ai-aerospace/ams_data_full_2000-2020 huggingface dataset:
419   Figure 7. Damaged motor bearings from shock event  

The g- level induced by pyroshock was initially analyzed with respect to the brittle components and  solder/ strain relief only, and did not consider the spring/mass system of the bearing/rotor combination. A  design change between COTS and flight had the front bearing hard mounted in both directions, allowing  the back bearing to float, resulting in the front bear ing absorbing all of the energy of the accelerated rotor.  

The M32 motors are used in multiple locations across the Mars 2020 rover and any changes would cause  a significant project schedule impact. In most locations, the pyroshock requirement was over specified; the  hardware was requalification tested to the new pyroshock requirements. The motor, and subsequent  gearmotor tests, all passed the new pyroshock exposure and long rotational life requirements. There  remained a single location where the motor had exposure to the high pyroshock environment but with  a  short rotational life. With the project considering the risk, the solution was to use the damaged motor to  qualify a deployable device with high shock and limited motor rotational life. The qualification demonstrated  that the damaged motor could achieve the limited revolutions needed. The flight unit was delivered and the  mechanism underwent qualification at the next level of assembly. The pyrotechnic deployment was  demonstrated four times  and did not pr oduce a “coffee- grinder” motor, showing that the either the  requirement or tunable beam test was over conservative. In both solution cases, no changes to the flight  motor were made.  

Overheat ing Concern   The coil -to-coil resistance for the M32 and M20 motor s are 14 ohms and 28 ohms , respectively. Additionally,  the copper wire is wound around two PEEK stator  caps at each end of the stator lamination stack, which  provides thermal isolation from the bulk stator mass. The current driving the motor causes significant Joule  heating due to the high resistance, quickly increas ing the temperature of the low mass of copper wire  insulated from the stator, which can lead to thermal runaway. Th e overheating worsens as the current  reaches 1A.  

Due to design principle requirements to demonstrate 2× torque at operational temperature limits, hot  dynamom eter tests are the highest risk activities to perform on the ground. The M32 has an additional  thermal risk due to the detent mechanism; the larger holding torque design was roughly equal to the nominal  performance of the motor, so twice as much current is needed during slow speeds (where the rotor has no  signifcant inertia to carry it over the detent peaks) compared to fast speeds. A thermally instrumented unit  was used to create a thermal model to understand this interaction. Guidelines for operation early on in the  Mars 2020 design and qualification program were related to current input and time —a current profile could  be estimated, the heat input was analyzed to confirm no overheat condition was achieved, and then a wait  period was enforced in order to ensure the test article reached steady state.
421 Table  4. M20 motor acceptance tests 

The acceptance tests listed in Table 5 were c ompleted at the M20 gearmotor level, which included thermal  environment exposure at non- operational temperature limits of - 135°C. 

Table  5. M20 gearmotor acceptance tests 

The majority of the M32 and M20 hardware subjected to acceptance testing met performance requirements  and were accepted as meeting  flight fidelity.  

JPL accepted the risk of qualifying the M20 actuator post  delivery of the flight units due to the schedule   criticality of needing the flight hardware assembly. Tests completed by JPL to fully qualify the design  (Table 6) exposed an actua tor to the flight acceptance test (with the exception of 2 minutes  per axis random  vibration exposure) and then completed a long-duration thermal bakeout, pyroshock , and humidity  exposure. The gearmotors underwent a long- duration thermal vacuum bakeout to replicate the maximum  expected exposure the assembled mechanisms may be subjected to in order to meet Planetary Protection  requirements. The gearmotors  were also subjected to 40 hours of accelerated corrosion exposure in an  80°C/80% relative humidity environment  meant to simulate any corrosion that may occur in the gearbox  while exposed to the Earth’s atmosphere before launch. The gearmotor was then subject to the rotary life  test. A separate motor -only unit was subject ed to the thermal life test .    M20 motor Acceptance Test  Test Conditions  Acceptance Criteria   Dielectric strength  500V AC, 60Hz for 60 ±5 seconds  I < 1 mA  Insulation resistance  500V DC, for 60 ±5 seconds  R > 100  MΩ  Electrical bonding  N/A R (flange to housing) < 25  mΩ; R (flange to shaft) < 100 Ω  Run-in 60 sec CW and 60 sec CCW at 28V  N/A  Hall effect (INL) sensor  Maxon defined test  <10° INL  Torque ripple CW and CCW over one revolution  N/A (characterization only) , validates torque is produced  at all rotor positions   No load  24V, CW and CCW, 22 °C N/A (characterization only)   M20 Gearmotor Acceptance Test  Test Conditions  Acceptance Criteria   Run-in  7.5 min utes CW and 7.5 min utes CCW at 24V  N/A  Back drive torque  Ramp to 75 mN •m, CW only  Static holding torque: > 50  mN•m   No load  24V, CW only, STP  N/A (characterization only)   “Launch” random vibration   Frequency (Hz)  Protoflight/Design  Verification (Qual/PF)  Test Level   20 0.15 g2/Hz  30–90 0.5 g2/Hz  166–450 0.08 g2/Hz  2,000  0.0040 g2/Hz  Grms 10  (in all 3 axes, 1 min each)  N/A (exposure only)   Non-operational thermal  One thermal cycle from +113°C  to -135°C   Two thermal cycles from +70 °C to -135°C Restrain static holding torque   (T > 50 mN•m)   Operational thermal   Start -up sensitivity   No load   Load  Each test performed at +70°C /22°C /-70°C    CW only   24V, CW only   12V/20V/28V, CW and CCW  Gearmotor torque constant =   1620 N •m/A ± 10%
422 The rotary life qualification unit completed its life test with minimal wear, and the design completed  qualification  for flight as of June 2018. The Thermal Life test is 1.5 years in duration and was successfully  completed in December 2019.  

Table 6. M20 gearmotor qualification tests.   M20 Gearmotor Qualification Test  Test Conditions  Acceptance Criteria   Thermal vacuum bake -out 113°C ±2°C at 1.3 × 10-3 Pa for  288 hours  N/A (exposure only)   Pyroshock  Tunable beam (2 hits in 2 axes)   Frequency (Hz)  Shock Response  Spectrum ( SRS)  Level (Q = 10)   100 5 g  100–3,500  +9.0 dB/Oct.   3,500 –10,000  1,000 g    N/A (exposure only) ,  in family with no- load  characterization   Humidity exposure  40 hours of 80% humidity/+80 °C N/A (exposure only)   Rotary life test Torque bins:   0.2 N •m for 40,400 revolutions   0.3 N •m for 12,400 revolutions   0.4 N •m for 2,100 revolutions   0.5 N •m for 1,500 revolutions   Split between CW and CCW and alternating between +70°C,  +22°C, -55°C , and -70°C  In fami ly with no- load  characterization  Thermal life test  3015 thermal cycles (4.5 Mars years) broken into seasonal  profiles ( -80°C to +85°C for summer and - 115°C to +50°C for  winter)  In family with no- load  characterization at interval  checks  

Final Delivery  

For the nine flight locations of the M32 motor  across the Mars 2020 rover , a total of 37 gearmotors were  needed to meet flight, flight spare, engineering model  (EM) , and qualification units. m axon required a  minimal build to efficiently use their production and manufacturing series, which resulted in three lot builds  (EM1, FM1, FM2) totaling 99 delivered M32 units. The 99 units consisted of 10 EM1s  (3× 10 mN •m and 7×   20 mN•m), 35 FM1s  (14× 10 mN•m and 24 × 20 mN•m), and 54 FM2 s (33× 10 mNm and 21×  20 mN•m).  Similarly, the M20 gearmotor had a JPL need of 6 units but a total of 40 units were delivered. The 40 units  consisted of 16 EMs  and 24  FMs. On the FM units, the carrier s with sun  gear were made out of two different  material t ypes, with 12 units of each type delivered. An attrition of over 50% during the manufacturing and  production phases had delivery effects but the flexibility and availability of the on- site facilities allowed  maxon to recoup schedule as best as possible and phase deliveries between the lots as needed.  

Due to the surplus of flight units available, JPL was able to cherry -pick motors with the optimum  performance and flight fidelity for the 10 flight units. The residual units are logged in the JPL flight hardware  logistics program for future NASA project us e. Additionally, maxon has a catalog flight -qualified specification  for the M32 and M20 flat motors.
424 sufficient time in mission planning for solution develop ment . It was  pivotal for JPL and maxon to leverage  experts at each location to address critical problems (along with using a 9- hour time difference between  Luzern and Los Angeles to accomplish an 18- hour project work day).   

Both the M20 and M32 designs have been successfully qualified. The M32 designs were integrated into  three different, separately sourced gearboxes and passed all qualification testing. The two versions of the  M32 motor are implemented into nine locations on the Mars 2020 rover and the M20 is integrated into a  single location. Flight model hardware has been built, acceptance tested, integrated into the target  gearboxes and mechanisms , and installed on the Mars 2020 rover awaiting launch to Mars in July 2020  (Figure 8).  

Figure 8. maxon hardware installed on Mars 2020  

Acknowledgements  

The successful collaboration and delivery of the maxon flight motors to Mars 2020 would not be possible  without many individuals at JPL and maxon over the three- year effort. In addition, the authors would like to  thank their families who endured years of middle -of-the-night teleconferences and nights away on  international trips.  

The research was carried out at the Jet Propulsion Laboratory, California Institute of Technology, under a  contract with the National Aeronautics and Space Administration.
425 References   1. NASA Procedural  Requirements , NPR 8705.4 (2004),  https://nodis3.gsfc.nasa.gov/npg_img/N_PR_870 5_0004_/N_PR_8705_0004_.pdf .  2. Braun, David and Don Noon, “Long Life”  DC brush motor for use on the Mars Surveyor Program,  32nd  Aerospace Mechanisms Symposium (1998).   3. See graphic on page 2 of  Ref. [5 ], or this direct link:  https://mars.nasa.gov/imgs/2013/07/mars - exploration- timeline- 07-2016 -hpfeat2.png .  4  Office of Inspector General, NASA’s Mars 2020 Project, Report no. IG -17-009 (2017) ,  https://oig.nasa.gov/docs/IG -17-009.pdf .  5. NASA, Mars 2020 Mission Overview , available at: https://mars.nasa.gov/mars2020/mission/overview/ .    6. Szwarc, Timothy, Jonathan Parker, and Johannes Kreuser, STIG: A Two- Speed Transmission Aboard  the Mars  2020 Coring Drill, 45th Aerospace Mechanisms Symposium (2020).   7. Chrystal, Kyle, Percussion Mechanism for the Mars  2020 Coring Drill,  45th Aerospace Mechanisms  Symposium (2020).  8. Barletta, Anthony, Design and Development of a Robust Chuck Mechanism for the Mars  2020 Coring  Drill, 45th Aerospace Mechanisms Symposium (2020).   9. Silverman, Milo and Justin Lin, Mars 2020 Rover Adaptive Caching Assembly: So Many Challenges,   45th Aerospace Mechanisms Symposium (2020).   10. Grimes -York, Jesse and Ken Glazebrook, Sealing Station Mechanisms for the Mars 2020 Rover Sample  Caching Subsystem, 45th Aerospace Mechanisms Symposium (2020).   11. Brown, Adrian, Mars Science Laboratory: the technical reasons behind its delay , The Space Science  Review  (2009), http://www.thespacereview.com/article/1319/1 .  12. Phillips, Robin R., M. Palladino,  and C. Courtois , Development of Brushed and Brushless DC Motors  for use in the ExoMars Drilling and Sampling Mechanism , Aerospace Mechanisms Symposium (2012).   13. McCoubrey, R., et al., C anada’s  Suspension and Locomotion Subsystem for ExoMars 2 018, 65th  International Astronautical Congress , Toronto, Canada (2014) .  14. Herman, Jason and Kiel Davis , Evaluation of Perflour opolyether Lubricant Lifetime in the High Stress  and High Stress -Cycle Regime for Mars Applications , 39th Aerospace Mechanisms Symposium (2008) .
431 Since flexure strength and stiffness were two self -competing design constraints and overall CDPR  compliancy in mobility deployment load absorption is entirely depending on the flexure stiffness, its physical  measurement would serve invaluable to  confirm FEA predictions.  Fig. 4 show s the images of the nominal  test set -up as well as deformed shapes under both tension and compression.  

Figure 4. Fully reversed range of deflection of the p rototype flexure at (a)  max extension from tensile  load, (b) nominal unloaded state, and (c) max compression from compressive load.  

Below are the key findings from this flexure prototype characterization test:    • Flexure’s uniaxial stiffness (load vs. deflection) appeared highly linear and repeatable at both room  (illustrated in Fig. 5 ), hot, and cold temperatures  with little to no signs of hysteresis over a set of 34  tension- compression cycles   • Temperature effects are minimal – within a maximum difference of 2% of the measured flexure’s  stiffness relative to the desired target stiffness   • Residual strength test achieved a tension load of approximately 46 kN, which demonstrated a 33%  more tension load capability than predicted by linear FEA   • No signs of material yielding observed throughout the entirety of the test  

This prototype testing sufficiently demonstrated that flexure critical stress concentrations exhibited morethan- adequate strength capability. Based on this, nonlinearity in material’s elastic -plastic behavior as well  as large displacement theory were incorporated in the FEA. This updated the predicted onset of yielding in  the flexure to be approximately 71 kN, which was more than twice the initial predicted ultimate failure load,  and 40% higher than the max load reached during the residual strength charact erization test. This  warranted the final round of flexure thinning to further reduce its stiffness and mass (Fig. 6).
432  

Figure 5. 15 overlaid cycles of raw force vs. displacement data at room temperature.  

Figure 6. The evolution of the flexure design: (a) pre- Initial Design Review design, (b) proposed Detailed  Design Review design, and (c) final delivered flight design.  

The trade- off being after this final round of flexure thinning out , was that  the flexure’s lowest structural  margin became buckling critical rather than based on local stress concentrations. Buckling test of the flexure  was subsequentially performed post -DDR on a flight like design.  Test results indicated no signs of structural  buckling up to the required 1.4 test factor over the FLL presented at DDR. Resulting from this f lexure  developmental design and testing, the final design was proposed at DDR and was hardly changed for the  final delivered flight flexure configuration post-DDR.  

Deployment Mechanism  

This new  architecture of combining the compliant  component  (the flexure) and the deployable  resulted in a  redesign of the elements of the mechanism to deploy the flexure.  More often than not, it required increasing  size as well as strength of these elements . While the Mars 2020 CDPR had more components, the different  approach of the mechanism resulted in approximately 50% less mass for the whole assembly while  actuating a 10 times more massive deployable. The MSL deployment mechanism used symmetrical  torsion  spring to deploy a 150-gram linkage. The Mars 2020 deploy s the 1500- gram flexure using larger  asymmetric torsion springs, initial motion kickoff s pring, and a magnetic latching hardstop to contain it in its  deployed state. This section details the development of these elements of mechanism in the Mars 2020  CDPR as well as several findings and key lessons learned throughout the development process.
433 Torsion Springs   The main components to deploy the fl exure are the torsions springs. At the time of DDR, it was required to  show by design twice the needed torque to deploy the flexure to overcome mechanism frictional losses in  dynamic environments. For Mars 2020, each torsion spring increased in  size and strength  compared to the  MSL CDPR torsion springs . These larger torsion springs led to a design challenge of how to positively  capture the two free ends of these larger torsion springs. For MSL, the torsion spring ends  were manually   bent to route the spring wire end through a hole in the t itanium linkage . For larger Mars 2020 torsion springs,  the spring wire was fabricated out of a stronger stainless -steel  material and from spring wire nearly twice   the diameter . There was concern if one could manually bend the spring wire accurately and without  damaging the spring wire. Thus , it was chosen to have the vendor make the torsion spring with prescribed   bends at each end.  

The style of bends at the end of the torsion spring to route through the flexure was inspired from linkage  rods found in automotive engines. There are t wo 90° bends at one end of the torsion spring and are route d  through a hole in flexure. Then once the torsion spring is installed on its arbor, each torsion spring would  be positively captured. The size and position of the pass -through  holes in the f lexure and the torsion spring  bends were designed together to achieve this positive retention of torsion spring end in the flexure. In the  end, this design proved to be easy to implement and less susceptible to damaging the hardware with a  manual bending process. See Fig.  7 for photos of wire end bends of each torsion spring.  

Figure 7. Views of torsion spring ends: (a) upper torsion spring, (b) lower torsion spring and the travel  limiters are protruding through the chassis bracket clevis with set gaps to the back shelf on the flexure.  

There were design space limitations  on the lower of the two torsion springs, which resulted in an  asymmetrical design. As a torsion spring is preloaded, the effective spring diameter is decreased and the  overall spring length is increased. The lower torsion spring had a fixed height from the confined space  between the chassis bracket bottom and top deck. There was concern that the lower torsion spring could  get to o long when the flexure is in its stowed state. The upper torsion spring did not have the same height  limitations and  could have more loops and delivery more torque for the same deflection. The combination  of the two asymmetric springs surpassed the DDR requirement of 100% torque margin, roughly 66% of the  combined delivered torque came from the upper torsion spring. However, the fault case of only the lower  torsion spring acting was enough still to deploy the flexure. To verify this release redundancy capability,  various prototype and Dynamic Test Model (DTM) tests have indeed demonstrated that the mechanism  could still deploy fully without the upper, stronger, torsion spring contributing to the deployment.
436 magnetic latch work, a magnetic flux circuit with the magnet housing and flexure needs to be closed. Most  flight acceptable CRES are austenitic , with the exception of 400 series CRES , which is a ferritic CRES . It  was learned that the part  of the housing in contact with the magnet should also be the part that is closest  to the flexure when the flexure  is in contact to the hardstop. This piece of the housing should also be a  ferritic steel  to complete the  magnetic flux circuit from the magnet , through the ferritic housing component,  through the f lexure , and return to the magnet  (Fig. 10) . This is what produced the latching force. The top of  the magnet housing should be an austenitic CRES. Otherwise, the magnetic flux would short through the  top part of the magnet housing and not through the flexure. The final design called out the back part of the  magnet housing to be 410 CRES and the top part to be 303 CRES. This allowed the magnetic latching  force to flow from the magnet through the flexure.  The delivered magnetic holding force for the flight unit  was measured to be an average of 24.4 N.  

Figure 10. Cross section of the magnet in its CRES housing with the flexure against the hardstop. Also  showing the magnetic flux circuit, which produces the holding force of the magnetic latch.  

There was a lesson learned that while magnetic latches could be used, some permanent magnet materials  cannot tolerate higher temperatures. To comply with mission planetary protection requirements, this  neodymium magnet would have seen temperatures during vacuum bakeout such that the magnet could  have lost some of i ts residual magnetic strength. As a result, the magnet and its housing were excluded  from vacuum bakeout as well as parts of other non -functional tests that went over + 80°C. Planetary  protection requirements were satisfied with other methods rather than a vacuum bakeout. In future magnetic  latches, it is recommended that the magnet can withstand higher temperatures without any risk of losing  magnetic strength over the whole non- operational temperature range.  

CDPR Design Validation and Verification  (V&V)  Testing  

Since t he Mars 2020 CDPR was non- MSL heritage , this new assembly was subjected to a full design V&V  campaign that involved a series of functional and structural testing at  various lev el of build integration.  As  of January 2020, all testing was  complete d. Rather than presenting the entirety of this lengthy test  campaign, two (2) significant tests and some key lessons learned are highlighted.  For a complete summary  of the test programs  in order of execution  (including the prototype testing during the engineering  development phase) , see Table 1 and Table 2  at the end of this section . 

Engineering Model (EM) “Dirty” Testing   When the CDPR mechanism is fully actuating,  in the power decent stage  of EDL, there is a lot of Martian  regolith that could get kicked up and lodged into the numerous control led gaps within  the mechanism.  This  raised a concern that the deployment of the flexure could be hindered by  this “dirty” condition. To address