. 

These two drives also take  advantage of commonality in  the areas of rotating shaft  seals, large diameter bearings,  lubrication, motor shielding and  motor electrical connections.  Internal and external surface  coatings are also the same. 

Note that the two motors are  located in the same proximity.  For both the MDD and the  Azimuth drive, there was a  choice to mount the motor on  the fixed part of the mechanism  or the moving part. By  mounting the motor on the  moving part of the MDD and  the fixed part of the Azimuth  Drive, the two motors could be  mounted adjacent to each  other. This, in turn, facilitated  the routing for the motor  electrical cables.   Mast Interface  Azimuth  Drive  Mast Deployment Drive  RED  Interface 

Figure 11.  Low er Driv e Assem bly  270
rotate. The issue was addressed by implementing dy namic braking in which the motor windings are  shorted so that the motor acts as a brake. Dynamic braking is turned off during operation.  

The bearings on the output shaft of the gear-motor we re sized to support the mirror and support structure  weight so that no additional support was required. 

System level testing of pointing to meet the 1 mrad (0.057 degree) requirement was carried out at a  number of specific pointing angles. The largest pointi ng error in tests prior to vibration was 0.035 degree,  and after vibration of 0.046 degree. 

Subsequent testing revealed a  potential additional line of sight  error of 0.03 degree when the  mast assembly was subjected to  lateral loads, however this error  was found to be a single time  error in the azimuth bearing  immediately after the load was  applied. Once the azimuth  bearing was run through at least  90 degrees of motion, the  alignment returned to its  previous state. 

In order to calibrate the  instrument after arriv al at Mars  and periodically during the mission, t he instrument is required to observe a calibration target of known  temperature and emissivity mounted in front of all mirrors. To accomplish this, the elevation mirror was  capable of rotating completely around to point at the inside of the back of the mast where the calibration  target was mounted in a recessed pocket. This target was constructed to have a high emissivity surface  by machining small groves across its surface and then applying high emissivity black paint. The  temperature of the  calibration target was  measured with redundant  sensors potted in a blind  hole in the back of the  target. When the elevation  mirror is rotated completely   around to view the target,  the Mirror Shroud  effectively seals the front  opening to exclude ambient  light. The shroud, shown in  Figure 15, is a thin graphite  structure mounted so that it  rotates with the mirror. It  prevents stray light from  entering the periscope and  allows the instrument to be  closed up somewhat when  not in use to limit  contamination. The contamination c oncern is significant in the Marti an environment. Martian dust is very  fine and is easily carried up to the mirror level by t he Martian wind. The air bags that surround the rover  during landing are separate and do not form a seal ed envelope around the rover. During landing, dust  that is kicked up by the impact and subsequent 5 minut es of bouncing and rolling across the surface can  easily pass between the air bags and affect the rover.   Fixed Fold M irrorInternal Calibration Target Mini-TES   Elevation  Mirror FFigure 14.  Mini-TES Elevation Mechanism   Internal Calibration Target (6063 Aluminum )MiniTES   Mirror Shroud  Figure 15.  Calibration Target  

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In order to facilitate the compressed test schedule, four separate test racks were built for the program.  This allowed multiple mechanisms to be under test at one time. The test systems were configured to  allow the use with any on the PMA mechanisms. These test systems had integrated commercial motor  controllers, power supplies and a PC loaded with a specialized LabView program. The LabView program  allowed many parameters to be controlled and also  record all the mechanism/test parameters.  Specialized thermal chambers were built that operat ed on liquid nitrogen to eliminate scheduling issues  with the company's thermal chambers. Electronic c ontrollers were used to maintain the specified  temperatures. The mechanisms were loaded using we ights and pulleys as shown in Figure 17. This  allowed incremental loading of the mechani sm to characterize the performance. 

Once the mechanism level testing was completed  and the PMA was integrated, the assembly was  optically aligned and ambient functional tests  performed. The PMA was then exposed to system  required vibration levels. Testing of the pyro pinpullers were also performed while on the shaker  table. This facilitated instrumentation of the unit  with small accelerometers to measure the shock  levels. 

In the stowed configuration, the PMA was  subjected to 3-axis sine sweeps, random vibration  and a quasi-static sine pulse. In addition, All, the  units were subjected to a pyro-shock test from the  pin-pullers. In the deployed configuration, the PMA  was subjected to 2-axis sin sweeps and a quasistatic sine pulse. 

Next, the PMA was mounted on a thermal fixture  and installed in a thermal vacuum chamber. A 10-5 torr vacuum was pulled on the chamber and a 3-hour  bake-out at 110°C was performed to remove any moisture in the hardware. The chamber was back-filled  with dry nitrogen to a pressure of 10 torr before proceedi ng with the remainder of the thermal tests. This  pressure of nitrogen is thermally equivalent to the CO 2 atmosphere on Mars. The thermal testing profile  exposed the hardware to the non-operational temperat ure limits before deploying the mast at the low  operational temperature extreme.    Figure 17.  Ty pical Test setup 

Functional testing was performed at several differ ent temperature plateaus as the temperature was  cycled between the operating limits. At the end of the thermal tests,  a 50-hour planetary protection bakeout was performed. Motor revolutions were re corded anytime the mechanisms were operated. 

One of the challenges in the test program was the reduc tion of the test data. In most cases, the data was  collected at a 10-Hz rate with tests lasting 10 to 20 hours for mechanism level testing, so the amount of  data collected was substantial. Many different MATLAB scripts were developed to automate the data  reduction.  These scripts enabled engineers to quickly and a ccurately locate critical parameters within the  large data files as well as to facilitate statistical analyses. 

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Lessons Learned 

Communication and cooperation is critical between t he contractor and the cu stomer. Many unexpected  issues were encountered over the duration of the program, which required both sides to understand the  issues and impacts to the hardware. Both Ba ll and JPL have a deep commitment to get the job done  right, so issues were resolved relatively qui ckly in order to keep the program on schedule. 

Requirements creep in a program is costly, both in  dollars and schedule. Many small changes add up to a  large impact. 

Define all interfaces early. Undefined requirements always need to be specified at some time and this can  lead to changes late in the program. 

Proven technologies sometimes aren't! Technology from  a prior program cannot be assumed to transfer  directly over to another program. The PMA program was initially envisioned as a rebuild of the Athena  program hardware with some changes to meet the new requirements. By the end of  the program, virtually  every drawing was changed. 

Testing is sometimes as complicated as the des ign and assembly. Good definition of the test  requirements are important to delivering a qualified product in a timely manner. The PMA test program  increased in scope over the length of the program to include unforeseen tasks. The finished assembly  performed very well and was characterized to the level that JPL required. 

Conclusion 

The three qualified Pancam Mast Assemblies were  delivered to JPL in September and October 2002.  JPL then integrated the PMA on the rover and perform ed extensive environmental testing before delivery  to Cape Canaveral in Florida. The Rovers were  integrated on a Delta rocket and launched on June 16th,  2003 and June 30th, 2003.  The rovers are scheduled to land on Mars in January 2004.  

Recognition 

The authors would like to thank all the many te chnicians, machinists, designers, engineers and managers  at Ball for their commitment to developing these state of  the art mechanisms. Also all the great folks at  Starsys for the gear-motor actuator s and Azimuth drive. And especially the entire Rover Team at JPL. 

This work was perform ed under a contract with Jet Propulsion Laboratory, Ca lifornia Institute of  Technology, for the National Aeronautics and Space Administration. Reference herein to any specific  commercial product, process or service by trade nam e, tradem ark, m anufacturer, or otherwise does not  constitute or im ply its endorsem ent by the United States Governm ent or the Jet Propulsion Laboratory. 

References 

1.  Athena Mars Exploration Rover home page, Cornell University site  http://athena.cornell.edu/  2.  Mars Exploration Rover Mission home page, Jet Propulsion Laboratory site  http://Marsrovers.jpl.nasa.gov/home/index.html  3.  Mars Exploration Rovers, Ball Aerospace site  http://www.ball.com/aerospace/mer.html 

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made to obtain information on the cryst allite si ze of the magneti cally ordered F e (III) compound(s) in the  dust. 

Figure 2.  Ro ck Abr asion Tool -- Butte rfly Switch S ection Vie w 

The surfa ce produced by  the grindin g operatio n will serve the science intere sts well. Typically, the  surfa ced is quite smooth  and in  some  case s it is as smooth  as a granite  count er top. This clean, smooth   surfa ce should enabl e the microsco pic imager to view grain struct ures and oth er feature s of the ro ck. 

Science Return 

In addition to  grindin g a clean, sm ooth surfa ce, the RAT ca n also provide i nformatio n characterizing th e  rock’ s physical prope rties.  By analyzin g the d ata set returned to  Earth, the RAT gene rated data will b e  comp ared to  informatio n from an extensive collection of ro ck grindin g data gene rated at Hon eybee   Robotics an d stored i n Honeybee’ s rock lib rary. Oth er methods of manipul ating the d ata will be u sed to  determi ne th e rock’s phy sical p roperties. If the am ount of m aterial remove d is known, th e amo unt of  requi red energy pe r unit volume  of m aterial  can be determi ned. This ratio  can then  be com pared to data  within Honey bee’s libra ry to determin e which type of  rock te sted on Earth mo st clo sely re presents th e  excavated M ars rock, thu s providing th e rock’s physi cal stren gth propertie s. 

Figure 3.  Ro ck Abr asion Tool -- Botto m View  281
In ord er to d etermine  the actual grinding torqu e based on true motor current, a calibrated  multimeter was  placed in  series with the  grind  moto r durin g dyn amomet er test ing. Thi s data was then used to conve rt  amplifier current to true  motor curre nt. Figure 5 shows the ave rage moto r current vs. ave rage a mplifier  curre nt obtai ned from dyn amomete r testing pe rform ed at 0ºC (sa me temperat ure at which grindi ng was  perfo rmed ). The vertical line sh ows th e 99th percen tile grind mot or amplif ier current obtain ed from th e  grind m otor h istogram. Th e hori zontal line sh ows th e corre sponding direct mot or current of 459.65 mA.   This is th e wo rst-ca se scen ario for moto r current du ring grindi ng. 

Figure 4.  Grind Motor Hi stogr am 

Figure 5.  Motor Curren t vs. Amplifier Curren t  284
Figure 6 sho ws the moto r current vs. torqu e for the  grind motor at 0ºC. This grap h wa s obtained u sing  averag ed data from the g rind moto r dyn amomete r testing. The horizontal line repre sents the  worst-ca se  motor current  durin g grin ding, and the v ertical li ne on  the left repre sents the corresp onding torque on the   motor (25.1 3 mNm). This is the worst-case to rque applied to  the grind  moto r during a grinding op eration.   This value was used to defi ne the torq ue requi reme nt for the RAT margi ns. 

Figure 6.  Motor Curren t vs. Torque at 0ºC 

Revolve Moto r Analysi s  The closed loop current / torqu e relati onship was determi ned f or the  revolv e moto r similar to how the  grind  moto r was analyze d. The worst-ca se amplifier current d rawn by th e revolve motor durin g a worst  case grind was cal culate d to be 28.29 mA (99th percentile). F rom this value it i s estimated  that the revolve  motor experi ences a worst-ca se to rque of 16 6.6 mNm during a worst-ca se gri nding operatio n. Th e  revolve moto r is heavily ge ared down m echanically, and t he majo rity of the grindi ng torq ue cre ated is d ue  to intern al losses. This causes moto r margin to be difficult to d etermine ba sed solely on torque.  In order to  demon strate that the actio n of gri nding  does not cause signifi cant torqu e loads on the revol ve motor, the   amount  of torque experi enced by the  motor with no load attach ed to it was calculate d. To achieve  this,   the revolve motor wa s run de-cou pled from the dynamom eter, and the no-l oad cu rrent was mea sured.  The am plifier curre nt was then inte rpolated bet ween 3 V and 6 V because the revolve motor ty pically run s  at 4 V d uring a g rinding ope ration. The h eavy geari ng of the revolve motor causes a torq ue of  approximatel y 156.8  mNm , which is onl y about  10 mNm less th an the to rque seen during a worst-ca se  grindi ng op eration.  

Z-Axis Motor Analysis   The cl osed l oop curre nt / torque relationship was determined f or the Z -axis motor simila r to ho w the   revolve moto r was analyzed si nce bot h motors a re heavily gea red do wn m echanically. The worst-ca se  amplifier current dra wn by the Z-ax is motor du ring a worst-ca se grind was calculated to be 23.8 mA (99th  percentile ). From thi s value it is e stimat ed that th e Z-axis m otor experie nces a worst-ca se force of 5.907   N du ring a worst -case gri nding o peration. From no-load current values, it is estimated th at the heavy   geari ng of th e Z-axi s moto r cau ses a force of ap proxi mately 4.04 N, whi ch is approxim ately 2 N le ss than   the force see n durin g a wo rst-ca se grinding ope ration.   285
Review  of International Space Stat ion Mechanical System Ano malies 

David S. McCann* 

Abstract 

The Intern ational Spa ce Station (ISS) when comp leted will co nsist of over 32 different  mecha nical  system s utili zed i n more than 15 0 appli cation s providin g fun ctionality for assem bly, rotation and  deployme nt of major co mponent s, and maintena nce. Currently, a majori ty of these system s have bee n  launched and either fu nctio ned o r checked out on orbit, and all h ave co mplete d preflight te sting an d flight  readi ness processing. Within the  ISS p rogram th e Structu res and Mechani sms (S&M) Syste m Probl em  Resolutio n Team (SP RT) is respon sible for the reso lution of p roblems asso ciated with all ISS mech anical  system  during qualificatio n, pre -flight checko ut and on-o rbit ope rations. This activity combi ned with the  shear number of compl ex mecha nical sy stems on ISS has expo sed the tea m to a great  numbe r of  probl ems of intere st to the Space M echanism commu nity. 

The Problem Rep orting an d Corrective Action (PRACA) system i s used to  track, d ocume nt, and ap prove   the investig ation an d analysis, root cause dete rminati on, and  corre ctive actio n and recurren ce co ntrol fo r  these problems. Fo rmal S PRT’s were formed in 2000. Since it s inceptio n the S &M SPRT ha s dip ositioned  over 35 0 gro und an d on-orbit pro blems. The PRACA  database is used to store  all probl em reco rds and  can b e used to gene rate re ports o n vario us metri cs. 

Figure 1.  ISS on orbit 

This paper will present the result s of a thorough review of the me chani cal system problems that have  been a ddressed by the S &M SPRT. Metrics on th e types an d freque ncy of  variou s pro blems will be  presented. Problem s with f astener installation an d secondary lo cking featu res as well a s tolerancing fo r  thermal d eformation and rigging a re example s of the most com mon problem s. The ro ot causes will be  categ orized and the mo st prevalent causes will b e discu ssed from a lessons le arned perspe ctive.  Determinatio n of root ca use is often th e most difficu lt part of a failure inve stigation and yiel ds the mo st                                                    * The Boeing Comp any, Ho uston, TX 

Procee dings of the 37th Aeros pace Mec hanisms Sy mpos ium, Joh nson S pace Ce nter, May 19-2 1, 200 4  291
insight for de terminin g proper re currence control. So me of th e recurre nce controls put i n place by the   SPRT for th ese problem s can be a pplied on ot her hardware program s. Exam ples i nclude p roper  appli cation o f drawin g ca ll-outs for fa stene r install ation, control  of lubricant s to prevent ing galling   situation s, and prope r inte rface  control  for me chanisms. A summ ary of the problem s that have ha d the   most impa ct on the ISS prog ram will a lso be reviewed. It is hope d that by pro viding the re sults of thi s  investigatio n that the most  commo n causes of ISS mech anical sy stems ano malies can be av oided in the  developm ent of future sp acecraft mech anisms. 

ISS O verview 

The Intern ational Spa ce Station whe n completed will be the larg est and most co mplex spa ce vehicle eve r  assembl ed. It  will be mad e up of mo re than 18 individually la unched m ajor ele ments. There will be more  than 35 different mech anical system s that will provide  a wide variet y of functions for the vehicl e. Figure 2   shows the m echanical sy stem topol ogy for the co mple ted ISS. Five differe nt attach sy stems provid e  remote  struct ural attach ment for th e major struct ural eleme nts. S eventeen  separate  Com mon Be rthing   Mech anisms provide structural atta chme nt and pressu rized access between the habitabl e ele ments. Five   Russian Docking System s provide the same type of  pressurized attachme nt betwee n the US elements  and the Ru ssian el ement s. Four Seg ment-to -Segment Attach  Systems an d five Rocke tdyne Tru ss  Attachment S ystems provi de st ructural  attachme nt betwee n the major t russ elements that make  up th e  backb one of the station.  Six Comm on Attach  System s provide lo cation s to attach l arge exteri or  unpressu rized element s and payload s to the ISS. Several differe nt mecha nical systems a re used in the   deployme nt of the sola r arrays, radiat ors, and  antenna. T wo Solar Al pha Rotary Joints, two Th ermal  Radiator Rotary Joint s, and the eight  Beta Gim bal Assemblie s (BGA) p rovide autonom ous pointing  of th e  solar array s and radiato rs to supp ort th e power g eneration and heat rejection sy stem s. Nu mero us oth er  mech anical systems li ke the Mo bil Transporter, Trai ling Umbili cal System a nd Hatches account for the  more tha n 150 appli cation s of complex mech anical systems o n ISS. 

Figure 2.  Mechanical Sy stem Topolog y  292