ai-aerospace/ams_data_full_2000-2020 huggingface dataset:
309 Preliminary Assessment of Seals for Dust  Mitigation of Mechanical Components  for Lunar Surface Systems 

Irebert R. Delgado* and Michael J. Handschuh** 

Abstract   Component tests were conducted on spring-loaded Teflon seals to determine their performance in  keeping lunar simulant out of mechanical component gearbox, motor, and bearing housings. Baseline tests were run in a dry-room without simulant for 10 ,000 cycles to determine wear effects of the seal  against either anodized aluminum or stainless steel shafts. Repeat tests were conducted using lunar simulants JSC-1A and LHT-2M. Finally, tests were co nducted with and without simulant in vacuum at  ambient temperature. Preliminary results indicate mini mal seal and shaft wear through 10,000 cycles, and  more importantly, no simulant was observed to pass through the seal-shaft interface. Future endurance tests are planned at rele vant NASA Lunar Surface System architecture shaft sizes and operating  conditions.   Introduction and Background   NASA’s Constellation program currently calls for an eventual return to the moon. During the Apollo  Missions astronauts sited multiple problems with luna r dust. This included accelerated visor wear, false  instrumentation readings, seal failures, abrasion of materials and degradation of mechanisms. Lunar dust has been characterized to be very abrasive with sharp angular features and ranging in diameter from tens  to hundreds of micrometers. 1 With NASA’s current plans for an ex tended stay on the lunar surface, dust  mitigation of gearbox, motor, and bearing housings is es pecially critical. One technology currently under  development is a spring-loaded Teflon seal which could potentially be used for dust mitigation of  mechanical housings. These types of seals have seen use as dust mitigation components in the Mars  Exploration Rover Instrumentation D eployment Device as shown in Fi gure 1. The device is responsible  for the deployment, placement, and control of vari ous measurement devices including a Mossbauer  Spectrometer, Alpha Particle X-ray Spectrometer , Microscopic Imager, and Rock Abrasion Tool.2 The  Rover uses canted, spring-preloaded sliding Teflon seals manufactured by Bal Seal® to keep small dust  particles out of the rover mechanisms.3 Because of the Rover’s continue d successful long-term operation  on the Martian surface, baseline experiments were run on this type of dust seal using lunar simulant to determine their potential performance on mitigating dust in lunar mechanisms. 

                                            Figure 1  Mars Exploration Rover                       Figure 2  Bal-Seal Cross-Section       Instrument Deployment Device.                                                    * NASA Glenn Research Center, Cleveland, OH  ** Ohio State University, Columbus, OH 

Proceedings of the 40th Aerospace Mechanisms Symposium, NASA Kennedy Space Center, May 12-14, 2010  NASA/CP-2010-216272
310 Discussion 

Experimental Procedure   A series of rotating shaft tests were run against spring-loaded Teflon seals to determine their performance in preventing lunar simulant from passi ng through the seal-shaft interface. Baseline tests  without simulant were run in ambient dry-room co nditions to determine wear of the Teflon seal against  both stainless steel and anodized aluminum shafts. Then, these tests were repeated using lunar simulants JSC-1A and LHT-2M. Finally, these tests were  repeated in vacuum. Shaft rotation was constant  at 20 RPM per lunar rover technology demonstrator r equirements. The number of cycles was limited to  10,000 to determine initial feasibility of the seals. Note that with NASA’s planned  extended operations on  the lunar surface, the seals are expected to last fo r millions of cycles. Thus, endurance tests on these  seals are planned for design validation. Table 1 shows the matrix of tests completed for this feasibility study. In addition, a secondary study on the initial wear  rate of the seal was performed for 0.375-in (9.52mm), 0.75-in (19.0-mm), and 1.5-in (38-mm) diameter  seals where the seals were weighed after 1000,  3000, and 10,000 cycles.  The simulants JSC-1A and LHT-2M were synthetically manufactured such that their physical and chemical properties, as well as composition, simulate lunar regolith. JSC-1A simulates lunar regolith found in the mare, or dark regions of the lunar surf ace while LHT-2M simulates lunar regolith found in the  highland, or light regions of the lunar surface. 4 

Table 1  Bal-Seal Test Matrix 

 Test Article Description

An example of the spring-loaded Teflon seal, manufacture d by Bal Seal, is shown in Figure 2. The seal is  composed of a PTFE - Polytetrafluoroethylene ring with a U-shaped cross-sect ion. A stainless steel  canted coil-spring is inserted into the U-sh aped cross-section thereby energizing the seal.5 Seal sizes of  0.375-in (9.52-mm), 0.75-in (19.0-mm) , and 1.5-in (38-mm) inner diameter were selected to test against  either stainless steel or anodized aluminum shafts of the same diameter.       Seal IDSeal Diameter mm (in)Atm Shaft Shaft ID Simulant A22, A23 9.52 (0.375) Dry Room anodized AI SA22, SA23 none B7, B20 19 (0.75) Dry Room anodized AI SB7, B20 none C17, C18 38 (1.5) Dry Room anodized AI SC17, SC18 none B12 19 (0.75) Dry Room stainless steel T6 none A12 9.52 (0.375) Dry Room anodized AI SA51 JSC-1A B10 19 (0.75) Dry Room anodized AI SB10 JSC-1A C6 38 (1.5) Dry Room anodized AI SC51 JSC-1A B13 19 (0.75) Dry Room stainless steel T8 JSC-1AA17 9.52 (0.375) Dry Room anodized AI SA55 LHT-2MA18 9.52 (0.375) Dry Room stainless steel S-10 LHT-2M A13 9.52 (0.375) 4x10ˆ-7 torr anodized AI SA52 none B11 19 (0.75) 4x10ˆ-7 torr anodized AI SB51 noneA15 0.375 3x10ˆ-7 torr anodized AI SA53 JSC-1A A16 0.375 4x10ˆ-7 torr anodized AI SA54 LHT-2M NASA/CP-2010-216272
311 Test Equipment Description   The test set-up is composed of a test stand, motor, seal housing, seal, and shaft as shown in Figure 3.  The assembly is arranged in a vertical orientation to  allow lunar simulant to enter the seal-shaft interface  through the top of the assembly. The top of the seal  housing is designed with a coned interior to funnel  simulant towards the seal-shaft interface.       Figure 3  Rotary seal rig test set-up and view of underside of seal within seal holder.   Procedure

Prior to testing, the seals were first cleaned with Alconox, rinsed with ethanol alcohol and dried to remove any residual oils or other residue. Pre-test photos were taken of the seals and weight and inner-diameter  measurements were recorded. Pre-test photos were taken of the shafts as well, and their surface  roughness profiles were recorded. Shaft roughness averaged 0.102 ± 0.013 µm (4.020 ± 0.514 µin) for the stainless steel shafts and 0.145 ± 0.019 µm (5.691 ± 0.758 µin) for the anodized aluminum shafts. A representative profilometer plot of the surface roughness measured around the shaft circumference is shown in Figure 4.  

Figure 4  Profilometer stylus measuring anodized aluminum shaft surface   and sample shaft roughness profile.   Prior to installation, the seal and shaft were allow ed to sit in the dry-room overnight to remove any  residual moisture. The seal was then installed in the s eal holder with a slight interference fit and oriented  with the U-shaped channel opening of the seal facing do wnward. Note that simulant would be introduced  at the top of the rotary seal rig.  The seal was match-marked with resp ect to its orientation in the seal  holder to determine if the seal rotated during testing. The top and bottom pieces next to the seal holder were installed using socket head cap screws, SHCS. At tention was given to ensure that no misalignment  occurred when tightening the SHCS that bolt the three pieces of the seal holder together. The assembled  seal holder was placed on the seal cartridge holder of the rotary seal rig, Figure 3. Depending upon the  test seal size, the appropriately sized coupling adapter was used to mate the test shaft to the motor. The  test shaft was carefully inserted into the top openi ng of the seal cartridg e and gently pushed downward  until seated inside the coupling adapter. After securing  the shaft to the coupling, match-marks were made  on the shaft and coupling adapter to determine if any s lippage occurred during testing. For baseline tests  with no lunar simulant added the motor was run for 10 ,000 cycles. For tests with either JSC-1A or LHTBal-Seal Seal-Housing Match MarkCanted-Coil Spring Anodized Aluminum Shaft MotorSeal Housing NASA/CP-2010-216272
312 2M, approximately 20 ml of simulant was added to the t op of the seal cartridge prior to test start-up. In  addition, a secondary platform was attached to the sha ft just below the seal cartridge holder to contain  any simulant that may pass thro ugh the seal-shaft interface. 

After testing was completed, the seal cartridge along with  the shaft were removed from the rotary seal rig.  Observations were made as to the amount of seal and shaft wear. Post-test analyses included shaft profilometry, seal weight loss, and microscopic examination of both seal and shaft surfaces. For tests with  lunar simulant, any simulant remaining at the top of  the seal cartridge was removed prior to disassembly  and examination of the test seal and shaft. More importantly, observ ations were made to determine if  simulant had passed through the seal-shaft interface. The seal cartridge was disassembled starting from the bottom of the cartridge to determine the extent to which simulant had passed through the seal-shaft  interface, if any.  For tests in vacuum, a vacuum-rated motor was used in  place of the dry-room motor. The rotary seal rig  was placed within a bell jar capable of 10 -7 Torr, Figure 5. The chamber was then pumped down overnight  until approximately 4 x 10-7 Torr was reached. Tests were also run at 10,000 cycles and 20 RPM.  Disassembly and examination procedures of both the test seal and shaft were identical to dry-room post-test procedures.            

Figure 5  Bell Jar vacuum chamber co ntaining rotary seal rig test set-up.        Results and Discussion   For all tests run in the dry-room or in vacuum with either JSC-1A or LHT-2M, no simulant was observed to  pass through the seal-shaft interface, as shown in Figures 6a and 6b. Note that only Teflon flakes were observed on the downstream side of the seal-shaft interface. This is indicative of some seal wear as will be quantified later in the discussion. Also note that the simulant was observed to go no further than approximately half-way down the in ner diameter of the Bal-Seal. 

NASA/CP-2010-216272
314 Table 2  Wear results for Bal-Seal Tests through 10,000 Cycles 

Figure 7  Seal percent weight loss versus cycles. 

  Seal ID Environ.Seal Diameter mm (in)% W t ChangeSimulant Shaft Shaft ID % ΔRa A22 Dry Room 9.52 (0.375) -0.24 none Anod. AI SA22 2.16 A23 Dry Room 9.52 (0.375) -0.33 none Anod. AI SA23 -1.25 B7 Dry Room 19 (0.75) -0.61 none Anod. AI SB7 -5.2 B20 Dry Room 19 (0.75) -0.87 none Anod. AI SB20 -7.34 C17 Dry Room 38 (1.5) -3.00 none Anod. AI SC17 0.04 C18 Dry Room 38 (1.5) -2.63 none Anod. AI SC18 2.72 B12 Dry Room 19 (0.75) -0.13 none SS T6 5.32 A12 Dry Room 9.52 (0.375) -0.27 JSC-1A Anod. AI SA51 12.14 B10 Dry Room 19 (0.75) -0.54 JSC-1A Anod. AI SB10 13.73 C6 Dry Room 38 (1.5) -1.46 JSC-1A Anod. AI SC51 19.55 B13 Dry Room 19 (0.75) -0.66 JSC-1A SS T8 -0.39 A17 Dry Room 9.52 (0.375) -0.17 LHT-2M Anod. AI SA55 14.81 A18 Dry Room 9.52 (0.375) -0.11 LHT-2M SS S-10 30.48 A13 4x10ˆ-7 torr 9.52 (0.375) -0.11 none Anod. AI SA52 -13.87 B11 4x10ˆ-7 torr 19 (0.75) -0.07 none Anod. AI SB51 5.70 A15 3x10ˆ-7 torr 9.52 (0.375) -0.11 JSC-1A Anod. AI SA53 -8.03 A16 4x10ˆ-7 torr 9.52 (0.375) -0.22 LHT-2M Anod. AI SA54 6.90 0.00.51.01.52.02.53.0 0 2000 4000 6000 8000 10000 12000Seal % Wt Loss Cycles0.375 in. Seals0.75 in. Seals1.5 in. Seals NASA/CP-2010-216272
315 Conclusions and Recommendations 

Seals of three different diameter s were tested:  9.5 mm, 19 mm, and 38 mm (0.375 in., 0.75 in., and 1.5  in). Tests were conducted at 20 RPM up to 10,000 cyc les in dry-room and vacuum conditions using lunar  simulants JSC-1A and LHT-2M. For the tests conducted:  • No simulant was observed to pas s through the seal-shaft interface.  • A minimal amount of wear was observed on both seal and shaft. Seal weight loss was minimal with  only Teflon ‘flakes’ observed on the downstream side of the seal.  • Shaft profilometery generally show a slight deterioration in shaft surface roughness with simulant  use. Inconsistencies between surface roughness and seal weight loss require further analysis.  •  Infrared microscopic analysis of the anodized alum inum shaft surface has revealed the presence of  Teflon which is beneficial as a lubricant between the seal and shaft during operation.   Based on these results, further tests are planned incl uding effects of temperature and extended cycles in  vacuum. Efforts are also underway to integrate th e seal tests with NASA Lunar Surface Systems  architectures.   Acknowledgements   The authors wish to acknowledge Dr. Paula Dempsey,  Dr. James Gaier, Dr. Ken Street, and Daniel Valco  for their guidance in testing and analyses of the s eals. We also acknowledge Mr. Mark Hyatt of the  Exploration Technology Development Program for his s upport of this study. Finally we also acknowledge  Ed Sechkar, Scott Panko, and Frank Lam for design, as sembly, fabrication of the rotary seal test rig  hardware, and programming of the vacuum motor controller.   References 

1.  Gaier, James R., “The Effects of Lunar Dust on EVA Systems During the Apollo Missions.”  NASA/TM—2005-213610  2.  Baumgartner, Eric T., et.al., “The Mars Ex ploration Rover Instrument Positioning System.”  Proceedings of the 2005 IEEE Aerospace C onference, Big Sky, MT, March 2005.  3.  Harrington, B.D and Voorhees, C.: The Challenges of Designing the Rocker-Bogie suspension for the  Mars Exploration Rover. Proceedings of the 37th Aerospace Mechanisms Symposium, May 19-21, 2004, Johnson Space Center, Houston, Texas.  4. Sibille, L, Carpenter, P., et.al., “Lunar Rego lith Simulant Materials: Recommendations for  Standardization, Production, and Us age.”  NASA/TP—2006-214605  5. Reciprocating Seal Catalog DM-6, Bal Seal Engineering       NASA/CP-2010-216272
317 Development of the Tri-ATHL ETE Lunar Vehicle Prototype 

Matt Heverly *, Jaret Matthews*, Matt Frost* and Chris McQuin* 

Abstract 

The Tri-ATHLETE vehicle is the second generation of a wheel-on-limb vehicle being developed to support  the return of humans to the lunar surface. This paper  describes the design, assembly, and test of the TriATHLETE robotic system with a specific emphas is on the limb joint actuators. The design and  implementation of the structural components is di scussed, and a novel a nd low cost approach to  approximating flight-like cabling is also presented. T he paper concludes with a discussion of the “second  system effect” and other lessons learned as well as resu lts from a three week long field trial of the vehicle  in the Arizona desert. 

Introduction   In order to establish a continual human presence on  the Moon we must develop assisting infrastructure  that can carry cargo as well as provide mobility to the astronauts for exploration and the development of a central, but not necessarily fixed, lunar base. The Tri-ATHLETE vehicle system is a new form of two  cooperative robotic vehicles that can act individually  or physically connect together through a structural  pallet to transport and manipulate cargo.   

Figure 1.  The second generation ATHLETE rover    The basis of the ATHLETE ( All Terrain Hex Limed Extra Terrestrial Explorer) robot is the wheel-on-limb  vehicle concept. This hybrid mobility platform enables  the vehicle to traverse at high speeds across  benign terrain, as well as enabling walking, by lock ing the wheels and using them as feet, on extreme  terrain. This vehicle architecture also allows for manipulation since the vehicle is stable on three or more wheels. Non-adjacent limbs can be lifted and used to interact with the environment. A tool mechanism at                                                    * Jet Propulsion Laboratory, California In stitute of Technology, Pasadena, CA 

Proceedings of the 40th Aerospace Mechanisms Symposium, NASA Kennedy Space Center, May 12-14, 2010   NASA/CP-2010-216272
319 Lander as outlined in the Constellation Lunar architecture [3]. This also allows for a 3.5-m diameter x 8-m  long cylindrical habitat that is potentially suitable as long duration living quarters for multiple crew members on the lunar surface [4]. While the ATHLETE vehicle concept is highly  scalable, the current ½  scale was chosen due to practical limitations of tr ansporting reasonable payloads in Earth gravity. Each  Tri-ATHELTE has a mass of 720 kg and the two Tri-ATHL ETEs combine to have a payload capacity of   500 kg. For the lunar environment, the payload / vehicle mass fraction would increase drastically allowing the mass of the entire ATHLETE mobility system to be between 15 – 20% of t he payload mass that it  transports.   

Figure 3. ATHLETE vehicle standing 4-m tall, shown crawling off of a ½ scale lunar lander mockup       Figure 4. Layout of Tri-ATHLETE limb  

Joint Details   The two, three legged Tri-ATHLETE vehicles are iden tical to each other. The limbs of the Tri-ATHELTE  vehicle are also all identical, making six copies of ea ch joint. Of the seven joints in each limb there are  only two truly unique joint designs (yaw/pitch, and ro ll). While each joint has a unique gear ratio and  torque capability, the basic joint designs are all the same. Every join t has a brushless DC motor with a  power-off holding brake and an incremental encoder. Ther e is a low power motor that is used in the Hip  Yaw, Knee Roll, Ankle Pitch, and Ankle Roll Joints and a high power motor used in the Hip Pitch, Thigh Pitch, and Knee Pitch joints. Having only two unique moto rs minimizes the different interfaces necessary  and allows for a significant amount of interchangeability between joints. Each motor has a planetary  Hip Pitch  Hip Yaw Knee Pitch  Knee Roll  Ankle Pitch  Ankle Roll  Wheel Thigh Pitch  NASA/CP-2010-216272
320 gearbox with a ratio ranging between 40:1 and 81:1, depending upon the joint configuration. The output  gear of each joint is a CSG style harmonic drive. This motor / brake / planetary / harmonic actuator  configuration allows for zero backlash at the out put of the joint and gives high torque capability in a  relatively low mass configuration. This also allows fo r the use of a relatively small brake attached to the  motor shaft that can hold a substantial output torque due to the large gear reduction. Figure 5 shows a  representative joint (Hip Pitch) in the Tri-ATHLETE system. The joint is configured such that the motor /  planetary body and the flexspline of the harmonic drive are stationary with respect to the previous joint in  the kinematic chain. The wave generator of the harm onic is connected to output  of the planetary gearbox  using an involute spline. The Oldham coupling that is standard in the harmonic component set is discarded to reduce mass. The involute spline coupler bolts directly to a custom six bolt pattern in the  wave generator. While this configuration saves mass over the use of the Oldham coupler, tight positional tolerances and alignment features are required to  ensure proper concentricity  between the planetary  output shaft and the rotational axis of the harmonic.  No-load current measurement s of the joint assembly  are no larger than no-load current of the motor / pl anetary alone indicating that the combination of the  interface and the tolerances on the indicating features is sufficient to align the mechanism without binding. A non-standard bearing was required in the output of the planetary gearbox to handle the thrust  loads generated by the harmonic wave generator. The ci rcular spline of the harmonic is the output of the  actuator and is connected to a thin walled output tube . This tube houses the outer race of a pair of thin  section radial bearings that are spread across the ent ire length of the actuator give large moment carrying  capability. The structure of the output of the joint is configured as a yoke, giving large side load capability,  connecting to each side of the output tube. The yoke structure then connects to each side of the next joint in the kinematic chain.  

Figure 5.  Cross section of the Hip Pitch joint actuator   The output torque capability of each joint is based upon the ratchet torque of the harmonic. This value is used as the design load for all the structural components in the joint. The machined parts are designed with a 1.5 Factor of Safety on yield strength and a 2.0 Factor of Safety on ultimate strength above the  harmonic ratchet torque. In this way, the harmonic acts  as a mechanical fuse. Loads larger than the rated  ratchet torque will cause a non-catastrophic ratchet event. This reduces the future torque carrying  Output Bearings  Planetary  Gearbox Motor Brake  Absolute  EncoderHarmonic  DriveCable  SpoolIncremental  Encoder NASA/CP-2010-216272
323 Avionics Packaging 

Each Tri-ATHLETE contains an identical set of avioni cs, which are packaged in a “drop tray” that can be  easily lowered down from the main rover chassis sh own in Figure 7. Central to the drop tray, and  dominating its design considerations, is a set of four Valence Technologies lithium iron magnesium phosphate batteries wired in series. The resulting 78 -kg pack has a voltage of 51.2 V and a capacity of   130 Ah. The batteries are capable of  providing 300 A max current for 30 seconds. The batteries are  charged by a pair of 1 kW / 60 V power supplies whose outputs are toggled by a battery management  system (also from Valence) that monitors the battery health and state of ch arge. The drop tray also  houses a 4U cPCI card cage t hat contains four XCalibur1000  800-MHz PowerPC computers from  Extreme Engineering. These computers were chosen beca use of the existence of a space flight qualified  computer with comparable capabilities. One of the four computers is devoted to controlling the rover’s motion and communicating with ground systems while the other three computers are each devoted to stereo vision processing for four of the twelve camera pairs on the rover. Seven of these camera pairs are housed in the drop tray along with a 1-W wireless r outer and other navigation/attitude control sensors  (MEMS IMU, GPS, and high resolution tilt sensor). 

Figure 7. Rover avionics, housed in a drop tray that can be lowered from the main chassis   With the exception of the leg cabling harness and di stributed motor controllers,  the drop tray contains  nearly all of the rover’s avionics and associated wi ring. This proved to be of  great convenience and  represents a significant advancement ov er the previous chassi s design. The ability to remove the avionics  in this monolithic way allowed for easy assembly, maintenance, and repair operations. In addition, the  drop tray provided a method for bench top testing all the major systems before they  were fully integrated  into the rover.   Cabling Design   A major innovation in the Tri-ATHLETE vehicle over the original ATHLETE is the implementation of the cable harness. The original vehicle used a round wire harness that was entirely external to the limb. This  created several “chopping” points where the cable c ould be cut by structural links passing over each  other. For the second generation of the robot, the cable harness is completely internal to the limb structure. This is im plemented using essentially a round wi re harness imbedded in a flat silicone  extrusion. This technique, developed by Cicoil ™, allows for the use of any combination of round wires,  packaged in a flexible, flat ribbon arrangement. T he harness is routed down along the limb and an  NASA/CP-2010-216272