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Materials, Processes, and Technologies for Advancing In-Space Electric Propulsion Thrusters

Description:

Scope Title:

High-Temperature, High-Voltage Electric Propulsion Harness Connectors and Cables

Scope Description:

In EP systems, power, commands, and telemetry are relayed between the power processing unit (PPU) and the thruster via dedicated electrical harness assemblies. These harnesses must support the voltage and current needs of the thruster, survive in-space conditions and the operational thermal environment, and not incur unacceptable line loss, radiated emissions, and mass and volume impacts to the spacecraft. Harnesses must also have sufficient flexibility and abrasion resistance, especially for thrusters that are integrated onto actuated gimbals. Individual EP technologies may have specific needs that must be addressed; for example, low-inductance harnesses are preferred in Hall-effect thrusters to reduce thruster discharge oscillations and to promote system stability.

Thermal management of EP systems is a persistent challenge and can be severe in both high-power (>10 kW) and high-power-density (e.g., compact sub-kilowatt) thrusters. This solicitation seeks advancements in connector and cable materials and designs to support harness assembly solutions addressing the following requirements set:

  • Voltages (after derating) of at least 600 VDC; extensibility to support the full range of Hall-effect thruster (up to 800 VDC) and gridded-ion thruster (up to 2 kVDC) operations is desirable.
  • Operating temperatures of at least 300 °C, survival temperatures down to at least -60 °C, and the ability to survive at least 10,000 on-off thermal cycles.
  • Direct currents (after derating) of at least 10 A; extensibility to support high-power EP thrusters (up to to 200 A) is desirable.
  • Deratings consistent with NASA Technical Standard MSFC-STD-3012A (Appendix A) for connectors and wiring.
  • Low outgassing materials consistent with the guideline (i.e., maximum total mass loss (TML) of 1% and maximum collected volatile condensable material (CVCM) deposition of 0.1%) in NASA Technical Standard MSFC-SPEC-1443B.
  • For connectors: features (e.g., venting of connectors and backshells) to mitigate Paschen or corona discharges due to materials or trapped volume outgassing at operating temperatures.
  • For cables: available lengths, flexibility (e.g., bend radius), and abrasion resistance comparable to or better than SOA.

Expected TRL or TRL Range at completion of the Project: 3 to 5

Primary Technology Taxonomy:

  • Level 1 01 Propulsion Systems
  • Level 2 01.2 Electric Space Propulsion

Desired Deliverables of Phase I and Phase II:

  • Analysis
  • Prototype
  • Hardware

Desired Deliverables Description:

Phase I:

  1. Final report containing test data characterizing key properties that address the critical gaps as well as the design and test plan for a component or assembly-level solution to be implemented in Phase II.
  2. Material samples that can be used for independent verification of claimed improvements over SOA.

Phase II:

  1. Final report containing test data verifying key functional and environmental requirements of the solution, including a functional demonstration in an operating thruster environment (in which partnering with EP developers may be necessary).
  2. Prototype harness component or assembly that can be used for independent verification of claimed improvements over SOA.

State of the Art and Critical Gaps:

Recent NASA EP harnesses have utilized stranded, plated copper wiring with multilayer, crosslinked fluoropolymer (e.g., polytetrafluoroethylene (PTFE) and ethylene tetrafluoroethylene (ETFE)) insulation consistent with MIL-W-22759/SAE Standard AS22759D. Commercial off-the-shelf (COTS) wiring rated to 600 VDC and 1,000 VDC exists but is limited to temperatures below ~260 °C. Meanwhile, COTS electrical connectors (such as MIL-SPEC circular connectors) typically have even lower temperature limits.

Temperature derating requirements for electrical connectors mating to SOA EP thrusters have been challenging for recent NASA missions and have complicated mechanical retention and strain relief at the interface. Custom connector solutions or extensive component testing to relax derating requirements are possible approaches, but they are unattractive as increased development costs would be incurred for each mission. Harness material and design improvements that increase the maximum allowable harness temperature would improve the thermal margin for derating purposes on SOA thrusters and facilitate the development of thrusters with higher powers or power densities relative to SOA.

SOA EP harnesses frequently employ custom insulation wraps on COTS wiring in order to support high thruster operating voltages. Such wraps can be mechanically fragile and complicate harness handling and installation. Harness material and design improvements that increase the voltage rating are desirable to improve system reliability and to reduce life-cycle costs.

Relevance / Science Traceability:

Both NASA's Science Mission Directorate (SMD) and Exploration System Development Mission Directorate (ESDMD) need spacecraft with demanding propulsive performance and greater flexibility for more ambitious missions requiring high duty cycles and extended operations under challenging environmental conditions. SMD spacecraft need the ability to rendezvous with, orbit, and conduct in situ exploration of planets, moons, and other small bodies (e.g., comets, asteroids, and near-Earth objects) in the solar system; mission priorities are outlined in the decadal surveys for each of the SMD divisions (https://science.nasa.gov/about-us/science-strategy/decadal-surveys). For ESDMD, higher power EP is a key element in supporting sustained crewed exploration of cislunar space and Mars.

This subtopic seeks innovations to meet future SMD and ESDMD propulsion requirements in EP systems related to such missions. The roadmap for such in-space propulsion technologies is covered under the 2020 NASA Technology Taxonomy (https://www.nasa.gov/offices/oct/taxonomy/index.html), with supporting information archived in the 2015 NASA Technology Roadmap TA-2 (https://www.nasa.gov/offices/oct/home/roadmaps/index.html).

References:

  1. Goebel, D. M., and Katz, I., “Fundamentals of Electric Propulsion: Ion and Hall Thrusters,” https://descanso.jpl.nasa.gov/SciTechBook/SciTechBook.html
  2. NASA Technical Standard MSFC-STD-3012A, “Electrical, Electronic, and Electromechanical (EEE) Parts Management and Control Requirements for MSFC Space Flight Hardware,” https://standards.nasa.gov/standard/msfc/msfc-std-3012
  3. NASA Technical Standard MSFC-SPEC-1443B, “Outgassing Test for Nonmetallic Materials Associated with Sensitive Optical Surfaces in a Space Environment,” https://standards.nasa.gov/standard/msfc/msfc-spec-1443
  4. NASA Technical Handbook NASA-HDBK-4007 (Change 3), “Spacecraft High-Voltage Paschen and Corona Design Handbook,” https://standards.nasa.gov/standard/nasa/nasa-hdbk-4007
  5. U.S. Military Specification MIL-W-22759/SAE Standard AS22759D, “Wire, Electrical, Fluoropolymer-Insulated, Copper or Copper Alloy.”
  6. Clark, S. D., et al., “BepiColombo Electric Propulsion Thruster and High Power Electronics Coupling Test Performances,” IEPC-2013-133, http://electricrocket.org/IEPC/e2cbw2a1.pdf
  7. Pinero, L. R., “The Impact of Harness Impedance on Hall Thruster Discharge Oscillations,” IEPC-2017-023, http://electricrocket.org/IEPC/IEPC_2017_23.pdf

Scope Title:

Cost-Effective, Wear-Resistant Electrodes for High-Power, High-Performance Gridded Ion Thrusters

Scope Description:

Gridded ion thruster technology offers high efficiency, high specific-impulse capabilities, and has been used successfully to support NASA science missions as well as commercial Earth-orbiting applications. The primary life limiter for these devices is typically erosion of the accelerator electrode due to bombardment by charge-exchange ions. While NASA gridded ion thrusters have achieved the necessary lifetimes in the past by operating at derated current densities, there is interest in operation at higher thrust and power densities that would increase mission capture and allow for more compact thruster designs. Higher power and current densities result in increased erosion rates of the accelerator electrode, such that the refractory metals used on previous designs may no longer be sufficient to meet demanding lifetime requirements.

Carbon electrodes have shown promise by offering significantly higher erosion resistance compared to refractory metals. Innovative solutions are desired that would result in manufacturing processes for carbon-based electrodes that are cost-effective relative to prior efforts, making them competitive with SOA electrode manufacturing using refractory metals. Alternative materials besides carbon that allow for improvements in wear resistance over refractory metals such as molybdenum are also desired. These solutions must be capable of producing electrodes with the following geometries, operating voltages, and thermomechanical properties:

  • Screen and accelerator electrode thicknesses of ~0.33 mm and ~0.50 to 0.75 mm, respectively.
  • Screen and accelerator electrode open area fractions of ~70% and ~25%, respectively.
  • Screen and accelerator aperture diameters of ~2 mm and ~1.25 mm, respectively.
  • Gap between the screen and accelerator electrode of ~0.50 to 0.75 mm.
  • A shallow spherical dome (i.e., dished) geometry for both screen and accelerator electrodes.

Note: Both dome and flat geometries are of interest to NASA. However, a dome geometry ensures sufficient electrode stiffness and first-mode natural frequency to withstand expected structural loading during launch as well as maintaining required electrode gaps and avoiding buckling due to compressive stresses caused by nonuniform temperature distributions along electrodes. Manufacturing solutions capable of producing only flat electrodes will also be considered but must demonstrate that structural loading during launch and potential buckling during operation will not be issues.

  • Extensibility to beam extraction (i.e., perforated) diameters of 40 cm or larger.
  • Tight tolerances on apertures’ locations (<0.1 mm) to facilitate proper alignment of apertures between screen and accelerator electrodes.
  • Minimum voltage standoff capability between screen and accelerator electrodes of 2 kV.
  • Peak operating temperatures of 450 °C.
  • Coefficients of thermal expansion less than or equal to that of molybdenum (4.8 x 10-6 K-1).
  • Low outgassing materials consistent with the guideline (i.e., maximum total mass loss (TML) of 1% and maximum collected volatile condensable material (CVCM) deposition of 0.1%) in NASA Technical Standard MSFC-SPEC-1443B.

Proposals are desired that offer solutions which are applicable for manufacturing of both screen and accelerator electrodes. However, proposals that focus only on accelerator electrodes will be considered if such solutions are shown to be compatible with screen electrodes made with heritage refractory metals.

Expected TRL or TRL Range at completion of the Project: 3 to 5

Primary Technology Taxonomy:

  • Level 1 01 Propulsion Systems
  • Level 2 01.2 Electric Space Propulsion

Desired Deliverables of Phase I and Phase II:

  • Analysis
  • Prototype
  • Hardware

Desired Deliverables Description:

Phase I:

  1. A final report detailing the material properties and the manufacturing processes for the electrodes, as well as an evaluation of the extensibility of the processes to sizes of interest (i.e., 40-cm perforated diameter or larger).
  2. A scaled-down sample of each electrode (either screen and accelerator or accelerator only, depending on the approach) representative of typical electrode thickness and open area fraction to be delivered to NASA for independent assessment and tests.

Phase II:

  1. A final report detailing final manufacturing processes and an updated evaluation of the extensibility of these processes to sizes of interest (i.e., 40-cm perforated diameter or larger).
  2. Screen and accelerator electrodes (or accelerator electrode only, depending on the approach) at least 30 cm in diameter that can be hot-fire tested with a gridded ion thruster (in which partnering with EP developers may be necessary).

State of the Art and Critical Gaps:

While extensive research and development of carbon electrodes have resulted in solutions that were technically adequate, the complexity and associated costs of manufacturing have been prohibitive toward widespread adoption into ion thruster technology. The material used for electrodes has historically been refractory metals, whose thermal and mechanical properties allow the electrodes to withstand the temperatures and launch loads they will experience while offering adequate erosion resistance. Fabrication using refractory metals such as molybdenum typically involves chemical etching to produce the apertures within the electrodes. Carbon-based solutions have been developed previously by several organizations and include carbon-carbon, amorphous graphite, and pyrolytic graphite (PG). Fabrication techniques for carbon electrodes have been rather varied and complex and have included methods such as chemical vapor deposition and carbonization. Apertures in carbon electrodes have been created using laser drilling, electric discharge machining (EDM), or conventional machining. As such, innovative solutions are desired that would result in manufacturing processes for carbon electrodes that are less complex and/or more cost effective than prior efforts. Alternatively, solutions are desired involving other materials that can provide improved erosion resistance while having comparable manufacturing cost or complexity compared to existing electrode materials such as molybdenum.

Relevance / Science Traceability:

Both NASA's Science Mission Directorate (SMD) and Exploration Systems Development Mission Directorate (ESDMD) need spacecraft with demanding propulsive performance and greater flexibility for more ambitious missions requiring high duty cycles and extended operations under challenging environmental conditions. SMD spacecraft need the ability to rendezvous with, orbit, and conduct in situ exploration of planets, moons, and other small bodies (e.g., comets, asteroids, and near-Earth objects) in the solar system; mission priorities are outlined in the decadal surveys for each of the SMD divisions (https://science.nasa.gov/about-us/science-strategy/decadal-surveys). For ESDMD, higher power EP is a key element in supporting sustained crewed exploration of cislunar space and Mars.

This subtopic seeks innovations to meet future SMD and ESDMD propulsion requirements in EP systems related to such missions. The roadmap for such in-space propulsion technologies is covered under the 2020 NASA Technology Taxonomy (https://www.nasa.gov/offices/oct/taxonomy/index.html), with supporting information archived in the 2015 NASA Technology Roadmap TA-2 (https://www.nasa.gov/offices/oct/home/roadmaps/index.html).

References:

  1. Goebel, D. M., and Katz, I., "Fundamentals of Electric Propulsion: Ion and Hall Thrusters," https://descanso.jpl.nasa.gov/SciTechBook/SciTechBook.html
  2. Sangregorio, M., Xie, K., Wang, N., Guo, N., and Zang, Z., "Ion engine grids: Function, main parameters, issues, configurations, geometries, materials and fabrication methods," Chinese Journal of Aeronautics, Vol. 31, No. 8, 2018, pp. 1635-1649.
  3. Snyder, J. S., "Review of Carbon-based Grid Development Activities for Ion Thrusters," 39th AIAA/ASME/SAE/ASEE Joint Propulsion Conference and Exhibit, AIAA-2003-4715, Huntsville, AL, July 20-23, 2003.
  4. Haag, T., "Mechanical Design of Carbon Ion Optics," 41st AIAA/ASME/SAE/ASEE Joint Propulsion Conference & Exhibit, AIAA-2005-4408, Tucson, AZ, July 10-13, 2005.
  5. De Pano, M. K., Hart, S. L., Hanna, A. A., and Schneider, A. C., "Fabrication and Vibration Results of 30-cm Pyrolytic Graphite Ion Optics," 40th AIAA/ASME/SAE/ASEE Joint Propulsion Conference and Exhibit, AIAA-2004-3615, Fort Lauderdale, FL, July 11-14, 2004.
  6. Polk, J. E., Goebel, D. M., Snyder, J. S., Schneider, A. C., Johnson, L. K., and Sengupta, A., "A high power ion thruster for deep space missions," Review of Scientific Instruments, Vol. 83, No. 7, 2012, pp. 073306-1–073306-14.
  7. Wallace, N. C., and Corbett, M., "Optimization and Assessment of the Total Impulse Capability of the T6 Ion Thruster," 30th International Electric Propulsion Conference, IEPC-2007-231, Florence, Italy, September 17-20, 2007.
  8. Wang, J., Polk, J., Brophy, J., and Katz, I., "Three-Dimensional Particle Simulations of NSTAR Ion Optics," 27th International Electric Propulsion Conference, IEPC-2001-085, Pasadena, CA, October 15-19, 2001.
  9. Christensen, J. A., Freick, K. J., Hamel, D. J., Hart, S. L., Norenberg, K. T., Haag, T. W., Patterson, M. J., Rawlin, V. K., Sovey, J. S., Anderson, J. R., Becker, R. A., and Polk, J. E., "Design and Fabrication of a Flight Model 2.3 kW Ion Thruster for the Deep Space 1 Mission," 34th AIAA/ASME/SAE/ASEE Joint Propulsion Conference and Exhibit, AIAA-1998-3327, Cleveland, OH, July 13-15, 1998.
  10. NASA Technical Standard MSFC-SPEC-1443B, “Outgassing Test for Nonmetallic Materials Associated with Sensitive Optical Surfaces in a Space Environment,” https://standards.nasa.gov/standard/msfc/msfc-spec-1443
  11. Snyder, J. S., Anderson, J. R., Van Noord, J. L., and Soulas, G. C. "Environmental Testing of NASA's Evolutionary Xenon Thruster Prototype Model 1 Reworked Ion Engine," Journal of Propulsion and Power, Vol. 25, No. 1, 2009, pp. 94-104.

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