ai-aerospace/ams_data_full_2000-2020 huggingface dataset:
The specific inlet conditions for the compressor had to be narrowed down to a single point for sizing and  testing purposes. The specific design condition was for a nominal inlet condition of 20°C and 7.0 torr gas,  comprising ~95% carbon dioxide, 3% nitrogen and 2% argon. The compressor had to be able to handle a  gas inlet temperature range of -45°C to +55°C. Since the original requirements document was created, the  inlet conditions have since changed to more accurately reflect the landing site of the rover. Instead of the  inlet pressure being 7 torr, it will more likely be around 5 torr for the inlet pressure.   The original outlet conditions required were to compress the gas up to 760 torr, but further testing of the  SOXE system showed that lower outlet pressures were permissible without degradation to health of the  system.  Development Plan  Prototype  For the MOXIE operating conditions, an ideal volume ratio of 44.26 is required, which is not practical for  any scroll machine, let alone in the footprint requirement imposed by JPL. Just for reference, if the design  utilized a volume ratio of 44.26, it would require approximately 29 scroll wraps and the footprint would  increase by 465 millimeters. Instead of using a 44.26 volume ratio, the design again fit the largest volume  ratio within the given 95-mm footprint.   The 5.82 volume ratio results in a pressure-volume (P-V) curve as shown in Figure 4. The purple curve is  the ideal P-V curve for the operating pressures required. The compression process starts at the far right of  the curve, when the scrolls seal off the suction volume pockets. As the scroll rotates, the volume of the gas  pocket decreases, traveling right to left along the purple line. If the compressor were to have a volume ratio  of 44.26, the process would follow along the purple line all the way to the upper left corner of the curve until  it reached its desired outlet pressure of 760 torr, when the scroll pockets open. Instead, the volume ratio  was decreased to 5.8 and the compression process jumps from the ideal compression process around 0.18  on the x-axis.  The scroll compressor then relies on a downstream orifice to continue the compression process and reach  the desired outlet pressure. The additional work required to do the compression downstream is represented  by the pink area in the P-V curve in Figure 4. The total compression power for the MOXIE scroll compressor  is the purple area and the pink area of the P-V curve added together.  Figure 4. MOXIE Compressor P-V Curve NASA/CP—2018-219887 4
Air Squared first provided a prototype compressor to JPL to permit demonstration of proof of concept. This  prototype design was based on previous Air Squared vacuum pumps, which also required a high built-in  volume ratio. The volume ratio settled on (5.84) was the largest volume ratio that could fit into the given  105-mm width. Air Squared used its patented delayed porting design to get this high of a volume ratio in such a small footprint. Delayed porting (Figure 5) is a small radius towards the end of the scroll involute which delays when the compression pocket opens by 180°. This additional 180° of rotation significantly increases the compression ratio of the compressor, which improves the performance for this application. Figure 5. Delayed Porting Figure 6. Prototype Scroll Compressor (P11H024A-A01) 105 mm150mm 175 mmMounting FaceDelayed Porting Outlet NASA/CP—2018-219887 5
Another unique design feature that the prototype scroll compressor utilized was a flat fixed scroll plate with  the outlet routed through the base and pointing backwards towards the motor. The reason for this design  feature was primarily due to the mounting requirements given by JPL and the location of the MOXIE  experiment on the rover. The compressor will be mounted face down on the bottom of the MOXIE baseplate,  which will be acting as the heat sink for the compressor. It wasn’t possible to route plumbing through the  baseplate of the MOXIE system, so instead the fixed scroll routes the outlet flow towards the back of the  unit.   One discovery during testing of the prototype at JPL was the pressure oscillations that naturally occur in a  scroll compressor. It was feared that these may be detrimental to the SOXE, as it had never been tested  with a pulsating input flow. To reduce these pressure oscillations, a check valve and viscous flow control  device were inserted downstream of the compressor. These additional components not only decreased the  pressure pulsations, they also help prevent backflow through the compressor.   Another potential issue discovered with the prototype scroll compressor was that the bearing grease  seemed to be pulled through the bearing shields during pressure cycles, as shown in Figure 7. If bearing  grease were to enter the operating fluid flow, it could be detrimental to the SOXE stack. The thermal  decomposition of the grease would deactivate the catalytic active sites on the cathode. It has been  estimated that approximately 100 micrograms of halogenated compounds would destroy the stack,  although the number has a lot of uncertainty. Certain design changes would be implemented on the flight  unit to try and prevent this from becoming a major issue.   Figure 7. Prototype Idler Bearing Grease Loss  The prototype was inserted into the MOXIE testbed at JPL, where it was able to meet the flow and pressure  requirements desired by JPL. Succesful testing of the prototype unit led to a flight unit development  agreement between JPL and Air Squared.   Bearing Grease  NASA/CP—2018-219887 6
Flight Design and Development  Scroll Geometry With the validity of the MOXIE architecture firmly established, flight design proceeded with tighter restriction  on the compressor’s footprint, mass, and power. With 10 millimeters less to work with on the width of the  compressor (95 mm), the involute needed to be resized from the prototype (105-mm width).   The decreased width of the compressor footprint drove the design to utilize a smaller involute pitch, which  decreased the suction volume of the compressor. Instead of keeping the same ratio of involute height to  involute spacing as the prototype, the height of the unit was increased to provide more margin on the  minimum mass flow rate requirement. The increased height added risk to the machining of the scrolls, but  after the first machining iteration, it was found not to be a concern.   Figure 8 shows the mass flow rate estimate for the flight compressor which was done at the start of the  design process. It shows that the major consequence to the performance was a lower overall mass flow  rate, which was reduced from 114 g/hr in the prototype to 84 g/hr for the flight unit for the 7 torr, 20°C inlet  condition.  Figure 8. MOXIE Compressor Sizing  NASA/CP—2018-219887 7
Figure 9. P09H026A-A01 Scroll Geometry Bearing Configuration  The bearing configuration for the flight scroll compressor varied slightly from the prototype compressor. The  flight compressor now had to be designed for launch loads as well as normal operational loads, which  means that the bearings needed to be preloaded for launch. To accomplish this, the bearings used were a  ground, matched pair bearing set. The spacers used for ventilation, as shown in Figure 10, had to be  matched as well to ensure the preload on the bearings did not change. The spacers had cross holes put in  them as well as a groove machined in the outside to allow air trapped between the bearings to vent through  the side holes instead of pulling out bearing grease, which was an issue on the prototype.  The bearing grease used switched from a Krytox to a Braycote 600 Micronic EF, which is a better grease  for the specified operating conditions. Another important study conducted at the beginning of the flight unit  testing was a power draw study over the operating temperature of the compressor. It was found the lower  grease fill of 10% gave the most torque margin for startup torque at when the compressor is at -55°C.   Figure 10. Flight Compressor Idler Bearing Venting Configuration NASA/CP—2018-219887 8
Motor and Thermal Performance  Another major change between the prototype compressor and flight compressor is the motor. The  compressor still uses a BLDC motor, but it is a housed motor design instead of a frameless motor which  was used in the prototype. The primary reason for the motor change was to use a motor vendor that has  flight heritage, which the prototype motor did not have. The major concerns with the new motor design are  heat transfer from the motor, increased motor mass, and additional tolerance stack-up with the motor shaft  alignment to the scroll pack.   The change to a flight motor came with many tradeoffs and design changes to accommodate the motor.  First, the motor was much heavier than the prototype motor, partly due to the torque margin required, but  also due to the stainless steel housing. The stainless steel housing was used to better match up the CTE  of the motor winding and bearings to the motor housing, which helps prevents losing the fit on the stator  and the bearings. Due to the heavier mass of the motor, additional weight needed to be removed from the  compressor and housing to get closer to the mass spec. The majority of the mass was removed from the  fixed scroll base, which needed to be thicker for the cross hole to be drilled through the base.   An important design factor of these compressors is the mounting interface, which were designed to be  mounted on the flat fixed scroll face. This is beneficial when it comes to removing the heat of compression  from the fixed scroll, but it is far from optimal for removing heat from the hall sensors, which are located at  the back of the motor. With the de-rated maximum temperature of the hall sensors set to 105°C, getting the  heat out of the motor through the housing and fixed scroll into the heat sink became a challenge. It was  discovered during the testing phase that only mounting the motor to the housing and having a small  interface to transfer heat out of the motor would not result in acceptable temperatures at the hall sensor  locations. The worst case operating condition, which is a 70°C fixed scroll, resulted in a hall sensor  temperature of approximately 135°C. Figure 11 shows a correlated thermal model of the original scroll  compressor, without a motor sleeve, which has a temperature differential across the motor alone of 37.2°C.  This is greater than the allowable temperature differential across the entire compressor, which led to the  housing-motor-sleeve design.   Figure 11. MOXIE Compressor Thermal Model  NASA/CP—2018-219887 9
The heat transfer concern is being mitigated by changing the housing design to use an aluminum thermal  sleeve, shown in Figure 12, which is bonded to both the motor housing using Nusil. This sleeve improved  the thermal path for the hottest part of the system, the stator, and better conducts it down through the  housing to the fixed scroll. The steady state temperature gradient across the compressor for the hot  operating condition (70°C) was ~35°C, or a decrease of 30°C from the non-sleeved motor.   Figure 12. Motor Thermal Sleeve (Original on Left, Updated on Right)  Compressor Performance  The scroll compressor was still tested at the original 7-torr design condition and the results are shown in  Figure 13. The prototype performance very clearly informed the performance estimates of the flight unit  (shown in blue) and even though the flight unit was scaled down, the efficiency estimates from the prototype  were still accurate, as the actual performance shows in yellow.   Figure 13. Performance Estimate vs Actual Motor Thermal Sleeve Hall SensorsOriginal Heat Path Hall Sensors NASA/CP—2018-219887 10
As the compressor design matured, so did the proposed landing sites for the rover. The landing site has  been narrowed down and there is now a better understanding of the inlet conditions the compressor will  see. A more likely inlet pressure for the compressor is now 5 torr, instead of the original 7-torr inlet pressure  that Air Squared designed around. This clearly will influence the mass flow rate of the unit, since the density  of the inlet fluid has been reduced by ~28%.   The compressor is still capable of reaching the minimum requirement of 60 g/hr, even with the 5-torr inlet  pressure. A complete flow curve can be seen in Figure 14 for the 5-torr inlet condition. At 3500 RPM, the  nominal speed for the compressor, it can produce ~63 g/hr of CO 2 flow at a discharge pressure of   ~660 torr.   Figure 14. Flight Unit Performance Curve - 5 Torr Inlet Conclusion  The major challenges that arose during the design and testing process was the heat removal and grease  retention. The heat removal from the motor was improved by ~30°C by adding an extension to the aluminum  compressor housing to increase the conductance from the motor housing to the MOXIE baseplate, or heat  sink. The grease loss concern was mitigated, and no significant grease has been detected on the  compressor outlet by providing vent holes for all trapped air pockets between bearings to vent during  pressure cycles.   The flight compressors have survived shock, random vibration, thermal cycling and a qualification unit  survived an 18-hour life test. The flight units will soon be delivered to JPL for integration into the MOXIE  system, which has an estimated launch of summer 2020.   NASA/CP—2018-219887 11
Lessons learned with using a Capstan Mechanism as a driving mechanism for linear feed and deployment stages. Proper isolation of temperature sensors to improve the measurement response of sensors embedded in the mechanism. Simplified Slip-ring design to help reduce production time and cost. Materials selection and design of a passive brush system to assist with sample transfer of regolith. Manufacturing approach for brazing Tungsten Carbide (WC) to Titanium (Ti). Replacement of the drill’s cable harnesses with three unique Flexible Printed Circuits (FPC) to reduce system mass and increase reliability. Capstan Mechanism  A capstan mechanism uses a driven spool and a  cable pulley system with two independent cables.  In the case of the TRIDENT linear feed  mechanism, one of the cables routes to the top  side of a carriage on a guide rail system and the  other cable routes to the bottom side of this  carriage. This allows the capstan mechanism to  apply force to the carriage in both directions. For  instance, when the spool rotates clockwise, the  cable attached to the bottom of the carriage will  be in tension and pull on the carriage. In this case,  the bottom cable gets wound on the spool while  the top cable is being unwound. When the spool  rotates in the counter-clockwise direction, the  opposite is true. In this case, the cable attached  to the top of the carriage is in tension and this  cable becomes wound on the spool while the  bottom cable is unwound.  In space flight applications, such as a drill, where  a relatively long linear feed stage is required, the  capstan feed approach offers some key benefits  over other drive mechanisms such as screws and  belt drives. The primary benefit is significant  mass savings. Figure 1. TRIDENT Drill and Deployment  System.  When compared to screw mechanisms, the cable pulley type mechanism is also much more robust to dusty  environments; can more easily accommodate coefficient of thermal expansion mismatches between the  rail and the driving mechanism (i.e. screw or cable); and because the mechanism is not as stiff, the capstan  mechanism attenuates more of the vibrations from the percussive drill that are traveling back towards the  rover. When compared to belt systems, the capstan offers better material options for cryogenic and vacuum  environments and because there are two cables fixed at both ends (one to the carriage and one to the  spool), cable slippage is not a possible failure.  The TRIDENT Engineering Test Unit (ETU) is a third generation one-meter class drill that utilizes the  capstan for driving linear feed and/or deployment stages. Previous drill designs (IceBreaker [4] and LITA  Drill [5]) used a stainless-steel wire rope cable to drive the carriages. These mechanisms worked quite well  NASA/CP—2018-219887 14
for both of these drilling systems. On the LITA Drill, there were some fatigue issues with the initial cable  selection as the bend radius of the pulleys in the drive system were too small for the cable. However,  switching to a more flexible wire arrangement (7 x 19 cable) eliminated any further failures due to fatigue.  For the TRIDENT ETU, however, a new challenge existed: selecting a drive cable that could survive not  only the vacuum environment, but also the cryogenic environment that can get as low as 40K. After a  preliminary study, Technora – a para-aramid fiber cable produced by Teijin Aramid - was selected as the  drive cable. This material offered great tensile strength properties and a broad operating temperature range.  A PTFE coating on this cable is used to provide some protection against UV degradation.   On paper, the Technora 600 cable appeared to  be an excellent choice to replace the stainlesssteel wire rope: a 2-mm-diameter cable was rated  to 2667 N (600 lbf) of static tension compared to  the 2001 N (450 lbf) rating of the wire rope, and  both had comparable allowable bend radii. In  testing, however, there were some subtle issues  that made this type of cable a relatively poor  choice for the capstan mechanism. A “finger trap”  loop splice at the ends of the cable (Figure 1)  created slight diametrical growth and increased  fatigue susceptibility within the last 10 cm of the  cable. This region is still nearly as flexible as the  rest of the cable and was originally designed to  pass over the end pulleys to minimize the length  of the deployment structures (z-axes). However,  this resulted in an early-life cable failure at the  end of the spliced area. Instead, this region of  each cable should be treated as a “dead zone”  (length at the end of a cable that cannot be  passed over a pulley), and in fact this length is  much larger for Technora than that created by  crimped terminations used for wire rope. Fraying  and diametrical growth are also observed in all  cable areas subjected to fatigue or abrasion: the  spool, the pulleys, and the tensioner mechanisms  in particular (Figure 2). Figure 1. Technora cable termination and spool. For the cables to wind neatly onto the capstan spool, it works well to use a single row and let the cables  route side by side. When using steel cables at low rotational speeds, this is not a problem because of the  relative hardness of the wire rope. With the Technora cable (and other fibrous cables) this is far from ideal,  however, as the rubbing friction between two cables in contact increases the rate of fraying. The Teijin  manufacturer data sheet reports that Technora can abrade up to 219% faster when moving in contact with  itself than it does when moving over a metal surface [6]. Once fraying begins, the damage to the cable is  amplified as the fibers that fray can get trapped under the wrap on the spool that is adjacent to the frayed  portion of the cable. When the cable is being spooled out, these trapped fibers are pulled on with the full  force of cable tension. This type of behavior significantly reduces the life of the cable.  NASA/CP—2018-219887 15
Figure 2. Fraying observed at various cable bend locations throughout the capstan mechanism ( left –  spool; center  – tensioner; right – pulley). In-house testing was conducted to better characterize the expected bending fatigue degradation of the  cables. With a minimum mission operational life of eight holes to a depth of one meter each (and an  expected “bite” size of 10 cm), the absolute minimum capstan system life is 80 up-and-down Z-stage moves,  excluding any factor of safety. A representative load test put different small sections of cable under  maximum expected tension (310 N, or 11.6% of the Technora static load rating) and cycled that section up  and down over a single 90° pulley bend. When the section tested included part of the loop end termination,  cables were found to only survive around 110 cycles and always failed at the end of the splice. When tested  in an otherwise featureless middle section, cables could survive more than 250 cycles. No testing was  conducted to characterize the wear rate on the spool, but it was believed to have a worse cycle life  expectancy than the cable middle segments on the pulleys due to the increased Technora-on-Technora  abrasion consideration.  With the failure of the Technora cable, an additional study on rope or cable options was conducted. Again,  the most important first consideration is that the cable maintains flexibility in cryogenic environments.  Though more tests are required, the cable down selection process thus far has indicated that the titanium  wire rope cable is the frontrunner for replacement. The titanium wire rope provides similar fatigue  performance characteristics as the stainless-steel wire rope but maintains its ductility properties better at  cryogenic temperatures.  NASA/CP—2018-219887 16
Figure 3. Cable fatigue characterization test results ( top – Technora 600 cable with a featureless middle  section cycled up and down 100 times over a 90° pulley bend; bottom  – the same cable after continued  testing to failure). Using the TRIDENT Drill for Temperature Measurements  Obtaining subsurface regolith  measurements on the Moon is  difficult because of the low thermal  conductivity of the regolith. The  thermal conductivity of the lunar  regolith (~0.01 W/mK) (Figure 4) is  expected to be a few orders of  magnitude lower than thermal  conductivity of a titanium drill  auger (~6.7 W/mK). Of course,  variables such as relative water /  ice concentrations in the regolith,  among other things (particle size,  concentration, density) can affect  the thermal conductivity of the  regolith. Relatively speaking, the  baseline concept of drilling  operations only allows for a short  time duration for temperature  measurement, as the drill string is  only ever in contact with regolith at  depth for a few minutes at a time.  This makes the task of measuring the subsurface lunar temperature even more challenging. Two design  concepts have been tested that involve embedding a Resistive Temperature Detector (RTD) in the drill  string. The first concept had the RTD embedded in the center of the drill bit (near the tungsten carbide  cutter - Figure 5, top). The RTD was thermally bonded to the metal structure (in this case 17-4 Stainless  Figure 4. Thermal Conductivity of JSC-1A vs. Atmospheric  pressure [7].  NASA/CP—2018-219887 17