ESMD Course Material : Fundamentals of Lunar and Systems Engineering for Senior Project Teams, with Application to a Lunar Excavator

Contact: David Beale,

Chapter X
Lunar Engineering Handbook
Chapter 1
Chapter 2
Chapter 3
Chapter 4
Chapter 5
Chapter 6
Chapter 7
Chapter 8
Chapter 9
Additional Resources
Contact Info
About the Authors

Chapter 6: Component and Material Selection

David Beale


  1. The Lunar Rovers
  2. Standards and References
  3. Flight Qualified
  4. Material Selection
    1. Structural Material Choices the Student Investigated
    2. Coating Material Choices the Student Investigated        
    3. Conveyor Material Choices the Student Investigated
    4. Bit Material Choices the Student Investigated
    5. Mechanical Component Material Choices the Student Investigated
    6. Insulation Material Choices the Student Investigated
    7. Miscellaneous Material Issues the Student Investigated
  5. Mechanical Components
    1. Fasteners
    2. Bearings
    3. Lubricants
  6. Motors - A Student's Trade Study
  7. Power Components - A Student's Trade Study
  8. References
  9. Appendix

Component and part design, selection, and material choice are driven by the application and the environment.  A component that is built and tested successfully (perhaps even subjected to test chamber conditions that mimic some lunar environmental conditions) may still fail or perform poorly during a mission.  This happens because of the unforeseen ("hindsight is always 20/20"), and it is difficult to recreate a test environment that mimics the exact environmental conditions of a space mission.  So component and part design and selection for the moon will be influenced by previous space and lunar successes and failures, and the knowledge base that has been built.  Much of this knowledge base comes from satellite systems and the Apollo lunar missions.  Hence Legacy should be an important criteria to consider when selecting and designing components for the moon.   From (NASA, 2007) legacy “refers to the original manufacturer’s level of quality and reliability that is built into the parts which have been proven by (1) time and service, (2) number of units in service, (3) mean time between failure performance, and (4) number of use cycles.”  If a candidate component or part has a successful legacy, then a designer should strongly consider using it.   Many times parts with legacy can be available as COTS (Commercial Off the Shelf) products.

 (Gies, 1996) listed the issues that need to be considered when designing lunar equipment:

·         Abrasion and wear on parts that contact regolith, such as for digging, cutting edges, material handling equipment, track systems, wheels, construction equipment, pulleys, sheaves and other machinery elements. 

·         Vacuum welding of metals, which may require special coating and treatments.

·         Electrostatic properties of regolith will cause it to adhere to and penetrate unprotected mechanical joints with bearings and structural connections.  It will also adhere to viewing surfaces, solar panels, radiators and antennas.

·         Strategies will be needed to create effective vacuum seals (e.g. for door locks) and effective bearings (including lubricants, filters, and seals for bearings).

Student Project Considerations

A student team designing a lunar excavator or any lunar machinery may have a mission to construct and test a prototype. Or perhaps the team would like to enter the Lunar Challenge Competition ( ).   In both case it is not necessary to purchase space qualified components or manufacture using materials that would be expected to be in a lunar mission-ready lunar excavator.  Nevertheless the team should be able to justify that their prototype, if tested successfully, could be a basis for further development beyond Phase B (remember that Phase C ends with component fabrication for the space mission).  This requires an awareness of components’ design and selection choices for a lunar mission.

Here the focus is on components that would be a concern of a mechanical designer.  This includes mechanical components (bearings, fasteners), lubricants, motors, materials and an overview of power systems.  This topic is too broad to consider in great detail here, so references are often cited instead.  Often the selection of a component is not clear, and many choices are possible.  In these situations a trade study may be appropriate, and trade studies performed by several students are presented.

1‑1  LRV, from (LPI, 2008a)


1‑2  LRV Dimensions

1‑3  LRV steering system


1‑4 LRV chassis


1‑5 LRV traction drive

1‑6  LRV Suspension System

1‑7  LRV Wheel

The Lunar Rovers (LPI, 2008b), (NASA, 1971), (Eckart, 1999)

An example of a system that demonstrate legacy is the LRV in Figure 1-1.  Three Lunar Roving Vehicles (LRV) were sent to the moon on Apollo 15, 16 and 17 with the specific mission of traversing the terrain of the moon to extend the operational distance that would otherwise be covered by an astronaut on foot. Many features were incorporated into each of the rovers in order to meet the specific demands of the lunar environment, while making it small and light enough to be transported to the moon.  The LRV weighed less than the two astronaut passengers.  Materials were chosen that could survive in the lunar environment, it included specialized power sources, and also barriers that protect from the regolith.  Its performance characteristics are shown below in Figure 1-8. 

1‑8 LRV performance characteristics (Heiken, Vaniman, & French, 1991)

The foldable frame of the LRV (Figure 1-4) was constructed from Aluminum 2219 tube-welded assemblies.  The suspension (Figure 1-6) was a double wishbone, each wishbone attached to a torsion bar and a damper between the chassis and upper wishbone. The wheels (Figure 1-7) consisted of an aluminum hub, tire made of zinc coated woven piano wires and titanium chevron treads, attached to the rim and discs of formed aluminum.  Dust guards were mounted about each wheel in Figure 1-1.  Front and rear wheel steer was accomplished by an Ackermann-geometry steering linkage system (Figure 1-3), driven by an electric motor servo-system that amplifies the left and right joystick motion from the astronaut.  The power system consisted of two 36 V silver-zinc potassium hydroxide non-rechargeable batteries used to drive four ¼ horsepower electric motors (Figure 1-5) located at each wheel.  This power system gave the rover an operating range of approximately 57 miles and enabled it to reach speeds up to 17 km/h.   The motors were speed reduced 80:1 with a harmonic drive gearing ( ), which are known for large gear ratios, light weight, compact size and no gear backlash when compared to a planetary gear system.  The motors and harmonic drive were hermetically sealed and pressurized to 7.5 psia to protect from lunar dust and for improved brush lubrication. Braking was both electodynamic by the motors and from brake shoes forced against a drum through a linkage and cable.  Designed to carry twice its own weight, the LRV had a clearance of 36 cm when fully loaded (Figure 1-2). 

 In order to protect the LRV against the thermal environment of the moon several different thermal control systems were incorporated into the LRV design. These systems consisted of MLI blankets (Multi-layer insulation) covered by Beta Cloth, space radiators, mass heat sinks, special surface coatings and finishes, and thermal straps. Most of these components performed well. However one of the main problems that the LRV encountered was an issue concerning the lunar dust. Degradation of thermal and electronic components was a problem as well as the wear and tear of components and other surfaces from the abrasive lunar dust. 

Standards and References


When reviewing or referencing the following standards, be aware that they may not have been intended for the lunar environment, so they may not be entirely applicable in the presence of regolith, and a different temperature and radiation environment than earth orbit.

·         AIAA S-114-2005, “Moving Mechanical Assemblies for Space and Launch Vehicles” covers the design of moving mechanical assemblies for orbit and launch.  This standard “specifies general requirements for the design, manufacture, quality control, testing, and storage of moving mechanical assemblies (MMAs) to be used on space and launch vehicles.”  It considers the mechanical or electromechanical devices that control the movement of mechanical parts, including gears, lubricants, bearings, fasteners, springs, dampers and motors. 

·         The Proceedings of the Aerospace Mechanism Symposium are published annually and papers are concerned with actuators, lubricants, latches, connectors, and other mechanisms. 

·         NASA/TP-1999-2069888  NASA Space Mechanisms Handbook.  The Handbook (including CD/DVD) is available only to US citizens who need the material.  It is restricted under ITAR (International Traffic in Arms Regulations).  

·         MIL-HDBK-5 Metallic Materials and Elements for Aerospace Structures, contains standardized mechanical property design values and other related design information for metallic materials, fasteners and joints.

Other Standards:

·         DOD-HDBK-343 Design, Construction, and Testing Requirements for One of a Kind Space Equipment

·         MIL-STD-100 Engineering Drawing Practices

·         MIL-STD-1539 Direct Current Electrical Power Space Vehicle Design Requirements

·         DOD-E-8983 General Specification for Extended Space Environment Aerospace Electronic Equipment

·         MIL-S-83576 General Specification for Design and Testing of Space Vehicle Solar Cell Arrays

·         DOD-STD-1578 Nickel-Cadmium Battery Usage Practice for Space Vehicles

 Reference Texts

A number of excellent reference texts are listed in the References section, including "Space Vehicle Mechanisms", "Space Vehicle Design" and  "Fundamentals of Space Systems". 

Flight Qualified

Any hardware or materials used for lunar missions will need to be of a special variety known as "Flight Qualified".  Therefore, any designer will have at his or her disposal only a limited number of parts and materials.  Care should always be taken that all materials and parts called out in a mechanical design are flight qualified.  NASA and commercial spacecraft vendors keep up to date lists of acceptable materials and components.  Any such list from a reputable source should always be sought at the start of a mechanical design.  Flight qualified materials and parts are always flight proven hardware with program heritage.  The process to get any new material or part flight qualified is an arduous and long task.  This is why most technologies in flight today are 10 to 15 years old.  Any good program manager will choose a flight proven workhorse solution over a newer, better more exotic technology option.  Materials and parts on spacecraft (and the lunar surface) are exposed to extreme conditions in space and failures cannot be easily serviced.  Therefore, most programs require all sub-systems to have double or even triple redundancy in failure threats.  For example, thermostats can fail open or closed and often backup thermostats are included to take over in the event that the primary thermostat fails.  Any part that inserts a single point failure in the system is unacceptable.  So when it comes to mechanical designs for lunar missions, simpler is better, all parts are flight qualified, and all possible failures have been identified and fully explored and remedied. 

Material Selection

Material selection is strongly affected by the lunar environment, which can severely degrade material properties.  Radiation will weaken exposed polymeric material, while regolith is very abrasive and invasive. Material properties can depend on temperature, for instance carbon steels are not a good choice because they become brittle at the low lunar temperatures, and poorly chosen epoxies in composites may soften at high temperatures.  The specific lunar application (temperature, loads, etc.) drives material selection.  Hence a trade study is called for.  Below is attached a portion of a trade study by a student to choose materials for his excavator components (excavator bit, structural members, and a conveyor belt).   The student's material choice is not given.

Structural Material Choices the Student Investigated

 1.1.1    Aluminum

  • pros: easily machinable, good strength to weight, low cost
  • Applications: frame materials of spacecraft (Conley, 1998), lunar rover frame
  • Among the conventional structural materials, aluminum is by far the most common.  A large variety of alloys exist providing a broad range of such characteristics as strength and weldability.  Thus, for applications at moderate temperature in which moderate strength and good strength-to-weight ratio are desirable, aluminum is still most often the material of choice.  This popularity is enhanced by ready availability and ease of fabrication.  A number of surface-coating processes exist to allow tailoring of surface characteristics for hardness, emissivity, absorbtivity, etc (Griffin, 1991).
  • Material Properties (MATWEB, 2008) (search: Aluminum 1199-O)
    • Density: 0.0975 lb/in³ (AA; Typical)
    • Ultimate Tensile Strength: 6530 psi
    • Tensile Yield Strength: 1450 psi
    • Modulus of Elasticity: 8990 ksi (In Tension; Compressive Modulus is about 2% higher)
    • Poisson’s Ratio: 0.330
    • Shear Strength: 4930 psi (Calculated value)
    • Shear Modulus: 3630 ksi
    • CTE, linear 20°C: 13.1 µin/in-°F (20-100ºC)
    • CTE, linear 250°C: 14.2 µin/in-°F (Average over the range 20-300ºC)
    • Specific Heat: 0.215 BTU/lb-°F
    • Thermal Conductivity: 1690 BTU-in/hr-ft²-°F
    • Melting Point: 1220 °F
    • Material Components:
      • Aluminum, Al             >= 99.99 %

 1.1.2    Beryllium

  • pros: good stiffness, low density
  • cons: machining takes time and care
  • Applications: infrared optics (high reflectivity), precision gimbal structures (low density & high modulus-to-density ratio), guidance systems
  • Beryllium offers the highest stiffness of any naturally occurring material along with low density, high strength, and high temperature tolerance.  Thermal conductivity is also good.  Beryllium has been used in limited applications where its desirable characteristics have been required.  The main limitation on more extensive use of this apparently excellent material is toxicity.  In bulk form, beryllium metal is quite benign and can be handled freely.  The dust of beryllium or its oxide, however, has very detrimental effects on the human respiratory tract.  This means that machining or grinding operations are subject to extensive safety measures to capture and contain dust and chips.  This renders normal fabrication methods unusable without resorting to these intensive (i.e., expensive) measures [2].
  • Properties (Conley, 1998) pg. 42-46
    • Young’s Modulus: ~303 GPa (-40C to 80C)
    • Poisson’s Ratio: 0.028
    • Tensile Yield: ~265GPa (-40C to 80C)
    • Tensile Strength: 380 to 480 GPa (depending on the grade)
    • Max Service Temp: 750C

 1.1.3    Titanium

  • pros: high strength
  • cons: high cost
  • difficult to machine or weld
  • titanium alloys are highly praised for their use in aerospace and high-temperature applications
  • Titanium is a lightweight, high-strength structural material with excellent high temperature capability.  It also exhibits good stiffness.  Some alloys are fairly brittle, which tends to limit their application, but a number of alloys with reasonable ductility exist.  Use of titanium is limited mostly by higher cost, lower availability, and fabrication complexity to applications that particularly benefit from its special capabilities (Griffin, 1991).
  • Properties (MATWEB, 2008)
    • Density: 0.163 lb/in³
    • Ultimate Tensile Strength: 31900 psi
    • Tensile Yield Strength: 20300 psi
    • Modulus of Elasticity: 16800 ksi
    • Poisson’s Ratio: 0.340
    • CTE, linear 20°C: 4.94 µin/in-°F (over the range 20-100ºC)
    • CTE, linear 1000°C: 5.61 µin/in-°F
    • Specific Heat: 0.126 BTU/lb-°F
    • Thermal Conductivity: 118 BTU-in/hr-ft²-°F
    • Melting Point: 3000 - 3040 °F
    • Emissivity: 0.630 (unoxidized; 650 nm)

 1.1.4    Graphite-epoxy

  • pros: lightweight
  • cons: outgassing
  • The use of high-strength and stiffness graphite fiber within an epoxy matrix offers an excellent high-strength structural material.  Proper selection of the cloth and/or unidirectional fibers offers the ability to tailor strength and stiffness directionally and to the desired levels to optimize it for the purpose.  The low density of graphite offers a weight advantage as well.  High-temperature characteristics are improved by use of graphite instead of glass, although the matrix is the final limiting factor (Griffin, 1991).
  • Properties (MATWEB, 2008) (search: Epoxy/Carbon Fiber Composite)
    • Density: 0.0455 - 0.0650 lb/in³
    • Ultimate Tensile Strength: 7250 – 305,000 psi (Average value: 845 MPa Grade Count:14)
    • Modulus of Elasticity: 1450 – 75,400 ksi (Average value: 153 GPa Grade Count:14)
    • Compressive Yield Strength: 7250 – 249,000 psi (Average value: 593 MPa Grade Count:14)
    • Compressive Modulus: 1190 – 17,400 ksi (Average value: 33.0 GPa Grade Count:5)
    • Shear Strength: 116 - 17400 psi (Average value: 60.1 MPa Grade Count:9)
    • Flexural Yield Strength: 930 - 18100 ksi (Average value: 47.2 GPa Grade Count:7)
    • Flexural Modulus: 16000 - 232000 psi (Average value: 542 MPa Grade Count:7)
    • Thermal Conductivity: 41.6 - 2780 BTU-in/hr-ft²-°F (Average value: 105 W/m-K Grade Count:9)
    • CTE, linear 20°C: -0.167 - 15.6 µin/in-°F (Average value: 8.87 µm/m-°C Grade Count:7)
    • Specific Heat: 0.239 - 0.287 BTU/lb-°F (Average value: 1.13 J/g-°C Grade Count:3)

 1.1.5    Stainless steel

  • cons: weight
  • Applications: structural members, although it is often replaced by lighter materials, power train components (Conley, 1998)
  • Stainless steel is typically used in applications requiring higher strength and/or higher temperature resistance.  Stainless is preferred because its use eliminates concern about rust and corrosion during fabrication and testing.  Also, if the part is exposed to low temperature, the low ductile-to-brittle (DBT) transition temperature is important (Griffin, 1991).
  • Properties (Conley, 1998) pg. 20-27

Coating Material Choices the Student Investigated

  • Properties (Conley, 1998) pg. 230

Conveyor Material Choices the Student Investigated

1.3.1    Linked metal 

  • Not likely feasible – too bulky

1.3.2    Teflon

  • pros: self-lubricated
  • cons: erosion due to radiation
  • Material Properties (MATWEB, 2008) (search: DuPont Teflon® Grade 30 Aqueous Dispersion)
    • Base Resin Density: 0.0325 lb/in³ (Weight of PTFE resin solids)
    • Specific Gravity: 0.0542 lb/in³
    • Solids Content: 60.0 % (Percent of PTFE resin solids)
    • Particle Size: 0.220 µm (Average dispersion particle size)
    • pH: 9.50
    • Viscosity: 20.0 cP (at 25°C)
    • Melting Point: 621 °F

 Teflon® PTFE 30 fluoropolymer resin is a negatively charged, hydrophobic colloid containing approximately 60% (by total weight) of 0.05 to 0.5 mm polytetrafluoroethylene (PTFE) resin particles suspended in water. A milky white liquid, Teflon® PTFE 30 also contains approximately 6% (by weight of PTFE) of a nonionic wetting agent and stabilizer. Viscosity at room temperature is approximately 20 cP. Nominal pH is 10 (MATWEB, 2008).

Compared with other grades of PTFE dispersions,Teflon® PTFE 30 is a general-purpose product, often preferred for impregnating woven goods and for some coating processes. It imparts some of the unique properties of PTFE resin to porous structures (MATWEB, 2008).

When properly processed, the PTFE resin in Teflon® PTFE 30 exhibits the superior properties typical of the fluoropolymer resins: retention of properties after service at 260°C (500°F), useful properties at –240°C (–400°F), chemical inertness to nearly all industrial chemicals and solvents, and low friction and antistick surfaces. Dielectric properties are outstanding and stable with frequency and temperature (MATWEB, 2008).


Teflon® PTFE 30 is used to impregnate packings made from braided fibers for severe chemical and thermal service; to coat glass fabric for industrial conveyor belting, nonadhesive separator sheets for laminating and press blankets, and gaskets; and as surface coatings for other substrates (MATWEB, 2008).

1.3.3    Glass Fabric

  • Pros: flexible; can be Teflon coated to improve; lightweight; can be manufactured for specific properties
  • Cons:
  • Applications: external surface of spacecraft thermal blankets (Griffin, 1991)
  • Fiberglass cloth, which is strong and flexible, has seen use as an insulator and as protective armor against micrometeoroids.  A commercially available cloth of fiberglass coated with Teflon called Betacloth has been used as the external surface of spacecraft thermal blankets for this purpose (Griffin, 1991).
  • Material Properties (MATWEB, 2008) (search: Industrial Laminates/Norplex NP511 Glass Fabric)
    • Specific Gravity: 0.0650 - 0.0686 lb/in³ (0.062"; ASTM D792)
    • Moisture Absorption at Equilibrium: 0.200 % (0.062"; ASTM D229)
    • Tensile Yield Strength: 37000 psi (0.062", CW; ASTM D638)           43000 psi (0.062", LW; ASTM D638)
    • Modulus of Elasticity: 2700 ksi (0.062", CW; ASTM D229)                3000 ksi (0.062", LW; ASTM D229)
    • Flexural Strength: 70000 psi (0.062", CW; ASTM D790)                   80000 psi (0.062", LW; ASTM D790)
    • Compressive Strength: 63000 psi (0.5"; ASTM D695)
    • Shear Strength: 22000 psi (0.62"; ASTM D732)
    • CTE, linear 20°C: 7.22 µin/in-°F (x-axis (0.062"); IPC-TM 650-2.4.24)
    • CTE, linear 20°C Transverse to Flow: 8.33 µin/in-°F (y-axis (0.062"); IPC-TM 650-2.4.24)
    • Maximum Service Temperature, Air: 365 °F
    • Glass Temperature: 329 °F (Tg)
    • Flammability, UL94: HB (0.062")
    • Bond Strength: 2200 lb (0.5", ASTM D229)
    • Color: Natural

 1.3.4    Kapton tape

  • Applications: outer layers of thermal blankets (Griffin, 1991)
  • A new polymeric film material with higher strength and the ability to withstand higher temperature than Mylar is the polyimide Kapton.  These characteristics have made Kapton a desirable choice for outer layers of thermal blankets.  A problem has arisen with the discovery that, in low Earth orbits, polymer surfaces undergo attack and erosion by atomic oxygen, which is prevalent at these altitudes. Kapton seems to be more susceptible to this sort of attack than Mylar.  In any case, for long life use in low orbit, metallization or coating with a more resistive polymer such as Teflon will probably be required.  This erosion rate is sufficiently low that, for shorter missions, the problem may not be serious (Griffin, 1991).

 1.3.5    Mylar

  • Application: MLI on spacecraft (Griffin, 1991)
  • By far the most commonly used plastic film material in space applications has been Mylar.  This is a strong, transparent polymer that lends itself well to fabrication into sheets or films as thin as 0.00025 in.  Coated with a few angstroms of aluminum to provide reflectivity, Mylar is well suited to the fabrication of the multilayer insulation extensively used on spacecraft (Griffin, 1991).

 Bit Material Choices the Student Investigated 

1.4.1    Hardened steel

  • pros: high toughness

 1.4.2    Diamond

  • pros: high hardness
  • cons: cost, brittle failure
  • Material Properties (MATWEB, 2008)
    • Density: 0.108 - 0.145 lb/in³
    • Compressive Yield Strength: 1900 - 6900 MPa
    • Poisson’s Ratio: 0.0700 - 0.200
    • Thermal Conductivity: 8330 BTU-in/hr-ft²-°F (Thick film diamond made by SP3)
    • Thermal Conductivity: 12500 BTU-in/hr-ft²-°F (De Beers thermal thick film synthetic diamond)
    • Fracture Toughness: 5.46 - 8.01 ksi-in½

Mechanical Component Materials the Student Investigated

 1.5.1    AISI 440C Stainless Steel

  • Application: typically used in space mechanism bearings (Conley, 1998)
  • heat treatable
  • high wear resistance
  • moderate corrosion resistance in mild environments
  • Material Properties (MATWEB, 2008)
    • Density: 0.282 lb/in³
    • Ultimate Tensile Strength: 254000 psi
    • Tensile Yield Strength: 186000 psi
    • Modulus of Elasticity: 29000 ksi
    • Charpy impact: 14.0 ft-lb
    • CTE, linear 20°C: 5.67 µin/in-°F (from 32-212°F)
    • Specific Heat: 0.110 BTU/lb-°F (from 32-212°F)
    • Max service temp, air: 1400 °F (Continuous Service);                                 1500 °F (Intermittent Service)
    • Material Components:
      • Carbon, C                    0.600 - 0.750 %
      • Chromium, Cr             16.0 - 18.0 %
      • Manganese, Mn           <= 1.00 %
      • Molybdenum, Mo       <= 0.750 %
      • Phosphorous, P           <= 0.0400 %
      • Silicon, Si                    <= 1.00 %
      • Sulfur, S                      <= 0.0300 %

Insulation Material Choices the Student Investigated


Miscellaneous Material Issues the Student Investigated

 1.7.1    Outgassing concerns

  • (Conley, 1998) pg. 224
  • Most composite materials will outgas their volatile content when taken into a low pressure environment

 1.7.2    Other Issues

  • Some temperature ranges are given for the material properties above as well as operable temperatures in sources
  • No data is available on Mylar, Kapton, etc., other than the general information.
  • Most of the structure will be built from aluminum, while the bit will be comprised of stainless steel
  • Efforts were made to select materials that are used in space applications (space rated).

 Mechanical Components


Space Fasteners design choices, with attention given to aerospace applications, materials and temperature ranges, are presented in the Fastener Design Manual (Barrett, 1990),  The document considers fastener material selection, platings, lubricants, locking methods, washers, inserts, rivets and lockbolts, and bolthead markings and design data tables.  Design calculations for loading conditions is also presented, including fatigue loading, fastener torque, combined shear and tension loads, pullout load for tapped holes, grip length, head styles, and fastener strengths.  MIL-HDBK-5 also contains allowable strengths for many fasteners.  Fasteners for MS (military standard) and NAS (national aerospace standard) can be found at

 From the table in Figure 1-9, notice that carbon steel bolts have a limited temperature range, becoming brittle at -65F.   The stainless steel bolts appears offer a range of choices that may be acceptable for lunar machinery.

 From the table in Figure 1-10, cadmium is the most common.  Zinc coatings can be considered if expected temperatures do not exceed the limit.

 From the table of Figure 1-11, molybdenum disulfide is most appropriate for the temperature and vacuum conditions as a thread lubricant.   

1‑9 Fastener materials, from (Barrett, 1990) 

1‑10 Fastener platings and coatings (Barrett, 1990)

1‑11  Thread lubricants, from (Barrett, 1990)

Bearings (Conley, 1998):

Rolling-element bearings for lunar applications must capably withstand the challenges of the lunar environment (temperature extremes, penetrating regolith and the vacuum environment) and be highly reliable to minimize repairs.  If not hermetically sealed they will be subject to vacuum pressures.  Temperature extremes could induce thermal stresses and thermal distortions.   Significant loads and vibrations often occur during launch are also a concern.  For space flights the AISI 440C (a high hardness, corrosion resistant steel) and AISI 52100 (not as hard or corrosion-resistant as AISI 440C, but better wear resistance) are the most common.  Hybrid bearings, a combination of ceramic ball and metallic race, reduce micro-welding at ball and race contact, may have advantages in some situations. 

 Shields and seals cover the rolling element so they are not exposed and protected to a certain degree from outside contaminates like regolith.  Shields and seals are attached on a bearing’s outer race, and move with the outer race.  A shield will not touch the inner race because of a small clearance gap.  Seals do rub against the inner race but will be less likely to allow regolith particles inside.  Thermal control is a concern in a lunar environment where convection is not an available heat transfer mechanism.  Thermal conductivity through a bearing is increased by the presence of a lubricant.  


Lubricant inadequacies have been implicated as a cause of a number of space mechanism failures. An ideal lubricant would retain the desired viscosity over a wide temperature range and be nonvolatile.  The ability of a lubricant to resist becoming a gas is related to its molecular weight.  Low molecular weight lubricants are more volatile in vacuum and heat than higher molecular weight lubricants.  The three types of lubricants are liquids (lubricating oils, lubricant greases) and solid films.  Fluorinated oils and greases have excellent vacuum characteristics.   Solid films, such as soft metal films, polymers and low-shear strength materials, find use in bearings, bushings, contacts and gears.   See (Conley, 1998) and (Fusaro, 1994) for details. 

Motors - A Student's Trade Study

A motor is considered to be a component in a system hierarchy.  The types that have been used in satellites include DC brush, DC brushless and stepper motors.  Below is a trade study performed by a student, comparing different motor types for a lunar excavator conveyor, rated for 100W available DC power.   The student's recommendations and conclusions are included, however they have not been verified.  You will need to do your own trade study.

DC- Brush Motor

-          Operation of brushed motor at reduced atmospheric pressures cause the motor to fail prematurely due to brush wear

-          In general the longest brush life for a motor occurs under room ambient conditions

-          Brush wear is dependent of brush material, motor design, and operating loads

-          Slow speed torque motors have good life, usually more expensive

-          Samarium Cobalt: Operating, Tmax 300-330 deg C without damage

-          On Earth moisture in the air helps to lubricate the brushes of the motor, in a vacuum there is no moisture to lubricate the brushes of the motor, which is the main reason for brush wear.

-          RECOMMENDATION- To find alternatives to brush motor because of short life, and high costs. For the conveyor the motor is running almost continually thus wearing down the brushes at an accelerate rate.

 DC- Stepper Motor

-          Natural companion for occasional duty, moderate rates and frequently precise pointing accuracy

-          This type of motor was used on Surveyor Lander in the 1960’s

-          Has the capability to be directly driven by a digital processor, and runs of traditional DC power

-          Attractive in applications where power is limited

-          Has high, detent torque, which allows the motor to be held in position without power. Detent torque is approximately 85% of powered torque

-          Hybrid stepper motor produces most torque for a given diameter and length

-          Almost all heat losses occur in the stator, so should have a good conduction path to other structures

-          RECOMMENDATION- Both the stepper and stepper hybrid motors are an attractive option for this project. Since power is limited for the lunar excavator a stepper motor would be a more sensible choice. Also the motor can be digitally-controlled which would allow easy control from earth.

DC-Brushless Motor

-          Most versatile general purpose of the electric motors

-          Windings are switched or communicated electronically rather than mechanically

-          Internal permanent-magnetic field is trapezoidal shape for max torque and smooth rotation

-          Careful thermal design can increase motor’s capacity

-          RECOMMENDATION- DC Brushless may be an appropriate option because of its versatility and ability for continual use. With high torque capability it may provide the best solution to run the conveyor belt. Thermal concerns can be reduced with a precision design of the motor.

Final Choice

 The final choice for the space-qualified motor was a stepper motor. The specific motor is:

 Phyrton Vacuum Sealed Stepper Motor, model VSH 125.200.10

 Please see appendix for detailed specifications. This motor was chosen because, despite having a max power rating above the allotted 100 watts the motor could operate at the power output of 37.7 watts (1/20 HP). 37.7 watts was the power rating on the first generation excavator. 

Power Components - A Student's Trade Study

Below is a trade study performed by a student to select a power system for a lunar excavator.   You will need to perform your own trade study. 

Basic Components of a Spacecraft Power System (Patel, 2004), (Gilmore, 2002)

Selection of a power system is based primarily on low mass and low cost components. 


  • Primary Energy Source – Solar radiation, radioisotopes, nuclear reactors, electrochemical and/or chemical fuel.
  • Energy Conversion – Photovoltaic, thermoelectric, dynamic alternator, fuel cell and thermo-ionic.
  • Power Regulation & Control – Battery charge and discharge converters, shunt dissipators, mode controller for bus voltage error signal
  • Rechargeable Energy Storage – Rechargeable batteries (Silver Zinc, Nickel Cadmium, Nickel Hydrogen, Nickel Metal Hydride, Lithium Ion).
  • Distribution & Protection – Structure to protect vital electrical components from the space environment.

Primary Energy Source

A primary energy source is applicable for low power, short-life satellites and spacecraft.  After the battery is drained, it is typically jettisoned for mass reduction.  Therefore, the primary energy source will be replaced by a rechargeable energy source for a longer life span.

Energy Conversion

The most widely used and cost efficient form of energy conversion is the photovoltaic solar array.  Solar arrays can provide power requirements from tens of watts to several kilowatts with a life span of a few months to fifteen years.  After fifteen years, the life of a solar array degrades due to the space environmental effects on the photovoltaic cells.  Ionized particles, micrometeoroid impact, debris and ultraviolet radiation are concerns when designing solar arrays.  A method to protect against radiation effects and to enhance solar energy absorption is to layer the array with fused silica or ceria-doped microsheet coated with silica monoxide. Single-crystal silicon, gallium arsenide, semi-crystalline and poly-crystalline, thin film, amorphous and multi-junction are types of photovoltaic cells used in the space environment. 

  • Single-Crystal Silicon Cells
    • Advantage:  Widely available and have been used as the “workhorse” of the space industry
    • Disadvantage:  Expensive manufacturing process for space qualified cells
  • Gallium Arsenide Cells
    • Advantage:  High conversion efficiency in comparison Single-crystal Silicon cells
    • Disadvantage:  Extremely expensive to manufacture
  • Semi-Crystalline & Poly-Crystalline Cells
    • Advantage:  Low cost of manufacture which gives a net reduction in the cost per watt
    • Disadvantage:  Low energy conversion efficiency
  • Thin Film Cells
    • Advantage:  Less expensive to manufacture
    • Disadvantage:  Has not been used widely in space applications (lack of data).
  • Amorphous Cells – Not enough data to be selected as a serious candidate for space applications (new technology).
  • Multi-Junction Cells – High efficiency and good manufacturability.

Power Regulation & Control

Power regulation and control is necessary in monitoring battery health and preventing over charge and discharge because batteries may be susceptible to damage.  Power regulation also includes the incorporation of shunt dissipators to dissipate power that is unwanted after meeting the load power and the battery charge requirements.  Thermal requirements for electrical components can also be contained in the power regulation and control because batteries and associated components typically require certain temperature operating ranges to perform efficiently.

Rechargeable Energy Sources

  • Silver Zinc Batteries – Although Silver Zinc batteries have a high specific energy, they consist of a low cycle life and are not suitable for lunar excavation applications.
  • Nickel Cadmium (NiCd)
    • Advantages:
      • Responsible for powering all satellites until mid-1980s
    • Disadvantages:
      • Low specific energy
      • Temperature sensitive (Operating range: 0-10°C)
      • Short life cycle
      • Cadmium is under environment regulatory scrutiny
  • Nickel Hydrogen (NiH2) – Currently used in place of Nickel Cadmium for space applications.
    • Advantages:
      • Most widely used in space applications in the last 20 years
      • Selected for use in the International Space Station
      • Can withstand some abuse due to overcharge and over discharging
      • Superior charge/discharge life cycle compared to Nickel Cadmium
      • Low internal resistance
      • Internal pressure can be measured by a mounting a strain gage on pressure vessel
    • Disadvantages:
      • Low energy density
      • Pressure vessel rupture, handling and safety concerns
      • Compression seal is typically coated with Teflon resulting in poor radiation tolerance in certain space applications
      • High self discharge rate
      • Temperature sensitive (Operating range: 0-20°C) – Adds mass to thermal control system
      • High loss of capacity on storage
      • Large support structure for battery systems (Increased mass)
  • Nickel Metal Hydride (NiMH)
    • Advantages compared to NiCd:
      • Electrode materials are non-toxic
      • Improvement in specific energy
      • Negligible memory effect
    • Advantages compared to NiH2:
      • Less sensitive to temperature – Lower thermal control cost
      • Improved energy density
      • Less support structure
      • Lower operating pressure
    • Disadvantages:
      • Less capable of producing high peak power
      • High self discharge rate
      • Adversely affected by high temperatures
      • Susceptible to damage due to overcharging
      • Very expensive
      • Few space applications use NiMH batteries
  • Lithium-Ion (Li-Ion)
    • Advantages:
      • Higher specific energy and energy density than NiH2
      • High charge efficiency
      • Potentially long cycle life at high depth of discharge
      • Low temperature sensitivity
      • Low internal impedance
      • Capable of delivering high short time peak power
      • Small battery footprint
    • Disadvantages:
      • Sensitive to overcharge and over discharge – Requires more elaborate charging circuitry with adequate protection
      • Requires trickle charging


The most common form of protection of electrical components and batteries is the design and construction of a “battery box” type structure to house all the components.  The structure is typically built out of aluminum or a similar type space rated material.  To protect against radiation, a MLI (Multilayer insulation) material is placed around the structure as a barrier.  MLI serves a dual purpose by preventing excess heat loss from a component and by preventing excessive heating to a component from environmental fluxes.  The simplest type of MLI is assembled from thin embossed Mylar sheets with a layer of aluminum on one side.  More efficient designs include the material Kapton.    Thermal design of the spacecraft primarily uses resources in the environment for heating and cooling needs.  However, in severe environments or extremely sensitive components, auxiliary insulators, heaters, radiators or louvers may be used to maintain operational temperatures.


Battery – Lithium Ion (Saft Battery Group) – See attached industrial brochure.

For the lunar excavator, the best choice for a rechargeable battery is the Lithium Ion.  The main disadvantage of the Lithium Ion is the sensitivity to overcharging and over discharging.  This can be dealt with by the implementation of an acceptable power regulation and control system.  The Lithium Ion is also a good choice for the excavator power system because of mass requirements.  Li-Ion batteries do not require a great deal of space and therefore the mass of the system can be reduced.  Lithium Ion batteries are currently being used on the Mars Rover  for approximately 4.5 years (Halpert, 1999).


Solar Cells/Array (Able Engineering) – See industrial brochure in the Appendix

Barrett, R. T. (1990). Fastener Design Manual: NASA Lewis Research Center.

Conley, P. L. (Ed.). (1998). Space Vehicle Mechanisms: Elements of a Successful Design: John Wiley & Sons.

Eckart, P. (1999). The Lunar Base Handbook: McGraw-Hill.

Fusaro, R. L. (1994). Lubrication of Space Systems. Paper presented at the Society of Tribologists and Lubrication Engineers Annual Meeting.

Gies, J. (1996). The Effects of the Lunar Surface upon Machinery. Paper presented at the Proceeding of Space 96.

Gilmore, D. G. (Ed.). (2002). Spacecraft Thermal Control Handbook, Volume 1: Fundamental Technologies (2nd ed.): The Aerospace Corporation.

Griffin, M. D., French, J.R. (1991). Space Vehicle Design: AIAA Education Series.

Halpert, G., Frank, H., Subbarao, S. (1999). Batteries and Fuel Cells in Space. The Electrochemical Society Interface, 25-30.

Heiken, G. H., Vaniman, D. T., & French, B. M. (Eds.). (1991). Lunar Sourcebook A User's Guide to the Moon: Cambridge University Press

LPI. (2008a). Lunar and Planetary Institute Website. from

LPI. (2008b). Lunar and Planetary Institute Website: The Apollo Roving Vehicle. from

MATWEB. (2008). MatWeb - Material Property Data from

NASA. (1971). Lunar Roving Vehicle Operations Handbook LS0006-002-2H.

NASA. (2007). Systems Engineering Handbook, NASA/SP-2007-6105 Rev1

 NASA Headquarters, Washington, D.C., 20546.

Patel, M. R. (2004). Spacecraft Power Systems: CRC Press.



 COTS solar panels and batteries.