ai-aerospace/ams_data_full_2000-2020 huggingface dataset:
364 The electronic control system architecture was adopted and leveraged from the WindSat BAPTA  electronics and software to perform the speed and position control tasks for GMI. The model includes an analog controller used to control SMA rate, a control plant  that includes models of the motor (transfer  function with current feedback loop) and disturbance torques, and feedback sensor that models position sensor with noise and velocity calculation via position differentiation. Further detail of the control system modeling and verification testing are beyond the scope of this paper.   Component, Material, and Lubrication Selections 

The design-driving components for the SMA were, not surprisingly, the challenging procurements of the  SRA, the resolver, rotary transformer, and the motor. However, since this design application requires such a long life (75million revolutions) and reliable performance over the entire life, the bearings, bellows, and lubrication are just as critical to the ultimate success of the drive. Figure 5 shows a cross-section of the SMA.  

Figure 5. Cross-section of the SMA d esign with drive component labels 

As previously mentioned, the SRA is a heritage design procured from Moog Components Group (formerly Poly-Scientific and Electro-Tech Corp.). The SRA configuration consists of an angular contact bearing pair, a gold-on-gold, double v-groove type brush to ring design with a total of 124 electrical circuits (rings) that are wet lubricated. The SRA brushes are a solid gold alloy and each brush pair rides in the hard gold over soft gold double v-groove rings. See Figure 6. This is a heritage design which dates back to DSCS II drive in 1980, and was similar to the WindSat BAPTA design with minor modifications. The PAO wet lubrication system is a BATC proprietary formulation and process [3] first demonstrated on the WindSat BAPTA SRA and will be life-tested in the GMI Life Test Unit (LTU) SRA discussed in section 5 of this paper.  The primary drive and control components; motor, resolver, and rotary transformer were all procured from Axsys Technologies, now owned by General Dynamics AIS. The motor is an external rotor 3-phase brushless DC torque motor with a skewed and redundant winding design. The motor design parameters; Kt = 320 oz-in/amp (2.26 N-m/amp), cogging torque = 13 in-oz (9.2 N-cm), peak torque = 639 oz-in  (4.51 N-m) (at 1.71 amps max. current), air gap = 0.023 in (0.58 mm), and weight = 7.02 lb (3.18 kg) near identical to the heritage motor used on WindSat BAPTA. See Figure 7.  The resolver and rotary transformer were procured as matched sets with redundant units mounted on common hub and sleeve designs. The resolver design is the BAPTA heritage 2-speed (1x/64x) resolver used for commutation and position feedback. The combined resolver/rotary transformer set was specified
365 to have a reduced 30 arcsec minimum accuracy (at room temperature) and a 6 arcsec maximum  electrical zero shift over temperature (±20oC), and a rotary transformer weight = 3.9 lb (1.8 kg). The  resolver mounting alignment requirement = 0.0005 in (0.01 mm), air gap = 0.016 in (0.41 mm), and resolver weight = 1.52 lb (689 g) were all heritage designs too. See Figure 8.  

Figure 6. Slip Ring Assembly (SRA) showing double gold brush pairs, double v-groove gold rings  

Figure 7. BAPTA 3-Phase DC torque motor  Figure 8. BAPTA 2-speed Resolver (1x/64x)  

It is probably becoming clear that ‘Flight Heritage Design’ does not always mean it is truly ‘flight’ heritage. There is no such thing as 100% heritage based design; materials are altered, processes change, suppliers change, and requirements, or specific applications change loads, environments, or other critical conditions that may affect the design performance in sometimes dramatic ways.  

The main SMA drive consists of a pair of preloaded angular contact bearings separated axially on an  AlBeMet TM (Aluminum Beryllium Alloy Material) shaft and housing. The bearings are ABEC 7 (ABEC 5  balls), 52100 steel bearings procured from NSK Corp. (formerly manufactured by RHP, England) and is a heritage component except for a manufacturing change by the new vendor to improve the inner raceway curvature and dam height. These changes were modeled in the structural analysis and taken into account in the lubrication analysis that still shows a slightly positive margin of safety (MS = 0.02) at End of Life (EOL) for the number of stress crossings and with the mean Hertzian contact stress of the 110 lb (489 N) preloaded bearings, which is acceptable. The bearings are precisely preloaded and ‘snubbed’ by means of a titanium diaphragm and snubber. Several axial deflection vs. force measurements were taken during assembly to confirm the correct preload and snubber gap were achieved in the assembly. The lubrication for the SMA main drive bearings is a Nye Synthetic Oil with Rheolube grease and is heritage BATC tribology elements [4].  

The SMA primary structural housings and inner shaft elements were fabricated out of AlBeMet TM which  contains 62% Beryllium and 38% Aluminum and was chosen because of the material properties; mainly  for high specific stiffness and heat conduction, compared to other more common aerospace structural materials. However, because of the Beryllium content in this alloy, it does require special handling to prevent human exposure to any particles that could be generated from the parts.
366 Finally, it should be mentioned here that a despin shaft with bellows coupling is used to provide the  mechanical ‘despin’ interface to the stationary upper instrument structure, calibration assembly, and to account for any minor misalignments between the bearing pair in the SMA main drive and the bearing pair in the SRA. There is also another preloaded angular contact pair that supports the Calibration Pedestal and the upper IBS structure, this bearing configuration is referred to as the Despin Mechanism in the GMI instrument. Those bearings are being life-tested as well but the design of that mechanism is not included or mentioned further in the discussions in this paper but similar issues and constraints apply. The bellows life test was a significant effort at Tara BelFab.   Integration Experience, Test Setups, and Test Planning   The SMA drive successfully completed performance and environmental (thermal vacuum and EMI) acceptance testing in Nov. 2010, and met all of the sub-system requirements that were to be verified via test. The objective of this testing was risk reduction for verifying the critical functional and performance requirements early in the integration and test program, so that any issues found in the SMA integration with the electronics, performance or environmental testing could be addressed at the sub-system level, months before the SMA is integrated into the GMI instrument.    The test setups for performing the acceptance testing are briefly described here. The electronics integration, commutation offsets, initial spin testing, and baseline performance tests were all run in the clean room with the SMA mounted within the Inertial Test Fixture (ITF), See Figure 9. The completed SMA is seen within the test fixture along with the circular instrument mass and inertia simulator plate mounted on the SMA. The square shaped torque tube plates around the SRA above the SMA are seen and are used to interface with the optical encoder, torque transducer, and brake during the torque margin and rate testing.    

Figure 9. SMA mounted in test fixture to prepare for acceptance test in the clean room.  

Once the ambient testing was successfully completed, the test setup was loaded into the Thermal Vacuum (TVAC) chamber and the article under test was taken thru one complete survival thermal cycle with 4 hour minimum dwells at each temperature extreme. Then all the acceptance tests were repeated in the TVAC at the operating temperature extremes, see Figure 10.  Performance Test Results   The results of the acceptance testing of the SMA, including measured rate control, position accuracy, pointing alignment, wobble, and power dissipation over operating temperature ranges in vacuum, are summarized in Table 2.
367  

Figure 10. SMA mounted in inertial test fixture and loaded into chamber for TVAC Testing  

The SMA torque margin was only tested in the cleanroom at ambient temperature but the drive current was monitored over temperature during operation. The torque margin can be calculated by knowing the motor constant Kt (which is essentially constant over temperature) and measuring the current at the operating temperature extremes. Then the change in current is due entirely to the change in bearing and SRA drag torque or the drive, within the measurement uncertainty of the torque transducer.       Table 2. SMA Tested Performance vs. Requirements 

The rate control was measured independently with an optical encoder and compared to the tachometer/ resolver feedback data. In this way, we measured and compared the rate control of the drive as commanded and the resultant rate, by performing an interpolation between any 2 points in the sample

data set, and calculating the error between the encoder and the resolver information for rate and position.  BAPTA 8 yrs. on-orbit to  date, SRA and SMA Bearing  Life Tests BegunDesign Life 38 months on-orbit op + 12 months  I&T + 2.4 months op storage (2 yr @  10%); total life reqmt 52.4 months20.3 kgMass 20.4 kg maximum 0.87 MS Torque Margin 1.5x on known's, 2x on variables32 + 0.1% RPM 0 – 33 RPMRotation Rate /Control Rate shall be 32 + / - 0.3% RPM Range shall be 0 – 33 RPM 10 to 15C OpOperating and Non Operating  Temperatures On-orbit Op: 0 to +50C Survival: -35 to +55C60.8 Hz (GMI level  test)> 7.6 Hz (GMI deploy  test) > 7.8 Hz (GMI deploy  test)Stiffness First mode > 50 Hz (in GMI assy)Deployed First Mode > 6 Hz De-spin torsion > 7.6 Hz< 15 arcsec (RSS) < 5 arcsec (RSS)Spin axisAlign to base shall be < 75 arcsec  Wobble shall be < 30 arcsec 3 sigma  50 arcsec (32 tach pulses / rev)Position Uncertainty Tach pulses/rate control position  accuracy < 60 arcsecMeasured / Tested  PerformanceRequirement BAPTA 8 yrs. on-orbit to  date, SRA and SMA Bearing  Life Tests BegunDesign Life 38 months on-orbit op + 12 months  I&T + 2.4 months op storage (2 yr @  10%); total life reqmt 52.4 months20.3 kgMass 20.4 kg maximum 0.87 MS Torque Margin 1.5x on known's, 2x on variables32 + 0.1% RPM 0 – 33 RPMRotation Rate /Control Rate shall be 32 + / - 0.3% RPM Range shall be 0 – 33 RPM 10 to 15C OpOperating and Non Operating  Temperatures On-orbit Op: 0 to +50C Survival: -35 to +55C60.8 Hz (GMI level  test)> 7.6 Hz (GMI deploy  test) > 7.8 Hz (GMI deploy  test)Stiffness First mode > 50 Hz (in GMI assy)Deployed First Mode > 6 Hz De-spin torsion > 7.6 Hz< 15 arcsec (RSS) < 5 arcsec (RSS)Spin axisAlign to base shall be < 75 arcsec  Wobble shall be < 30 arcsec 3 sigma  50 arcsec (32 tach pulses / rev)Position Uncertainty Tach pulses/rate control position  accuracy < 60 arcsecMeasured / Tested  PerformanceRequirement
368 The rate control and position knowledge of the drive was very good and well within the specified required  accuracy. See Table 2 for SMA test results.  The current was monitored during GMI instrument level before and after environmental tests including vibration, acoustics, EMC/EMI, and TVAC testing. Figure 11 below shows a summary of the SMA motor current monitoring and trending during GMI instrument level testing. Characteristics to note are that the motor current is higher during ambient environment testing due to the air drag on the main reflector and reflector deployment assembly in the cleanroom. The motor current measurements also show that the bearing drag torque in vacuum is higher at cold temperature as expected and predicted. The variation in motor current at cold is due to cycling of the SMA operation heaters at or near cold operating temperature. 

Figure 11. Summary of SMA motor current monitoring and trending thru GMI environmental test    Additionally, the SMA bearings and Slipring Assembly (SRA) have each begun a life-test program. The SMA bearings (3 preloaded bearing pairs) began testing Dec. 2010, see Figure 12. To date, the bearings have completed a survival temperature cycle, followed by 30 days of ambient operation, then 8 operational temperature cycles (1 day each at operating temperature extremes each cycle), 60 days back at ambient, and 30 days at each temperature extreme, all under vacuum and rotating at 32 rpm. That is just over 3.5 million revolutions or 3.4% of the total 104 million cycles (1.25x Flight Model total life) test objective. The test articles and test setups were all performing nominally for all the specified test conditions as initially planned.       

Figure 12. SMA main bearings Life Test setup and preloaded bearing capsules (3 pairs)
369  

Figure 13. SRA life test Setup with the SRA Life-Test Unit (LTU) inside a small chamber ready for  TVAC test 

The SRA was configured in a vacuum and temperature controlled (TVAC) setup (Figure 13) and began a  life test in May 2011. This SRA assembly will also see one month of ambient operation, then one survival temperature cycle with 8 hour minimum dwells at the temperature extremes, followed by 8 thermal cycles to the operating temperature extremes (1 day dwell duration at each extreme) while rotating, and then 1 month at each temperature extreme before returning to 60 days running at ambient temperature before repeating.    However, during the first cycle cold it was discovered that the dewar/chamber being used to chill the SRA assembly was not functioning properly and could not achieve the cold survival temperature, or remain stable enough to meet requirements at the cold operating temperature for the duration of the test. So the decision was made to purchase another chiller for the chamber with more cooling capacity to meet the requirements with adequate margin for desired stability at both cold survival and operating temperatures. The new chiller has been installed and the SRA assembly has been taken cold and successfully completed the first cold survival cycle and 7 cold operating cycles to date.   

The temperature profile and durations were designed to give roughly 80% of the revolutions at the onorbit predicted temperature (essentially room temperature, in this case), 10% of the cycles at hot operating, and 10% of the cycles at cold operating by the EOL of the Life Test. It was also found to be very important to monitor and control the temperature gradient across the bearings and rotating to spinning elements in both the bearings and SRA life test units under test. The Life Tests are expected to run continuously for the next 5 years and accumulate approximately 104 million cycles or revolutions on each test article. It is possible that the Life Test may be extended beyond the projected life if the life test is successfully completed, in order to gain qualified Life Test data for similar longer life missions.  In October, 2011 the SMA bearing life test was briefly stopped for a few weeks due to an apparent increase in bearing torque of one of the bearing pairs. After further investigation, it was found that the strain gage use to measure the bearing pair torque external to the housing was found to have become misaligned and broken, giving a falsely high torque reading (Figure 14). The decision was made to break vacuum, rework GSE setup, replace the strain gage, and restart the test. The Life-Test was restarted in late November 2011 and has been running without incident since then.
370        Figure 14. SMA main bearings life test temperature, bearing torque baseline over temperature, and  total number of revolutions just prior to torque rise latter attributed to GSE strain gage  

Figure 14 shows the average bearing temperature, total bearing torque, and torque measured going into the life test assembly over the revolutions completed to date [5], as well, as the faulty torque readings later concluded to have originated from a misaligned bracket and broken GSE strain gage in the SMA bearing life test setup.   Table 3. GMI Instrument and SMA vibration test and design analysis for launch environments 

   Qualification Acceptance 20 0.010 0.010 20-50 +2.83 dB/Oct +0.56 dB/Oct 50-800 0.024 0.012 800-2000 -2.83 dB/Oct -0.56 dB/Oct 2000 0.010 0.010 Overall 6.0 G rms 4.7 G rmsASD Level (g2/Hz) Frequency  (HZ) Frequency Limit Level Protoflight/  Qualification Testing Acceptance  Testing  5-50 Hz 4.8g 6g 4.8g   Random Vibration Sine Vibration Design Limit Loads  8.0 Lateral Z8.0 Lateral Y10.0 Axial XDesign Limit  Loads (g)Launch Vehicle  DirectionGMI  Axis 8.0 Lateral Z8.0 Lateral Y10.0 Axial XDesign Limit  Loads (g)Launch Vehicle  DirectionGMI  AxisQualification Acceptance 20 0.010 0.010 20-50 +2.83 dB/Oct +0.56 dB/Oct 50-800 0.024 0.012 800-2000 -2.83 dB/Oct -0.56 dB/Oct 2000 0.010 0.010 Overall 6.0 G rms 4.7 G rmsASD Level (g2/Hz) Frequency  (HZ) Frequency Limit Level Protoflight/  Qualification Testing Acceptance  Testing  5-50 Hz 4.8g 6g 4.8g   Random Vibration Sine Vibration Design Limit Loads  8.0 Lateral Z8.0 Lateral Y10.0 Axial XDesign Limit  Loads (g)Launch Vehicle  DirectionGMI  Axis 8.0 Lateral Z8.0 Lateral Y10.0 Axial XDesign Limit  Loads (g)Launch Vehicle  DirectionGMI  Axis
371 Conclusions and Future Work 

The SMA was integrated into the GMI instrument assembly, continued through the rest of the instrument final integration, and successfully completed all instrument-level performance and environmental testing including vibration testing, see Figure 15, to levels outlined in Table 3 above, at the end of January 2012.   The GMI instrument level vibration testing verified the stowed 1 st mode frequency requirement. The SMA  deployed stiffness was also measured and verified during the GMI Instrument level test program in the final flight configuration with the main reflector and Reflector Deployment Assembly installed and deployed on the IBS, see Figure 16.  

Figure 15.  GMI Instrument during vibration testing on vibr ation test table in stowed configuration    The GMI instrument will also successfully completed a static and dynamic spin balance measurement and balancing program both in the cleanroom and in vacuum, to separate and eliminate the effects of air drag on the spin balance results. There has been extensive work done in this area and this subject alone is worthy of another paper in itself [6]. The GMI instrument was successfully completed, delivered to Goddard Space Flight Center, and integrated onto the GPM spacecraft in early March 2012.  Finally, the SMA Bearing and SRA Life Tests will continue for the next 4-5 years and updated status reports will be completed at least once a year until the completion of the tests, assuming the test data remains nominal until the tests are complete.    

Figure 16.  GMI Instrument during initial SMA Spin testing in Cleanroom with the reflector deployed
372 A special thanks to Sergey Krimchansky of GSFC for his input over-seeing the technical work on GMI and  for his support in writing these technical papers.   References 

1.  Newell, D., Rait, G., Ta, T., Berdanier, B., Draper, D., Kubitschek, M., Krimchansky, S. “GPM  Microwave Imager Design, Predicted Performance and Status”, Proceedings of the 2010 IEEE  International Geoscience and Remote Sensing Symposium (IGARSS)  Publication Date: July 25, 2010  2.  Koss, S., Woolaway, S.  “Lessons Learned From the WindSat BAPTA Design and On-Orbit  Anomalies”, Proceedings of 38 th NASA Aerospace Mechanisms Symposium , Langley Research  Center, May 2006.  3.  Pierre, W. “GMI Spin Mechanism Structural Analysis”, BATC Internal SER Report No. 2307821B , Ball  Aerospace and Technologies Corp., Nov. 2008.  4.  Dayton, C. “Slipring and SMA Bearing Lubrication Analysis”, BATC Internal SER Report No.  2242913D & 2241045D , Ball Aerospace and Technologies Corp., Oct. 2008.  5.  Hoffman, C. “Status of the GPM/GMI Bearing Life Test”, 541/Materials Engineering Branch  Memorandum No. MEB-2011-025 , Ball Aerospace and Technologies Corp., Apr. 2011.  6.  Ayari, L. “GMI Rotating Mass Balance Uncertainty”, BATC Internal SER Report No. 2324813A , Ball  Aerospace and Technologies Corp., Jun. 2009.
373 Lessons Learned from the TIRS Instrument Mechanisms Development 

Jason Budinoff*, Richard Barclay*, James Basl**, Konrad Bergandy**, Thomas Capon*, Bart Drake**,  Michael Hersh*, Chris Hormann**, Edwin Lee**, Adam Matuszeski*, Armani Nerses+, Kenneth Pellak++,  Kermit Pope**, Joseph Schepis* and Ted Sholar++ 

Abstract   This paper describes the many lessons learned during the design and development of several mechanisms for the Landsat Data Continuity Mission (LDCM) Thermal Infrared Sensor (TIRS) instrument, built by an engineering team at NASA Goddard Space Flight Center (GSFC). Several mechanisms were developed for TIRS including an arc-second precise mirror positioning system, a launch lock for a 90-lbm (41-kg) cryo-cooler assembly, and a large deployable earth shield. These mechanisms were developed over a 2 year period, and several obstacles were encountered and subsequently solved prior to delivery.    Introduction   The LDCM satellite will launch into a low polar orbit in late 2012. LDCM will provide earth resources data continuity between the currently operational Landsat 5 and Landsat 7 missions and the Joint Polar Satellite System (JPSS) Missions. It will provide a high spatial resolution complement to the lower spatial resolution, higher temporal sampling JPSS data set. LDCM will be carrying TIRS, an actively cooled, nadir-looking, mid-infrared imager. TIRS on LDCM is a 100-meter spatial resolution push-broom imager whose two spectral channels, centered near 10.8 and 12 microns, split the spectral range of the Thematic Mapper (TM) and Enhanced Thematic Mapper (ETM+) instruments 1.  

Figure 1. LDCM will launch in late 2012 carrying the TIRS Instrument. The LDCM spacecraft is  shown in orbit in the left figure. The TIRS instrument is shown on the right with the large white  earth shield panel deployed.                                                              * NASA Goddard Space Flight Center, Greenbelt MD  ** Stinger Ghaffarian Technologies Incorporated, Greenbelt MD  + Orbital Sciences Corporation, Dulles VA  ++ Vantage Systems Incorporated, Lanham MD 

Proceedings of the 41st Aerospace Mechanisms Symposium, Jet Propulsion Laboratory, May 16-18, 2012
374 The TIRS instrument has 3 mechanisms and a set of mechanism control electronics, all built at GSFC. All  of the mechanisms were developed on very compressed schedules, resulting in increased development risks. Each mechanism had unique development problems whic h were successfully overcome.   Scene Select Mechanism (SSM)  TIRS requires multi-scene calibration every orbit, so a flat scene mirror is used to switch the instrument firld of view between nadir, cold space, and a warm black body calibration target. The Scene Select Mechanism will rotate and hold the scene mirror in position within ± 9.7 µradians using closed-loop digital control. The location of the SSM within the TIRS instrument is shown in Figure 2. 

Figure 2. The SSM is located in the heart of the TIRS instrument, just above the cryogenic  telescope. The quarter section view on the left shows the location of the SSM within the TIRS  instrument structure. The figure on the right shows the SSM above the telescope. Baffles and  secondary structures have been removed for clarity. Note the close proximity of the edge of  Scene Mirror to the cold telescope, which radiatively drives the mirror temperature down.