ai-aerospace/ams_data_full_2000-2020 huggingface dataset:
Table 5.  Magnetomotiv e force of the friction plate in the spare brake 

 0.051-mm shim 2x 0.051-mm shim  PLATEmmf (ampere-turns) 13.4 18.4 

Table 6.  Magnetomotiv e force of the solenoid coil v ersus v oltage 

Voltage (VDC) 1 2 3 4 5  Ni (ampere-turns) 7.9 15.8 23.7 31.6 39.5 

Conclusions and Lessons Learned 

Given the close tolerances of the parts in the brake assemblies, there was more unit-to-unit difference in  drop-out voltage than expected. This was likely due to a larger than expected variation in the magnetic  properties of the 15-5 PH stainless st eel friction plates, which resulted in  a significant differences in the  amount of residual magnetism in the magnetic circuit from brake to brake. These variations must be  carefully accounted for when using 15-5 PH stainless steel in a magnetic circuit. 

Although residual magnetism was recognized as an iss ue in the design (hence the existence of the nonmagnetic 0.051 mm shim), the lack of an appropria tely specified lower bound on drop-out voltage  prevented brakes with inadequate margin from bei ng discovered and reworked earlier in the flight  schedule. More specifically, drop-out should have been treated like any deployment, guaranteeing a  minimum forc e fac tor of s afety of  2.0 that was verified by test.   Acknow ledgements  The work described in this paper was performed by t he Jet Propulsion Laboratory, California Institute of  Technology, under contract with the Nati onal Aeronautics and Space Administration. 

The effort described in this paper encompassing the discove ry of the failure, the failure investigation, the  assessment of all LPAs, the rework of two LPAs, and the removal and reinstallation of LPAs on the flight  landers, all in a three week period of time, required the combined talents of m any people. In particular,  the author gratefully acknowledges the contributions of Mike Johnson,  Pablo Narvaez, and Don Sevilla of  JPL, and Ed Joscelyn and Boz Shar if of Aeroflex Laboratories.  References 

1.  Gillis-Smith, Greg R.  "Mars Path finder Lander Deployment Mechanisms."  Proceedings of the 30th  Aerospace Mechanism  Sym posium , (May 1996), pp. 255-271.  2. Hanselman, Duane. Brushless Perm anent-Magnet Motor Design.  New York: McGraw-Hill, Inc., 1994,  pp. 13-59.  3. Wangsness, Roald K. Electrom agnetic Fields.  New York: John Wiley & Sons, 1979, pp. 320-333.  4. Slocum, Alexander H. Precision Machine Design.  Englewood Cliffs: Prentice Hall, 1992, p. 663.  235
High Gain Antenna Gimbal for the 2003-2004 Mars Exploration Rover Program 

Jeff Sokol*, Satish Krishnan** and Laoucet Ayari*  

Abstract 

The High Gain Antenna Assemblies built for the 2003-2004 Mars Exploration Rover (MER) missions  provide the primary communication link for the Rovers  once they arrive on Mars. The High Gain Antenna  Gimbal (HGAG) portion of the assembly is a two-axis gimbal that provides the structural support, pointing,  and tracking for the High Gain Antenna (HGA). The MER mission requirements provided some unique  design challenges for the HGAG. This paper describes a ll the major subsystems of the HGAG that were  developed to meet these challenges, and the requirements that dr ove their design. 

Introduction 

MER Overview   The 2003-2004 Mars Exploration Rover mission consists of two long-range rovers that carry a payload of  science instruments. The goal of t he mission is to determine the histor y of climate and water at sites on  Mars where conditions may once have been favorable to life. Two landing sites were chosen (Gusev  Crater and Meridiani Planum) because they offer ev idence that liquid water was once present. The  science instruments will be used to read the geologic record at each site and to determine how suitable  the conditions may have been for supporting life. The two rovers were named “Spirit” and “Opportunity”  and were launched in June and July of 2003 with a scheduled Mars landing in January of 2004. Each  rover has a mass of approximately 180 kg and has a range up to 100 meters per Martian day (~24 hours  and 39 minutes) [1].  

Figure 1 shows the rover in its deployed configurati on. The topic of this paper, the High Gain Antenna  Gimbal, can be seen on the top deck of the rover along with the Low Gain Antenna and the Panoramic  Camera Mast Assembly. The HGAG provides t he pointing and tracking capability for the High Gain  Antenna RF communication system. 

Figure 1.  Mars Exploration Rov er [1]                                                    * Ball Aerospace & Technologies Corp., Boulder, CO  ** Jet Propulsion Laboratory, Pasadena, CA 

Proceedings of the 37th Aerospace Mechanisms Symposium, Johnson Space Center, May 19-21, 2004  237
Table 1.  MER HGAG significant driv ing requirements 

Requirement Value Unit HGAG mass alloc ation ( no HG A) 4.5 kg Distance from HGA bor esight to Az imuth Ax is 0.33 m Azimuth Dr ive Angular  Rotation 280 deg Elevation Dr ive A ngular  Rotation 234 deg Minim um mechanic al control and k nowledge 0.5 deg Drive oper ational s tep size 0.2 deg Minim um drive rotation s peed 0.78 RPM Operating voltage r ange 22.5 to 34 V Maximum MER inc line w ith respect to gr avity  during oper ation40 deg Quasi-static design load ( stowed) 45 g Stow ed first modal f requenc y (minimum) 50 Hz Deploy ed first modal f requenc y (minimum) 14 Hz Contam ination c ontrol bak e out tem perature 110 C Survival tem perature range ( non-bake out) -120 to 60 C Operating tem perature range -70 to 45 C Motor  revolution lim it (testing / f light) 2.5E+06 revs RF loss betw een HG A and Rover  bulk head 1.8 dB

Three HGAA assemblies, 2 flight and one Engineering Model (EM), were delivered to JPL during the third  quarter of 2002 for integration into the rover asse mblies. The flight assemblies were successfully  integrated and launched aboard the two Mars Rover mi ssions, “Spirit” and “Opportunity”, during the  summer of 2003 for arrival on Mars in January 2004.  The engineering model was integrated on the EM  Rover that is currently being used for mission control testing at Jet Propulsion Laboratory. 

Design Discussion 

Drive Train   The gear train and RF system were designed conc urrently to accommodate packaging and performance  requirements. While the expected maximum external torsion load on the azimuth drive (2.9 Nm starting  and 2.1 Nm running) is larger than the elevation driv e (1.3 Nm starting and 0.6 Nm running), both drives  were designed to accommodate the higher azimuth drive loads. In order to simplify the design, analysis,  manufacturing, and assembly of the gi mbal, both drives use identical gear  trains that consist of a gear  motor, spur gear stage, and a harm onic drive that combine to produc e a 5424.7:1 gear reduction. A cross  section view of the system is depicted in Figure 3. Early design trades included a more modular design,  with the RF coax routed external to the main gim bal structure. This approach simplified the gear train  design (and should therefore be considered for futu re gimbal designs) but required additional envelope  external to the gimbal. Because of the tight pa ckaging requirements on the ME R, the decision was made  to route the coax cables internally. 

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By assuming worst-case performance for all drive train components in the mechanism analysis, the need  to pre-screen components prior to assembly was elimi nated. Overall, simple and robust design principles  were used to minimize the gimbal’s sensitivity to variability in the performance of each component. 

Figure 3.  HGAG cross-section 

Drive Bearings   The HGAG drives each contain two pairs of ba ck-to-back mounted angular contact bearings. Both  bearing pairs were manufactured by Timken Super  Precision (MPB) and were lubricated at Ball  Aerospace & Technologies Corp. The larger drive output bearings are required to carry the axial,  radial, and moment loads generated during launch,  landing, and MER maneuvering. The smaller  drive input bearings are required to support the harm onic drive input loads as well as the radial and  moment loads from the gear motor to harmonic dr ive spur gear stage. The bearings are made from  440C stainless steel (which helps to match the CTE of the surrounding titanium housings) and have  phenolic retainers. Because of the operating envir onment on Mars as well as the mechanism life  requirements, all the bearings are lubricated with Brayco Oil Company 815Z oil along with a 10%  grease fill of Braycote 601. Although this lubricat ion allows the HGAG to survive the operating  conditions and meet its life require ments, it does significantly in crease the starting and running torque  of the system at the cold oper ating temperatures. The increase in cold temperature starting and  running torque made it critical that the minimum di ameter bearings were chosen that could meet our  structural load and packaging require ments. In both cases, the beari ng pairs employ a “hard preload”  assembly that utilizes precision shims to ac hieve the desired preload in the bearing set. After  assembly, the bearing set preload and stiffness we re verified by running torque and load tests. 

Gear m otor  In an effort to minimize the number of differ ent drive components used on the overall MER system,  the majority of mechanisms employ  one of two brush DC motor sizes supplied by Maxon Corporation.  The large gear reduction in the HGAG mechanism allows the gimbal to use the smaller Maxon  REO20 motor (20-mm diameter), in lieu of the larger Maxon RE25 motor (25-mm diameter). By using  the smaller motor, the HGAG design saves mass and reduces power consumption. Nearly all of the  gear train calculations assumed the worst-case  performance REO20 motor (except for mechanism  240
the rotary joint. As the drive turns, the torque tube pr ovides the force required to turn the rotary joint.  Because of the clearance fit required between the tor que tube and the rotary joint for assembly reasons,  the RF cable is required to accommodate a small am ount of rotation prior to the torque tube engaging.   However, this small amount of torsion is within the capability of the semi-rigid coax cable. Figure 4 shows  the torque tube arrangement. 

Figure 4.  RF sy stem torque tube 

A seal system is used to prevent contamination from entering the driv e train at the point where the RF  system exits the drive. An O-ring is slipped over the RF cable prior to final assembly. During final  assembly, the O-ring is located into the groove of a two-piece seal clamp that is  mounted to the output of  each drive. Figure 5 indicates the seal location and design. 

Figure 5.  RF coax contamination seal 

Flex Cable System   The gimbal design requires that electrical power  and signals be passed across both the rotating azimuth  and elevation drives. To accomplish this, flex cable tw ist capsules were used at the output of each drive.  The flex cables are wound inside the capsules “clock spring” style, and wind and unwind as the drive  rotates. Care was taken to size the flex cable lengt hs to prevent the system from binding up at either end  244
bearing. The two horizontal mounting pads of the f our-strut mount are then shimmed and fastened. The  final step is to adjust the elevation drive contact pad until it just touches t he feature on the HGA arm and  then lock it in place using the two jam nuts. Du ring subsequent re-stowing operations, the four-strut  mount is not readjusted. The drives  are simply driven back to their launch positions and the pyrotechnic  pin puller is reinstalled. This mount  provides the system with the additi onal structural stiffness required to  achieve the 50 Hz first mode fundamental frequency requirement during launch and landing. Figure 7  shows the launch locking features on the HGAG. 

Figure 7.  HGAG launch lock features 

Pyrotechnic restraint devices are used in a num ber of mechanisms throughout the MER system and  underwent a number of development tests throughout t he life of the program. One significant design  issue that was discovered during this testing is t hat the shock load induced by  the pyrotechnics has the  ability to damage the brush DC motors (as well as ot her sensitive components) that are being used on all  the MER mechanisms. Locating the py rotechnic device a sufficient dist ance away from any motor, or  attempting to try and isolate the pyrotechnic mount from the rest of the system mi tigates the risk. On the  HGAG, there was concern that the pyrotechnic device was located too close to the azimuth drive motor  and that the four-strut mount pr ovided a direct path between the py ro-shock and the motor. Analysis  results were inconclusive, so it was decided to verify the design through additional testing. Final design  validation was performed on the Engineering Model  (EM). The EM was fitted with a redundant  pyrotechnic pin puller whose two charges were rated at 120% those of the fli ght design. Both charges  were fired simultaneously and the motor survived the test. Additional pyrotechnic event tests were  performed by JPL at the MER system integration level, including test s at the cold operating temperatures.  In all cases, no performance effects were seen on the HGAG drive motors. 

Internal Hard Stops and Azimuth Drive “Flipper” Hard Stop   Both the azimuth and elevation drives employ internal hard stops at the end of  travel for each drive.  These stops provide a consistent mechanical “home”  for the drive electronics and prevent the mechanism  from exceeding its dynamic envel ope and damaging any of the other hardware on the MER. The hard  stops consist of an integrally machined titanium boss on the rotating output housing of each drive and a  stationary Nitronic 60 hard-stop that is pinned and fast ened to the stationary portion of the drive. The  Nitronic 60 piece is removable to allow for continuous  drive rotation during subassembly testing, but the  interface is designed to remain repeatable after consec utive installations. The stop materials were chosen  to prevent galling in the system. Figure 8 shows the internal hard-stop features. 

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Figure 8.  Driv e internal hard stop features 

In addition to the internal hard stops, the azimuth dr ive must also provide a secondary hard stop that is  engaged after the first azimuth drive deployment. This  secondary stop is necessary because the dynamic  sweep envelope of the High Gain Antenna on the elevation drive can violate the envelope of the Low  Gain Antenna if the azimuth drive is in its launch locked position. To avoid this, during its first sweep  away from the launch position, a pin feature on t he azimuth drive passes through a spring-loaded “flipper  stop” that closes behind the drive once the pin feature has passed. The new “flipper” hard stop will only  allow the azimuth drive to return within 15 degrees of the original launch lock position. At this spot, the  elevation drive is free to rotate the HGA thr ough its full dynamic envelope without endangering the Low  Gain Antenna. The flipper stop employs two redundant  torsion springs that are sized to independently  provide enough closing force to keep the flipper se cure against the housing during launch and landing.  The materials of the pin feature and flipper as well as  the shape of the sliding surfaces were chosen to  minimize the torsion load required to pass thr ough the spring-loaded gate. Drive level testing was  performed to verify the starting / running torque requir ed to pass through the gate as well as to verify the  repeatability and strength of the stop under numerous hard stop contacts. Figure 9 shows the secondary  azimuth drive “flipper” hard stop. 

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HGAG Test Discussion 

The HGAG Mechanism test verification plan included individual drive testing and full assembly level  testing under all operating conditions. Each individual drive was characterized for performance (including  torque speed curves, current limits, pointing accuracy, and stiffness) pr ior to being integrated together to  form a complete HGAG assembly. System level veri fication tests included struct ural, RF system loss,  pointing accuracy, and dimensional compliance. Ev ery attempt was made to simulate the Martian  operating environment for design verification.  Therma l vacuum chambers were used to achieve the low  operating temperatures and low atmospheric pressures required. Many tests were run with the gimbal  mounted on an incline or working against weight and pu lley systems to simulate the possible operating  orientations and drive loads aboard the rover. After each significant testi ng milestone, a standardized  ambient functional test was perform ed on the entire system to monito r the health of the mechanisms and  ensure that no performance degradation was seen. At the end of its testing pr otocol, the Engineering  Model (EM) underwent a rigorous life test where each drive was cycled under the maximum expected  load for the total required number of motor revolutions . In the end, the HGAG units were shown to meet  or exceed all design requirements. Figure 10 shows one of the flight HGAA assemblies just prior to  shipment to JPL for Rover integration. The ther mocouple wire bundles that can be seen coming from  different locations on the HGAG were used by JPL during system level testing and were removed or  clipped by JPL prior to flight. 

Figure 10.  Flight HGAA show n just prior to shipment to JPL 

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Conclusion 

The design, assembly, test, and delivery of the HG AG mechanisms were a cooperative effort between  JPL and BATC. It is difficult to capture all of the lessons learned on a development program such as this,  but it is safe to say that all personnel involved wi th the HGAG program received a valuable education in  what it takes to deliver qualified flight hardware fo r a time sensitive planetar y exploration mission. One  design lesson that does stick out because of its risk to the HGAG program is the potential damaging  effects of the shock load created by  pyrotechnic launch lock devices lik e the pin puller. Care must be  taken to properly isolate sensitive components (electronics, motors, etc. ) whenever possible. In the case  of the HGAG, the problem was discovered late and c ould not be shown to be acceptable analytically. This  required additional verification testing, which in turn  introduced additional risk, schedule, and cost late in  the program. Fortunately, the design was proven to be robust through testing and any potential redesign  was avoided. The HGAG successfully met the design and schedule challenges that it faced, and is on its  way to play an important role in the communication system of the MER mission. 

Acknow ledgements 

The authors would like to thank all the HGAG progr am technicians, machinists, designers, engineers and  managers at BATC, and the entire ME R team at JPL, for their commitment to developing these  mechanisms. 

This work was perform ed under a contract with Jet Propulsion Laboratory, Ca lifornia Institute of  Technology, for the National Aeronautics and Space Adm inistration.  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. “Mission Overview”. Athena Mars Exploration Rove rs website. Cornell University. Date of access  October 24, 2003. http://athena.cornell.edu   

2. R. N. Taylor, “BEAR3D Users Manual,” Inte rnal Report, Ball Aerospac e & Technologies Corp.,  Boulder, Colorado, 1993. 

3. A. B. Jones, New Departure Engineering H andbook, Analysis of Stresses and Deflections, New  Departure, Division of Generals Moto rs Corp., (out of print) 1946. 

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Figure 1.  Simulated  limb tran smittanc e spectrum  (0.0 5 cm-1  spec tral re solution) at 300 km tang ent alti tude  in the atmo sphere o f  Titan, sho wing the abso rption co ntributions o f ethane and v arious  isotopome rs of meth ane. This 6  cm-1 interv al represen ts less than  0.2% of the full PAOS bandpas s  Most sp aceb orne, hig h-resolutio n Fourier tran sform spe ctrometers to date h ave been Earth-orbitin g  instru ment s constructe d using traditio nal electromag netic act uators in the form of dire ct driv e voice coils  or brushless DC motors with lead scre ws or belt drive system s to drive th e transl ation sta ge (Persky,  1995 ). The se mech anisms are eithe r too massive or don’t have th e req uisite spectral resoluti on for future  exploratio n of planetary atmosph eres. Figure 2 bel ow sh ows a  compa rison of the PAO S instrume nt  requi rements with other F ourie r Tran sform Spe ctrometers by plotting the overall instrum ent mass ve rsus  the numb er of spectral channels (simult aneo us spectral ba ndwidth  divided by spectral resol ution) (T oon,  2001 ). 

Figure 2.  Comparison o f PAOS ma ss and s pectral resolv ing  power with other F TS. Colored symbols repre sent pla netary  instruments.  Filled s ymbols repr esent launched in struments.    252
Figure 4.  Left End of Tr avel Figure 5.  Right End of T ravel  Linea r Piezo Motor   Ultimately the design g oals and the dev elopmental n ature of the pro ject  led to the d ecision to u se new te chnology in the form of an ultra sonic  linear pie zo motor d eveloped by Nanomotion Ltd, F igure 6. Thi s motor  allows for drastic ma ss savings over mo re tra ditional system s whil e, at  the sa me tim e, pro ducing very co nstant velocity moti on. Piezoele ctric  material s allow me chanical strai n to be dire ctly cou pled to electrical   potential. Th e moto r works by si multaneo usly exciting  both a   longitudin al (Figure 7 ) and tra nsverse (Fig ure 8 ) bendin g mode in a   piezo elem ent using t wo sinusoidal vo ltage in puts. T he re sult of t hese  combi ned modes i s an elliptical motio n at the tip of the pie zo element  that drives a l inear stage th rough a prelo aded fri ction i nterface.  Figure 6.  HR-4 (Four  Element ) Linear Piezo                      Figure 7.  Longitudin al Extension Mode   Figure 8.  Transv erse Be nding Mode  

The pie zo ele ments have a  hard ce ramic alumina fing er tip attache d to the element body, whi ch interact s  with an al umina stri p mou nted to the m echanism’s moving sl ed as shown in Figure 9. Thi s provide s an  optimum coefficient of fri ction an d signifi cantly redu ces wea r. During extensi on and be nding of the pie zo  element the  sled i s actively force d in the de sired directio n, whil e durin g contraction and bendin g in the  other di rection, the piezo either slides on, or sepa rates slig htly from t he ce ramic drive strip. This motio n  results in a net forwa rd movement of the sle d over one cy cle of t he pie zo element. This ell iptical motio n  takes pla ce at a frequ ency of 40 kHz. Because of th e speed of th is oscillation,  the ma ss of the moving  object mech anically filters any jitter, resulting in  extremely smoot h linear moti on. The mag nitude of the   displacement  during any o ne cycle i s proportio nal to the voltage applied a cross the piezo el ement. This  voltage can be cha nged rapidly to co ntrol the veloci ty of the moving sle d. Since the piezo ele ment  displacement s are on th e nano meter scal e, these motors ca n also be u sed to a chieve position al  accuraci es of  less tha n 10 nano meters.  A stan dard servo  controll er with any  standard fee dback d evice  can close the  control lo op and servo the  applied volta ge in ord er to follow a de sired traje ctory profile. 

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