SCIENTIFIC AND TECHNICAL AEROSPACE REPORTS
A Biweekly Publication of the National Aeronautics and Space Administration
VOLUME 43, ISSUE 16 - AUGUST 12, 2005
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.
20050192553 NASA Kennedy Space Center, Cocoa Beach, FL, USA
Rocket Science: The Shuttle’s Main Engines, though Old, Are Not Forgotten in the New Exploration Initiative
Covault, Craig; Aviation Week and Space Technology; July 11, 2005; ISSN 0005-2175; Volume 163, No. 2, pp. 59; In English; Original contains color illustrations; Copyright; Avail: Other Sources
The Space Shuttle Main Engine (SSME), developed 30 years ago, remains a strong candidate for use in the new Exploration Initiative as part of a shuttle-derived heavy-lift expendable booster. This is because the Boeing-Rocket- dyne man-rated SSME remains the most highly efficient liquid rocket engine ever developed. There are only enough parts for 12-15 existing SSMEs, however, so one NASA option is to reinitiate SSME production to use it as a throw-away, as opposed to a reusable, powerplant for NASA s new heavy-lift booster. Derived from text
Liquid Propellant Rocket Engines; Space Shuttle Main Engine
20050194570 NASA Glenn Research Center, Cleveland, OH, USA, NASA Johnson Space Center, Houston, TX, USA, NASA Kennedy Space Center, Cocoa Beach, FL, USA
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Power System Options Evaluated for the Radiation and Technology Demonstration Mission
Kerslake, ThomasW.; Benson, ScottW.; Research and Technology 1999; March 2000; 3 pp.; In English; No Copyright;Avail: CASI; A01, Hardcopy
The Radiation and Technology Demonstration (RTD) Mission is under joint study by three NASA Centers: the NASA Johnson Space Center, the NASA Goddard Space Flight Center, and the NASA Glenn Research Center at Lewis Field. This Earth-orbiting mission, which may launch on a space shuttle in the first half of the next decade, has the primary objective of demonstrating high-power electric thruster technologies. Secondary objectives include better characterization of Earth’s Van Allen trapped-radiation belts, measurement of the effectiveness of the radiation shielding for human protection, measurement of radiation effects on advanced solar cells, and demonstration of radiation-tolerant microelectronics.
During the mission, which may continue up to 1 year, the 2000-kg RTD spacecraft will first spiral outward from the shuttle-deployed, medium-inclination, low Earth orbit. By the phased operation of a 10-kW Hall thruster and a 10-kW Variable Specific Impulse Magneto-Plasma Rocket, the RTD spacecraft will reach a low-inclination Earth orbit with a radius greater than five Earth radii. This will be followed by an inward spiraling orbit phase when the spacecraft deploys 8 to 12 microsatellites to map the Van Allen belts. The mission will conclude in low Earth orbit with the possible retrieval of the spacecraft by the space shuttle.
A conceptual RTD spacecraft design showing two photovoltaic (PV) array wings, the Hall thruster with propellant tanks, and stowed microsatellites is presented. Early power system studies assessed five different PV array design options coupled with a 120-Vdc power management and distribution system (PMAD) and secondary lithium battery energy storage. Array options include (1) state-of-the-art 10-percent efficient three-junction amorphous SiGe thin-film cells on thin polymer panels deployed with an inflatable (or articulated) truss, (2) SCARLET array panels, (3) commercial state-of-the-art, planar PV array rigid panels with 25-percent efficient, three-junction GaInP2/GaAs/Ge solar cells, (4) rigid panels with 25-percent efficient, three-junction GaInP2/GaAs/Ge solar cells, in a 2 -concentrator trough configuration, and (5) thin polymer panels with 25-percent efficient, three-junction GaInP2/GaAs/Ge solar cells deployed with an inflatable (or articulated) truss.
To assess the relative merits of these PV array design options, the study group developed a dedicated Fortran code to predict power system performance and estimate system mass. This code also modeled Earth orbital environments important for accurately predicting PV array performance. The most important environmental effect, solar cell radiation degradation, was calculated from electron-proton fluence input from the industry standard AE8/AP8 trapped radiation models and the concept of damage equivalence. Power systems were sized to provide 10 kW of thruster power and approximately 1 kW of spacecraft power at end of life. Of the five PV array design options, the option 1 (thin-film cells) power system was the most massive 590 kg, whereas the option 4 (trough concentrator) power system was the lightest 260 kg. Arguably, the lowest cost would come from the option 3 (commercial array panels) power system with an acceptable, albeit greater, system mass of 320 kg. Predicted power system performance during the spiral-out mission phase is shown the preceding graph for the option 5 (flexible-panel) array.
From the results, the radiation-induced power loss over time is evident as the spacecraft slowly spirals outward through the trapped proton belt. The importance of the spiral trip time is also evident in the two curves representing 74-day and 182-day spiral-out periods. The longer spiral time introduces a beginning-of-life power oversizing penalty greater than 1 kW. Future studies will analyze power system performance and mass with a 50-Vdc power management and distribution architecture favorable to the VASIMR thruster and longer missions. Author
Space Missions; Radiation Measurement; Technology Utilization; Spacecraft Propulsion; Power Conditioning
20050194572 NASA Glenn Research Center, Cleveland, OH, USA
Power Systems Evaluated for Solar Electric Propulsion Vehicles
Kerslake, Thomas W.; Gefert, Leon P.; Research and Technology 1999; March 2000; 3 pp.; In English; No Copyright; Avail: CASI; A01, Hardcopy
Solar electric propulsion (SEP) mission architectures are applicable to a wide range NASA missions including the robotic exploration of the outer planets in the next decade and the human exploration of Mars within the next 2 decades. SEP enables architectures that are very mass efficient with reasonable power levels (1-MW class) aerobrake and cryogenic upper-stage transportation technologies are utilized. In this architecture, the efficient SEP stage transfers the payload from low Earth orbit (LEO) High Energy Elliptical Parking Orbit (HEEPO) within a period of 6 to 12 months. highthrust, cryogenic upper stage and payload then separate from the SEP vehicle for injection to the planetary target, allowing for fast heliocentric trip times. This mission architecture offers a potential reduction in mass to LEO in comparison to alternative all-chemical nuclear propulsion schemes.
Mass reductions may allow launch vehicle downsizing enable missions that would have been grounded because of cost constraints. The preceding figure illustrates a conceptual SEP stage design for a human Mars mission. Researchers at the NASA Glenn Research Center at Lewis Field designed conceptual SEP vehicle, conceived the mission architecture to use this vehicle, and analyzed the vehicle's performance. This SEP stage has a dry mass of 35 metric tons (MT), 40 MT of xenon propellant, and a photovoltaic array that spans 110 m, providing power to a cluster of eight 100-kW Hall thrusters.
The stage can transfer an 80-MT payload and upper stage to the desired HEEPO. Preliminary packaging studies show this space-station-class SEP vehicle meets the proposed ‘Magnum’ launch vehicle and volume requirements with considerable margin. An SEP vehicle for outer planetary missions, such as the Europa Mapper Mission, would be dramatically smaller than human Mars mission SEP stage. In this mission architecture, the SEP power system with the payload to provide spacecraft power throughout the mission. Several photovoltaic array design concepts were considered for the SEP vehicle power system for the human mission to Mars. These include a space station derivative, a SCARLET (Solar Concentrator Arrays with Refractive Linear Element Technology) derivative, and a hybrid inflatable-deployable thin polymer membrane array with thin-film solar cells (as shown in the concept illustration). This concept is based on a design developed for the Next Generation Space Telescope Sun shield. The array is divided into 16 independent electrical sections with 500-V, negative-grounded solar cell strings.
The power system employs a channelized, 500-Vdc power management and distribution (PMAD) architecture with lithium ion batteries for energy storage for vehicle and payload secondary loads (the high-power Hall thrusters do not operate in eclipse periods). The 500-V PMAD voltage permits ‘direct-drive’ thruster operation, greatly reducing the power processing unit size, complexity, and power loss. Similar power system architecture, designs, and technology are assumed for the Europa Mapper Mission SEP vehicle. The primary exceptions are that the photovoltaic array is assumed to consist of two rectangular wings and that the power system rating is 15 kW in Earth orbit and 200 W at Europa. To size the SEP vehicle power system, a dedicated Fortran code was developed to predict detailed power system performance, mass, and thermal control requirements. This code also modeled all the relevant Earth orbit environments; that is, the particulate radiation, plasma, meteoroids and debris, ultraviolet radiation, contamination, and thermal conditions. Analysis results for the Human Mars Mission SEP vehicle show a power system mass of 9-MT and photovoltaic array area of 5800-square meters for the thin-membrane design concept with CuInS2 thin-film cells. Power processing unit input power for a thin-membrane array design with three-junction, amorphous SiGe solar cells is shown in the graph. Power falls off rapidly inhe first weeks of the mission because of light-induced (Staebler-Wronksi) solar cell losses. During the next 200 days, power decreases steadily as the SEP stage spirals through the proton belts and sustains the bulk of the mission radiation damage. Once the vehicle apogee is above approximately four Earth radii, little additional degradation is incurred.
From 400 to 800 days, a 1100-km ‘parking’ orbit is maintained to await the next payload transfer opportunity. This orbit is below the main proton belt, and thus, little radiation dose is accumulated during this time period. During the second LEO-to-HEEPO transfer, power degrades somewhat further, but power requirements are still met. In comparison, the Europa Mapper SEP vehicle power system had a mass of 150 kg and a thin membrane array area of 100 square meters. Author
NASA Programs; Solar Electric Propulsion; Spacecraft Design; Power Conditioning; Launch Vehicles
20050195834 NASA Glenn Research Center, Cleveland, OH, USA
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Thermal Barriers Developed for Solid Rocket Motor Nozzle Joints
Steinetz, Bruce M.; Dunlap, Patrick H., Jr.; Research and Technology 1999; March 2000; 5 pp.; In English; Original contains color and black and white illustrations; No Copyright; Avail: CASI; A01, Hardcopy
Space shuttle solid rocket motor case assembly joints are sealed with conventional O-ring seals that are shielded from 5500 F combustion gases by thick layers of insulation and by special joint-fill compounds that fill assembly splitlines in the insulation. On a number of occasions, NASA has observed hot gas penetration through defects in the joint-fill compound of several of the rocket nozzle assembly joints. In the current nozzle-to-case joint, NASA has observed penetration of hot combustion gases through the joint-fill compound to the inboard wiper O-ring in one out of seven motors. Although this condition does not threaten motor safety, evidence of hot gas penetration to the wiper O-ring results in extensive reviews before resuming flight. The solid rocket motor manufacturer (Thiokol) approached the NASA Glenn Research Center at Lewis Field about the possibility of applying Glenn’s braided fiber preform seal as a thermal barrier to protect the O-ring seals. Glenn and Thiokol are working to improve the nozzle-to-case joint design by implementing a more reliable J-leg design and by using a braided carbon fiber thermal barrier that would resist any hot gases that the J-leg does not block. Author (revised)
O Ring Seals; Thermal Insulation; Space Shuttle Boosters; Joints (Junctions); Rocket Engine Cases; Braided Composites
20050196156 Air Force Research Lab., Edwards AFB, CA USA
Materials Modeling for Rocket Propulsion
Boatz, Jerry; Apr. 2005; 23 pp.; In English Contract(s)/Grant(s): Proj-2303 Report No.(s): AD-A435038; No Copyright; Avail: CASI; A03, Hardcopy
OUTLINE: 1. Introduction; 2. Technical challenges in propellant design; 3. Modeling and Simulation (M&S) techniques & tools: a) Quantum chemistry; b) Molecular dynamics; c) QSPR; d) High Performance Computing (HPC); 4. Examples: a) Identification of suitable target compounds; b) Determination of viable intermediates; c) Confirmation of successful synthesis; 5. Summary and Conclusions. DTIC
Models; Propulsion; Rocket Propellants; Simulation
20050196551 NASA Glenn Research Center, Cleveland, OH, USA
Coupled-Flow Simulation of HP-LP Turbines Has Resulted in Significant Fuel Savings
Veres, Joseph P.; Research and Technology 2000; March 2001; 2 pp.; In English; Original contains color illustrations; No Copyright; Avail: CASI; A01, Hardcopy
Our objective was to create a high-fidelity Navier-Stokes computer simulation of the flow through the turbines of a modern high-bypass-ratio turbofan engine. The simulation would have to capture the aerodynamic interactions between closely coupled high- and low-pressure turbines.
A computer simulation of the flow in the GE90 turbofan engine’s high-pressure (HP) and low-pressure (LP) turbines was created at GE Aircraft Engines under contract with the NASA Glenn Research Center. The three-dimensional steady-state computer simulation was performed using Glenn’s average-passage approach named APNASA. The areas upstream and downstream of each blade row mutually interact with each other during engine operation. The embedded blade row operating conditions are modeled since the average passage equations in APNASA actively include the effects of the adjacent blade rows. The turbine airfoils, platforms, and casing are actively cooled by compressor bleed air. Hot gas leaks around the tips of rotors through labyrinth seals. The flow exiting the high work HP turbines is partially transonic and, therefore, has a strong shock system in the transition region.
The simulation was done using 121 processors of a Silicon Graphics Origin 2000 (NAS 02K) cluster at the NASA Ames Research Center, with a parallel efficiency of 87 percent in 15 hr. The typical average-passage analysis mesh size per blade row was 280 by 45 by 55, or approx.700,000 grid points. The total number of blade rows was 18 for a combined HP and LP turbine system including the struts in the transition duct and exit guide vane, which contain 12.6 million grid points. Design cycle turnaround time requirements ran typically from 24 to 48 hr of wall clock time. The number of iterations for convergence was 10,000 at 8.03x10(exp -5) sec/iteration/grid point (NAS O2K). Parallel processing by up to 40 processors is required to meet the design cycle time constraints. This is the first-ever flow simulation of an HP and LP turbine. In addition, it includes the struts in the transition duct and exit guide vanes. Author
Turbines; Turbofan Engines; Ducted Flow; Bypass Ratio; Computerized Simulation; Navier-Stokes Equation; High Pressure; Low Pressure
20050196554 NASA Glenn Research Center, Cleveland, OH, USA
Laser High-Cycle Thermal Fatigue of Pulse Detonation Engine Combustor Materials Tested
Zhu, Dong-Ming; Fox, Dennis S.; Miller, Robert A.; Research and Technology 2000; March 2001; 5 pp.; In English; Original contains color and black and white illustrations; No Copyright; Avail: CASI; A01, Hardcopy
Pulse detonation engines (PDE’s) have received increasing attention for future aerospace propulsion applications. Because the PDE is designed for a high-frequency, intermittent detonation combustion process, extremely high gas temperatures and pressures can be realized under the nearly constant-volume combustion environment.
The PDE’s can potentially achieve higher thermodynamic cycle efficiency and thrust density in comparison to traditional constant-pressure combustion gas turbine engines (ref. 1). However, the development of these engines requires robust design of the engine components that must endure harsh detonation environments. In particular, the detonation combustor chamber, which is designed to sustain and confine the detonation combustion process, will experience high pressure and temperature pulses with very short durations (refs. 2 and 3). Therefore, it is of great importance to evaluate PDE combustor materials and components under simulated engine temperatures and stress conditions in the laboratory.
In this study, a high-cycle thermal fatigue test rig was established at the NASA Glenn Research Center using a 1.5-kW CO2 laser. The high-power laser, operating in the pulsed mode, can be controlled at various pulse energy levels and waveform distributions. The enhanced laser pulses can be used to mimic the time-dependent temperature and pressure waves encountered in a pulsed detonation engine. Under the enhanced laser pulse condition, a maximum 7.5-kW peak power with a duration of approximately 0.1 to 0.2 msec (a spike) can be achieved, followed by a plateau region that has about one-fifth of the maximum power level with several milliseconds duration. The laser thermal fatigue rig has also been developed to adopt flat and rotating tubular specimen configurations for the simulated engine tests.
More sophisticated laser optic systems can be used to simulate the spatial distributions of the temperature and shock waves in the engine. Pulse laser high-cycle thermal fatigue behavior has been investigated on a flat Haynes 188 alloy specimen, under the test condition of 30-Hz cycle frequency (33-msec pulse period and 10-msec pulse width including a 0.2-msec pulse spike; ref. 4). Temperature distributions were calculated with one-dimensional finite difference models. The calculations show that that the 0.2-msec pulse spike can cause an additional 40 C temperature fluctuation with an interaction depth of 0.08 mm near the specimen surface region. This temperature swing will be superimposed onto the temperature swing of 80 C that is induced by the 10-msec laser pulse near the 0.53-mm-deep surface interaction region. Author
Pulse Detonation Engines; High Temperature Gases; Carbon Dioxide Lasers; Engine Tests; Gas Turbine Engines; Thermodynamic Cycles; Thermal Fatigue
20050196574 Ohio Aerospace Inst., OH, USA
High-Power Magnetoplasmadynamic Thruster Being Developed
LaPointe, Michael R.; Research and Technology 2000; March 2001; 2 pp.; In English; Original contains color illustrations; No Copyright; Avail: CASI; A01, Hardcopy
High-power electromagnetic thrusters have been proposed as primary in-space propulsion options for several of the bold new interplanetary and deep space missions envisioned by the Human Exploration and Development of Space (HEDS) Strategic Enterprise. As the lead center for electric propulsion, the NASA Glenn Research Center is actively involved in the design, development, and testing of high-power electromagnetic technologies to meet these demanding mission requirements. One concept of particular interest is the magnetoplasmadynamic (MPD) thruster, shown schematically in the preceding figure. In its basic form, the MPD thruster consists of a central cathode surrounded by a concentric cylindrical anode. A high-current arc is struck between the anode and cathode, which ionizes and accelerates a gas (plasma) propellant. In the self-field version of the thruster, an azimuthal magnetic field generated by the current returning through the cathode interacts with the radial discharge current flowing through the plasma to produce an axial electromagnetic body force, providing thrust. In applied field-versions of the thruster, a magnetic field coil surrounding the anode is used to provide additional radial and axial magnetic fields that can help stabilize and accelerate the plasma propellant. The following figure shows an experimental megawatt-class MPD thruster developed at Glenn. The MPD thruster is fitted inside a magnetic field coil, which in turn is mounted on a thrust stand supported by thin metal flexures. A calibrated position transducer is used to determine the force provided by the thruster as a function of thrust stand displacement. Power to the thruster is supplied by a 250-kJ capacitor bank, which provides up to 30- MW to the thruster for a period of 2 msec. This short period of time is sufficient to establish thruster performance similar to steady-state operation, and it allows a number of thruster designs to be quickly and economically evaluated. In concert with this experimental research, Glenn is also developing and using advanced numerical simulations to predict the performance of self-field and applied-field MPD thrusters. Author
Magnetoplasmadynamic Thrusters; Interplanetary Space; Deep Space; Space Missions; Electric Propulsion
20050196577 NASA Glenn Research Center, Cleveland, OH, USA
1000 Hours of Testing Completed on 10-kW Hall Thruster
Mason, Lee S.; Research and Technology 2000; March 2001; 2 pp.; In English; Original contains color illustrations; No Copyright; Avail: CASI; A01, Hardcopy
Between the months of April and August 2000, a 10-kW Hall effect thruster, designated T- 220, was subjected to a 1000-hr life test evaluation. Hall effect thrusters are propulsion devices that electrostatically accelerate xenon ions to produce thrust. Hall effect propulsion has been in development for many years, and low-power devices (1.35 kW) have been used in space for satellite orbit maintenance. The T-220, shown in the photo, produces sufficient thrust to enable efficient orbital transfers, saving hundreds of kilograms in propellant over conventional chemical propulsion systems. This test is the longest operation ever achieved on a high-power Hall thruster (greater than 4.5 kW) and is a key milestone leading to the use of this technology for future NASA, commercial, and military missions. Author
Hall Thrusters; Propulsion System Configurations; Chemical Propulsion; Hall Effect; Propellants
20050196585 NASA Glenn Research Center, Cleveland, OH, USA
Numerical Propulsion System Simulation: A Common Tool for Aerospace Propulsion Being Developed
Follen, Gregory J.; Naiman, Cynthia G.; Research and Technology 2000; March 2001; 3 pp.; In English; Original contains color illustrations; No Copyright; Avail: CASI; A01, Hardcopy
The NASA Glenn Research Center is developing an advanced multidisciplinary analysis environment for aerospace propulsion systems called the Numerical Propulsion System Simulation (NPSS). This simulation is initially being used to support aeropropulsion in the analysis and design of aircraft engines. NPSS provides increased flexibility for the user, which reduces the total development time and cost. It is currently being extended to support the Aviation Safety Program and Advanced Space Transportation. NPSS focuses on the integration of multiple disciplines such as aerodynamics, structure, and heat transfer with numerical zooming on component codes. Zooming is the coupling of analyses at various levels of detail. NPSS development includes using the Common Object Request Broker Architecture (CORBA) in the NPSS Developer’s Kit to facilitate collaborative engineering. The NPSS Developer’s Kit will provide the tools to develop custom components and to use the CORBA capability for zooming to higher fidelity codes, coupling to multidiscipline codes, transmitting secure data, and distributing simulations across different platforms. These powerful capabilities will extend NPSS from a zero-dimensional simulation tool to a multifidelity, multidiscipline system-level simulation tool for the full life cycle of an engine. Derived from text
Propulsion System Configurations; Propulsion System Performance; Spacecraft Propulsion
20050196606 NASA Glenn Research Center, Cleveland, OH, USA
Assessment of Stirling Technology Has Provided Critical Data Leading Toward Flight Readiness of the Stirling Converter
Thieme, Lanny G.; Research and Technology 2000; March 2001; 3 pp.; In English; No Copyright; Avail: CASI; A01, Hardcopy
The NASA Glenn Research Center is supporting the development of a Stirling converter with the Department of Energy (DOE, Germantown, Maryland) for an advanced Stirling Radioisotope Power System (SRPS) to provide spacecraft onboard electric power for NASA space science missions.
A key technology assessment completed by Glenn and DOE has led to the SRPS being identified as a high-efficiency power source for such deep space missions as the Europa Orbiter and the Solar Probe. In addition, the Stirling system is now being considered for unmanned Mars rovers, especially where mission profiles may exclude the use of photovoltaic power systems, such as exploration at high Martian latitudes or for missions of long duration.
The SRPS efficiency of over 20 percent will reduce the required amount of radioisotope by more than a factor of 3 in comparison to current radioisotope thermoelectric generators. This significantly reduces radioisotope cost, radiological inventory, and system cost, and it provides efficient use of scarce radioisotope resources. In support of this technology assessment, Glenn conducted a series of independent evaluations and tests to determine the technology readiness of a 55-We Stirling converter developed by Stirling Technology Company (Kennewick, Washington) and DOE.
Key areas evaluated by Glenn included:
1) Radiation tolerance of materials;
2) Random vibration testing of the Stirling converter in Glenn’s Structural Dynamics Lab to simulate operation in the launch environment;
3) Electromagnetic interference and compatibility (EMI/EMC) of the converter operating in Glenn’s EMI lab; Independent failure modes, effects, and criticality analysis, and life and reliability 4.
Independent failure modes, effects, and criticality analysis, and life and reliability assessment; and 5) SRPS cost estimate. The data from these evaluations were presented to NASA Headquarters and the Jet Propulsion Laboratory mission office by a joint industry/Government team consisting of DOE, Glenn, and Lockheed Martin Astronautics. This team concluded that there are no technical reasons that would rule out using the Stirling converter for deep space missions. As a direct result of the successful testing at Glenn, the DOE/Stirling Technology Company 55-We Stirling converter has been baselined for the SRPS. Glenn is now continuing an in-house project to assist in developing the Stirling converter for readiness for space qualification and mission implementation. As part of this effort, the Stirling converter will be further characterized under launch environment random vibration testing, methods to reduce converter EMI will be developed, and an independent performance verification will be completed. Converter life assessment and permanent magnet aging characterization tasks are also underway. Substitute organic materials for the linear alternator and piston bearing coatings for use in a high-radiation environment have been identified and have now been incorporated in Stirling converters built by Stirling Technology Company for Glenn. Electromagnetic and thermal finite element analyses for the alternator are also being conducted. Author
Stirling Cycle; Technology Assessment; NASA Space Programs; Space Missions
20050196614 NASA Glenn Research Center, Cleveland, OH, USA
Pulse Detonation Engine Modeled
Paxson, Daniel E.; Research and Technology 2000; March 2001; 2 pp.; In English; No Copyright; Avail: CASI; A01, Hardcopy
Pulse Detonation Engine Technology is currently being investigated at Glenn for both airbreathing and rocket propulsion applications. The potential for both mechanical simplicity and high efficiency due to the inherent near-constant-volume combustion process, may make Pulse Detonation Engines (PDE’s) well suited for a number of mission profiles.
Assessment of PDE cycles requires a simulation capability that is both fast and accurate. It should capture the essential physics of the system, yet run at speeds that allow parametric analysis. A quasi-one-dimensional, computational-fluid-dynamics-based simulation has been developed that may meet these requirements. The Euler equations of mass, momentum, and energy have been used along with a single reactive species transport equation, and submodels to account for dominant loss mechanisms (e.g., viscous losses, heat transfer, and valving) to successfully simulate PDE cycles. A high-resolution numerical integration scheme was chosen to capture the discontinuities associated with detonation, and robust boundary condition procedures were incorporated to accommodate flow reversals that may arise during a given cycle. The accompanying graphs compare experimentally measured and computed performance over a range of operating conditions for a particular PDE.
Experimental data were supplied by Fred Schauer and Jeff Stutrud from the Air Force Research Laboratory at Wright-Patterson AFB and by Royce Bradley from Innovative Scientific Solutions, Inc. The left graph shows thrust and specific impulse, Isp, as functions of equivalence ratio for a PDE cycle in which the tube is completely filled with a detonable hydrogen/air mixture. The right graph shows thrust and specific impulse as functions of the fraction of the tube that is filled with a stoichiometric mixture of hydrogen and air. For both figures, the operating frequency was 16 Hz. The agreement between measured and computed values is quite good, both in terms of trend and magnitude. The error is under 10 percent everywhere except for the thrust value at an equivalence ratio of 0.8 in the left figure, where it is 14 percent. The simulation results shown were made using 200 numerical cells.
Each cycle of the engine, approximately 0.06 sec, required 2.0 min of CPU time on a Sun Ultra2. The simulation is currently being used to analyze existing experiments, design new experiments, and predict performance in propulsion concepts where the PDE is a component (e.g., hybrid engines and combined cycles). Author
Pulse Detonation Engines; Performance Prediction; Reaction Kinetics; Rocket Thrust; Specific Impulse; Heat Transfer
20050196660 NASA Glenn Research Center, Cleveland, OH, USA
Single-String Integration Test Measurements of the NEXT Ion Engine Plume
Snyder, Aaron; Kamhawi, Hani; Patterson, Michael; Britton, Melissa; June 2005; 25 pp.; In English; 40th Joint Propulsion Conference and Exhibit, 11-14 Jul. 2004, Fort Lauderdale, FL, USA Contract(s)/Grant(s): WBS 22-800-50-01 Report No.(s): NASA/TM-2005-213196; E-14697; AIAA Paper 2004-3790; No Copyright; Avail: CASI; A03, Hardcopy
Measurements were made of a 40 cm ion-thruster plume as part of the single-string-integration-test (SSIT) activity of Phase I of the NASA’s Evolutionary Xenon Thruster (NEXT) project. The NEXT ion engine incorporates design improvements that extend NSTAR power levels and efficiencies. During SSIT, an engineering model (EM2) 40 cm engine was operated using an advanced xenon propellant system in combination with either a GRC power console or advanced power processing unit. Integral goals of the single-string phase were to characterize engine performance over the full input power range and to detail thruster operation within the specification of the NEXT throttle table. Plume diagnostics measurements of relative Xe(+) and Xe(++) currents were made using near-field and far-field ExB probes. Planar geometry faraday probes were used to obtain beam current density profiles. This paper reports on the characterization of the EM2 plume over a range of SSIT operating conditions, first with the advanced propellant management system teamed with the GRC power console and then with the power-processing unit. Autho
rIon Engines; Diagnosis; Plumes; Xenon
20050196671 NASA Glenn Research Center, Cleveland, OH, USA
Solid Hydrogen Particles Analyzed for Atomic Fuels
Palaszewski, Bryan A.; Research and Technology 2000; March 2001; 2 pp.; In English; No Copyright; Avail: CASI; A01, Hardcopy
Solid hydrogen particles have been selected as a means of storing atomic propellants in future launch vehicles (refs. 1 to 2). In preparation for this, hydrogen particle formation in liquid helium was tested experimentally. These experiments were conducted to visually characterize the particles and to observe their formation and molecular transformations (aging) while in liquid helium. The particle sizes, molecular transformations, and agglomeration times were estimated from video image analyses. The experiments were conducted at the NASA Glenn Research Center in the Supplemental Multilayer Insulation Research Facility (SMIRF, ref. 3). The facility has a vacuum tank, into which the experimental setup was placed. The vacuum tank prevented heat leaks and subsequent boiloff of the liquid helium, and the supporting systems maintained the temperature and pressure of the liquid helium bath where the solid particles were created. As the operation of the apparatus was developed, the hydrogen particles were easily visualized. The figures (ref. 1) show images from the experimental runs. The first image shows the initial particle freezing, and the second image shows the particles after the small particles have agglomerated. The particles finally all clump, but stick together loosely. The solid particles tended to agglomerate within a maximum of 11 min, and the agglomerate was very weak. Because the hydrogen particles are buoyant in the helium, the agglomerate tends to compact itself into a flat pancake on the surface of the helium. This pancake agglomerate is easily broken apart by reducing the pressure above the liquid. The weak agglomerate implies that the particles can be used as a gelling agent for the liquid helium, as well as a storage medium for atomic boron, carbon, or hydrogen. The smallest particle sizes that resulted from the initial freezing experiments were about 1.8 mm. About 50 percent of the particles formed were between 1.8 to 4.6 mm in diameter. These very small particle sizes are encouraging for future formation experiments, where simpler operations will reduce the costs of production. Author
Propellants; Solid Cryogens; Hydrogen; Fuels; Elementary Particles; Liquid Helium; Multilayer Insulation
20050196705 NASA Glenn Research Center, Cleveland, OH, USA
Validation of the NCC Code for Staged Transverse Injection and Computations for a RBCC Combustor
Ajmani, Kumud; Liu, Nan-Suey; July 2005; 36 pp.; In English Contract(s)/Grant(s): WBS 22-714-20-22 Report No.(s): NASA/TM-2005-213647; E-15135; No Copyright; Avail: CASI; A03, Hardcopy
The NCC code was validated for a case involving staged transverse injection into Mach 2 flow behind a rearward facing step. Comparisons with experimental data and with solutions from the FPVortex code was then used to perform computations to study fuel-air mixing for the combustor of a candidate rocket based combined cycle engine geometry. Comparisons with a one-dimensional analysis and a three-dimensional code (VULCAN) were performed to assess the qualitative and quantitative performance of the NCC solver.
Author
Injection; Combustion Chambers; Backward Facing Steps
20050196823 NASA Glenn Research Center, Cleveland, OH, USA
REP Concept Feasibility Study
Edwards, Daryl A.; Ensworth, Clinton B. F.; Goodnight, ThomasW.; Sheehe, Charles J.;Wiersma, Stephen C.; Adamsen, Paul B., II; Frank, Larry; [2004]; 9 pp.; In English; 55th International Astronautical Congress, 4-8 Oct. 2004, Vancouver, Canada Contract(s)/Grant(s): 22-972-30-07 Report No.(s): E-14845; IAC-04-IAA.3.6.P.01; Copyright; Avail: CASI; A02, Hardcopy
Radioisotope Electric Propulsion (REP) may have the potential to provide certain advantages, over conventional chemical propulsion, for outer planetary exploration involving small bodies and long term investigations for medium class missions requiring power comparable to past outer planetary exploration missions. This paper describes a study that investigates the concept s feasibility by performing a preliminary conceptual design of an REP-based spacecraft for a design reference mission. The mission utilizes a spacecraft with a radioisotope power supply less than one kilowatt while operating for a minimum of 10-years. A key element of the REP spacecraft is to ensure sustained science return by orbiting or flying in formation with selected targets. Utilizing current and impending technological advances, this study finds that at a conceptual design level a small body REP orbiter/explorer appears to be feasible for the design reference mission selected for this study. Author
Electric Propulsion; Nuclear Electric Propulsion; Spacecraft Propulsion; Radioactive Isotopes; Feasibility Analysis; Interplanetary Spacecraft
20050198947 NASA Glenn Research Center, Cleveland, OH, USA
Photovoltaic Cell Operation on Mars
Landis, Geoffrey A.; Kerslake, Thomas; Jenkins, Phillip P.; Scheiman, David A.; [2004]; 4 pp.; In English; 19th European Photovoltaic Solar Energy Power Conference, 7-11 Jun. 2004, Paris, France Contract(s)/Grant(s): WBS 22-759-20-01 Report No.(s): E-14870; No Copyright; Avail: CASI; A01, Hardcopy
The Martian surface environment provides peculiar challenges for the operation of solar arrays: low temperature, solar flux with a significant scattered component that varies in intensity and spectrum with the amount of suspended atmospheric dust, and the possibility of performance loss due to dust deposition on the array surface. This paper presents theoretical analyses of solar cell performance on the surface of Mars and measurements of cells under Martian conditions. Author
Photovoltaic Cells; Spacecraft Power Supplies; Mars Roving Vehicles; Mars Environment; Operational Problems; Mars Missions
Source: NASA.
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