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Search indexed NASA NTRS and DOE OSTI research on propulsion, heat transfer, battery materials and energy systems. Follow report and document links to the original sources.

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NASA NTRSDate not supplied

In Situ Small Spacecraft Missions Utilizing Heatshield for Extreme Entry Environments Technology

There is considerable interest in utilizing Small Spacecraft beyond low Earth orbit. In November of 2018, successful data relay operations of the MarCO CubeSats during the entry, descent, and landing (EDL) of the Mars InSight mission showed the viability of using CubeSats for interplanetary missions. Additional testament to the promise of Small Spacecraft class rideshare missions is the upcoming Artemis-1 flight test that will launch thirteen 6U CubeSats, as well as the establishment of NASA’s SIMPLEx program which will conduct stand-alone planetary science missions that launch with a primary payload. It is anticipated that continued innovations in Small Spacecraft capabilities combined with small EDL systems will expand the range of potential missions to allow for in situ investigations. Recently, NASA invested in the development of a new, efficient and capable ablative thermal protection system (TPS), utilizing 3-D Weaving. The new dual layer TPS, Heatshield for Extreme Entry Environment Technology (HEEET), is tailorable, scalable, robust, mass efficient and capable of supporting in situ missions across the solar system. Combining the HEEET entry system with innovative Small Spacecraft technology will substantially expand the range of Small Spacecraft mission applications by offering the capability for high speed entry or aerocapture at destinations with atmospheres. This paper will describe the HEEET aeroshell system and highlight various mission concepts including a dual technology demonstration mission that is under development and other concepts to deliver instruments for planetary science.

Small Spacecraft↗
NASA NTRSDate not supplied

In Situ Small Spacecraft Missions Utilizing Heatshield for Extreme Entry Environments Technology

There is considerable interest in utilizing Small Spacecraft beyond low Earth orbit. In November of 2018, successful data relay operations of the MarCO CubeSats during the entry, descent, and landing (EDL) of the Mars InSight mission showed the viability of using CubeSats for interplanetary missions. Additional testament to the promise of Small Spacecraft class rideshare missions is the upcoming Artemis-1 flight test that will launch thirteen 6U CubeSats, as well as the establishment of NASA’s SIMPLEx program which will conduct stand-alone planetary science missions that launch with a primary payload. It is anticipated that continued innovations in Small Spacecraft capabilities combined with small EDL systems will expand the range of potential missions to allow for in situ investigations. Recently, NASA invested in the development of a new, efficient and capable ablative thermal protection system (TPS), utilizing 3-D Weaving. The new dual layer TPS, Heatshield for Extreme Entry Environment Technology (HEEET), is tailorable, scalable, robust, mass efficient and capable of supporting in situ missions across the solar system. Combining the HEEET entry system with innovative Small Spacecraft technology will substantially expand the range of Small Spacecraft mission applications by offering the capability for high speed entry or aerocapture at destinations with atmospheres. This paper will describe the HEEET aeroshell system and highlight various mission concepts including a dual technology demonstration mission that is under development and other concepts to deliver instruments for planetary science.

Heatshield↗
NASA NTRS2008

Applications of Optical Fiber Assemblies in Harsh Environments, the Journey Past, Present and Future

Over the past ten years, NASA has studied the effects of harsh environments on optical fiber assemblies for communication systems, lidar systems, and science missions. The culmination of this has resulted in recent technologies that are unique and tailored to meeting difficult requirements under challenging performance constraints. This presentation will focus on the past mission applications of optical fiber assemblies including; qualification information, lessons learned and new technological advances that will enable the road ahead.

Ott, Melanie N.↗
NASA NTRS1990

Upcoming planetary missions and the applicability of high temperature superconductor bolometers

Planetary missions to Mars and beyond can last 11 years and longer, making impractical the use of stored cryogens. Passive radiative coolers and single-stage mechanical coolers remain possibilities. CRAF and CASSINI, both using the newly developed Mariner Mark 2 spacecraft, will be the next outer planet missions after Galileo; they are intended to provide information on the origin and evolution of the solar system. CRAF is a cometary rendezvous mission slated for a 1994 launch. CASSINI has been chosen by ESA and will be launched by a Titan 4/Centaur in 1996. It will fly by Jupiter in 2000, inject an ESA-supplied probe into Titan in 2002, and take data in Saturn orbit from 2002 to 2006. NASA/Goddard is currently developing a prototype Fourier transform spectrometer (CIRS) that will be proposed for the CASSINI mission. The baseline infrared detectors for CIRS are HgCdTe to 16 microns and Schwarz-type thermopiles from 16 to 1000 microns. The far infrared focal plane could be switched from thermopiles to high temperature superconductor (HTS) bolometers between now and 1996. An HTS bolometer could be built using the kinetic inductance effect, or the sharp resistance change at the transition. The transition-edge bolometer is more straightforward to implement and initial efforts at NASA/Goddard are directed to that device. A working device was made and tested in early 1989. It also has somewhat elevated noise levels below 100 Hz. Upcoming efforts will center on reducing the time constant of the HTS bolometer by attempting to deposit an HTS film on a diamond substrate, and by thinning SrTiO3 substrates. Attempts will be made to improve the film quality to reduce th 1/f noise level, and to improve the thermal isolation to increase the bolometer sensitivity. An attempt is being made to deposit good-quality HTS films on diamond films using an MOCVD technique.

Brasunas, J.↗
NASA NTRS1991

Upcoming planetary missions and the applicability of high temperature superconductor bolometers

Planetary missions to Mars and beyond can last 11 years and longer, making impractical the use of stored cryogens. Passive radiative coolers and single-stage mechanical coolers remain possibilities. Cassini and Comet Rendezvous/Asteroid Fly-by (CRAF), both using the newly developed Mariner Mark 2 spacecraft, will be the next outer planet missions after Galileo; they are intended to provide information on the origin and evolution of the solar system. CRAF is slated for a 1994 launch. Cassini was chosen by ESA and will be launched by a Titan 4/Centaur in 1996. It will fly by Jupiter in 2000, inject an ESA-supplied probe into Titan in 2002, and take data in Saturn's orbit from 2002 to 2006. NASA/Goddard is currently developing a prototype Fourier transform spectrometer, the Composite Infrared Spectrometer (CIRS), for the Cassini mission. The baseline infrared detectors for CIRS are HgCdTe to 16 microns and Schwarz-type thermopiles from 16 to 1000 microns. The far infrared focal plane could be switched from thermopiles to high temperature superconductor (HTS) bolometers between now and 1996. An HTS bolometer could be built using the kinetic inductance effect, or the sharp resistance change at the transition. The transition-edge bolometer is more straightforward to implement, and initial efforts at NASA/Goddard are directed to that device. A working device was made and tested in early 1989. It also has somewhat elevated noise levels below 100 Hz. Upcoming efforts will center on reducing the time constant of the HTS bolometer by attempting to deposit an HTS film on a diamond substrate, and by thinning SrTiO3 substrates. Attempts will be made to improve the film quality to reduce the 1/4 noise level, and to improve the thermal isolation to increase the bolometer sensitivity. An attempt is being made to deposit good-quality HTS films on diamond films using a metal-organic chemical vapor deposition (MOCVD) technique.

Brasunas, J.↗
NASA NTRS1987

Deep space optical communications

Preliminary concepts and designs of a deep space optical link for planetary and deep space science are described. As a mission application, attention is given to a spacecraft optical transceiver package (OPTRANSDAC) attached to a Mars rover vehicle. Also considered are a preliminary concept for a 1000-AU mission, and to a tentative long-range plan for NASA's deep space optical communications program.

Sokoloski, Martin M.↗
NASA NTRS1991

Electric propulsion system technology

The work performed on the Ion Propulsion System Technology Task in FY90 is described. The objectives of this work fall under two broad categories. The first of these deals with issues associated with the application of xenon ion thrusters for primary propulsion of planetary spacecraft, and the second with the investigation of technologies which will facilitate the development of larger, higher power ion thrusters to support more advanced mission applications. Most of the effort was devoted to investigation of the critical issues associated with the use of ion thrusters for planetary spacecraft. These issues may be succinctly referred to as life time, system integration, and throttling. Chief among these is the engine life time. If the engines do not have sufficient life to perform the missions of interest, then the other issues become unimportant. Ion engine life time was investigated through two experimental programs: an investigation into the reduction of ion engine internal sputter erosion through the addition of small quantities of nitrogen, and a long duration cathode life test. In addition, a literature review and analysis of accelerator grid erosion were performed. The nitrogen addition tests indicated that the addition of between 0.5 and 1.0 percent of nitrogen by mass to the xenon propellant results in a reduction in the sputter erosion of discharge chamber components by a factor of between 20 and 50, with negligible reduction in thruster performance. The long duration test of a 6.35-mm dia. xenon hollow cathode is still in progress, and has accumulated more than 4,000 hours of operation at an emission current of 25 A at the time of this writing. One of the major system integration issues concerns possible interactions of the ion thruster produced charge exchange plasma with the spacecraft. A computer model originally developed to describe the behavior of mercury ion thruster charge exchange plasmas was resurrected and modified for xenon propellant. This model enables one to calculate the flow direction and local density of the charge exchange plasma, and indicates the degree to which this plasma can flow upstream of the thruster exhaust plane. A continuing effort to investigate the most desirable throttling technique for noble gas ion thrusters concentrated this year on experimentally determining the fixed flow rate throttling range of a 30-cm dia. thruster with a two-grid accelerator system. These experiments demonstrated a throttling capability which covers a 2.8 to 1 variation in input power. This throttling range is 55 percent greater than expected, and is due to better accelerator system performance at low net-to-total voltage ratios than indicated in the literature. To facilitate the development of large, higher power ion thrusters several brief studies were performed. These include the development of a technique which simulates ion thruster operation without beam extraction, the development of an optical technique to measure ion thruster grid distortion due to thermal expansion, tests of a capacitance measurement technique to quantify the accelerator system grid separation, and the development of a segmented thruster geometry which enables near term development of ion thrusters at power levels greater than 100 kW. Finally, a paper detailing the benefits of electric propulsion for the Space Exploration Initiative was written.

Brophy, John R.↗
NASA NTRS2002

Status of Brayton Cycle Power Conversion Development at NASA GRC

The NASA Glenn Research Center (GRC) is pursuing the development of Brayton cycle power conversion for various NASA initiatives. Brayton cycle power systems offer numerous advantages for space power generation including high efficiency, long life, high maturity, and broad scalability. Candidate mission applications include surface rovers and bases, advanced propulsion vehicles, and earth orbiting satellites. A key advantage is the ability for Brayton converters to span the wide range of power demands of future missions from several kilowatts to multi-megawatts using either solar, isotope, or reactor heat sources. Brayton technology has been under development by NASA since the early 1960's resulting in engine prototypes in the 2 to 15 kW-class that have demonstrated conversion efficiency of almost 30% and cumulative operation in excess of 40,000 hours. Present efforts at GRC are focusing on a 2 kW testbed as a proving ground for future component advances and operational strategies, and a 25 kW engine design as a modular building block for 100 kW-class electric propulsion and Mars surface power applications.

Mason, Lee S.↗
NASA NTRS1990

SP-100 nuclear space power systems with application to space commercialization

The technology of the SP-100 space nuclear power system program is compared to that of more familiar solar-power systems. The SP-100 program develops, validates, and demonstrates the technology for space nuclear power systems in the range of 10 to 1000 kilowatts electric for use in future military and civilian space missions. Mission applications, including earth orbiting platforms and lunar/Mars surface power, are enhanced or made possible by SP-100 technology. Attention is given to the SP-100 reference flight system design, the SP-100 nuclear reactor and nuclear-reactor shield, the platform-mounted, tethered, and free-flying reactors, and installation, operation, and disposal options, as well as lunar-Mars surface applications. The SP-100 is presented as one of the nuclear energy sources needed for long-life, compact, lightweight, continuous high power independent of solar orientation, specific orbits, or missions.

Smith, J. M.↗
DOE OSTI2022

Mathematical Foundations for Nonlocal Interface Problems: Multiscale Simulations of Heterogeneous Materials (Final LDRD Report)

Nonlocal models provide a much-needed predictive capability for important Sandia mission applications, ranging from fracture mechanics for nuclear components to subsurface flow for nuclear waste disposal, where traditional partial differential equations (PDEs) models fail to capture effects due to long-range forces at the microscale and mesoscale. However, utilization of this capability is seriously compromised by the lack of a rigorous nonlocal interface theory, required for both application and efficient solution of nonlocal models. To unlock the full potential of nonlocal modeling we developed a mathematically rigorous and physically consistent interface theory and demonstrate its scope in mission-relevant exemplar problems.

97 MATHEMATICS AND COMPUTING↗
NASA NTRS1983

Airborne reconnaissance in the civilian sector - Agricultural monitoring from high-altitude powered platforms

Design concepts and mission applications for unmanned high-altitude powered platforms (HAPPs) are discussed. A chemically powered HAPP (operating altitude 18-21 km, wingspan 26 m, payload 91 kg, endurance 2-3 days) would use current turboprop technology. A microwave-powered HAPP (operating altitude around 21 km, wingspan 57.9 m, payload 500 kg, endurance weeks or months) would circle within or perform boost-glide maneuvers around a microwave beam of density 1.1 kw/sq m. Of two solar-powered-HAPP designs presented, the more promising uses five vertical solar-panel-bearing fins, two of which can be made horizontal at night, (wingspan 57.8/98.3 m, payload 113 kg, endurance weeks or months). The operating altitude depends on the latitude and season: this HAPP design is shown to be capable of year-round 20-km-altitude flights over the San Joaquin Valley in California, where an agricultural-monitoring mission using Landsat-like remote sensors is proposed. Other applications may be better served by the characteristics of the other HAPPs. The primary advantage of HAPPs over satellites is found to be their ability to provide rapidly available high-resolution continuous or repetitive coverage of specific areas at relatively low cost.

Youngblood, J. W.↗
NASA NTRS2006

Demonstration of a Non-Toxic Reaction Control Engine

Three non-toxic demonstration reaction control engines (RCE) were successfully tested at the Aerojet Sacramento facility under a technology contract sponsored by the National Aeronautics and Space Administration s (NASA) Marshall Space Flight Center (MSFC). The goals of the NASA MSFC contract (NAS8-01109) were to develop and expand the technical maturity of a non-toxic, on-orbit auxiliary propulsion system (APS) thruster under the auspices of the Exploration Systems Mission Directorate. The demonstration engine utilized Liquid Oxygen (LOX) and Ethanol as propellants to produce 870 lbf thrust. The Aerojet RCE s were successfully acceptance tested over a broad range of operating conditions. Steady state tests evaluated engine response to varying chamber pressures and mixture ratios. In addition to the steady state tests, a variety of pulsing tests were conducted over a wide range of electrical pulse widths (EPW). Each EPW condition was also tested over a range of percent duty cycles (DC), and bit impulse and pulsing specific impulse were determined for each of these conditions. White Sands Test Facility (WSTF) in April 2005 for incorporation into a cryogenic Auxiliary Propulsion System Test Bed (APSTB). The APSTB is a test article that will be utilized in an altitude test cell to simulate anticipated mission applications. The objectives of this APSTB testing included evaluation of engine performance over an extended duty cycle map of propellant pressure and temperature, as well as engine and system performance at typical mission duty cycles over extended periods of time. This paper provides acceptance test results and a status of the engine performance as part of the system level testing. Subsequent to acceptance testing at Aerojet, these three engines were delivered to the NASA

Robinson, Philip J.↗
NASA NTRS2007

Demonstration of a Non-Toxic Reaction Control Engine

T:hree non-toxic demonstration reaction control engines (RCE) were successfully tested at the Aerojet Sacramento facility under a technology contract sponsored by the National Aeronautics and Space Administration's (NASA) Marshall Space Flight Center (MSFC). The goals of the NASA MSFC contract (NAS8-01109) were to develop and expand the technical maturity of a non-toxic, on-orbit auxiliary propulsion system (APS) thruster under the auspices of the Exploration Systems Mission Directorate. The demonstration engine utilized Liquid Oxygen (LOX) and Ethanol as propellants to produce 870 lbf thrust. The Aerojet RCE's were successfully acceptance tested over a broad range of operating conditions. Steady state tests evaluated engine response to varying chamber pressures and mixture ratios. In addition to the steady state tests, a variety of pulsing tests were conducted over a wide range of electrical pulse widths (EPW). Each EPW condition was also tested over a range of percent duty cycles (DC), and bit impulse and pulsing specific impulse were determined for each of these conditions. Subsequent to acceptance testing at Aerojet, these three engines were delivered to the NASA White Sands Test Facility (WSTF) in April 2005 for incorporation into a cryogenic Auxiliary Propulsion System Test Bed (APSTB). The APSTB is a test article that will be utilized in an altitude test cell to simulate anticipated mission applications. The objectives of this APSTB testing included evaluation of engine performance over an extended duty cycle map of propellant pressure and temperature, as well as engine and system performance at typical mission duty cycles over extended periods of time. This paper provides acceptance test results and a status of the engine performance as part of the system level testing.

Robinson, Philip J.↗
NASA NTRS1987

Mars manned transportation vehicle

A viable power system technology for a surface transportation vehicle to explore the planet Mars is presented. A number of power traction systems were investigated, and it was found that a regenerative hydrogen-oxygen fuel cell appears to be attractive for a manned Mars rover application. Mission requirements were obtained from the Manned Mars Mission Working Group. Power systems weights, power, and reactants requirements were determined as a function of vehicle weights for vehicles weighing from 6,000 to 16,000 lb (2,722 to 7,257 kg), (Earth weight). The vehicle performance requirements were: velocity, 10 km/hr; range, 100 km; slope climbing capability, 30 deg uphill for 50 km; mission duration, 5 days; and crew, 5. Power requirements for the operation of scientific equipment and support system capabilities were also specified and included in this study. The concept developed here would also be applicable to a Lunar based vehicle for Lunar exploration. The reduced gravity on the Lunar surface, (over that on the Martian surface), would result in an increased range or capability over that of the Mars vehicle since many of the power and energy requirements for the vehicle are gravity dependent.

Perez-Davis, Marla E.↗
NASA NTRS1989

Mars manned transportation vehicle

A viable power system technology for a surface transportation vehicle to explore the planet Mars is presented. A number of power traction systems were investigated, and it was found that a regenerative hydrogen-oxygen fuel cell appears to be attractive for a manned Mars rover application. Mission requirements were obtained from the Manned Mars Mission Working Group. Power systems weights, power, and reactants requirements were determined as a function of vehicle weights for vehicles weighting from 6,000 to 16,000 lb (2,722 to 7,257 kg), (Earth weight). The vehicle performance requirements were: velocity, 10 km/hr; range, 100 km; slope climbing capability, 30 deg uphill for 50 km; mission duration, 5 days; and crew, 5. Power requirements for the operation of scientific equipment and support system capabilities were also specified and included in this study. The concept developed here would also be applicable to a Lunar based vehicle for Lunar exploration. The reduced gravity on the Lunar surface, (over that on the Martian surface), would result in an increased range or capability over that of the Mars vehicle since many of the power and energy requirements for the vehicle are gravity dependent.

Perez-Davis, Marla E.↗
NASA NTRSDate not supplied

Development of ASCENT Propellant Thrusters and Propulsion Systems

Since the 1960s, hydrazine technology has dominated the market of propulsion technology by providing reliable performance on a wide variety of spacecraft. Efforts to replace hydrazine with a higher-performance and safer alternative have engaged numerous emerging technologies in the industry to mature. ASCENT propellant, also known as AF-M315E, is an example of one such green propellant. Initially developed to provide a safer handling environment than hydrazine, ASCENT offers higher overall performance than a hydrazine equivalent by creating more heat during ignition. The next generation of propulsion technology, based on ASCENT propellant, produces a higher impulse and energy density, expanding the realm of applicable missions from LEO to GEO and beyond. Building on a strong background in small satellite technology and expertise in thermal and oxidation resistant materials, Plasma Processes began several thruster programs for variety of small satellite applications using ASCENT-based technology. Plasma Processes’ background expertise in thermal and oxidation resistant materials provides the necessary foundation to create ASCENT-based technology hardware.

Green Propellant↗
NASA NTRS2018

Transiting Exoplanet Survey Satellite (TESS) Flight Dynamics Commissioning Results and Experiences

The Transiting Exoplanet Survey Satellite (TESS) will perform the first-ever spaceborne all-sky exoplanet transit survey and is the first primary-mission application of a lunar-resonant orbit. Launched on April 18, 2018, TESS completed a two-month commissioning phase consisting of three phasing loops followed by a lunar flyby and a final maneuver to achieve resonance. During the mission orbit, no further station-keeping maneuvers are planned or required. NASA Goddard Space Flight Center is performing flight dynamics operations for the mission. This paper covers the design, implementation, and results from TESS commissioning, including the projected performance of the final mission orbit.

Satellite↗
NASA NTRS2018

Transiting Exoplanet Survey Satellite (TESS) Flight Dynamics Commissioning Results and Experiences

The Transiting Exoplanet Survey Satellite (TESS) will perform the first-ever spaceborne all-sky exoplanet transit survey and is the first primary-mission application of a lunar-resonant orbit. Launched on April 18, 2018, TESS completed a two-month commissioning phase consisting of three phasing loops followed by a lunar flyby and a final maneuver to achieve resonance. During the mission orbit, no further station-keeping maneuvers are planned or required. NASA Goddard Space Flight Center is performing flight dynamics operations for the mission. This paper covers the design, implementation, and results from TESS commissioning, including the projected performance of the final mission orbit.

Parker, Joel J. K.↗