ai-aerospace/ams_data_full_2000-2020 huggingface dataset:
274 supply, the cyclone body and all other metal components contained within the secondary containment  box are connected to the electrical ground. 

Figure 10: An eductor produces a dense flow of dusty gas by creating a partial vacuum via the  Venturi effect that entrains lunar regolith feedstock material into a flow of clean dry air.   Figure 11 shows the actual testing of  the pneumatic transfer hardware during the reduced gravity flight as  Tephra was transferred pneumatically into the Re golith Discharge Container, which served as an ISRU  reactor mockup. Although nearly all of the Tephra was transferred during the reduced gravity flight, dust  adherence to the walls of the acrylic container prev ented an exact determination of the total time needed  to complete the pneumatic transfer of Tephra into the discharge container. The same hardware setup was  used to conduct the terrestrial testing discussed ne xt. The RGF results are discussed later in further  detail. 

Figure 11: Photo of the RGF experiment showing Tephra that fell from cyclone A into the Regolith  Discharge Container. The Tephra flowed vertically  upward along a convey pipe from the eductor  located at the bottom of the Supply Container (n ot visible in photo). The convey pipe had a  vertical-to-horizontal transition which connected it to the inlet of cyclone A in Fig. 9. The gas  exhaust from cyclone A became the input for the smaller cyclone B, which discharged dust into  the blue hose shown in the picture. The exhaust from cyclone B passed through HEPA filter bags  before exiting into the transparent containment box, which had a secondary HEPA filter that  allowed the filtered air to enter into the aircraft cabin.  NASA/CP-2010-216272
276 15 kg of Tephra was easily transferred pneumatical ly from the supply container in less than the ~15  minute duration in simulated lunar gravity.    When the NU-LHT-2M transfer process began under 1/ 6-g conditions, this granular material was also  transferred easily from the supply container to the discharge container despite being more compactable than Tephra. However, after ten parabolas (~250 sec), it was observed that the transfer process had stopped. Although the NU-LHT-2M was assumed to hav e been sieved, a post analysis showed that a  ~0.5-cm-diameter “rock” had become clogged in the eductor’s solids inlet which prevented regolith simulant particles from entering the eductor and bei ng entrained in the air flow. Nevertheless, it was  observed that 8.8 kg of NU-LHT-2M could be transferred ve rtically resulting in a mass transfer rate of 2.1  kg/min. The Tephra mass transfer rate is believed to be probably greater than this value given that Tephra is less dense than NU-LHT-2M.  The efficiency of our particular series cyclone filtrati on system in separating part icles, including the use of  an electrocyclone, could not be determined due to the dense flow of dusty air having overwhelmed the cyclones in 1-g and in the 1/6-g environment. The d ense flow must be factored into a re-design of the  cyclone system and 1/6-g effects must  be included in the design parameters. It is likely that different  cyclone systems will be requi red for the terrestrial 1-g system and for the lunar 1/6-g system due to the  expected higher mass transfer rate of  regolith in 1/6-g (Sullivan 1992, 1994).   A pneumatic alternative to the venturi eductor for transferring regolith is to use the jet-lift method, which was successfully demonstrated at lo w pressure (Zacny). The jet-lift method selectively excites particles  through transfer of gas momentum into the particles in the vicinity of a gas injection head. The resulting dusty gas flow is subsequently transported along a tube containing the particles as the gas attempts to ultimately escape into a low pressure environment. Experimental tests in a vacuum chamber at ~5 torr  demonstrated regolith-mass-to-gas-mass transport effici encies exceeding 1000:1 utilizing this technique.   

Figure 12: Jet-Lift Regolith Transfer method.   Finally, the compressor requirements for pneumatic regolith transfer are 5 SCFC at 20 psi (138 kPa) of gaseous hydrogen. Since the temper ature of the regolith transferr ed is approximately 700 degrees  Celsius, compressors that meet these requirements are highly specia lized and few vendors are capable  or willing to tackle this challenge. Consequently, the se arch is continuing for a compressor that is suited  to the special conditions imposed on the pneumatic re golith feed system by the ISRU reactor’s working  environment.      NASA/CP-2010-216272
280 Each wheel pair is driven by a differential which is powered by a shaft that extends vertically down  through the steering gears from a motor-gearbox mount ed on top. This gearbox is the subject of this  paper. Furthermore, each pod has its own combination of  passive and active suspension. The active part  allows for adjusting the chassis height, including lo wering it to the ground for easy egress of the  astronauts, and easier loading and unloading of cargo.   The vehicle can be operated by an on-board astronaut (Fig. 1), a remote astronaut (possibly while stationed in a lunar habitat), or remotely from ear th. Some limited autonomous operation is also likely.  The prototype has been demonstrated with an astronaut  operation station. A suited astronaut stands on a  platform which holds the operator controls, an in strument panel, and a surround railing for astronaut  safety. This entire structure, contro lled by the astronaut, rota tes to provide the full range of visibility that a  constraining suit would not otherwise easily allow.  

Figure 1.  Chariot with Astronaut Operation Pod (Dec-07) 

Figure 2.  Chariot Wheel Pods 

More recently, the Chariot has been outfitted with a pr essurized crew cabin that  occupies most of the  available mounting space. In this configuration, the vehicle is called the Lunar Electric Rover (LER). The  cabin allows for shirt-sleeve operation for up to four astronauts, and includes berthing for extended  occupations of up to 14 days (Fig. 3, 12, 14). Normal access to the cabin is through a door on one side of  NASA/CP-2010-216272
281 the cabin. Using Chariot’s sideways crab motion abili ty, the cabin door can be ‘docked’ to a mating door,  thus maintaining a pressurized environment.    There are two externally mounted luna r walking suits located on the back side of the cabin. By way of  clever “suit ports”, an astronaut can enter and occupy  either suit from inside the cabin, and become  sealed-off and disconnected from the cabin, all with no help. Following the completion of their extraterrestrial activities, the astronaut can return to the cabin using the reverse pr ocedure. This system is  intended to minimize the invasive moon regolith from entering the pressurized volume.   

Figure 3.   LER Starboard Center Wheel Pod with New Gearbox      The Original Gearbox 

As already mentioned, one gearbox on each pod dr ives the differential that turns both wheels. The  original gearbox was designed and built around readily  purchased off-the-shelf gears, bearings, electric  clutches, and shafting. The design utilizes a total of 12  gears, 17 bearings, and 7 different shafts (Fig. 4).  The design also incorporates an electric brake that  prevents motion when it is de-energized, and an RPM  encoder. Two selectable speed ratios are provided. The low speed provides a 16:1 gear reduction for a  vehicle speed of about 4.8 km/hr (3 mph), while the hi gh speed provides a 4:1 gear reduction for a vehicle  top speed of 19 km/hr (12 mph). The speed selection is accomplished with two electric clutches.    Two concurrently operating electric motors drive the first large gear, both to generate adequate power and for motor redundancy (Fig. 5). The motors used are rated for 300 VDC, and each one draws up to 8 amps to produce 2.1 kW (2.8 hp) at  peak efficiency. The re sulting gearbox works very well, and is still in  use on the LER prototypes today.        NASA/CP-2010-216272
283 Second Generation Gearbox, Concept 

The design goal for the new gearbox was to provide a configuration which would be better suited to send  to the moon. By using helicopter-grade gearing, the new design would have fewer gears, and be both  smaller and lighter.    Drive Train Design

Reducing the number of gears inherently requires that  each meshing gear pair accomplish more of a gear  reduction than might be possible or practical when c onstrained by the use of catalog gears. The original  design used off-the-shelf gears to meet the time and budg et constraints for the fabrication of the original  gearboxes. For the second generation gearbox, after cons idering many different gearing configurations,  the simplest approach that still met the speed ra nge criteria was adopted. Custom-made high-strength  gears allowed the use of fewer (7 instead of 12) and light er gears. Also, the gear train consists of just two  parallel axes, resulting in fewer shafts (3 instead of 7) and bearings (13 instead of 17). Two electric  clutches, similar to those in the original gearbox, were mounted in series on the same shaft (Fig. 6). This also allows the gear casing profile to be reduced in size, thus contributing to making a more compact and lighter final product.   

Figure 6.  Chariot Gearbox Cross-Section Labeled Illustration         NASA/CP-2010-216272
286 .   Figure 8.  New Gearbox on Bench 

Figure 9.  Rapid Prototype Gears in Clear RP Lid 

 Rapid prototypes (RP) were made of all the casing parts and gears before the design was finalized (Fig.  9) to allow for practicing the assembly sequence. The transparent casing parts were helpful during the trial assemblies. Some important changes to the final case design details resulted, and a detailed assembly procedure was initiated.    NASA/CP-2010-216272
287 Grease Retention Shrouds   Enclosures around the gears were added to help hold the applied grease close to the teeth to extend the  re-grease service lengths. Referred to as grease re tention shrouds (GRS), similar enclosures have been  used in certain commercial products such as heavy- duty hand drills. Of the three gear meshing planes of  the design, a GRS was designed for two of them; the upper area where the two motor pinions engage gear #1 (Fig. 10), and at the lower output end where the low speed gear #4 engages with the output gear #5. The center gear plane of gear #2 and #3 was not  shrouded because it would have made final  assembly difficult, and the space between gear #3 and the housing was prohibitively small.  

Figure 10.  New Gearbox, Motor Pinions and Gear #1, with GRS    For quick and inexpensive fabrication, the GRS’s were  made by rapid prototyping. Exposure tests were  conducted using coupons of three different RP mate rials and two kinds of grease because of concern  whether commonly available RP materials might  degrade when in contact with the anticipated  hydrocarbon based greases. The results indicated t hat the material used by a laser stereolithography  machine showed excellent property resistance to the grease exposure, therefore, work continued using the same material and process.  The Lower GRS is positioned over gears #4 and #5, and  is screwed to the Short Case. This occurs just  before the Lower Cover is installed and the lower four  case parts are fastened (sandwiched) together by  long screws. The Upper GRS was originally designed similarly, to be screwed to the Upper Cover, because this worked well with the topside-down order of  assembly that is used. However, as a result of  trial practice assemblies with the rapid prototyp e case, it was demonstrated that there is a great  advantage in being able to remove the Upper Cover a fter the gearbox assembly was otherwise complete.   Removing the Upper Cover is the only way that, after the assembly is otherwise complete, a final check of whether the top gear meshing plane of gear #1 to th e motor pinions is adequately centered (usually done  by way of adding shaft shims). The Upper Cover is fastened on separately from the other four casing  parts, making its post-assembly removal possible. Removal of the Upper Cover also serves as a convenient way to grease the gears initially, and to oc casionally inspect the area for grease distribution  and gear wear. The moto rs can even be run with this cover off, although only slowly and unloaded  because all three shafts are not properly supported since their secondary support bearings are pressed into the Upper Cover.  NASA/CP-2010-216272
288 With the GRS screwed to the Upper Cover, removing this cover required that the motor pinions and Gear  #1 also come off, trapped and pulled by the GRS. While th is is theoretically possi ble, in practice, these  gears usually do not remove easily, and since the GRS is  not very strong, it could easily break while  attempting this. Therefore, instead of using screws , the design was changed so that holes in the GRS  would fit onto pins from the case. Because it was desired to either remove the Upper Case and have the  GRS remain in place on the Transition Case, or during initial assembly, place the GRS on the Upper Case and add the Transition Case to it, locating pins  were provided on both cases (Fig. 11). Supports  were designed into both case parts to help hold the GRS in its final trapped position, although a precise final position is actually not required.  This approach appeared to work well. 

Figure 11.  New Gearbox, Lid Removed 

Gearbox Fabrication and Assembly

Two functional gearboxes of the same design were fabricated in-house and assembled. One was ultimately adapted and mounted to the Chariot prototy pe while the other was kept for dynamics testing  and modeling. A combination of 6061 and 7075 aluminum alloys were used due to their availability in the required starting stock sizes, and since either one su ited the structural need s of the housing. The end  result was that the Upper Cover,  Transition Case and Deep Case were made of 7075, while the Short  Case, Lower Cover, and the later Adapter Spacers were made of 6061.  Three case parts were primarily machined using a CNC 2-axis milling machine, while two were completed on a more sophisticated CNC machine that directly used the solid model data. After all five parts were close to completion, the assembly was placed into a boring machine to accurately locate and precisely precision machine the bearing pockets. Two additional parts were designed to allow the gearbox to easily adapt to Chariot; a new steering-box top plate on which the gearbox mounts, and a spacing plate to raise the mounted height of the new gearbox to accommoda te the more shallow mating female hex.    NASA/CP-2010-216272
289 The “Assembly Procedure” that was developed during trial assemblies with the RP casings was used for  assembly, but with many changes and a great deal mo re added detail. One of the main lessons learned  during the build was the need to pay attention to how the gears were to be axially aligned. Shaft shims  were used to accomplish the aligning, but deci sions of how to determine the shim size, what  measurements were needed, and in what order to per form all the operations proved to be a challenge.  Although the final assembly procedure  does capture the process very well, it was more time consuming  than originally thought.   Conclusions, Gearbox Mounted to Chariot 

The new gearbox was completed and mounted on the c enter starboard pod of the Chariot, in LER form,  for an inaugural drive (Fig. 3, 12-15). Software adjus tments within the Chariot main drive controller were  required to accommodate the slightly different gear  ratio outputs. Also, the motor forward and reverse  functions, due to the different gearing, had to be swapped for the High Speed function to work correctly. Beyond these issues, the gearbox appears to have operated well.    

Figure 12.  Chariot in LER Form, Starboard Center Wheel Pod with New Gearbox  NASA/CP-2010-216272
291 Lessons Learned 

1. Document the complete “Design Requirements” ea rly, and revisit them regularly. While the project  intent was to supply a smaller, lighter “drop-in replacement” gearbox with all the capabilities of the  original, in the casing design process, both the ro tary encoder and electric brake were dropped. JSC  agreed to allow us to not provide these features, si nce they use the encoders integral to the motors,  and not every gearbox has to have an electric brake.  

2. While assembly is one thing, serviceability is qui te another. Although it was possible to fully assemble  the gearbox as designed, it would be highly beneficia l if the Upper Cover could be removed without  disassembling most of the unit. This led to an alteration in the way the Upper GRS was secured, allowing removal of just the Upper Cover.   3. Even in high precision work, account for tolerance stack-up by design. For instance, leave room on  drive axles for spacing shims. Given the tolerance stack-up of any gearbox assembly, shims may be needed. 

4. When possible, build contingency time into t he schedule. Even in-house manufacturing shops may  miss an estimated completion date, or number of hour s to completion, as occurred with the gearbox  casing parts. Due to Constellation Program driven priorities, the in-house shop started in August 2008 and was to be complete in October, but was not completed until December. 

5. When schedule and budget allow, particularly for “high visibility” activities, minor resources devoted to  “quality finish” of a product can be beneficial. No t all viewers appreciate seeing the “engineering  details” hanging out. In this case, in order to matc h the original parts' finish, the original gearbox  anodizing vendor was utilized after local vendors proved unable to match the required polished color finish. 

Summary   A second generation “drop-in replacement” gearbox was designed and built for the concept vehicle called Chariot. State-of-the-art aerospace gear design and manufacturing practice were taken advantage of to  produce optimally light weight gear sets, which were incorporated into a smaller, lighter gearbox with otherwise conventional components. Th e “lessons learned” in this gearbox build are broadly applicable.         NASA/CP-2010-216272
295 measured at 6.71 g/cc and is about 15% lower than  440C stainless steel. Table I displays known and  estimated properties of 60NiTi.  

Figure 1.  Nitinol 60 grade 5 test balls. 

Table I. Nominal Comparative properties for conventional bearing alloys and NiTi55 and 60NiTi  Nitinol 60 grade 5 test balls.  Property  NiTi-60 NiTi-55 440C Si3N4 M-50  Density 6.7 g/cc 6.5 g/cc 7. 7 g/cc 3.2 g/cc 8.0 g/cc  Hardness 56-62 Rc 35-40 Rc 58-62 Rc 1300-1500Hv 60-65Rc  Thermal Conductivity W/m-°K 18 9 24 33 ~36  Thermal Expansion ~10x10-6/°C ~10x10-6 /°C 10x10-6/°C 2.6x10-6 ~11 10-6/°C  Magnetic Non Non Magnetic Non Magnetic  Corrosion Resistance Excellent Excellent Marginal Excellent Poor  Tensile/Flexural Strength ~1000MPa ~900 MPa 1900 MPa 600-1200MPa  (Bend  Strength) 2500 MPa  Young’s Modulus ~114 GPa ~100 GPa 200 GPa 310 GPa 210 GPa  Poisson’s Ratio TBD ~.3 .3 .29 .30  Fracture Toughness TBD TBD 22 MPa/ √m 5-7 MPa/ √m 20-23  MPa/√m  Maximum Use Temp ~400° C ~400° C ~400° C ~1100° C ~400° C  Electrical Resistivity ~80x10-6 Ω - cm ~80x10-6 Ωcm ~36x10-6 Ωcm Insulator ~60x10-6 Ωcm  *TBD means “to be determined”  Figure 2 shows the cross sectional microstructure of the 60NiTi ball specimens in the final hardened and polished condition. Microhardness meas urements indicate values in the range of 58-62 on the Rockwell C  scale in the hardened condition. As with most HIPed or sintered powder compacts, the prior particle  boundaries are quite evident and are delineated by oxi des and other tramp phases. Despite containing  only Ni and Ti, Ni-rich NITINOL microstructures can be very complex due to a series of metastable intermetallic phases that could exist depending on thermal history [9]. Analysis of the 60NiTi ball  specimens reveals multiple discrete phases (NiTi,) with NiTi dominating at approximately a 77% volume fraction. The balance is made up from Ni 4Ti3 at 11%, Ni 3Ti at 10%, and NiTi 2 at 2%. 

NASA/CP-2010-216272
296   Figure 2.  Cross-section electron micrograph of Nitinol 60 ball. 

Preliminary differential scanning calorimetry results for the bearing blanks and finished bearings suggest that the 60NiTi bearing blanks may have a slight amount of martensite that could form if cooled below   -15°C but in the hardened condition are microstructurally stable down to at least -100 degrees centigrade.  This indication is corroborated by our physical meas urements of the bearing balls and plates that have  been dimensionally stable througho ut the processing, handling and re search activities undertaken.  Buehler also reported excellent stability during the early alloy development [3].   To assess its tribological properties, Spiral Orbit Tribometer (SOT) tests were conducted using a single 60NiTi ball loaded between two rotating disks, lubricated wi th synthetic oil and tested in a vacuum. In this  test, the 60NiTi ball is marginally lubricated with a small quantity (~25 micrograms) of oil and undergoes  mostly rolling contact interrupted by  a scrubbing/sliding event every disk  rotation. The SOT is depicted in  Figure 3 and described in detail in references 6 and 10. It is basically a thrust bearing with one ball and  flat races (plates). One of the plates is stationary and the other rotates to drive the ball into an orbit that is an opening spiral. The ball contacts a “guide plate” at the end of each orbit, which forces the ball back  into its initial orbital radius. A piezoelectric forc e transducer supporting the guide plate senses the  frictional force developed on the ball as it slides on the rotating plate during the contact of the ball with the  guide plate. During this contact, the coefficient of fr iction is obtained from this force and the load imposed  on the system. The tribometer is housed in a stainless steel chamber t hat can be evacuated by a  turbomolecular pump to ≤2x10 -8 Torr. It can be operated either in this vacuum environment or at  atmospheric pressure.  The plate specimens were 440C stainless steel or 60N iTi. They were lapped flat and their final polish  resulted in an arithmetic mean surface roughness,  Ra, <25 nm (1 min) determined by optical  interferometry. The 60NiTi balls were 12.7 mm (.5 in.) diameter, Grade 5. The final surface cleaning procedure for all ball and plate specimens was by lightly  rubbing with aqueous slurries of silicon carbide  polishing powders, followed by sonication in deionized water. This preparation results in a surface on  which water exhibits ze ro contact angle (spreads) and which exhibits an XPS spectrum a) devoid of  impurities other than a small feat ure due to adventitious carbon.      NASA/CP-2010-216272