ai-aerospace/ams_data_full_2000-2020 huggingface dataset:
55 • The same number of load direction changes as for a life test at nominal velocity  had to be  performed. This implied that that a dynamic profile  with one acceleration phase and one  deceleration phase per revolution had to be performed.   • The acceleration/deceleration during the life test  was kept the same as  in the nominal scan profile  in order not to over -stress the cages of the bearings   • The ratio between constant velocity phases and acceleration/deceleration phases  was kept  constant  (approx imately  69% constant speed, 15.5% acceleration, 15.5% deceleration)   • Controller capability and power availability had to be given for  the accelerated test case (the peak  power consumption of the profile was optimized by inverting the acceleration and dec eleration  phases)  

Figure 13. Scanner life test profile for acceleration factor 3  

Life test results   Both the Scanner and the Derotator life tests are still running. The environmental campaigns (vibration,  shock, thermal -vacuum c ycling) have been finalized  successfully.  

The life tests are continuously monitored automatically  by EGSE and the set current is recorded. Once per  day, the accelerated profile is interrupted  and current measurements are performed at pre- defined constant  speeds. Motor current is used as an indicator of bearing friction. The motor current during the first 55 million  revolutions of the Scanner life test is shown in Figure 14. The low-speed motor current slightly decreased  during the initial phase of the test. This is attributed to a smoothening of the bearing internal surfaces. The  motor set current readings became noisier during summer (approx imately  day 230 onwards) . This behavior  is attributed  to the fact that – besides bearing heating – no active temperature control of the TV chamber is  performed. Temperature variations of the facility, whic h are more pronounced during summer when the air  conditioning is working with high effort to keep the room temperature within the specified limits, directly  impact the temperature in the TV chamber.  

The spikes at the beginning of the test and around day 90 are related to automatic switch- offs of the setup.  The industrial encoder that is used  for the life tests instead of the flight encoders provided erroneous  position feedback to the control loop, which reacted by  increasing the current beyond the limit defined for  automatic switch- off. The initial root -cause analysis at the beginning of the test came to the conclusion that  the analog- to-digital converter of the encoder, which had been placed inside the TV chamber,  was not  vacuum- compatible. After replacing the original analog- to-digital converter and mounting it outside the  chamber, the test was re- started and the position spike issue seemed to be solved. When it returned around  day 90, a more thorough assessment had to be performed.  After initial assembly of the mechanism, the  industrial encoder had been calibrated using its auto- calibration & reference search feature and its 0100200300400500600700 0 0.25 0.5 0.75 1 1.25 1.5 1.75 2Velocity [deg/s] Time [s]Nominal  profileLife test  profile
56 performance had been demonstrated in a test with an external reference encoder. The encoder stator had  then been temporarily dismounted for the vibration test. After re- integration, encoder performance was  again demonstrated using an external reference encoder. Subsequently, TV testing was started and the  mechanism including encoder was operate d successfully under thermal -vacuum conditions. No position  spikes were observed in any of the tests. The root -cause analysis finally led to the conclusion that the  industrial encoder should have been re- calibrated using its auto- calibration & reference se arch feature after  the second integration to the mechanism. AIT personnel had integrated the encoder stator twice in almost  identical positions and orientations, i.e. within some microns; this is why the encoder worked at all after the  vibration test. However, the missing proper  calibration led to a spontaneous loss of reference from time to  time, which caused the position spikes and, consequently, triggered the overcurrent protection. During the  encoder performance test after vibration and during the TV test, this issue had – for either good or bad luck  – simply not occurred. Based on this assessment, the TV chamber was opened and the industrial encoder  was re- calibrated.  The encoder spikes did not re- appear.  

Figure 14. Scanner motor current during first 55  million revolutions  

Summary and Conclusion  

The METimage Scanner and Derotator mechanism designs were established by modifying a standard drive  unit design to project -specific requirement s. Breadboard-  and component -level testing was performed in  order to reduce the design risk at an early stage. Flight -representative life test models (except for the  industrial encoder) were built up. Environmental testing of the life test models has been completed  Set current at accelerated velocity and scan profile Measured current at constant velocity (CV) 6 deg/s 156 deg/s 480 deg/s 642 deg/s
57 successfully. The accelerated life tests of both models under thermal -vacuum conditions have been started  and are ongoing.  

The following major lessons learned have been derived from the project so far :  • Despite efforts to minimize the development and qualification activities,  significant modifications  of  the standard drive unit design were required  in order to meet  the specific requirements  of the  Scanner and the Derotator, respectively.   • The tribological performance (friction and lifetime) of 20- year-old Maplub SH051 -a grease was  demonstrated to be no worse than that of a freshly manufactured grease batch.   • The actual outgassing behavior of the used batch of Nye 2001a is far better than expected and,  except for the CVCM, was in the same order of magnitude as the one of Fomblin Z25.   • Design for long- term storage comes with very specific cha llenges concerning material choice.  For  programmatic reasons, SmCo magnets were selected instead of  the technically more feasible  NdFeB  ones. A secondary moisture barrier had to be added on top of the traditional Nickel plating  on the motor stator in order to guarantee corrosion resistance at all locations.   • Representative accelerated life testing of fluid lubricated bearings  is (still) a major challenge.  The  widely accepted approach of maintaining oil film thickness by increasing the test temperature i s  considered conservative but comes with uncertainties (oil film thickness calculation, lubrication  regime assessment , validity for greased bearings). Dynamic scan profiles require particular  attention in order to ensure that the cages are stressed in a re presentative manner  during the life  test.  • A surface functionalization process that increases friction between adjacent parts was implemented  to prevent bolt slippage. The qualification of this process yielded excellent results: the friction  coefficient for  laser -treated surfaces of titanium on titanium (Ti6Al4V) is >0.5, the one for titanium  on aluminum 7075 is >0.8.  

Acknowledgements  

The work described was performed on behalf of the German Space Administration with funds from the  German Federal Ministry of Transport and Digital Infrastructure and co- funded by EUMETSAT under DLR  Contract No. 50EW1521.
58 References  

1.  Wallner, O., Ergenzinger, K.  and Schmülling, F.  “METimage Instrument Development Status .”  Proceedings of ICSO  2018  (2018)   2.  Buttery, M.  “An Evaluation of Liquid, Solid, and Grease Lubricants for Space Mechanisms Using a  Spiral Orbit Tribometer .” Proceedings of the 40th Aerospace Mechanisms Symposium (2010)   3.  ESR Technology Ltd.   “ESA Space Tribology Handbook ”, 5th edition (2013)   4.  Süss , M., Strobel , V., Zapata, A.   “Black functionalised surfaces by using laser technologies .”  European Congress on Advanced Materials and Processes  (2017)   5.  Buttery, M.  “GEN -ESTL -TM-0215 Airbus MetOp- SG Grease Assessment.”  Issue 2 (2016)   6.  de Heij, P.  “TEC-QTE-8772 Dynamic outgassing test and vapour pressure measurement of Nye  2001a vacuum lubricant oil – test report .” Issue 1  (2016)   7.  Busquet, M., et al   “Space grease tribological behavior for reformulation: numerical and experimental  investigations .” Proceedings of ESMATS 2017 (2017)   8.  Lewis, S.  “ESTL/TM/199  Guidelines for Accelerated Testing of Liquid Lubricated Mechanisms .” (1997)   9.  Lewis, S., et al  “Accelerated Testing of Tribological Components -  Uncertainties and Solutions .“  Proceedings of the 44th Aerospace Mechanisms Symposium (2018)   10. Orlandi, M.  "QM 5633 VBQC -  Kinetic Outgassing of Fomblin Z25 -  Test Report." Issue 1, revision 1  (2009)   11. Süss,  M., Strobel , V.  “Laser Surface Modification – An E2E success story for cost efficient bonding  preparation. ” Deutscher Luft - und Raumfahrtkongress  (2018)
59 Challenges of the D evelopment of a C ompliant Focus Mechanism Submitted to  the Harsh Martian  Environment for the ExoMars  Rover Mission  

A. Verhaeghe *, G. Perruchoud , P. Schwab* *, M. Gumy *, J. Rouvinet * and L. Kiener * 

    Abstract  

The CLUPI ( CLose- UP Imager) instrument is a high -resolution camera mounted onto the Drill of the  ExoMars 2020 rover mission carried out by the European Space Agency (ESA) and Roscosmos. The  CLUPI development is under the responsibility of Thales Alenia Space Switzerl and whereas the Principal  Investigator is Dr. Jean- Luc Josset from the Space Exploration Institute.  

For the development CLUPI instrument, the CSEM developed and delivered three models of a flexurebased Focus Mechanism. The CLUPI Focus Mechanism (CFM) des ign utilizes flexure guides to allow very  accurate frictionless adjustment of the focal distance of the imager. Such design must also comply with  very stringent requirement from the ExoMars mission, especially regarding the low Martian temperatures  and the launch/landing load environment .  

This article presents the three main challenges encountered during the development of the mechanism and  how these were addressed: resilience, performance and reliability. This article then draw s the lessons  learnt  from t his development  including potential design improvements for a similar design and general rules to  applicable to any development involving compliant mechanism.  

Introduction  

ExoMars 2020 is an ESA -Roscosmos led mission which will investigate the presence of past and present  life on Mars. Equipped with a drill and a chemical analysis laboratory, it will be the first mission to sample  and analyses  the Martian underground down to two meters in depth.  

Figure 1: CLUPI Instrument mounted onto the ExoMars flight model  

* CSEM SA, Neuchâtel, Switzerland ;  antoine.verhaeghe@csem.ch  

Proceedings of the 45th Aerospace Mechanisms Symposium, NASA Johnson Space Center, 2020
60 The Close- up Imager (CLUPI) on board the ExoMars 2020 mission is a high- resolution camera with the  primary objective of providing close -up images of the collected samples before their chemical analysis.  Detailed information on the CLUP I instrument was provided in a previous publication [ 1]. The instrument is  developed by Thales Alenia Space Switzerland, under the PI -ship of Pr. Josset of the Space Exploration  Institute. It is equipped with a focus mechanism which extends it capabilities and allows to acquire highresolution images of the surrounding geological environment.  

    Figure 2: CLUPI Focus mechanism FM              Figure 3: CLUPI Instrument FM  

The CLUPI Focus Mechanism is developed by CSEM based on flexible structure technology and the use  of Off- The-Shelf components. This approach was chosen to meet the tight schedule constrains of the  mission without a significant impact on the mechanism reliability.  

Mechanism Design  

The CLUPI Focus Mechanism illustrated in Figure 4, needs to accurately position a mobile set of lenses  with respect to a fixed one. It maintains the alignment between the two optics while adjusting  the distance  between them to change the instrument’s focal length. The main requirements driving the design of the  mechanism are:   • mass under 220 g  • operational stroke from - 4.3 mm to +4.3 mm  • concentricity and co- alignment at reference position better than 50 µm and 0.1 de gree  • stability of the concentricity and co- alignment during operation better than 20 µm and 0.1 degree   • compatibility with the ExoMars environment  

Figure 4. CLUPI Focus mechanism design overview
61   Flexible Guiding Structure   The guiding function of the mechanism is carried out using a flexible structure system composed of three  flexible guide  plates.  

Each plate is a deformable parallelogram, having four flexure blades acting as joints as illustrated by  Figure 5. By combining these three plates in an equilateral prism as illustrated in Figure 6, quasi -isostatic  linear guiding is obtained. This guiding allows linear movement along the Z axis (optical axis) and blocks  all other degrees of freedom. The obtained guiding has a non- linear rigidity along the main axis which  increases from 34  N/m at rest position to 95 N/m at operational end- of-stroke  

The displacement stroke is limited by flexible end- stops made of a stack of two blades. These end- stop  contact points are at ±4.6 mm an d they dampen excessive displacement of the mobile stage up to ±5.0 mm. 

 Figure 5. Flexible guiding plate         Figure 6: Mechanism guiding structure  

COTS Actuator and S ensor   In order to drive the mechanism, compact and contactless solutions have been implemented. To save costs  and development time, off -the-shelf components were chosen:   • a Voice -Coil Motor from Moticont   • a Linear Variable Differential Transformer (LVDT) sensor f rom Singer Instrument & Control Ltd  

Figure 7: LVDT sensor & voice- coil actuator
62 Launch Locking System   To prevent mechanism movement and to avoid excessive stresses in the guiding flexure blades during the  launch and Mars landing phases of the mission, a launch locking system has been implemented. This  system is illustrated in Figure  8.  

The launch lock consists of a non- explosive actuator (NEA)  from Eaton  actuating  a grooved ring and locking  stages. In the locked position, the locking stages k eep the mobile stage under pressure exerted by the ring.  When the NEA is released, it pulls on the grooved ring which rotates , aligning recesses in the ring above  each of the three locking stages . Theses recesses allow the locking stage to retract thus freeing the mobile  stage.  

Figure 8: Launch locking system  

Encountered Challenges  

During the mechanism development a great variety of challenges were encountered Many issues were  resolved with solutions compatible primarily with  cost and schedule constraints. These challenges can be  classed into three main domains: resilience, performance and reliability.  

Resilience   The biggest challenge of the CLUPI CFM development was to design a mechanism resilient to the ExoMars  harsh mechan ical environment. Such an environment can be summarized by the following points:   • Random vibration up to 33  Grms   • Sinusoidal vibration up to 25 g (between 30 Hz and 100  Hz)  • Shock up to 1500 G’s  When designing flexible structure, meeting strong vibration requirement s is always a trial due to the  intentional removal of  friction and thus very low damping.  For the CLUPI development , this task was made  even more complex due  a significant change in the random  vibration specification which occurred dur ing  the detailed  design phase (after PDR). Therefore,  to meet the resilience required,  multiple iterations  of  design were needed with associated analysis  and tests. The complete process is described  in Figure 9.  

The mechanical load specifications for random vibration were updated multiple times  after PDR due to  parallel progress in the overall instrument design. The specifications, their rationales  and their random  vibration levels are shown in the Table 1 .
63   Figure 9: Improving resilience throughout the complete CLUPI f ocus mechanism development  

Table 1: CFM random vibration specifications   Specification  Rationale  Random Levels   Early specification  Very early specification – Only quasi -static  -  PDR  From the last issue of instrument specification  which followed a  mission configuration change (Pasteur Payload IRD v8)  X: 19.05 Grms   Y: 16.90 Grms   Z: 18.00 Grms   1st CFM  specification  Derived from PDR: the level at CFM interface are amplified by   the instrument structure  All: 36 .04 Grms   2nd CFM  specification  Following random test failures, a formal request was issued to  reduce the instrument and CFM random levels. This led to the  issue of the 2nd CFM specification.  X: 17.56 Grms   Y: 17.86 Grms   Z: 27.49 Grms   3rd CFM  specification  Testing of the rover drill box on which the CLUPI is mounted  showed higher random level than expected. Along with  instrument design update this led to the 3rd CFM specification.  X: 19.27 Grms   Y: 18.03 Grms   Z: 33.44 Grms  

While the specification evolved, the design was improved and tested many times to e nsure that the CFM  would survive  the mission mechanical environment . These design iterations were presented in detail in [ 2]  and can be summarized in the following steps ill ustrated by Figure 10.   1. Initial design with titanium blades . This design was implemented in the first breadboard (BB).   2. Random vibration test of the breadboard (BB) led to rupture by fatigue. Investigation of the rupture  is described in [ 3].  3. Updates  of the d esign  which were implemented in the Updated Breadboard. This update includes  the following changes:   a. Change blade material to stainless steel Marval X12 which is known to have good fatigue  resilience as stated in [ 4]  b. Implementation of anti -buckling pins whi ch limit the off -plane displacements of the guiding  plate, thus limiting the stresses in the blades. These anti -buckling pins are described in [ 3]  4. Second random vibration test led to rupture due to the guiding plate arm vibration mode   5. Updates  to the design which were implemented in the Second Updated Breadboard. This update  implement ed transverse blades on the pivot flexures of the guiding plates  to shift the first guiding
64 plate mode from 830 Hz to above 2500 Hz. Random vibration analysis of this solution showed  very conservative  (low)  stresses.   6. Shock test and analysis showed higher acceleration levels on the mobile stage, which may lead  to excess ive buckling of the transverse blades.  While the test was successful , the design needed  to be improved.   7. The l ast update to  the design consist ed of a refinement to the transverse blades: shortened and  thickened to raise their bucking limit from 9 N to 30.8  N. This update was implemented on  Engineering Qualification and Flight Models (EQM/FM).   8. Successful random vibration test  on EQM.  

Figure 10: CFM guiding design update  

While m aking the CFM resilient to the launch and re- entry loads of the ExoMars mission was an  arduous  task, CSEM managed to meet the ever -increasing mechanical load requirements.  

Performance   The CLUPI instrument is a close- up imaging system. On the Martian  ground , it aims to replace the  geologist ’s eye and magnifying glass. To do so it is equipped with an opto- electronical system capable of  acquiring high resolution color images  of targeted object s from 10 cm to infinity. Furthermore, the instrument  is equipped with a processing unit able to perform z -stacking of multiple images  of the same object. This  algorithm is used to compensate the short depth of focus of the instrument: CLUPI can acquire a set of  images with various working distances  and reconstruct a completely sharp image (along with a 3D -map) of  the imaged object [ 1]. In order to fulfil these instrument  capabilities, the CLUPI Focus mechanism has  quite  stringent requirements with respect to positioning precision and guiding performance. Performance  requirement s are lis ted in Table 2 along with their  potential impact on the instrument performance:
67 The first observation was made during the EQM qualification test to random vibration loads. During this  test, it was observed that the mechanism sine sweep responses before and after a random cycle did not  match as they should according to the applicable ECSS standard. Such phenomenon usually indicates a  mechanical failure inside the tested item. After careful inspection of the mechanism , no damage was found.  Further testing and investigation revealed that , under the random vibration environment, components of the  launch lock system settled into a different position,  thus impacting the sine sweep response of the  mechanism. In the end, during the qualification test, the launch lock resetting  procedure  had to be revised.  In this update the final pressure applied onto the locking ring is along the unlocking direction. Wh en doing  so, the ring moves back about 0.05  mm and is then in more stable configuration. While this observation did  not impact the reliability of the mechanism, it illustrates how sensitive the mechanism is to the resetting  procedure.  

The second and more critical observation was made during the thermal -vacuum qualification and  acceptance test of the EQM and FM. In this test, after the first non- operational cycl e, the units  were placed   at the low operational temperature in order to perform a functionality check . At the beginning of the EQM  functional ity check, a representative current was applied through the NEA leads to actuate it and release  the launch lock. When doing so it was observed that the NEA fuse wire burned (electrical discontinuity).  However, t he launch lock did not release immediately. Following the  failure, it was decided to warm up the  thermal vacuum chamber  to prepare it for opening and investigation. During the warming up, about 1h after  the NEA firing, the EQM launch lock released as illus trated in Figure  11.  

Figure 11: Temperatures & LVDT reading during firing & release of the EQM launch lock  

The opening of the thermal vacuum chamber  was then interrupted and the chamber placed again at low  operational temperature in order to test the FM model similarly. In the case of the FM model the unlocking  occurred 74 s after firing,  and without any temperature change. Such an observation is even more complex  to interpret  since  the observed delay s are both very different and far from the expected behavior (<2 s).  Following these observations, the thermal vacuum qualification test continued without any further issue and  the unlocking failure was in vestigated.  

In order to investigate the issue, many unlocking tests at low temperature were conducted on the EQM.  The first tests performed were done with a mechanical release system instead of the fuse spool. Such test s  all had unlocking times  under 1s.  Then representative test s with fuse spools were done in order to  reproduce the issue. In those tests, similar behavior was observed with very dissimilar unlocking times. The  unlocking times  for all the troubleshooting tests are described by Table 4.