SCIENTIFIC AND TECHNICAL AEROSPACE REPORTS
A Biweekly Publication of the National Aeronautics and Space Administration
VOLUME 43, ISSUE 19 - SEPTEMBER 23, 2006
20 SPACECRAFT PROPULSION AND POWER
Includes main propulsion systems and components, e.g., rocket engines; and spacecraft auxiliary power sources.
For related information see also 07 Aircraft Propulsion and Power, 28 Propellants and Fuels, 15 Launch Vehicles and Launch Operations, and 44 Energy Production and Conversion.
20050210090 Alabama Univ., Huntsville, AL, USA
Uncertainty Analysis of Heat Transfer to Supercritical Hydrogen in Cooling Channels
Locke, Justin M.; Landrum, D. Brian; June 14, 2005; 14 pp.; In English; AIAA 41st Joint Propulsion Conference and Exhibit, 10-14 Jul. 2005, Tucson, AZ, USA; Original contains color illustrations Contract(s)/Grant(s): NCC8-200 Report No.(s): AIAA Paper 2005-4304; No Copyright; Avail: CASI; A03, Hardcopy
Sound understanding of the cooling efficiency of supercritical hydrogen is crucial to the development of high pressure thrust chambers for regeneratively cooled LOX/LH2 rocket engines. This paper examines historical heat transfer correlations for supercritical hydrogen and the effects of uncertainties in hydrogen property data. It is shown that uncertainty due to property data alone can be as high as 10%. Previous heated tube experiments with supercritical hydrogen are summarized, and data from a number of heated tube experiments are analyzed to evaluate conditions for which the available correlations are valid. Author
Cooling; Heat Transfer; Hydrogen Oxygen Engines; Regenerative Cooling; Rocket Engines; Thrust Chambers; Correlation
20050210109 NASA Marshall Space Flight Center, Huntsville, AL, USA
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Development of Ionic Liquid Monopropellants for In-Space Propulsion
Blevins, John A.; Osborne, Robin; Drake, Gregory W.; [2005]; 1 pp.; In English; 53rd JPM/2nd LPS/SP Joint Meeting, 5-8 Dec. 2005, Monterey, CA, USA; No Copyright; Avail: Other Sources; Abstract Only
A family of new, low toxicity, high energy monopropellants is currently being evaluated at NASA Marshall Space Flight Center for in-space rocket engine applications such as reaction control engines.
These ionic liquid monopropellants, developed in recent years by the Air Force Research Laboratory, could offer system simplification, less in-flight thermal management, and reduced handling precautions, while increasing propellant energy density as compared to traditional storable in-space propellants such as hydrazine and nitrogen tetroxide. However, challenges exist in identifying ignition schemes for these ionic liquid monopropellants, which are known to burn at much hotter combustion temperatures compared to traditional monopropellants such as hydrazine.
The high temperature combustion of these new monopropellants make the use of typical ignition catalyst beds prohibitive since the catalyst cannot withstand the elevated temperatures. Current research efforts are focused on monopropellant ignition and burn rate characterization, parameters that are important in the fundamental understanding of the monopropellant behavior and the eventual design of a thruster.
Laboratory studies will be conducted using alternative ignition techniques such as laser-induced spark ignition and hot wire ignition. Ignition delay, defined as the time between the introduction of the ignition source and the first sign of light emission from a developing flame kernel, will be measured using Schlieren visualization. An optically-accessible liquid monopropellant burner will be used to determine propellant burn rate as a function of pressure and initial propellant temperature. The burn rate will be measured via high speed imaging through the chamber s windows. Author
Monopropellants; Rocket Engines; Temperature Control; High Temperature; Combustion Temperature; Burning Rate; Spark Ignition
20050212136 NASA Goddard Space Flight Center, Greenbelt, MD, USA
Space Technology-5 Lithium-Ion Battery Design, Qualification and Integration and Testing
Rao, Gopalakishna M.; Stewart, Karen; Ameen, Syed; Banfield, Peter K.; [2005]; 21 pp.; In English; 3rd International Energy Conversion Engineering Conference, 15-18 Aug. 2005, San Francisco, CA, USA; No Copyright; Avail: CASI; A03, Hardcopy
This document is a viewgraph presentation that reviews the Lithium Ion Battery for the Space Technology-5 (ST-5) mission. Included in the document is a review of the ST-5 Mission, a review of the battery requirements, a description of the battery and the battery materials. The testing and the integration and qualification data is reviewed. CASI
Performance Tests; Lithium Batteries; Spacecraft Power Supplies
20050212149 NASA Marshall Space Flight Center, Huntsville, AL, USA
Transient Three-Dimensional Analysis of Nozzle Side Load in Regeneratively Cooled Engines
Wang, Ten-See; March 29, 2005; 12 pp.; In English; 41st AIAA/ASME/SAE/ASEE Joint Propulsion Conference and Exhibit, 10-13 Jul. 2005, Tucson, AZ, USA Report No.(s): AIAA Paper 2005-3942; No Copyright; Avail: CASI; A03, Hardcopy
Three-dimensional numerical investigations on the start-up side load physics for a regeneratively cooled, high-aspect-ratio nozzle were performed. The objectives of this study are to identify the three-dimensional side load physics and to compute the associated aerodynamic side load using an anchored computational methodology. The computational methodology is based on an unstructured-grid, pressure-based computational fluid dynamics formulation, and a transient inlet condition based on an engine system simulation. Computations were performed for both the adiabatic and cooled walls in order to understand the effect of boundary conditions. Finite-rate chemistry was used throughout the study so that combustion effect is always included. The results show that three types of shock evolution are responsible for side loads: generation of combustion wave; transitions among free-shock separation, restricted-shock separation, and simultaneous free-shock and restricted shock separations; along with oscillation of shocks across the lip.Wall boundary conditions drastically affect the computed side load physics: the adiabatic nozzle prefers free-shock separation while the cooled nozzle favors restricted-shock separation, resulting in higher peak side load for the cooled nozzle than that of the adiabatic nozzle. By comparing the computed physics with those of test observations, it is concluded that cooled wall is a more realistic boundary condition, and the oscillation of the restricted-shock separation flow pattern across the lip along with its associated tangential shock motion are the dominant side load physics for a regeneratively cooled, high aspect-ratio rocket engine. Author
Rocket Engines; Rocket Nozzles; Regenerative Cooling; Combustion; Computational Fluid Dynamics; High Aspect Ratio
20050212178 NASAMarshall Space Flight Center, Huntsville, AL, USA, ATK-Thiokol Propulsion, Brigham City, UT, USA
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RSRM TP-H1148 Main Grain Propellant Crack Initiation Evaluation
Earnest, Todd E.; July 2005; 9 pp.; In English; 41st AIAA/ASME/SAE/ASEE Joint Propulsion Conference and Exhibit, 11-13 Jul. 2005, Tucson, AZ, USA Contract(s)/Grant(s): NAS8-97238; No Copyright; Avail: CASI; A02, Hardcopy
Pressurized TP-HI 148 propellant fracture toughness testing was performed to assess the potential for initiation of visually undetectable cracks in the RSRM forward segment transition region during motor ignition. Two separate test specimens were used in this evaluation. Testing was performed in cold-gas and hot-fire environments, and under both static and dynamic pressurization conditions. Analysis of test results demonstrates safety factors against initiation of visually undetectable cracks in excess of 8.0. The Reusable Solid Rocket Motor (RSRM) forward segment is cast with PBAN propellant (TP-HI 148) to form T an 1 1-point star configuration that transitions to a tapered center perforated bore (see Figure 1). The geometry of the transition region between the fin valleys and the bore causes a localized area of high strain during horizontal storage. Updated analyses using worst-case mechanical properties at 40 F and improved modeling techniques indicated a slight reduction in safety margins over previous predictions. Although there is no history of strain induced cracks or flaws in the transition region propellant, a proactive test effort was initiated to better understand the implications of the new analysis, primarily the resistance of TP-H1148 propellant to crack initiation’ during RSRM ignition. Author
Solid Propellant Rocket Engines; Rocket Propellants; Cracks
20050212270 Pennsylvania State Univ., University Park, PA USA
Meso and Micro Scale Propulsion Concepts for Small Spacecraft
Yetter, R. A.; Yang, V.; Aksay, I. A.; Dryer, F. L.; Jun. 2005; 26 pp.; In English; Original contains color illustrations Contract(s)/Grant(s): FA9550-04-1-0088 Report No.(s): AD-A436438; AFRL-SR-AR-TR-05-0307; No Copyright; Avail: CASI; A03, Hardcopy
Ameso/micro scale liquid propellant thruster is being developed for small spacecraft. In this thruster, the liquid propellant was injected directly into the chamber tangentially along the combustor wall. Heat feedback from the flame, once ignited, decomposed and gasified the liquid propellant. The gasification process of the propellant produced a voflex flow in the combustor, which stabilized the gas- phase flame. Modeling studies with gas-phase injection of reactants have shown that the vortex flow creates various flow recirculation patterns depending upon the Reynolds numbers of the flows. Although the propellant of interest was a HAN based liquid propellant, initial studies were conducted with liquid nitromethane. Operation of the combustion chamber with pure nitromethane at chamber pressures as high as 30 atm was studied. To gasify and initially ignite a HAN based propellant, techniques that make use of the electrolytic character of the propellant were investigated. In particular, a continuous flow igniter with titanium electrodes and a gap spacing of 500 Inn was developed and studied. Results showed that a small voltage drop across the electrodes ignites the continuous flow, which can then be removed to achieve a steady-state decomposition process. Methods to further miniature, fabricate, and package the entire system will be investigated in the future. DTIC
Liquid Rocket Propellants; Propulsion; Propulsion System Configurations; Propulsion System Performance
20050212321 Phillips Lab., Edwards AFB, CA USA
Spectroscopic Emission Measurements of a Pulsed Plasma Thruster Plume
Markusic, Thomas E.; Spores, Ronald A.; Jul. 1997; 12 pp.; In English Contract(s)/Grant(s): Proj-2308 Report No.(s): AD-A436554; No Copyright; Avail: CASI; A03, Hardcopy
The optical emission spectra of plasmas produced by a Teflon (polytetrafluoroethylene (PTFE)) solid propellant pulsed plasma thruster (XPPT-1) were measured and analyzed. A broad temporally and spatially integrated survey of the emission from 3500 to 7500 angstrom is reported. It is shown that in the PPT discharge energy range surveyed (15 to 45 joule), the species formed do not vary; that is, no new species are formed as the discharge energy is increased. Furthermore, relative line ratios do not vary significantly with energy, suggesting that the bulk thermodynamic properties of the discharge plasma may not be affected by changing the initial stored capacitor energy. A spatially integrated but time- resolved measurement was carried out which shows the single species emission as a function of PPT discharge time. The magnitude of the emission was found to track closely with the PPT current and depend on the discharge energy. Relative line emission intensities are used to determine the degree of excitation equilibrium and establish bounds on the plasma temperature. Spatially and temporally resolved images are used to estimate the plasma streaming velocity from time-of-flight and Doppler-shift techniques. DTIC
Emission Spectra; Exhaust Gases; Measurement; Plasma Engines; Plumes; Pulsed Plasma Thrusters; Spectroscopy
20050212322 Cornell Univ., Ithaca, NY USA
Comparisons of Hydrogen Atom Measurements in an Arcjet Plume with DSMC Predictions
Wysong, Ingrid J.; Pobst, Jeffrey A.; Jul. 1996; 15 pp.; In English Contract(s)/Grant(s): F49620-94-1-0328; Proj-2308 Report No.(s): AD-A436555; No Copyright; Avail: CASI; A03, Hardcopy
We present comparisons between hydrogen atom number density, velocity, and temperature measurements with DSMC predictions at nozzle exit for two power levels of a hydrogen arcjet thruster: 1.34 kW and 800W. Comparisons between model and experiment are also presented in two downstream plume regions. Atomic density peaks at each power are not seen to vary substantially while velocity and temperature maximums decrease with lower-power operation. Changes in profile shapes are noticed at the different powers and similar profile changes are also observed in the model predictions. Differences in overall magnitude are observed at some locations between measured data and predicted results, but in general the model matches the measured data quite well. DTIC
Arc Jet Engines; Computerized SIMulation; Exhaust Gases; Hydrogen Atoms; Measurement; Monte Carlo Method; Plumes; Prediction Analysis Techniques; SIMulation
20050212323 Michigan Univ., Ann Arbor, MI USA
Performance Characteristics of a 5 kW Laboratory Hall Thruster
Haas, James M.; Gulczinski, Frank S., III; Gallimore, Alec D.; Spanjers, Gregory G.; Spores, Ronald A.; Jul. 1996; 9 pp.; In English Contract(s)/Grant(s): Proj-2308 Report No.(s): AD-A436556; No Copyright; Avail: CASI; A02, Hardcopy
The University of Michigan and USA Air Force Research Laboratory have jointly developed a 5 kW class Hall effect thruster. This thruster was developed to investigate, with a variety of diagnostics, a thruster similar to that specified by IHPRPT goals. The configuration of this thruster is adjustable so that diagnostic access to the interior of the thruster can be provided as necessary, and to allow for the exploration of various thruster geometries. At nominal conditions, the thruster was designed to operate at 5 kW with a predicted specific impulse of 2200 s. The actual operating parameters at 5 kW were 2326 s specific impulse, with 246 mN of thrust at an efficiency of 57%. These conditions are comparable to those of thrusters under commercial development, making the information learned from the study of this thruster applicable to the understanding of its commercial counterparts. DTIC
Hall Thrusters; Propulsion; Research Facilities
20050212329 Phillips Lab., Edwards AFB, CA USA
Arcjet Diagnostics for Measuring Velocity, Density and Temperature
Spores, Ronald A.; Hargus, William A., Jr; Pobst, Jeffrey A.; Schilling, John H.; Lutfy, Frederick M.; Erwin, Daniel A.; Jun. 1994; 16 pp.; In English Contract(s)/Grant(s): F49620-93-1-01213; Proj-2308 Report No.(s): AD-A436570; No Copyright; Avail: CASI; A03, Hardcopy
New diagnostic techniques to measure the fundamental plasma properties of an arcjet are being developed collaboratively between the Air Force Phillips Laboratory and the University of Southern California. These diagnostics are for measuring each of the principal parameters defining the plasma state within an arcjet: velocity, temperature, and density. Velocity measurements are being obtained from a new technique we have named Current Modulation Velocimetry (CMV) . A current spike applied to the arcjet input current generates an optical emission event that is observed to travel downstream with the propellant flow. Observing this event at two axial locations, a fixed distance apart, provides an average bulk velocity of the gas flowing out of the arcjet nozzle. Species density measurements are to be obtained using Pulsed Electron Beam Fluorescence. A pulsed beam of focused electrons bombard a gas sample causing the resulting emitted fluorescence to be proportional to the local species density. Proper calibration with known samples can then provide absolute density values. Emission spectroscopy is employed to measure propellant excitation, rotational and vibrational temperatures inside the arcjet nozzle. Small (0.020’) holes have been drilled through the arcjet nozzle to provide internal optical access. Velocity and density measurements are being conducted on the standard NASA Lewis 1kW arcjet while spectroscopic measurements are being taken on a 30kW class arcjet. DTIC
Arc Jet Engines; Diagnosis; Electron Density (Concentration); Measurement; Temperature Measurement
20050212330 Pennsylvania State Univ., University Park, PA USA
Interior Spectroscopic Investigation of the Propellant Energy Modes in an Arcjet Nozzle
Hargus, William A., Jr; Micci, Michael M.; Spores, Ronald A.; Jun. 1994; 12 pp.; In English Contract(s)/Grant(s): Proj-2308 Report No.(s): AD-A436571; No Copyright; Avail: CASI; A03, Hardcopy
Internal emission spectroscopy measurements were performed in the nozzle expansion region of a 26 kW class ammonia arcjet. A series of three optical access ports (0.020-inch diameter) were equally spaced along the center line of the expansion nozzle. Atomic and ionic excitation temperatures of H and NII were determined through the use of Boltzmann plots. Electron density measurements were also taken based upon Stark broadening of the hydrogen beta Balmer transition. It is believed that a change in the mechanism governing the electron density occurred just downstream of the constrictor between the first two access ports lending evidence that arc attachment is occurring in this region. Although the trends for excitation temperatures and electron density were similar to those previously observed in 1 kW arcjets, the temperature values were found to be significantly less. In addition, vibrational and rotational temperatures of NH were measured and appeared to be frozen throughout the expansion nozzle. DTIC
Arc Jet Engines; Emission Spectra; Propellants; Rocket Nozzles; Spectroscopy
20050212331 Phillips Lab., Edwards AFB, CA USA
A Comparison of Theory and Measurements in the Anode Region of a Self-Field Cylindrical MPD Thruster
Tilley, D. L.; Castillo, S.; Jolly, M. S.; Niewood, E.; Martinez-Sanchez, M.; Jun. 1994; 13 pp.; In English Contract(s)/Grant(s): Proj-2308 Report No.(s): AD-A436572; No Copyright; Avail: CASI; A03, Hardcopy
The need to validate and improve numerical models of the MPD thruster flow, using detailed comparisons with experimentally determined plasma properties, is essential in the effort to improve MPD thruster efficiency. This research represents an attempt to experimentally identify trends, in the variation of plasma properties inside of a self-field MPD thruster, which could be meaningfully compared with numerical results. Experimental results, consistent with the numerical model in trend, show an axial variation of the electron temperature along the anode which exhibits a very distinctive double-peaked profile. Experimental axial profiles of the electron number density and current density were also consistent with the numerical model in trend. Such evidence, along with results of the numerical model, suggest that an oblique shock is present near the inlet, and is required to establish the radial pressure gradient so prevalent in the MPD thruster. This work marks one of the first non-trivial, successful comparisons between a numerical model and experimentally determined plasma properties inside the MPD thruster. DTIC
Anodes; Cylindrical Bodies; Magnetoplasmadynamic Thrusters; Magnetoplasmadynamics; Measurement; Plasmas (Physics)
20050212332 Stanford Univ., Stanford, CA USA
Plasma and Cathode Emission from a High Power Hydrogen Arcjet Thruster
Berns, Darren H.; Storm, P. V.; Hargus,William A., Jr; Cappelli, Mark A.; McFall, K. A.; Spores, Ronald A.; Jul. 1996; 16 pp.; In English Contract(s)/Grant(s): Proj-2308 Report No.(s): AD-A436573; No Copyright; Avail: CASI; A03, Hardcopy
An experimental study of the measurement of cathode temperature, current distribution, and near-cathode electron number density in a high power hydrogen arcjet is presented. This study is motivated by the desire to better understand arc-electrode interactions in arcjet thrusters, which in many cases, is the main determinate of arcjet lifetime. Measurements such as these may also provide the needed boundary conditions for numerical arcjet simulations, presently under development.We describe in this paper the application of a non-intrusive in-situ measurement technique for on-axis, spectral imaging of the electrode region of arcjets, and the application of this technique to the measurement of the cathode and anode temperatures, cathode spot size, and current distribution in a 30kW hydrogen arcjet thruster. A relatively large field of view (twice the throat diameter) and high spatial resolution (9 micrometers) are achieved. DTIC
Arc Jet Engines; Cathode Glow; Cathodes; Electric Rocket Engines; Emission; Emission Spectra; Hydrogen; Measurement; Plasma Engines; Plasma Radiation
20050212340 Phillips Lab., Edwards AFB, CA USA
Investigation of Propellant Inefficiencies in a Pulsed Plasma Thruster
Spanjers, Gregory G.; McFall, Keith A.; Gulczinski III, Frank S.; Spores, Ronald A.; Jul. 1996; 10 pp.; In English Contract(s)/Grant(s): Proj-2308 Report No.(s): AD-A436594; No Copyright; Avail: CASI; A02, Hardcopy
A Pulsed Plasma Thruster (PPT) benefits from the inherent engineering simplicity and reduced tankage fraction gained by storing the propellant as a solid. The solid is converted to the gaseous state during an electric surface discharge. Previous research has concluded that the bulk of the propellant expands gas-dynamically from the chamber at low directed velocity, with possibly as little as 10% ionized and efficiently accelerated to high velocity using electromagnetic forces. The two velocity components result in a low propellant utilization efficiency. Critical to improving the PPT efficiency is preventing the formation of the low-velocity propellant and/or developing a means of accelerating it electromagnetically. In the present work measurements are made of the solid propellant conversion to the gaseous state with the intent of better understanding the formation process. By better understanding the propellant conversion it is hoped that future PPTs can be designed with significantly increased propellant efficiencies. DTIC
Plasma Engines; Propellants; Pulsed Plasma Thrusters; Solid Rocket Propellants
20050212341 University of Southern California, Los Angeles, CA USA
Arcjet Diagnostics Using Current Modulation Velocimetry and Pulsed Electron Beam Fluorescence
Pobst, Jeffrey A.; Schilling, John H.; Lutfy, Frederick M.; Erwin, Daniel A.; Spores, Ronald A.; Jun. 1994; 12 pp.; In English Contract(s)/Grant(s): F49620-93-1-01213; F49620-93-1-0373; Proj-2308 Report No.(s): AD-A436595; No Copyright; Avail: CASI; A03, Hardcopy
New diagnostic techniques to measure the fundamental plasma properties of an arcjet are being developed collaboratively between the University of Southern California and Air Force Phillips Laboratory. The diagnostics described here are for measuring velocity and density, principal parameters that describe the plasma state within an arcjet. Velocity measurements are being obtained from a new technique we have named Current Modulation Velocimetry (CMV). A current spike applied to the arcjet input current generates an optical emission event that is observed to travel downstream with the propellant flow. Observing this event at two axial locations, a fixed distance apart, provides an average bulk velocity of the gas flowing out of the arcjet nozzle. Species density measurements are to be obtained using Pulsed Electron Beam Fluorescence. A pulsed beam of focused electrons bombard a gas sample causing the resulting emitted fluorescence to be proportional to the local species density. Proper calibration with known samples can then provide absolute density values. Velocity measurements are being conducted on the standard NASA Lewis 1kW arcjet using hydrogen propellant. DTIC
Arc Jet Engines; Diagnosis; Electric Current; Electron Beams; Fluorescence; Modulation; Velocity Measurement
20050212379 Phillips Lab., Edwards AFB, CA USA
Laser Induced Fluorescence of Ground State Hydrogen Atoms at Nozzle Exit of an Arcjet Thruster
Pobst, Jeffrey A.; Wysong, Ingrid J.; Spores, Ronald A.; Jun. 1995; 7 pp.; In English Contract(s)/Grant(s): Proj-2308 Report No.(s): AD-A436728; No Copyright; Avail: CASI; A02, Hardcopy
We report the first observation of a two-photon laser-induced fluorescence (LIF) technique in an arcjet plume. Ground state hydrogen atoms are detected with high spatial resolution near the thruster nozzle exit. Number density, axial and radial velocity, and translational temperature distributions are obtained in the expansion plume of a 1 kW arcjet operating on hydrogen propellant. Comparison of the ground state properties with previously measured excited state hydrogen data and recent computational data is discussed. DTIC
Arc Jet Engines; Atoms; Electric Rocket Engines; Ground State; Hydrogen Atoms; Laser Induced Fluorescence; Rocket Nozzles
20050214043 United Technologies Research Center, USA
Ceramic Matrix Composite Vane Rig Testing and Design
Linsey, Gary; NASA Glenn Research Center UEET (Ultra-Efficient Engine Technology) Program: Agenda and Abstracts; [2001], pp. 19; In English; See also 20050214031; No Copyright; Abstract Only; Available from CASI only as part of the entire parent document
Under the NASA-funded Ultra-Efficient Engine Technology (UEET) program, the United Technologies Research Center (UTRC) has begin designing Ceramic Matrix Composite (CMC) turbine vanes with an Environmental Barrier Coating (EBC) capable of withstanding 2700 F surface temperature and 2400 F interface temperature for over 1,000 hours in turbine engine environments. The CMC turbine vanes consist of silicon carbide fiber-reinforced silicon carbide (SiC/SiC) material with a Celsian-based EBC similar to the system which has demonstrated over 27,000 hours of successful engine field testing under various DOE- funded programs. Author
Fiber Composites; Silicon Carbides
20050214201 Texas Univ., Austin, TX USA
Homopolar Motor and Brush Development Studies
McNab, Ian; Aug. 2005; 73 pp.; In English Contract(s)/Grant(s): N00014-04-1-0064 Report No.(s): AD-A436704; IAT-R0419; No Copyright; Avail: CASI; A04, Hardcopy
Substantial benefits in ship design and operating economies are expected to accrue to the Navy if electric ship propulsion techniques can be successfully developed and introduced in the future Navy fleet. One area of interest toward this end is in superconducting homopolar motors. This report surveys prior data, reports, papers, and other studies that relate to superconducting homopolar motors. It also includes summaries of discussions held by the author with General Atomics, the Naval Surface Warfare Center, and the Office of Naval Research on the subject. DTIC
Brushes; Electric Propulsion; Marine Propulsion; Metal Fibers; Ships; Superconductivity
20050214625 Brookhaven National Lab., Upton, NY USA
Low NOX Burner Development
Krishna, C. R.; Butcher, T.; Sep. 2004; 16 pp.; In English Report No.(s): DE2005-15010732; No Copyright; Avail: Department of Energy Information Bridge
The objective of the task is to develop concepts for ultra low NOx burners. One approach that has been tested previously uses internal recirculation of hot gases and the objective was to how to implement variable recirculation rates during burner operation. The second approach was to use fuel oil aerosolization (vaporization) and combustion in a porous medium in a manner similar to gas-fired radiant burners. This task is trying the second approach with the use of a somewhat novel, prototype system for aerosolization of the liquid fuel. NTIS
Air Pollution; Burners; Combustion Chambers; Nitrogen Oxides; Pollution Control
20050214691 NASA Glenn Research Center, Cleveland, OH, USA
Pulsed Plasma Thruster (PPT) Technology: Earth Observing-1 PPT Operational and Advanced Components Being Developed
Pencil, Eric J.; Benson, Scott W.; Arrington, Lynn A.; Frus, John; Hoskins, W. Andrew; Burton, Rodney; Research and Technology 2002; March 2003; 2 pp.; In English; No Copyright; Avail: CASI; A01, Hardcopy
In 2002 the pulsed plasma thruster (PPT) mounted on the Earth Observing-1 spacecraft was operated successfully in orbit. The two-axis thruster system is fully incorporated in the attitude determination and control system and is being used to automatically counteract disturbances in the pitch axis of the spacecraft.
The first tests conducted in space demonstrated the full range of PPT operation, followed by calibration of control torques from the PPT in the attitude control system. Then the spacecraft was placed in PPT control mode.
To date, it has operated for about 30 hr. The PPT successfully controlled pitch momentum during wheel de-spin, solar array acceleration and deceleration during array rewind, and environmental torques in nominal operating conditions. Images collected with the Advanced Landsat Imager during PPT operation have demonstrated that there was no degradation in comparison to full momentum wheel control. In addition, other experiments have been performed to interrogate the effects of PPT operation on communication packages and light reflection from spacecraft surfaces.
Future experiments will investigate the possibility of orbit-raising maneuvers, spacecraft roll, and concurrent operation with the Hyperion imager. Future applications envisioned for pulsed plasma thrusters include longer life, higher precision, multiaxis thruster configurations for three-axis attitude control systems or high-precision, formationflying systems. Advanced components, such as a ‘dry’ mica-foil capacitor, a wear-resistant spark plug, and a multichannel power processing unit have been developed under contract with Unison Industries, General Dynamics, and C.U. Aerospace.
Over the last year, evaluation tests have been conducted to determine power processing unit efficiency, atmospheric functionality, vacuum functionality, thruster performance evaluation, thermal performance, and component life. Author
Pulsed Plasma Thrusters; Landsat Satellites
20050214748 NASA Glenn Research Center, Cleveland, OH, USA
CORBASec Used to Secure Distributed Aerospace Propulsion Simulations
Blaser, Tammy M.; Research and Technology 2002; March 2003; 2 pp.; In English; No Copyright; Avail: CASI; A01, Hardcopy
The NASA Glenn Research Center and its industry partners are developing a Common Object Request Broker (CORBA) Security (CORBASec) test bed to secure their distributed aerospace propulsion simulations.
Glenn has been working with its aerospace propulsion industry partners to deploy the Numerical Propulsion System Simulation (NPSS) object-based technology. NPSS is a program focused on reducing the cost and time in developing aerospace propulsion engines. It was developed by Glenn and is being managed by the NASA Ames Research Center as the lead center reporting directly to NASA Headquarters’ Aerospace Technology Enterprise.
Glenn is an active domain member of the Object Management Group: an open membership, not-for-profit consortium that produces and manages computer industry specifications (i.e., CORBA) for interoperable enterprise applications. When NPSS is deployed, it will assemble a distributed aerospace propulsion simulation scenario from proprietary analytical CORBA servers and execute them with security afforded by the CORBASec implementation.
The NPSS CORBASec test bed was initially developed with the TPBroker Security Service product (Hitachi Computer Products (America), Inc., Waltham, MA) using the Object Request Broker (ORB), which is based on the TPBroker Basic Object Adaptor, and using NPSS software across different firewall products. The test bed has been migrated to the Portable Object Adaptor architecture using the Hitachi Security Service product based on the VisiBroker 4.x ORB (Borland, Scotts Valley, CA) and on the Orbix 2000 ORB (Dublin, Ireland, with U.S. headquarters inWaltham, MA). Glenn, GE Aircraft Engines, and Pratt & Whitney Aircraft are the initial industry partners contributing to the NPSS CORBASec test bed. The test bed uses Security SecurID (RSA Security Inc., Bedford, MA) two-factor token-based authentication together with Hitachi Security Service digital-certificate-based authentication to validate the various NPSS users. The test bed is expected to demonstrate NPSS CORBASec-specific policy functionality, confirm adequate performance, and validate the required Internet configuration in a distributed collaborative aerospace propulsion environment. Author
Aerospace Engineering; Spacecraft Propulsion; Test Stands; SIMulation
Source: NASA.
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