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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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All sourcesNASA NTRSDOE OSTI
NASA NTRS2001

Development of Aluminum Composites for a Rocket Engine's Lightweight Thrust Cell

The Aerospike liquid fueled rocket engine for the X-33 aerospace vehicle consists of several thrust cells, which can comprise as much as 25% of the engine weight. The interior wall of the thrust cell chamber is exposed to high temperature combustion products and must be cooled by using liquid hydrogen. Ultimately, reducing engine weight would increase vehicle performance and allow heavier payload capabilities. Currently, the thrust cell's structural jacket and manifolds are made of stainless steel 347, which can potentially be replaced by a lighter material such as an Aluminum (Al) Metal Matrix Composites (MMC). Up to 50% weight reduction can be expected for each of the thrust cell chambers using particulate SiC reinforced Al MMC. Currently, several Al MMC thrust cell structural jackets have been produced, using cost-effective processes such as gravity casting and plasma spray deposition, to demonstrate MMC technology readiness for NASA's advanced propulsion systems.

Lee, Jonathan A.↗
NASA NTRS2009

Neutralizer Hollow Cathode Simulations and Validation with Experiments

The fidelity of electric propulsion physics-based models depends largely on the validity of their predictions over a range of operating conditions and geometries. In general, increased complexity of the physics requires more extensive comparisons with laboratory data to identify the region(s) that lie outside the validity of the model assumptions and to quantify the uncertainties within its range of application. This paper presents numerical simulations of neutralizer hollow cathodes at various operating conditions and orifice sizes. The simulations were performed using a two-dimensional axisymmetric model that solves numerically a relatively extensive system of conservation laws for the partially-ionized gas in these devices. The results for the plasma are compared directly with Langmuir probe measurements. The computed keeper voltages are also compared with the observed values. Wherever model inputs and/or specific physics of the cathode discharge are uncertain, additional sensitivity calculations have been performed to quantify the uncertainties.

partially ionized gases↗
NASA NTRS2018

International Space Station Spacecraft Charging Environments: Modeling, Measurement and Implications for Future Human Space Flight Programs

Spacecraft charging analysis and migration is an interdisciplinary subject combining aspects of electrostatics, plasma physics, ionizing radiation, and materials science, as well as electronic system electromagnetic interference and compatibility (EMI/EMC) effects. Spacecraft charging hazards are caused by the accumulation of electrical charge on spacecraft and spacecraft components produced by interactions with space plasmas, energetic charged particles, and solar UV photons as well as spacecraft electrical power and propulsion systems operations. Spacecraft charging hazard effects include both hard and soft avionics and electrical power system anomalies and have led to the partial or complete loss of numerous spacecraft. The International Space Station (ISS) orbital altitude and inclination (~400 km and 51.6o) determined the dominant natural environment factors affecting ISS spacecraft charging; high speed flight through the geomagnetic field and electrical power system interaction with the cold, high-density ionospheric plasma. In addition ISS is exposed to energetic auroral electrons at high latitude. In this paper we present the results of ISS spacecraft charging modeling and measurements and compare the measurements with numerical modeling of ISS charging processes. ISS is a large metallic structure and flight through the geomagnetic field at orbital speed dominates ISS charging. Collection of ionospheric electrons by the large 160V PV arrays is the next largest contributor. Charging by auroral electrons is detectable but makes a relatively minor contribution. Finally we report the observation of short duration (~ 1 sec) rapid charging peaks associated with shunt/un-shunt operations of the 160V PV arrays, a phenomena not predicted before flight. ISS spacecraft charging environments are radically different from those encountered at higher altitudes in Earth?s magnetosphere and in cis-Lunar space. We present a brief review of those charging environments and an assessment of the applicability of ISS spacecraft charging management and experience to future human spaceflight programs beyond LEO.

Koontz, Steven L.↗
NASA NTRSDate not supplied

Characterizing Electrical Arc Behavior in Ion Engines

Gridded ion engines represent a class of electric thrusters characterized by high propulsive efficiency and high total impulse capabilities. These engines operate by applying high voltage to multi-aperture electrodes, thereby electrostatically accelerating plasma to generate thrust. Within the course of normal operations, periodic occurrences of high voltage breakdowns manifest within the thruster. These breakdowns constitute an inherent aspect of operation, with observations indicating a diminishing frequency and total count over the thruster's operational lifespan. To safeguard the propulsion system during such occurrences, the power processing unit (PPU) employs an automated arc extinction sequence, termed "recycle," which is triggered upon detection of an over-current condition. The initial phase of this sequence involves momentarily deactivating the high voltage and reducing the discharge current to a predetermined level, minimizing plasma impingement as the ion beam is defocused. The high voltage is subsequently reapplied, and the discharge current is returned to its nominal value. The entire sequence typically spans approximately 800 milliseconds. Throughout a recycle event, the propulsion system experiences various current and voltage transients as the arc current traverses the system and different power supplies are toggled on and off.

electric propulsion↗
NASA NTRS2002

Antimatter Driven P-B11 Fusion Propulsion System

One of the major advantages of using P-B11 fusion fuel is that the reaction produces only charged particles in the form of three alpha particles and no neutrons. A fusion concept that lends itself to this fuel cycle is the Magnetically Insulated Inertial Confinement Fusion (MICF) reactor whose distinct advantage lies in the very strong magnetic field that is created when an incident particle (or laser) beam strikes the inner wall of the target pellet. This field serves to thermally insulate the hot plasma from the metal wall thereby allowing thc plasma to burn for a long time and produce a large energy magnification. If used as a propulsion device, we propose using antiprotons to drive the system which we show to be capable of producing very large specific impulse and thrust. By way of validating the confinement propenies of MICF we will address a proposed experiment in which pellets coated with P-B11 fuel at the appropriate ratio will be zapped by a beam of antiprotons that enter the target through a hole. Calculations showing the density and temperature of the generated plasma along with the strength of the magnetic field and other properties of the system will be presented and discussed.

Kammash, Terry↗
NASA NTRS1990

The NASA-Lewis program on fusion energy for space power and propulsion, 1958-1978

An historical synopsis is provided of the NASA-Lewis research program on fusion energy for space power and propulsion systems. It was initiated to explore the potential applications of fusion energy to space power and propulsion systems. Some fusion related accomplishments and program areas covered include: basic research on the Electric Field Bumpy Torus (EFBT) magnetoelectric fusion containment concept, including identification of its radial transport mechanism and confinement time scaling; operation of the Pilot Rig mirror machine, the first superconducting magnet facility to be used in plasma physics or fusion research; operation of the Superconducting Bumpy Torus magnet facility, first used to generate a toroidal magnetic field; steady state production of neutrons from DD reactions; studies of the direct conversion of plasma enthalpy to thrust by a direct fusion rocket via propellant addition and magnetic nozzles; power and propulsion system studies, including D(3)He power balance, neutron shielding, and refrigeration requirements; and development of large volume, high field superconducting and cryogenic magnet technology.

Schulze, Norman R.↗
NASA NTRS1989

Electric Propulsion for Manned Mars Exploration

Advanced high-power electric propulsion systems can significantly enhance piloted Mars missions. An increase in the science payload delivered to Mars and the reduction of the total Earth-departure mass are the major system-level benefits of electric propulsion. Other potential benefits are the return of the cargo vehicle to Earth orbit and the availability of high power in Mars orbit for high-power science and communications. Parametric analyses for sizing the cargo mission vehicle for Mars exploration missions are presented. The nuclear-electric propulsion system thruster size, power level, mass, propellant type and payload mass capability are considered in these system-level trade studies. Descriptions of the propulsion system selection issues for both ion and MPD thruster technologies are also discussed. On a manned Mars mission, the total launch mass for an unmanned cargo vehicle in low earth orbit (LEO) can be reduced by up to 50 percent over the baseline oxygen/hydrogen propulsion system. Because the cargo vehicle is sent to Mars prior to the manned mission, the trip time for the vehicle is not a critical factor. By taking advantage of the high specific impulse (I sub sp) of an ion or a Magneto-Plasma-Dynamic (MPD) thruster system, the total LEO mass is reduced from 590,000 kg for the oxygen/hydrogen propulsion system to 309,000 kg for the MPD system and 295,000 kg for the ion system. Many factors must be analyzed in the design of a electric propulsion Mars cargo vehicle. The propellant selection, the number of thrusters, the power level and the specific impulse are among the most important of the parameters. To fully address the electric propulsion system design, trade studies for the differing ion and MPD propulsion system configurations (thruster power levels, number of thrusters, propellants and power systems) must be conducted.

Palaszewski, Bryan↗
NASA NTRS2001

Langmuir Probe Spacecraft Potential End Item Specification Document

This document describes the Langmuir Probe Spacecraft Potential (LPSP) investigation of the plasma environment in the vicinity of the ProSEDS Delta II spacecraft. This investigation will employ a group of three (3) Langmuir Probe Assemblies, LPAs, mounted on the Delta II second stage to measure the electron density and temperature (n(sub e) and T(sub e)), the ion density (n(sub i)), and the spacecraft potential (V(sub s)) relative to the surrounding ionospheric plasma. This document is also intended to define the technical requirements and flight-vehicle installation interfaces for the design, development, assembly, testing, qualification, and operation of the LPSP subsystem for the Propulsive Small Expendable Deployer System (ProSEDS) and its associated Ground Support Equipment (GSE). This document also defines the interfaces between the LPSP instrument and the ProSEDS Delta II spacecraft, as well as the design, fabrication, operation, and other requirements established to meet the mission objectives. The LPSP is the primary measurement instrument designed to characterize the background plasma environment and is a supporting instrument for measuring spacecraft potential of the Delta II vehicle used for the ProSEDS mission. Specifically, the LPSP will use the three LPAs equally spaced around the Delta II body to make measurements of the ambient ionospheric plasma during passive operations to aid in validating existing models of electrodynamic-tether propulsion. These same probes will also be used to measure Delta II spacecraft potential when active operations occur. When the electron emitting plasma contractor is on, dense neutral plasma is emitted. Effective operation of the plasma contactor (PC) will mean a low potential difference between the Delta II second stage and the surrounding plasma and represents one of the voltage parameters needed to fully characterize the electrodynamic-tether closed circuit. Given that the LP already needs to be well away from any near-field disturbances around the Delta II, it is possible to use the same probe with a simple reconfiguration of the electronics to measure potential with respect to the ambient plasma. The LP measurement techniques are outlined in the following text and discussed in detail in the Appendix. The scientific goals of the investigation, the physical and electrical characteristics of the instrument, and the on-orbit measurement requirements are also discussed in this document.

Gilchrist, Brian↗
NASA NTRS2003

High-Energy Two-Stage Pulsed Plasma Thruster

A high-energy (28 kJ per pulse) two-stage pulsed plasma thruster (MSFC PPT-1) has been constructed and tested. The motivation of this project is to develop a high power (approximately 500 kW), high specific impulse (approximately 10000 s), highly efficient (greater than 50%) thruster for use as primary propulsion in a high power nuclear electric propulsion system. PPT-1 was designed to overcome four negative characteristics which have detracted from the utility of pulsed plasma thrusters: poor electrical efficiency, poor propellant utilization efficiency, electrode erosion, and reliability issues associated with the use of high speed gas valves and high current switches. Traditional PPTs have been plagued with poor efficiency because they have not been operated in a plasma regime that fully exploits the potential benefits of pulsed plasma acceleration by electromagnetic forces. PPTs have generally been used to accelerate low-density plasmas with long current pulses. Operation of thrusters in this plasma regime allows for the development of certain undesirable particle-kinetic effects, such as Hall effect-induced current sheet canting. PPT-1 was designed to propel a highly collisional, dense plasma that has more fluid-like properties and, hence, is more effectively pushed by a magnetic field. The high-density plasma loading into the second stage of the accelerator is achieved through the use of a dense plasma injector (first stage). The injector produces a thermal plasma, derived from a molten lithium propellant feed system, which is subsequently accelerated by the second stage using mega-amp level currents, which eject the plasma at a speed on the order of 100 kilometers per second. Traditional PPTs also suffer from dynamic efficiency losses associated with snowplow loading of distributed neutral propellant. The twostage scheme used in PPT-I allows the propellant to be loaded in a manner which more closely approximates the optimal slug loading. Lithium propellant was chosen to test whether or not the reduced electrode erosion found in the Lithium Lorentz Force Accelerator (LiLFA) could also be realized in a pulsed plasma thruster. The use of the molten lithium dense plasma injector also eliminates the need for a gas valve and electrical switch; the injector design fulfills both roles, and uses no moving parts to provide, in principle, a highly reliable propellant feed and electrical switching system. Experimental results reported in this paper include: second-stage current traces, high-speed photographic and holographic imaging of the thruster exit plume, and internal mapping of the discharge chamber magnetic field from B-dot probe data. The magnetic field data is used to create a two-dimensional description of the evolution of the current sheet inside the thruster.

Markusic, Tom↗
NASA NTRS1989

Electrostatic Plasma Accelerator (EPA)

The Electrostatic Plasma Accelerator (EPA) is a thruster concept which promises specific impulse levels between low power arcjets and those of the ion engine while retaining the relative simplicity of the arcjet. The EPA thruster produces thrust through the electrostatic acceleration of a moderately dense plasma. No accelerating electrodes are used and the specific impulse is a direct function of the applied discharge voltage and the propellant atomic mass. The goal of the present program is to demonstrate feasibility of the EPA thruster concept through experimental and theoretical investigations of the EPA acceleration mechanism and discharge chamber performance. Experimental investigations will include operating the test bed ion (TBI) engine as an EPA thruster and parametrically varying the thruster geometry and operating conditions to quantify the electrostatic plasma acceleration effect. The theoretical investigations will include the development of a discharge chamber model which describes the relationships between the engine size, plasma properties, and overall performance. For the EPA thruster to be a viable propulsion concept, overall thruster efficiencies approaching 30% with specific impulses approaching 1000 s must be achieved.

Brophy, John R.↗
NASA NTRS1995

Plasma Sprayed Bondable Stainless Surface (BOSS) Coatings for Corrosion Protection and Adhesion Treatments

Plasma-sprayed Bondable Stainless Surface (BOSS) coatings are being developed under the Solid Propulsion Integrity Program's (SPIP) Bondlines Package. These coatings are designed as a steel case preparation treatment prior to insulation lay-up. Other uses include the exterior of steel cases and bonding surfaces of nozzle components. They provide excellent bondability - rubber insulation and epoxy bonds fail cohesively within the polymer - for both fresh surfaces and surfaces having undergone natural and accelerated environmental aging. They have passed the MSFC requirements for protection of inland and sea coast environment. Because BOSS coatings are inherently corrosion resistant, they do not require preservation by greases or oils. The reduction/elimination of greases and oils, known bondline degraders, can increase SRM reliability, decrease costs by reducing the number of process steps, and decrease environmental pollution by reducing the amount of methyl chloroform used for degreasing and thus reduce release of the ozone-depleting chemical in accordance with the Clean Air Act and the Montreal Protocol. The coatings can potential extend the life of RSRM case segments and nozzle components by eliminating erosion due to multiple grit blasting during each use cycle and corrosion damage during marine recovery. Concurrent work for the Air Force show that other BOSS coatings give excellent bondline strength and durability for high-performance structures of aluminum and titanium.

Davis, G. D.↗
NASA NTRS2000

Gasdynamic Mirror Fusion Propulsion Experiment

A gasdynamic mirror (GDM) fusion propulsion experiment is currently being constructed at the NASA Marshall Space Flight Center (MSFC) to test the feasibility of this particular type of fusion device. Because of the open magnetic field line configuration of mirror fusion devices, they are particularly well suited for propulsion system applications since they allow for the easy ejection of thrust producing plasma. Currently, the MSFC GDM is constructed in three segments. The vacuum chamber mirror segment, the plasma injector mirror segment, and the main plasma chamber segment. Enough magnets are currently available to construct up to three main plasma chamber segments. The mirror segments are also segmented such that they can be expanded to accommodate new end plugging strategies with out requiring the disassembly of the entire mirror segment. The plasma for the experiment is generated in a microwave cavity located between the main magnets and the mirror magnets. Ion heating is accomplished through ambipolar diffusion. The objective of the experiment is to investigate the stability characteristics of the gasdynamic mirror and to map a region of parameter space within which the plasma can be confined in a stable steady state configuration. The mirror ratio, plasma density, and plasma "b" will be varied over a range of values and measurements subsequently taken to determine the degree of plasma stability.

Emrich, Bill↗
NASA NTRS1999

Enabling Exploration of Deep Space: High Density Storage of Antimatter

Portable electromagnetic antiproton traps are now in a state of realization. This allows facilities like NASA Marshall Space Flight Center to conduct antimatter research remote to production sites. MSFC is currently developing a trap to store 10(exp 12) antiprotons for a twenty-day half-life period to be used in future experiments including antimatter plasma guns, antimatter-initiated microfusion, and the synthesis of antihydrogen for space propulsion applications. In 1998, issues including design, safety and transportation were considered for the MSFC High Performance Antimatter Trap (HiPAT). Radial diffusion and annihilation losses of antiprotons prompted the use of a 4 Tesla superconducting magnet and a 20 KV electrostatic potential at 10(exp -12) Torr pressure. Cryogenic fluids used to maintain a trap temperature of 4K were sized accordingly to provide twenty days of stand-alone storage time (half-life). Procurement of the superconducting magnet with associated cryostat has been completed. The inner, ultra-high vacuum system with electrode structures has been fabricated, tested and delivered to MSFC along with the magnet and cryostat. Assembly of these systems is currently in progress. Testing under high vacuum conditions, using electrons and hydrogen ions will follow in the months ahead.

Smith, Gerald A.↗
NASA NTRS2004

Compact Plasma Accelerator

A plasma accelerator has been conceived for both material-processing and spacecraft-propulsion applications. This accelerator generates and accelerates ions within a very small volume. Because of its compactness, this accelerator could be nearly ideal for primary or station-keeping propulsion for spacecraft having masses between 1 and 20 kg. Because this accelerator is designed to generate beams of ions having energies between 50 and 200 eV, it could also be used for surface modification or activation of thin films.

Foster, John E.↗
DOE OSTI2024

Sunbeam: Near-sun statites as beam platforms for beam-driven rockets

Here, we outline a method of beamed power for propulsion that utilizes relativistic electron beams. The physics of charged particle beam propagation in the space plasma environment is discussed and the long-range (> 100 AU) advantage of relativistic electron beams is emphasized. A preliminary statite-based beam emitter for powering probes to 0.1c is proposed and the challenges in beamed-power uses are explored.

33 ADVANCED PROPULSION SYSTEMS↗