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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 127 records · Page 7
All sourcesNASA NTRSDOE OSTI
NASA NTRS2003

Investigation of Impinging Stream Vortex Chamber Concepts for Liquid Rocket Engine Applications

This paper addresses the progress of technology development of a laser ignition system at NASA Marshall Spac Flight Center (MSFC). Initial hot-fire tests in a small-scale rocket chamber at MSFC have demonstrated the dual pulse laser-induced spark (DPLIS) technique, which has an advantage over existing single-pulse laser ignition techniques in that it can be optimized in its laser pulse format to maximize the initial plasma volume, the plasm: lifetime, as well as the flame kernel growth rate. The distribution of the total laser energy into two separate pulse also lowers the peak power that would need to be sent through fiber optics to the combustion chamber, making the implementation of this technique more practical than other single-pulse techniques. A first generation prototype of an optic fiber-coupled laser ignition system will be tested a rocket chamber with RP-1/GOX and GH2/GOX propellants systems. Other relevant technology, such as optical windows, flight-qualified laser system etc. will be discussed in this paper.

Trinh, Huu P.↗
NASA NTRS2000

Development of a Methodology for Conducting Hall Thruster EMI Tests in Metal Vacuum Chambers of Arbitrary Shape and Size

While the closed-drift Hall thruster (CDT) offers significant improvement in performance over conventional chemical rockets and other advanced propulsion systems such as the arcjet, its potential impact on spacecraft communication signals must be carefully assessed before widespread use of this device can take place. To this end, many of the potentially unique issues that are associated with these thrusters center on its plume plasma characteristics and the its interaction with electromagnetic waves. Although a great deal of experiments have been made in characterizing the electromagnetic interference (EMI) potential of these thrusters, the interpretation of the resulting data is difficult because most of these measurements have been made in vacuum chambers with metal walls which reflect radio waves emanating from the thruster. This project developed a means of assessing the impact of metal vacuum chambers of arbitrary size or shape on EMI experiments, thereby allowing for test results to be interpreted properly. Chamber calibration techniques were developed and initially tested at RIAME using their vacuum chamber. Calibration experiments were to have been made at Tank 5 of NASA GRC and the 6 m by 9 m vacuum chamber at the University of Michigan to test the new procedure, however the subcontract to RIAME was cancelled by NASA memorandum on Feb. 26. 1999.

Gallimore, Alec D.↗
NASA NTRS1978

Temperature distributions of a cesium-seeded hydrogen-oxygen supersonic free jet

The hydrogen-oxygen plasma was generated at combustion chamber pressures ranging from 0.5 to 2.0 megapascals and for various seed ratios (1 to 10 percent). The plasma was observed as the atmospheric exhaust from a Mach 2 rocket test facility. Transverse profiles of the absolute integrated intensity were measured with the optically thin CsI lines (0.5664 and 0.5636 microns) at a range of axial positions downstream of the 5-cm-diameter combustor nozzle exit. Radial profiles of the emission coefficient were obtained from the measured transverse profiles of intensity by Abel inversion. Temperatures were then determined from the emission coefficients for conditions of local thermodynamic equilibrium using particle densities generated by a two-dimensional free jet computer program. Temperature results show the inherent effects of compression and expansion pressure waves characteristic of a free jet exiting from a supersonic nozzle.

Wang, S. Y.↗
NASA NTRS1962

PYROLYTIC GRAPHITE ROCKET THRUST CHAMBER INVESTIGATION PROJECT. MONTHLY LETTER

This is the Monthly Letter Progress Report for the eighth month of a twelve month technical study directed toward the following objectives: A. Determining analytically and experimentally the feasibility and attractiveness of using free-standing pyrolytic graphite for a radiation cooled liquid rocket thrust chamber and exit nozzle for space application. B. Establishing through analysis and experimental demonstration, a set of design principles to facilitate the design of a reliable pyrolytic graphite thrust chamber assembly for a specific range of operating conditions. C. Demonstrate by means of test firings of complete thrust chamber assemblies, the validity of the design data and principles developed during the program. Preparations have been completed for test firing the two 100 lb. thrust Boron Pyralloy chambers at Marquardt during the first week in March 1962. These two chambers will first be static pressure checked with water to a pressure of 135 psig. The chambers will then be test fired with N204/N2H4-MMH propellants . Test samples of eleven different pyrolytic carbide and nitride materials have been ordered for high temperature oxidation tests and should be delivered by the end of March for evaluation in the plasma torch facility. The specific materials ordered in free- standing condition in thicknesses from 0.030 inch to 0.065 inch were as follows: (a) Pyrolytic graphite (on hand) (b) Pyrolytic graphite-boron alloy (on hand) (c) Pyrolytic graphite-tungsten alloy (d)Boron nitride (e) Titanium nitride (f) Silicon carbide (g) Titanium carbide (h) Zirconium carbide (i) Hafnium carbide (j) Niobium carbide (k) Tantalum carbide (l) High density (1 1-90) graphite (m) High density graphite (heat treated) These materials were ordered from Raytheon, HTM, and American Meta 1 Products . A summary of the available data on the oxidation rates of pyrolytic materials under slow air and rocket motor conditions is shown in Figure 1. High Temperature Materials, Inc. has been contacted relative to providing Pyrographite thrust chambers with extra-thick walls . They believe they can deliver chambers with substantially thicker walls than were possible a short time ago. Such a configuration, with acceptable residual stresses, would provide a longer firing life since it is ultimately limited by the oxidation rate of the chamber wall material. Stress and structural studies have continued. The analytical stress studies point out some of the problems associated with the experimental determination of the elastic constants of this highly anisotropic material. Further analysis may shed light on improved techniques for making these experimental determinations. The large potential weight saving available with the use of pyrolytic graphite as a thrust chamber wall material in a radiation cooled rocket motor has been calculated and is shown graphically in Figure 2. Figure A presents the program progress schedule and the actual expenditures as of 7 February 1962.

ROCKET ENGINE↗
NASA NTRS2003

Simulation of Mini-Magnetospheric Plasma Propulsion (M2P2) Interacting with an External Plasma Wind

Substantial progress has been made over the last year in the development of the laboratory Mini-Magnetospheric Plasma Propulsion (M2P2) prototype. The laboratory testing has shown that that the plasma can be produced at high neutral gas efficiency, at high temperatures (a few tens of eV) with excellent confinement up to the point where chamber wall interactions dominate the physics. This paper investigates the performance of the prototype as it is opposed by an external plasma acting as a surrogate for the solar wind. The experiments were performed in 5ft diameter by 6ft long vacuum chamber at the University of Washington. The solar wind source comprised of a 33 kWe arc jet attached to a 200 kWe inductively generated plasma source. The dual plasma sources allow the interaction to be studied for different power levels, shot duration and production method. It is shown that plasma from the solar wind source (SWS) is able to penetrate the field of the M2P2 magnetic when no plasma is present. With operation of the M2P2 plasma source at only 1.5 kWe, the penetration of the SWS even at the highest power of operation at 200 kWe is stopped. This deflection is shown to be greatly enhanced over that produced by the magnet alone. In addition it is shown that with the presence of the SWS, M2P2 is able to produce enhanced magnetized plasma production out to at least 10 magnet radii where the field strength is only marginally greater than the terrestrial field. The results are consistent with the initial predictions that kWe M2P2 systems would be able to deflect several hundred kWe plasma winds to produce enhanced propulsion for a spacecraft.

Winglee, R. M.↗
NASA NTRS2010

Coaxial Injectors for Liquid Oxygen/Methane (LOX/CH4) Technology Development

Since late 2005, NASA s Marshall Space Flight Center (MSFC) has been developing and demonstrating technology for liquid oxygen and methane (LOX/CH4) engine designs. Efforts were undertaken to help advance technology that might benefit NASA s Exploration Technology Development Program. The propellant combination has gained interest as a potential option for the ascent stage main propulsion system on the Altair lunar lander vehicle. The propellants are also attractive for eventual Mars missions in future development activities. MSFC s efforts focus on evaluating the performance potential of both liquid and gaseous methane for specific injector designs. Previous JANNAF papers reported the performance observed for LOX and methane with an impinging injector. More recent efforts have focused on testing coaxial injector designs. Hot-fire testing performed at MSFC with thrust levels close to 5000 lbf demonstrated high performance with coaxial injectors of different element densities and various fuel film cooling levels. Test data provided results on performance, chamber wall compatibility, and heat flux profiles for different injectors. A variety of igniters were also demonstrated, including a torch igniter and a microwave (or plasma) igniter.

Elam, Sandra↗
NASA NTRSDate not supplied

Designing the PLANET Chamber for Lunar Environment Ground Testing

The Planetary, Lunar, and Asteroid Natural Environments Testbed (PLANET) is a new 2m x 3m high-vacuum chamber facility currently being designed and built at NASA’s Marshall Space Flight Center. When completed, it will be one of the highest-fidelity space/surface environment simulation chambers in existence. Instrumented with equipment to produce charged particle radiation, full spectrum ultraviolet, low-density plasma, and thermal extremes, PLANET can be configured to recreate the most challenging space environments. Additionally, a large regolith bed and distribution system will expose hardware to lunar regolith simulants and dust. This paper will describe the process of designing and optimizing PLANET and discuss some common considerations for performing ground-based environmental testing for lunar exploration.

Erin G Hayward↗
NASA NTRS1989

Method and apparatus for maintaining thermal control in plasma conditions

An apparatus and method are disclosed for determining the effects of exposure of oxygen plasma on a thin film polymer whose bulk is maintained at a predetermined temperature. The apparatus includes a chamber having a specimen therein. A plasma environment is provided in the chamber. A closure member is provided for sealing the chamber after the specimen is introduced into the chamber. The closure member also serves as a support for the test apparatus which includes a cooling coil. A platform having the test specimen thereon is supported on the cooling coil to be cooled by coolant flowing through the cooling coils. A thermoelectric module is supported on the platform to assist in maintaining a low test temperature for the test samples. The temperature of the sample is monitored by a thermocouple probe which is in contact with the sample. Any change in bulk sample temperature caused by the thermocouple probe is quickly adjusted to maintain the initial bulk temperature by the thermoelectric module.

Whitaker, Ann F.↗
NASA NTRS2015

Propulsion System Testing for the Iodine Satellite (iSAT) Demonstration Mission

CUBESATS are relatively new spacecraft platforms that are typically deployed from a launch vehicle as a secondary payload, providing low-cost access to space for a wide range of end-users. These satellites are comprised of building blocks having dimensions of 10x10x10 cm cu and a mass of 1.33 kg (a 1-U size). While providing low-cost access to space, a major operational limitation is the lack of a propulsion system that can fit within a CubeSat and is capable of executing high delta v maneuvers. This makes it difficult to use CubeSats on missions requiring certain types of maneuvers (i.e. formation flying, spacecraft rendezvous). Recently, work has been performed investigating the use of iodine as a propellant for Hall-effect thrusters (HETs) 2 that could subsequently be used to provide a high specific impulse path to CubeSat propulsion. 3, 4 Iodine stores as a dense solid at very low pressures, making it acceptable as a propellant on a secondary payload. It has exceptionally high ρIsp (density times specific impulse), making it an enabling technology for small satellite near-term applications and providing the potential for systems-level advantages over mid-term high power electric propulsion options. Iodine flow can also be thermally regulated, subliming at relatively low temperature (less than 100 C) to yield I2 vapor at or below 50 torr. At low power, the measured performance of an iodine-fed HET is very similar to that of a state-of-the-art xenon-fed thruster. Just as importantly, the current-voltage discharge characteristics of low power iodine-fed and xenon-fed thrusters are remarkably similar, potentially reducing development and qualifications costs by making it possible to use an already-qualified xenon-HET PPU in an iodine-fed system. Finally, a cold surface can be installed in a vacuum test chamber on which expended iodine propellant can deposit. In addition, the temperature doesn't have to be extremely cold to maintain a low vapor pressure in the vacuum chamber (it is under 10(exp -6) torr at −75 C), making it possible to 'cryopump' the propellant with lower-cost recirculating refrigerant-based systems as opposed to using liquid nitrogen or low temperature gaseous helium cryopanels. An iodine-based system is not without its challenges. The primary challenge is that the entire feed system must be maintained at an elevated temperature to prevent the iodine from depositing (transitioning from the gas phase directly back into the solid phase), which will block the propellant feed lines. Furthermore, deposition will occur unless the temperature in the lines is not greater than the temperature of the propellant reservoir. The flow rate can be controlled by adjusting the heating applied to the reservoir, but as with any thermal control there is a relatively slow response to changes in the heating rate. In the present paper, we describe the propulsion and propellant feed system for the iodine satellite (iSAT) flight demonstration mission. The system is based around the Busek BHT-200 Hall thruster, which has been modified for chemical compatibility with iodine vapor. While the gross propellant flow rate is maintained by the heated propellant reservoir, the flow to the anode and cathode are adjusted using two heated Vacco proportional flow control valves (PFCV), which provide very fast response on the flow rate adjustment. The flight mission design layout will be presented, showing how the system will be packaged into the overall 12-U spacecraft and the techniques being employed to protect the remaining spacecraft hardware from the propulsion system (e.g., plasma impingement, iodine deposition, thermal loads). In addition to the flight system design, results of testing the thruster and cathode with both operating on iodine propellant are presented. The tests are conducted on a thrust stand (see Fig. 1) in a large vacuum chamber containing a beam dump chilled to below -100 C to 'cryopump' the propellant. The thruster performance during these tests is presented, with these data used to evaluate the feed system and guide further refinements. Results of relatively long duration testing are presented to demonstrate the capability to operate for the length of the iSAT mission and to perform a number of re-starts as will be required by the mission concept of operations.

Polzin, Kurt A.↗
NASA NTRS1976

Measurements of uranium mass confined in high density plasmas

An X-ray absorption method for measuring the amount of uranium confined in high density, rf-heated uranium plasmas is described. A comparison of measured absorption of 8 keV X-rays with absorption calculated using Beer Law indicated that the method could be used to measure uranium densities from 3 times 10 to the 16th power atoms/cu cm to 5 times 10 to the 18th power atoms/cu cm. Tests were conducted to measure the density of uranium in an rf-heated argon plasma with UF6 infection and with the power to maintain the discharge supplied by a 1.2 MW rf induction heater facility. The uranium density was measured as the flow rate through the test chamber was varied. A maximum uranium density of 3.85 times 10 to the 17th power atoms/cu cm was measured.

Stoeffler, R. C.↗
DOE OSTI2024

H - Ion Source RF Plasma Testing with 13 MHz and 27 MHz at the Spallation Neutron Source (SNS)

The baseline RF-driven H⁻ ion source configuration at the Spallation Neutron Source (SNS) facility uses a continuous wave (CW) 600 W 13 MHz RF system to ignite and maintain a low-density plasma inside the ion source vacuum chamber. After the continuous low-density 13 MHz plasma has been established, a pulsed (typical 1 ms pulse width and 60 Hz pulse repetition rate) 80 kW 2 MHz RF system is used to increase the plasma density to produce the pulsed H⁻ ion beam. Incremental upgrades and improvements to the SNS ion source systems have resulted in the ability to reliably operate an H⁻ ion source for SNS neutron production run cycles that can last up to four months. Conditions inside the SNS H⁻ ion source evolve throughout a four-month run cycle due to changes in impurity levels, sputtering, and erosion. As the internal ion source conditions change during the run cycle, there can also be changes in plasma stability and the 13 MHz RF power level required to ignite the plasma. This paper presents the preliminary results of testing performed on the SNS Ion Source Test Stand (ISTS) system where we looked at plasma ignition and plasma stability using 27 MHz RF in place of the baseline 13 MHz RF system.

70 PLASMA PHYSICS AND FUSION TECHNOLOGY↗
NASA NTRS1981

Advanced space propulsion thruster research

Experiments showed that stray magnetic fields can adversely affect the capacity of a hollow cathode neutralizer to couple to an ion beam. Magnetic field strength at the neutralizer cathode orifice is a crucial factor influencing the coupling voltage. The effects of electrostatic accelerator grid aperture diameters on the ion current extraction capabilities were examined experimentally to describe the divergence, deflection, and current extraction capabilities of grids with the screen and accelerator apertures displaced relative to one another. Experiments performed in orificed, mercury hollow cathodes support the model of field enhanced thermionic electron mission from cathode inserts. Tests supported the validity of a thermal model of the cathode insert. A theoretical justification of a Saha equation model relating cathode plasma properties is presented. Experiments suggest that ion loss rates to discharge chamber walls can be controlled. A series of new discharge chamber magnetic field configurations were generated in the flexible magnetic field thruster and their effect on performance was examined. A technique used in the thruster to measure ion currents to discharge chamber walls is described. Using these ion currents the fraction of ions produced that are extracted from the discharge chamber and the energy cost of plasma ions are computed.

Wilbur, P. J.↗
DOE OSTI2024

Design of the Electromagnetic Particle Injector (EPI) for Tokamak Deployment

Both predicting and controlling disruptions are critical and urgent issues for ITER as some disruptions with a short warning time may be unavoidable. For these cases, a rapid response disruption mitigation system (DMS) referred to as the electromagnetic particle injector (EPI) is being developed. The primary advantages of the EPI are its fast response time and high velocity, which have been demonstrated in offline experiments (Raman et al., 2021). The EPI is capable of accelerating a metallic sabot electromagnetically using a rail gun to the required velocities ( > 2 km/s) within 2 ms. Two high-field racetrack magnets able to generate fields over 2 T are positioned above and below the rails to permit high velocity at low rail currents, a requirement to minimize electrode erosion. At the end of the acceleration phase, a sabot capture mechanism retains the spent sabot inside the vacuum chamber that houses the EPI. At this point, it releases well-defined microspheres, or a shell pellet, of a radiative payload into the disrupting plasma. A remotely operated sabot loading system positioned behind the injector contains several pre-equipped sabots that can be loaded by an operator from the tokamak control room. The injector is interfaced to the tokamak through a guide tube attached to the front of the EPI vacuum chamber. Finally, the advantages of the EPI system over other DMSs under consideration are described in conjunction with the details of an EPI system designed for near-term test on an existing large tokamak.

70 PLASMA PHYSICS AND FUSION TECHNOLOGY↗
NASA NTRS1999

Numerical Thermal Model of a 30-cm NSTAR Ion Thruster

A thermal computer model of the NSTAR (Nasa Solar Electric Propulsion Technology Applications Readiness) xenon ion thruster has been produced using a lumped parameter thermal nodal network scheme. This model contains 104 nodes on the thruster and was implemented using SINDA (Systems Improved Numerical Differencing Analyzer) and TRASYS (Thermal Radiation Analyzer System) on various UNIX workstations. The model includes radiation and conduction heat transfer, the effect of plasma interaction on the thruster, and an account for finely perforated surfaces. The model was developed in conjunction with an NSTAR thruster outfitted with approximately 20 thermocouples for thermal testing at the NASA Lewis Research Center. The results of these experiments were used to calibrate and confirm the computer model first without and then with the plasma interaction. The calibrated model was able to predict discharge chamber temperatures to within 10 C of measured temperatures. To demonstrate the ability of the model under various circumstances the heat flux was examined for a thruster operating in the environment of space.

VanNoord, Jon↗
NASA NTRS2016

Filtration of Carbon Particulate Emissions from a Plasma Pyrolysis Assembly

NASA is investigating plasma pyrolysis as a candidate technology that will enable the recovery of hydrogen from the methane produced by the ISS Sabatier Reactor. The Plasma Pyrolysis Assembly (PPA) is the current prototype of this technology which converts the methane product from the Carbon Dioxide Reduction Assembly (CRA) to acetylene and hydrogen with 90% or greater conversion efficiency. A small amount of solid carbon particulates are generated as a side product and must be filtered before the acetylene is removed and the hydrogen-rich gas stream is recycled back to the CRA. We discuss developmental work on several options for filtering out the carbon particulate emissions from the PPA exit gas stream. The filtration technologies and concepts investigated range from fibrous media to monolithic ceramic and sintered metal media. This paper describes the different developed filter prototypes and characterizes their performance from integrated testing at the Environmental Chamber (E-Chamber) at MSFC. In addition, characterization data on the generated carbon particulates, that help to define filter requirements, are also presented.

methane↗
NASA NTRS1976

Characteristics of a 30-cm thruster operated with small hole accelerator grid ion optics

Small hole accelerator grid ion optical systems have been tested as a possible means of improving 30-cm ion thruster performance. The effects of small hole grids on the critical aspects of thruster operation including discharge chamber performance, doubly-charged ion concentration, effluent beam characteristics, and plasma properties have been evaluated. In general, small hole accelerator grids are beneficial in improving thruster performance while maintaining low double ion ratios. However, extremely small accelerator aperture diameters tend to degrade beam divergence characteristics. A quantitative discussion of these advantages and disadvantages of small hole accelerator grids, as well as resulting variations in thruster operation characteristics, is presented.

Vahrenkamp, R. P.↗
NASA NTRS1989

Development of a high power microwave thruster, with a magnetic nozzle, for space applications

This paper describes the current development of a high-power microwave electrothermal thruster (MET) concept at the NASA Lewis Research Center. Such a thruster would be employed in space for applications such as orbit raining, orbit maneuvering, station change, and possibly trans-lunar or trans-planetary propulsion of spacecraft. The MET concept employs low frequency continuous wave (CW) microwave power to create and continuously pump energy into a flowing propellant gas at relative high pressure via a plasma discharge. The propellant is heated to very high bulk temperatures while passing through the plasma discharge region and then is expanded through a throat-nozzle assembly to produce thrust, as in a conventional rocket engine. Apparatus, which is described, is being assembled at NASA Lewis to test the MET concept to CW power levels of 30 kW at a frequency of 915 MHz. The microwave energy is applied in a resonant cavity applicator and is absorbed by a plasma discharge in the flowing propellant. The ignited plasma acts as a lossy load, and with optimal tuning, energy absorption efficiencies over 95 percent (based on the applied microwave power) are expected. Nitrogen, helium, and hydrogen will be tested as propellants in the MET, at discharge chamber pressures to 10 atm.

Power, John L.↗