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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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At least 109 records · Page 6
All sourcesNASA NTRSDOE OSTI
NASA NTRS1989

Advanced solar-propelled cargo spacecraft for Mars missions

At the University of Washington, three concepts for an unmanned, solar powered, cargo spacecraft for Mars-support missions have been investigated. These spacecraft are designed to carry a 50,000 kg payload from a low Earth orbit to a low Mars orbit. Each design uses a distinctly different propulsion system: a solar radiation absorption (SRA) system, a solar-pumped laser (SPL) system, and a solar powered mangetoplasmadynamic (MPD) arc system. The SRA directly converts solar energy to thermal energy in the propellant through a novel process developed at the University of Washington. A solar concentrator focuses sunlight into an absorption chamber. A mixture of hydrogen and potassium vapor absorbs the incident radiation and is heated to approximately 3700 K. The hot propellant gas exhausts through a nozzle to produce thrust. The SRA has an I(sub sp) of approximately 1000 sec and produces a thrust of 2940 N using two thrust chambers. In the SPL system, a pair of solar-pumped, multi-megawatt, CO2 lasers in sun-synchronous Earth orbit converts solar energy to laser energy. The laser beams are transmitted to the spacecraft via laser relay satellites. The laser energy heats the hydrogen propellant through a plasma breakdown process in the center of an absorption chamber. Propellant flowing through the chamber, heated by the plasma core, expands through a nozzle to produce thrust. The SPL has an I(sub sp) of 1285 sec and produces a thrust of 1200 N using two thrust chambers. The MPD system uses indium phosphide solar cells to convert sunlight to electricity, which powers the propulsion system. In this system, the argon propellant is ionized and electromagnetically accelerated by a magnetoplasmadynamic arc to produce thrust. The MPD spacecraft has an I(sub sp) of 2490 sec and produces a thrust of 100 N. Various orbital transfer options are examined for these concepts. In the SRA system, the mother ship transfers the payload into a very high Earth orbit and a small auxiliary propulsion system boosts the payload into a Hohmann transfer to Mars. The SPL spacecraft releases the payload as the spacecraft passes by Mars. Both the SRA-powered spacecraft and the SPL-powered spacecraft return to Earth for subsequent missions. The MPD-propelled spacecraft, however, remains at Mars as an orbiting space station. A patched conic approximation was used to determine a heliocentric interplanetary transfer orbit for the MPD propelled spacecraft. All three solar-powered spacecraft use an aerobrake procedure to place the payload into a low Mars parking orbit. The payload delivery times range from 160 days to 873 days (2.39 years).

Auziasdeturenne, J.↗
NASA NTRS1999

NASA Breakthrough Propulsion Physics Workshop Proceedings

In August 1997, NASA sponsored a 3-day workshop to assess the prospects emerging from physics that may eventually lead to creating propulsion breakthroughs -the kind of breakthroughs that could revolutionize space flight and enable human voyages to other star systems. Experiments and theories were discussed regarding the coupling of gravity and electromagnetism, vacuum fluctuation energy, warp drives and wormholes, and superluminal quantum tunneling. Because the propulsion goals are presumably far from fruition, a special emphasis was to identify affordable, near-term, and credible research tasks that could make measurable progress toward these grand ambitions. This workshop was one of the first steps for the new NASA Breakthrough Propulsion Physics program led by the NASA Lewis Research Center.

Millis, Marc G.↗
NASA NTRS1963

CYCLOTRON RESONANCE PROPULSION SYSTEM

Plasma-acceleration system in which power is transferred from an rf electromagnetic field to a plasma by means of cyclotron-resonance coupling

PLASMA ACCELERATOR↗
NASA NTRS2022

Development of a Loop Antenna Deployment System for TEM-Based Subsurface Mars Water Detection

The subsurface of Mars has the potential to harbor existing deposits of liquid water, which are of great interest both scientifically (in the search for life) and as resources for future astronauts living on the red planet. However, the depth to a potential subsurface aquifer may be kilometers deep, which is well beyond what current surface or orbital approaches can survey while also confirming the unique signature for liquid water. Inspired by critical technology for exploring deep aquifers on Earth, we are developing the Transient H2O Reconnaissance (TH2OR) instrument at NASA’s Jet Propulsion Laboratory (JPL). TH2OR operates on the principle of transient electromagnetics (TEM), which leverages induction and electromotive force to induce a current in a subsurface water body using a loop-shaped antenna that provides both transmit and receive functions when placed on the surface. On Mars, TEM may be even more effective given the relative dryness of the subsurface compared to a more conductive, saline, liquid, water body on Earth. However, to probe deep within the subsurface (below 5 km) a large enclosed antenna is needed – specifically, a 100-m diameter effective loop or a shape with roughly equivalent area that can transmit at low frequencies (kHz-Hz). The deployment of a large-scale structure on the surface is complicated by the fact that the transmit wire must be both lightweight and robust to contact with the surface. Further, TH2OR may be delivered to the surface by a non-mobile vehicle, so it is desired that the deployment can be activated from a static location. This paper provides an overview of the deployment trade study, focusing on our current, favored approach, using a projectile wire launcher. Building on past approaches in the literature, we have developed and fielded an Earth-based, gas-projectile prototype for launching a triangle-shaped antenna onto analogue terrain. Our results compare simulated launch performance to actual field tests conducted under Earth gravity and pressure conditions. We discuss how Earth performance maps to a prospective Mars deployment under reduced gravity and pressure. We also provide lessons learned and next steps towards the development of an integrated TH2OR instrument for finding water on Mars.

Stamenkovic, Vlada↗
NASA NTRSDate not supplied

EMC Lessons on the Path to Flight

This slide presentation summarizes Electromagnetic Compatibility lessons learned over a sampling of space and aerospace hardware development projects.

Electromagnetic Interference↗
NASA NTRS1998

NASA Breakthrough Propulsion Physics Program

In 1996, NASA established the Breakthrough Propulsion Physics program to seek the ultimate breakthroughs in space transportation: propulsion that requires no propellant mass, propulsion that attains the maximum transit speeds physically possible, and breakthrough methods of energy production to power such devices. Topics of interest include experiments and theories regarding the coupling of gravity and electromagnetism, vacuum fluctuation energy, warp drives and worm-holes, and superluminal quantum effects. Because these propulsion goals are presumably far from fruition, a special emphasis is to identify affordable, near-term, and credible research that could make measurable progress toward these propulsion goals. The methods of the program and the results of the 1997 workshop are presented. This Breakthrough Propulsion Physics program, managed by Lewis Research Center, is one part of a comprehensive, long range Advanced Space Transportation Plan managed by Marshall Space Flight Center.

Millis, Marc G.↗
NASA NTRS1990

Space Station Freedom propulsion activities

The technical highlights and accomplishments made at NASA LeRC in the development of the Space Station Freedom (SSF) propulsion system are discussed. The objectives are as follows: develop and characterize resistojet-thruster components and assemblies; develop and characterize hydrogen-oxygen thruster components; and conduct system trade studies. The research projects primarily characterize propulsion performance and life. Other tests include environmental impacts, such as exhaust gas profiles and electromagnetic interference. The technical activities that are highlighted are being conducted at LeRC within the Aerospace Technology and Space Station Freedom directorates. These activities include the following: derivation of design analysis models; trade studies of design options; propulsion system impact studies; and component testing for characterization and design verification.

Spera, David A.↗
NASA NTRS2017

Low and High-Power Inductive Pulsed Plasma Thruster Development Testing at NASA-MSFC

The inductive pulsed plasma thruster (IPPT) is an electromagnetic plasma accelerator that has been identified in NASA roadmaps as an enabling propulsion technology for some niche low-power missions and for high-power in-space propulsion needs. The IPPT is an electrodeless space propulsion device where a capacitor is charged to an initial voltage and then discharged producing a high current pulse through a coil. The field produced by this pulse ionizes propellant, inductively driving current in a plasma located near the face of the coil. Once the plasma is formed it can be accelerated and expelled at a high exhaust velocity by the electromagnetic Lorentz body force arising from the interaction of the induced plasma current and the magnetic field produced by the current in the coil. Thrusters of this type possess many demonstrated and potential benefits that make them worthy of continued investigation. The electrodeless nature of these thrusters eliminates the lifetime and contamination issues associated with electrode erosion in conventional electric thrusters. Also, a wider variety of propellants are accessible when compatibility with metallic electrodes in no longer an issue. IPPTs have been successfully operated using propellants like ammonia, hydrazine, and CO2, and there is no fundamental reason why they would not operate on other in situ propellants like H2O. It is well-known that pulsed accelerators can maintain constant specific impulse (I(sub sp)) and thrust efficiency (eta(sub t)) over a wide range of input power levels by adjusting the pulse rate to hold the discharge energy per pulse constant. It has also been demonstrated that an inductive pulsed plasma thruster can operate in a regime where eta(sub t) is relatively constant over a wide range of I(sub sp) values (3000-8000 s). Finally, thrusters in this class have operated in single-pulse mode at high energy per pulse, and by increasing the pulse rate they offer the potential to process very high levels of power using a single thruster. There has been significant previous research on IPPTs designed around a planar-coil (flat-plate) geometry. The most notable of these was the Pulsed Inductive Thruster (PIT), with the PIT MkV presently representing the state-of- the-art in pulsed high-power IPPT technological development. In this paper, we focus on two planar-geometry devices that operate at significantly different power levels. Most work performed at NASA-Marshall Space Flight Center (MSFC) has, to date, focused on lower power thruster operation (approx. = 10s to 100s of J/pulse, up to 2-2.5 kW average power throughput) and previously described. The most recent work aimed to assemble a device that could be tested in cyclic mode on a thrust-stand, and which could augment the existing data set for IPPTs. In addition, the thruster was designed to serve as a test-bed for solid state switching circuitry and pulsed gas valves, with the modular design of the device allowing for variation in or upgrades to test configuration. Recently, MSFC obtained on loan from the Georgia Institute of Technology (Atlanta, GA) the PIT MkVI, successor to the PIT MkV. The MkV and MkVI are similar in design with much of the hardware from the former, specifically the capacitors and spark-gap switches, being reused in the latter. The coil is similar in geometry but has bent copper rods used in the latest iteration in place of the Litz wire windings found in the MkV. The MkVI master switch for the spark gaps is located in the vacuum chamber contained within a sealed, pressurized vessel fastened to the back of the thruster. This is different from the MkV where many capacitor charging lines and spark gap-triggering delay lines ran to the thruster from a master trigger located outside the vacuum chamber. The MkVI was damaged during testing soon after its fabrication was completed. The thruster arrived at MSFC still-damaged and mostly disassembled into many individual pieces. The device has been repaired, with a few additional design changes implemented after discussions with the late Prof. Lovberg regarding the initial testing results and issues encountered. In the present work, we present results from testing of both the small IPPT and the larger MkVI thruster. The smaller device (Fig. 1) is tested on a thrust stand on multiple gases to demonstrate its capability to operate in a repetition-rate mode and serve as a IPPT technology-development testbed. The larger MkVI (Fig. 2) is operated for the first time in its newly reconstituted state, demonstrating full-power pulsed operation and, for the first time, repetition-rate operation of a high-power IPPT. The additional upgrades required for synchronous operation of all the pulsed systems in single-pulse and repetition-rate mode are described in detail.

Polzin, Kurt A.↗
NASA NTRS2018

Inductive Pulsed Plasma Thruster Development Testing at NASA-MSFC

The inductive pulsed plasma thruster (IPPT) is an electromagnetic plasma accelerator that has been identified in NASA roadmaps as an enabling propulsion technology for some niche low-power missions and for high-power in-space propulsion needs. The IPPT is an electrodeless space propulsion device where a capacitor is charged to an initial voltage and then discharged producing a high current pulse through a coil. The field produced by this pulse ionizes propellant, inductively driving current in a plasma located near the face of the coil. Once the plasma is formed it can be accelerated and expelled at a high exhaust velocity by the electromagnetic Lorentz body force arising from the interaction of the induced plasma current and the magnetic field produced by the current in the coil. Thrusters of this type possess many demonstrated and potential benefits that make them worthy of continued investigation. The electrodeless nature of these thrusters eliminates the lifetime and contamination issues associated with electrode erosion in conventional electric thrusters. Also, a wider variety of propellants are accessible when compatibility with metallic electrodes in no longer an issue. IPPTs have been successfully operated using propellants like ammonia, hydrazine, and CO2, and there is no fundamental reason why they would not operate on other in situ propellants like H2O. It is well-known that pulsed accelerators can maintain constant specific impulse (I(sub sp)) and thrust efficiency (η(sub t)) over a wide range of input power levels by adjusting the pulse rate to hold the discharge energy per pulse constant. It has also been demonstrated that an inductive pulsed plasma thruster can operate in a regime where η(sub t) is relatively constant over a wide range of I(sub sp) values (3000-8000 s). Finally, thrusters in this class have operated in single-pulse mode at high energy per pulse, and by increasing the pulse rate they offer the potential to process very high levels of power using a single thruster. There has been significant previous research on IPPTs designed around a planar-coil (flat-plate) geometry. The most notable of these was the Pulsed Inductive Thruster (PIT), with the PIT MkV presently representing the state of- the-art in pulsed high-power IPPT technological development. In this paper, we focus on two planar-geometry devices that operate at significantly different power levels. Most work performed at NASA-Marshall Space Flight Center (MSFC) has, to date, focused on lower power thruster operation (≈ 10s to 100s of J/pulse, up to 2-2.5 kW average power throughput) and previously described in Refs. [5,6]. The most recent work aimed to assemble a device that could be tested in cyclic mode on a thrust-stand, and which could augment the existing data set for IPPTs. In addition, the thruster was designed to serve as a test-bed for solid state switching circuitry and pulsed gas valves, with the modular design of the device allowing for variation in or upgrades to test configuration. Recently, MSFC obtained on loan from the Georgia Institute of Technology (Atlanta, GA) the PIT MkVI, successor to the PIT MkV. The MkV and MkVI are similar in design with much of the hardware from the former, specifically the capacitors and spark-gap switches, being reused in the latter. The coil is simliar in geometry but has bent copper rods used in the latest iteration in place of the Litz wire windings found in the MkV. The MkVI master switch for the spark gaps is located in the vacuum chamber contained within a sealed, pressurized vessel fastened to the back of the thruster. This is different from the MkV where many capacitor charging lines and spark gap-triggering delay lines ran to the thruster from a master trigger located outside the vacuum chamber. The MkVI was damaged during testing soon after its fabrication was completed. The thruster arrived at MSFC still-damaged and mostly disassembled into many individual pieces. The device has been repaired, with a few additional design changes implemented after discussions with the late Prof. Lovberg regarding the initial testing results and issues encountered.

Polzin, Kurt A.↗
NASA NTRS2011

Comparison Between MAD-IPA Thrust Stand Measurements and Performance Modeling

The high specific impulse values associated with electric propulsion (EP) that allow for higher payload fractions than conventional, chemical propulsion are achieved by the acceleration of ionized propellant (plasma) by electromagnetic body forces. The lifetime of many EP devices is limited by electrode erosion caused by direct plasma-electrode interaction, while the efficiency is often limited by, among other thruster-specific factors, the available power in space. The efficiency can be increased during higher power operation since the amount of power required to ionize the propellant is fixed and decreases as a percentage of the increased input power. Pulsed inductive plasma accelerators [1-3] are a potentially elegant solution to the problems of high power demands and electrode erosion. The former is addressed in the pulsed nature of these devices, which allows for instantaneous operation on the order of megawatts while requiring a continuous supply on the order of only kilowatts, while the latter is addressed by inductive coupling of the thruster to ionized propellant. These devices operate by rapidly (on the order of microseconds) discharging stored energy at a given pulse repetition rate through a coil creating time-varying magnetic and electric fields that cause ionization, current formation and acceleration of propellant away from the coil. While the inductive (contact-less) nature of the energy transfer from the thruster to the propellant alleviates the problem of erosion and enables the use of in-situ and storable propellants incompatible with metallic electrodes, it places high voltage demands (on the order of tens of kilovolts) on the energy storage system to achieve propellant ionization.

Hallock, Ashley K.↗
NASA NTRS2016

Design of an Electric Propulsion System for SCEPTOR

The rise of electric propulsion systems has pushed aircraft designers towards new and potentially transformative concepts. As part of this effort, NASA is leading the SCEPTOR program which aims at designing a fully electric distributed propulsion general aviation aircraft. This article highlights critical aspects of the design of SCEPTOR's propulsion system conceived at Joby Aviation in partnership with NASA, including motor electromagnetic design and optimization as well as cooling system integration. The motor is designed with a finite element based multi-objective optimization approach. This provides insight into important design tradeoffs such as mass versus efficiency, and enables a detailed quantitative comparison between different motor topologies. Secondly, a complete design and Computational Fluid Dynamics analysis of the air breathing cooling system is presented. The cooling system is fully integrated into the nacelle, contains little to no moving parts and only incurs a small drag penalty. Several concepts are considered and compared over a range of operating conditions. The study presents trade-offs between various parameters such as cooling efficiency, drag, mechanical simplicity and robustness.

SCEPTOR↗
NASA NTRS2003

Efficient Plasma Production in Low Background Neutral Pressures with the M2P2 Prototype

Mini-Magnetospheric Plasma Propulsion (M2P2) seeks the creation of a large-scale (10 km radius) magnetic wall or bubble (i.e. a magnetosphere) by the electromagnetic inflation of a small-scale (20 cm radius) dipole magnet. The inflated magnetosphere will intercept the solar wind and thereby provide high-speed propulsion with modest power and fuel requirements due to the gain provided by the ambient medium. Magnetic field inflation is produced by the injection of plasma onto the dipole magnetic field eliminating the need for large mechanical structures and added material weight at launch. For successful inflation of the magnetic bubble a beta near unity must be achieved along the imposed dipole field. This is dependent on the plasma parameters that can be achieved with a plasma source that provide continuous operation at the desired power levels of 1 to 2 kilowatts. Over the last two years we have been developing a laboratory prototype to demonstrate the inflation of the magnetic field under space-like conditions. In this paper we will present some of the latest results from the prototype development at the University of Washington and show that the prototype can produce high ionization efficiencies while operating in near space like neutral background pressures producing electron temperatures of a few tens of electron volts. This allows for operation with propellant expenditures lower than originally estimated.

Ziemba, T.↗
NASA NTRS1973

Electromagnetic interference onboard an electrically propelled spacecraft

Electric and magnetic interference generated by a power conditioner and ion thruster were measured at JPL. Test results were judged by comparing the measured levels to the levels defined by the solar electric propulsion (SEP) specification that was derived from the existing mission oriented Viking-Orbiter 1975 and Mariner-Jupiter-Saturn 1977 electromagnetic compatibility specifications. In many areas the measured interference was excessive. In some cases, the cause of interference was identified. It was concluded that it should be possible on future designs to improve the EMI characteristics greatly at a modest penalty to weight and efficiency.

Macie, T. W.↗
NASA NTRS1984

Measurements of energy distribution in microwave plasmas of N2 and He and comparisons with results for H2

An electrothermal propulsion concept utilizing a microwave plasma system as the mechanism to convert electromagnetic energy into translational energy of the flowing gas is being investigated. Specifically, this study compares the energy transfer characteristics of three different gases, H2, N2, and He, to gain some insight as to the dominant energy transfer processes present in a microwave plasma. A calorimetric experimental system has been designed and built enclosing the microwave plasma system to accurately determine the net energy transferred to the flowing gas. Results are obtained for N2 and He discharges and compared with previously reported experimental results for H2.

Chapman, R.↗
NASA NTRS1985

Measurements of energy distribution and wall temperature in flowing hydrogen microwave plasma systems

An electrothermal propulsion concept utilizing a microwave plasma system as the mechanism to convert electromagnetic energy into translational energy of the flowing gas is being investigated. A calorimetric experimental system has been designed and built enclosing the microwave plasma system to accurately determine the net energy transferred to the flowing gas. For a flow rate of 8900 micromoles/sec, a pressure of 7.4 torr, and an absorbed power level of 80 W, an energy transfer efficiency of 50 percent has been measured. A heat transfer model that characterizes the energy transfer processes in the plasma is developed. A wall temperature for the plasma system is calculated.

Chapman, R.↗
NASA NTRS1985

Measurement of energy distribution in flowing hydrogen microwave plasmas

An electrothermal propulsion concept utilizing a microwave plasma system as the mechanism to convert electromagnetic energy into kinetic energy of a flowing gas is investigated. A calorimetry system enclosing a microwave plasma system has been developed to accurately measure the energy inputs and outputs of the microwave plasma system. The rate of energy transferred to the gas can be determined to within + or - 1.8 W from an energy balance around the microwave plasma system. The percentage of the power absorbed by the microwave plasma system transferred to the hydrogen gas as it flows through the system is found to increase with the increasing flow rate, to decrease with the increasing pressure, and to be independent of the absorbed power. An upper bound for the hydrogen gas temperature is estimated from the energy content, heat capacity, and flow rate of the gas stream. A lower bound for an overall heat-transfer coefficient is then calculated, characterizing the energy loss from the hydrogen gas stream to the air cooling of the plasma discharge tube wall. The heat-transfer coefficient is found to increase with the increasing flow rate and pressure and to be independent of the absorbed power. This result indicates that a convective-type mechanism is responsible for the energy transfer.

Chapman, R.↗