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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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Operation of Hall thrusters with bismuth propellant has been shown to be a promising path for development of high-power (140 kW per thruster), high performance (8000s I(sub sp at >70% efficiency) electric propulsion systems.
There is considerable interest in expanding the applicability of cubesat spacecraft into lightweight, low cost missions beyond Low Earth Orbit. A conceptual design was done for a 6-U cubesat for a technology demonstration to demonstrate use of electric propulsion systems on a small satellite platform. The candidate objective was a mission to be launched on the SLS test launch EM-1 to visit a Near-Earth asteroid. Both asteroid fly-by and asteroid rendezvous missions were analyzed. Propulsion systems analyzed included cold-gas thruster systems, Hall and ion thrusters, incorporating either Xenon or Iodine propellant, and an electrospray thruster. The mission takes advantage of the ability of the SLS launch to place it into an initial trajectory of C3=0. Targeting asteroids that fly close to earth minimizes the propulsion required for fly-by/rendezvous. Due to mass constraints, high specific impulse is required, and volume constraints mean the propellant density was also of great importance to the ability to achieve the required deltaV. This improves the relative usefulness of the electrospray salt, with higher propellant density. In order to minimize high pressure tanks and volatiles, the salt electrospray and iodine ion propulsion systems were the optimum designs for the fly-by and rendezvous missions respectively combined with a thruster gimbal and wheel system For the candidate fly-by mission, with a mission deltaV of about 400 m/s, the mission objectives could be accomplished with a 800s electrospray propulsion system, incorporating a propellant-less cathode and a bellows salt tank. This propulsion system is planned for demonstration on 2015 LEO and 2016 GEO DARPA flights. For the rendezvous mission, at a ΔV of 2000 m/s, the mission could be accomplished with a 50W miniature ion propulsion system running iodine propellant. This propulsion system is not yet demonstrated in space. The conceptual design shows that an asteroid mission is possible using a cubesat platform with high-efficiency electric propulsion.
Mini-Magnetospheric Plasma Propulsion (M2P2) seeks to create a plasma-inflated magnetic bubble capable of intercepting significant thrust from the solar wind for the purposes of high speed, high efficiency spacecraft propulsion. Previous laboratory experiments into the M2P2 concept have primarily used helicon plasma sources to inflate the dipole magnetic field. The work presented here uses an alternative plasma source, the cascaded arc, in a geometry similar to that used in previous helicon experiments. Time resolved measurements of the equatorial plasma density have been conducted and the results are discussed. The equatorial plasma density transitions from an initially asymmetric configuration early in the shot to a quasisymmetric configuration during plasma production, and then returns to an asymmetric configuration when the source is shut off. The exact reasons for these changes in configuration are unknown, but convection of the loaded flux tube is suspected. The diffusion time was found to be an order of magnitude longer than the Bohm diffusion time for the period of time after the plasma source was shut off. The data collected indicate the plasma has an electron temperature of approximately 11 eV, an order of magnitude hotter than plasmas generated by cascaded arcs operating under different conditions. In addition, indirect evidence suggests that the plasma has a beta of order unity in the source region.
Efficiency of two-stage repetitively pulsed coaxial plasma engine calculated from engine thrust and total mass flow
The purpose of this paper is to generate specific power and efficiency projections through the year 2050 for electric machines for aircraft applications. A general literature review was performed to identify the types of electric machines that are commonly used and which types have the biggest potential for future aircraft applications due to their high specific power and efficiency. A database with historical data was built to include parameters such as weight [kg], rated power [kW], specific power [kW/kg], RPM, efficiency, year, motor cooling type, application type and motor type to allow for trend identification and accurate projections. Once the data was gathered, multiple curve fits on the historical data were generated and extrapolated to produce the projections for specific power according to conservative, nominal and aggressive projection scenarios. A different process was followed for the efficiency projections due to the scattered nature of the data. A state of the art (SoA) value for efficiency was identified through literature review and was used to create the conservative, nominal and aggressive projections for the time frames of 2030, 2040, and 2050. The efficiency and the specific power projections of EMs for 2050 are 0.989 and 50kW/kg respectively. This paper will also be examining circuit protection as it is an additional component of electric powertrains.
The topics covered include the following: (1) Advanced Launch System (ALS); (2) life cycle cost/lb payload; (3) current operational cost; (4) major cost driver; (5) fully automated operations; (6) technology impacts on vehicle dry mass efficiency; (7) single stage to orbit (SSTO) approach; and (8) efficient SSTO propulsion system operations.
The power train performance of load leveled electric vehicles can be compared with that of nonload leveled systems by use of a simple mathematical model. This method of measurement involves a number of parameters including the degree of load leveling and regeneration, the flywheel mechanical to electrical energy fraction, and efficiencies of the motor, generator, flywheel, and transmission. Basic efficiency terms are defined and representative comparisons of a variety of systems are presented. Results of the study indicate that mechanical transfer of energy into and out of the flywheel is more advantageous than electrical transfer. An optimum degree of load leveling may be achieved in terms of the driving cycle, battery characteristics, mode of mechanization, and the efficiency of the components. For state of the art mechanically coupled flyheel systems, load leveling losses can be held to a reasonable 10%; electrically coupled systems can have losses that are up to six times larger. Propulsion system efficiencies for mechanically coupled flywheel systems are predicted to be approximately the 60% achieved on conventional nonload leveled systems.
Efficiency of two-stage repetitively pulsed coaxial plasma engine calculated from engine thrust and total mass flow
The efficiency, capability, and evolution of a lunar base will be largely dependent on the transportation system that supports it. Beyond the space station in low Earth orbit, a lunar-derived propellant supply could provide the most important resource for the transportation infrastructure. The key to an efficient lunar base propulsion system is the degree of lunar self-sufficiency and reasonable propulsion system performance. Lunar surface propellant production requirements must be accounted in the measurement of efficiency of the entire space transportation system. Of all chemical propellant/propulsion systems considered, hydrogen/oxygen (H/O) OTVs appear most desirable, while both H/O and aluminum/oxygen propulsion systems may be considered for the lander. Aluminized-hydrogen/oxygen and silane/oxygen propulsion systems are also promising candidates. Lunar propellant availability and processing techniques, chemical propulsion/vehicle design characteristics, and the associated performance of the total transportation infrastructure are reviewed, conceptual propulsion system designs and vehicle/basing concepts, and technology requirements are assessed.
NASA's X-57 all-electric aircraft was a research project aimed at investigating lightweight and efficient electric propulsion components. The general approach was to utilize a distributed electric propulsion (DEP) design. An essential component of this design was the High Lift Motor Controller (HLMC), a motor drive which provided power to the High Lift Motors (HLMs) and High Lift Propellers (HLPs) responsible for providing additional thrust for take-off and landing. This paper presents the detailed design, test results, and outcomes from the development of the HLMC, a 14 kW, 1kg, 98.3% efficient, outer mold line (OML) cooled, silicon carbide (SiC) MOSFET-based inverter and controller.
NASA's X-57 all-electric aircraft was a research project aimed at investigating lightweight and efficient electric propulsion components. The general approach was to utilize a distributed electric propulsion (DEP) design. An essential component of this design was the High Lift Motor Controller (HLMC), a motor drive which provided power to the High Lift Motors (HLMs) and High Lift Propellers (HLPs) responsible for providing additional thrust for take-off and landing. This paper presents the detailed design, test results, and outcomes from the development of the HLMC, a 14 kW, 1kg, 98.3% efficient, outer mold line (OML) cooled, silicon carbide (SiC) MOSFET-based inverter and controller.
In this paper we present ion thruster design concepts created using the new computer codes that model performance limiting and erosion mechanisms. Presently, the codes model extraction grid ion optics and both discharge and neutralizer hollow cathodes.
The Solar Array System contracts awarded by NASA's Space Technology Mission Directorate are developing solar arrays in the 30 kW to 50 kW power range (beginning of life at 1 AU) that have significantly higher specific powers (W/kg) and much smaller stowed volumes than conventional rigid-panel arrays. The successful development of these solar array technologies has the potential to enable new types of solar electric propulsion (SEP) vehicles and missions. This paper describes a 30-kW electric propulsion vehicle built into an EELV Secondary Payload Adapter (ESPA) ring. The system uses an ESPA ring as the primary structure and packages two 15-kW Megaflex solar array wings, two 14-kW Hall thrusters, a hydrazine Reaction Control Subsystem (RCS), 220 kg of xenon, 26 kg of hydrazine, and an avionics module that contains all of the rest of the spacecraft bus functions and the instrument suite. Direct-drive is used to maximize the propulsion subsystem efficiency and minimize the resulting waste heat and required radiator area. This is critical for packaging a high-power spacecraft into a very small volume. The fully-margined system dry mass would be approximately 1120 kg. This is not a small dry mass for a Discovery-class spacecraft, for example, the Dawn spacecraft dry mass was only about 750 kg. But the Dawn electric propulsion subsystem could process a maximum input power of 2.5 kW, and this spacecraft would process 28 kW, an increase of more than a factor of ten. With direct-drive the specific impulse would be limited to about 2,000 s assuming a nominal solar array output voltage of 300 V. The resulting spacecraft would have a beginning of life acceleration that is more than an order of magnitude greater than the Dawn spacecraft. Since the spacecraft would be built into an ESPA ring it could be launched as a secondary payload to a geosynchronous transfer orbit significantly reducing the launch costs for a planetary spacecraft. The SEP system would perform the escape from Earth and then the heliocentric transfer to the science target.
Weight, efficiency and control requirements of subsystems of electric propulsion applied to Mars missions, selecting power-conditioning modules for ion propulsion engines
NASAs current focus is on the Journey to Mars sometime around the mid-to-late 2030s. However, it is also supporting the development of commercial cargo and crew delivery to the ISS (e.g., SpaceX, Orbital Sciences, SNC, Boeing) where inflatable habitation technology (e.g., Bigelow Aerospaces BEAM) is currently being tested Significant private sector interest in commercial lunar activities has also been expressed by Bigelow Aerospace, Golden Spike Company, Shackleton Energy Company (SEC), and most recently by United Launch Alliance (ULA) in their Cislunar-1000 plan Lunar-derived propellant (LDP) production specifically LLO2 and LLH2 offers significant mission leverage and are central themes of both SECs and ULAs plans for commercial lunar development. An efficient, proven propulsion technology with reuse capability like NTP offers the potential for affordable access through space essential to realizing commercial lunar missions.This presentation examines the performance potential of an evolutionary lunar transportation system (LTS) architecture using NTR initially, then transitioning to LANTR as LDPs(e.g., LLO2 from regolith or volcanic glass, LLO2 and LLH2 from lunar polar ice deposits) become available in lunar orbit (LO) Mission applications range from cargo delivery, to crewed landing, to routine commuter flights to and from transportation system nodes located in both lunar equatorial and lunar polar orbits. This presentation examines the performance potential of an evolutionary lunar transportation system (LTS) architecture using NTR initially, then transitioning to LANTR as LDPs (e.g., LLO2 from regolith or volcanic glass, LLO2 and LLH2 from lunar polar ice deposits) become available in lunar orbit (LO) Mission applications range from cargo delivery, to crewed landing, to routine commuter flights to and from transportation system nodes located in both lunar equatorial and lunar polar orbits.
Propulsion system efficiency increases as operating temperatures are increased. Some very high-temperature materials are being developed, including refractory metal alloys, carbides, borides, and silicides. System design requires data for materials properties at operating temperatures. Materials property data are not available for many materials of interest at the desired operating temperatures (up to approx. 3000 K). The objective of this work is to provide important physical property data at ultra-high temperatures. The MSFC Electrostatic levitation (ESL) facility can provide measurements of thermophysical properties which include: creep strength, density and thermal expansion for materials being developed for propulsion applications. The ESL facility uses electrostatic fields to position samples between electrodes during processing and characterization studies. Because the samples float between the electrodes during studies, they are free from any contact with a container or test apparatus. This provides a high purity environment for the study of high-temperature, reactive materials. ESL can be used to process a wide variety of materials including metals, alloys, ceramics, glasses and semiconductors. The MSFC ESL has provided non-contact measurements of properties of materials up to 3400 C. Density and thermal expansion are measured by analyzing digital images of the sample at different temperatures. Our novel, non-contact method for measuring creep uses rapid rotation to deform the sample. Digital images of the deformed samples are analyzed to obtain the creep properties, which match those obtained using ASTM Standard E-139 for Nb at 1985 C. Data from selected ESL-based characterization studies will be presented. The ESL technique could support numerous propulsion technologies by advancing the knowledge base and the technology readiness level for ultra-high temperature materials. Applications include non-eroding nozzle materials and lightweight, high-temperature alloys for turbines and structures.
Solar sails reflect photons streaming from the sun and transfer momentum to the sail. The thrust, though small, is continuous and acts for the life of the mission without the need for propellant. Recent advances in materials and ultra-low mass gossamer structures have enabled a host of useful missions utilizing solar sail propulsion. The team of L'Garde, Jet Propulsion Laboratories, Ball Aerospace, and Langley Research Center, under the direction of the NASA In-Space Propulsion office, has been developing a scalable solar sail configuration to address NASA s future space propulsion needs. The baseline design currently in development and testing was optimized around the 1 AU solar sentinel mission. Featuring inflatably deployed sub-T(sub g), rigidized beam components, the 10,000 sq m sail and support structure weighs only 47.5 kg, including margin, yielding an areal density of 4.8 g/sq m. Striped sail architecture, net/membrane sail design, and L'Garde's conical boom deployment technique allows scalability without high mass penalties. This same structural concept can be scaled to meet and exceed the requirements of a number of other useful NASA missions. This paper discusses the interim accomplishments of phase 3 of a 3-phase NASA program to advance the technology readiness level (TRL) of the solar sail system from 3 toward a technology readiness level of 6 in 2005. Under earlier phases of the program many test articles have been fabricated and tested successfully. Most notably an unprecedented 4-quadrant 10 m solar sail ground test article was fabricated, subjected to launch environment tests, and was successfully deployed under simulated space conditions at NASA Plum Brook s 30m vacuum facility. Phase 2 of the program has seen much development and testing of this design validating assumptions, mass estimates, and predicted mission scalability. Under Phase 3 a much larger 20 m square test article including subscale vane has been fabricated and tested. A 20 m system ambient deployment has been successfully conducted after enduring Delta-2 launch environment testing. The program will culminate in a vacuum deployment of a 20 m subscale test article at the NASA Glenn s Plum Brook 30 m vacuum test facility to bring the TRL level as close to 6 as possible in 1 g. This focused program will pave the way for a flight experiment of this highly efficient space propulsion technology.
Electrified aircraft propulsion (EAP) research is a priority of the National Aeronautics and Space Administration (NASA) for its potential to increase propulsion system efficiency, performance, and operability at the subsystem and vehicle levels while decreasing emissions. These EAP systems demand more advanced control algorithms due to increased complexity. NASA has developed a reconfigurable, hardware-in-the-loop rig to verify control algorithm performance using a sub-scale electro-mechanical system. A novel capability of this rig is the ability to test full scale EAP control algorithms on a sub-scale representation of the electro-mechanical system without turbomachinery/rotors. A novel feature is the use of a physical energy storage device within the electro-mechanical system. A dual spool, parallel hybrid-electric turbofan architecture and energy management control system is tested with the goal of verifying the ability to obtain turbomachinery model operability benefits while controlling sub-scale electro-mechanical hardware. Pre-test predictions of the turbofan model, control, and rig performance were obtained through simulation using a software model of the rig. Theoretical results showing the true performance of the turbofan model were obtained through a software simulation using full-scale mechanical shaft models. The paper compares theoretical, predicted, and actual test results from the turbofan model, energy management control and rig perspectives. The results show that the presence of sub-scale electro-mechanical hardware did not inhibit the energy management algorithm from achieving turbomachinery operability benefits.