SCIENTIFIC AND TECHNICAL AEROSPACE REPORTS
A Biweekly Publication of the National Aeronautics and Space Administration
VOLUME 43, ISSUE 20 - OCTOBER 07, 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.
20050214857 NASA Glenn Research Center, Cleveland, OH, USA
Explicit Finite Element Techniques Used to Characterize Splashdown of the Space Shuttle Solid Rocket Booster Aft Skirt
Melis, Matthew E.; Research and Technology 2002; March 2003; 2 pp.; In English; No Copyright; Avail.: CASI: A01, Hardcopy
NASA Glenn Research Center s Structural Mechanics Branch has years of expertise in using explicit finite element methods to predict the outcome of ballistic impact events. Shuttle engineers from the NASA Marshall Space Flight Center and NASA Kennedy Space Flight Center required assistance in assessing the structural loads that a newly proposed thrust vector control system for the space shuttle solid rocket booster (SRB) aft skirt would expect to see during its recovery splashdown. Derived from text
Finite Element Method; Space Shuttle Boosters; Structural Analysis
20050215029 NASA Glenn Research Center, Cleveland, OH, USA
International Space Station Power System Telemetry Compared With Analytically Derived Data for Shadowed Cases
Fincannon, H. James; Research and Technology 2002; March 2003; 3 pp.; In English; No Copyright; Avail.: CASI: A01, Hardcopy
This article highlights fiscal year 2002 work performed by NASA Glenn Research Center personnel to validate algorithms and data developed in-house to predict shadowing effects on the International Space Station (ISS) solar arrays power generation. The validation effort utilized video footage and on-orbit telemetry for cases spanning a 1-yr period. Validation was required because of the uncertainty of various aspects involved in shadowing analysis. Results show that a good comparison exists between actual and predicted shadowed power system performance for solar array front and backside shadowing. Derived from text
International Space Station; Power Supplies; Telemetry
20050215032 NASA Glenn Research Center, Cleveland, OH, USA
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6-ft High-Power Electric Propulsion Test Port, EPL Tank 5 Installed
Grisnik, Stanley P.; Theman, Elmer R.; Swiatek, Michael W.; Speicer, Henry J.; Jankovsky, Robert S.; Research and Technology 2002; March 2003; 2 pp.; In English; No Copyright; Avail.: CASI: A01, Hardcopy
High-power electric propulsion is a critical component of NASA s proposed missions to the outer planets. Mission studies have shown that high-power, high-specific-impulse propulsion systems can deliver 2000 kg of scientific payload to Pluto with trip times on the order of 10 years. Of greater significance is the ability of these propulsion systems to place this science payload in orbit around the planet, rather than making the fast fly-bys associated with traditional chemical propulsion systems. Significant ground test programs are required to develop the new technologies needed for thrusters operating at power levels exceeding 20 kW, an order of magnitude above the state of the art. Derived from text
Electric Propulsion; Ground Tests; Propulsion System Configurations; Propulsion System Performance
20050215033 NASA Glenn Research Center, Cleveland, OH, USA
Research Combustion Lab Facility Capabilities and Throughput Enhanced by New Test Stands Research and Technology 2002
March 2003; 3 pp.; In English; No Copyright; Avail.: CASI: A01, Hardcopy
The original test stand location has a small copper rocket engine mounted on the stand. The new stand, located about 4 feet to the left, has a long pulse detonation combustion engine mounted on it. To the rear of the two stands can be seen a bulkhead with feed line outlets that can be switched at common valves behind the bulkhead to supply either stand. A gauge panel is visible through a doorway in the bulkhead at which various purge pressures are set. A connection panel for instrumentation wiring can be seen above the stands. Derived from text
Research Facilities; Combustion Chambers; Engine Design; Measuring Instruments; Pulse Detonation Engines; Test Stands
20050215041 NASA Glenn Research Center, Cleveland, OH, USA
Liquid Propellant Manipulated Acoustically
Oeftering, Richard C.; Chato, David J.; Mann, Adin, III; Research and Technology 2002; March 2003; 2 pp.; In English; No Copyright; Avail.: CASI: A01, Hardcopy
Fluids are difficult to manage in the space environment.Without gravity, the liquid and gas do not always remain separated as they do in the 1g environment of Earth. Instead the liquid and gas volumes mix and migrate under the influence of surface tension, thermodynamic forces, and external disturbances. As a result, liquid propellants may not be in a useable location or may even form a chaotic mix of liquid and gas bubbles. In the past, mechanical pumps, baffles, and a variety of specialized passive devices have been used to control the liquid and gas volumes. These methods need to be carefully tuned to a specific configuration to be effective. With increasing emphasis on long-term human activity in space there is a trend toward liquid systems that are more flexible and provide greater control. We are exploring new methods of manipulating liquids by using the nonlinear acoustic effects achieved by using beams of highly directed high-intensity acoustic waves. Derived from text
Liquid Rocket Propellants; Sound Waves; Interfacial Tension; Gravitation
20050215095 NASA Glenn Research Center, Cleveland, OH, USA
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Integrated Propulsion/Vehicle System Structurally Optimized
Hunter, James E.; McCurdy, David R.; Research and Technology 2002; March 2003; 3 pp.; In English; No Copyright; Avail.: CASI: A01, Hardcopy
Ongoing research and testing are essential in the development of air-breathing hypersonic propulsion technology, and this year some positive advancement was made at the NASA Glenn Research Center. Recent work performed for GTX, a rocket-based combined-cycle, single-stage-to-orbit concept, included structural assessments of both the engine and flight vehicle. In the development of air-breathing engine technology, it is impractical to design and optimize components apart from the fully integrated system because tradeoffs must be made between performance and structural capability.
Efforts were made to control the flight trajectory, for example, to minimize the aerodynamic heating effects. Structural optimization was applied to evaluate concept feasibility and was instrumental in the determination of the gross liftoff weight of the integrated system.
Achieving low Earth orbit with even a small payload requires an aggressive approach to weight minimization through the use of lightweight, oxidation-resistant composite materials. Assessing the integrated system involved investigating the flight trajectory to determine where the critical load events occur in flight and then generating the corresponding environment at each of these events. Structural evaluation requires the mapping of the critical flight loads to finite element models, including the combined effects of aerodynamic, inertial, combustion, and other loads.
NASA s APAS code was used to generate aerodynamic pressure and temperature profiles at each critical event. The radiation equilibrium surface temperatures from APAS were used to predict temperatures through the thickness. Heat transfer solutions using NASA's MINIVER code and the SINDA code (Cullimore & Ring Technologies, Littleton, CO) were calculated at selective points external to the integrated vehicle system and then extrapolated over the entire exposed surface. FORTRAN codes were written to expedite the finite element mapping of the aerodynamic heating effects for the internal structure. Author
Air Breathing Engines; Hypersonic Flight; Propulsion; Rocket-Based Combined-Cycle Engines; Systems Integration
20050215161 NASA Glenn Research Center, Cleveland, OH, USA
Cryogenic Nitrogen Thermosyphon Developed and Characterized
Plachta, David W.; Christie, Robert; Research and Technology 2003; May 2004; 3 pp.; In English; No Copyright; Avail.: CASI: A01, Hardcopy
A two-phase nitrogen thermosyphon was developed at the NASA Glenn Research Center to efficiently integrate a cryocooler into an insulated liquid-nitrogen-filled tank as part of an advanced development zero-boiloff (ZBO) ground test. NASA Marshall Space Flight Center's (MSFC) Advanced Space Transportation Program supported this test to improve the performance of in-space propulsion system concepts. Recent studies (ref. 1) have shown significant mass reductions and other advantages when incorporating active cooling in a ZBO configuration, enabling consideration of high-performing cryogenic propellants for long-duration applications in space. Active cooling was integrated via a thermosyphon, made of copper, 42 in. (1070 mm) long with an inner diameter of 0.436 in. (11 mm). It was charged with nitrogen to 225 psia at 300 K, which provided a fill ratio of 15 percent. The temperatures and heat flows through the thermosyphon were monitored during the startup phase of the ZBO test, and steady-state tests were conducted over a range of increasing and decreasing heat flows. The results also were compared with the initial design calculations and with results for a similar thermosyphon. They show that the thermal resistance of the thermosyphon was one-half of that expected--0.2 K/W at a heat flow of 8.0 W. The design calculations also showed that this resistance can be made relatively constant over a wider range of heat flows by making the ratio of evaporator area to condenser area 3:1. The better-than-expected results will translate into reduced integration loss for the ZBO concept. Author
Cryogenic Cooling; Liquid Nitrogen; Thermosiphons; Ground Tests; Propulsion System Configurations; Cryogenic Rocket Propellants
20050215170 NASA Glenn Research Center, Cleveland, OH, USA
Pulsed Ejector Thrust Amplification Tested and Modeled
Wilson, Jack; Research and Technology 2003; May 2004; 3 pp.; In English; See Document ID 20050019494; No Copyright; Avail.: CASI: A01, Hardcopy
There is currently much interest in pulsed detonation engines for aeronautical propulsion. This, in turn, has sparked renewed interest in pulsed ejectors to increase the thrust of such engines, since previous, though limited, research had indicated that pulsed ejectors could double the thrust in a short device. An experiment has been run at the NASA Glenn Research Center, using a shrouded Hartmann-Sprenger tube as a source of pulsed flow, to measure the thrust augmentation of a statistically designed set of ejectors. A Hartmann- Sprenger tube directs the flow from a supersonic nozzle (Mach 2 in the present experiment) into a closed tube. Under appropriate conditions, an oscillation is set up in which the jet flow alternately fills the tube and then spills around flow emerging from the tube. The tube length determines the frequency of oscillation. By shrouding the tube, the flow was directed out of the shroud as an axial stream. The set of ejectors comprised three different ejector lengths, three ejector diameters, and three nose radii. The thrust of the jet alone, and then of the jet plus ejector, was measured using a thrust plate. The arrangement is shown in this photograph. Thrust augmentation is defined as the thrust of the jet with an ejector divided by the thrust of the jet alone. The experiments exhibited an optimum ejector diameter and length for maximizing the thrust augmentation, but little dependence on nose radius. Different frequencies were produced by changing the length of the Hartmann-Sprenger tube, and the experiment was run at a total of four frequencies. Additional measurements showed that the major feature of the pulsed jet was a starting vortex ring. The size of the vortex ring depended on the frequency, as did the optimum ejector diameter. Author
Pulse Detonation Engines; Ejectors; Thrust Augmentation; Hartmann-Sprenger Tubes; Jet Flow; Supersonic Nozzles
20050215304 Florida Inst. of Tech., Melbourne, FL, USA
Modeling and Testing of Non-Nuclear, Highpower Simulated Nuclear Thermal Rocket Reactor Elements
Kirk, Daniel R.; The 2004 NASA Faculty Fellowship Program Research Reports; January 2005, pp. XXIV-1 - XXIV-7; In English; See also 20050215300; Original contains color illustrations; No Copyright; Avail.: CASI: A02, Hardcopy
When the President offered his new vision for space exploration in January of 2004, he said, ‘Our third goal is to return to the moon by 2020, as the launching point for missions beyond,' and, ‘With the experience and knowledge gained on the moon, we will then be ready to take the next steps of space exploration: human missions to Mars and to worlds beyond.' A human mission to Mars implies the need to move large payloads as rapidly as possible, in an efficient and cost-effective manner. Furthermore, with the scientific advancements possible with Project Prometheus and its Jupiter Icy Moons Orbiter (JIMO), (these use electric propulsion), there is a renewed interest in deep space exploration propulsion systems. According to many mission analyses, nuclear thermal propulsion (NTP), with its relatively high thrust and high specific impulse, is a serious candidate for such missions. Nuclear rockets utilize fission energy to heat a reactor core to very high temperatures. Hydrogen gas flowing through the core then becomes superheated and exits the engine at very high exhaust velocities. The combination of temperature and low molecular weight results in an engine with specific impulses above 900 seconds. This is almost twice the performance of the LOX/LH2 space shuttle engines, and the impact of this performance would be to reduce the trip time of a manned Mars mission from the 2.5 years, possible with chemical engines, to about 12-14 months. Derived from text
Electric Propulsion; Exhaust Velocity; Hydrogen Oxygen Engines; Nuclear Reactors; Propulsion System Performance; Rocket Thrust; Thermal Reactors
20050215311 State Univ. of New York, NY, USA
The Technology and Art of Aerocapture
Bangs, Mary C.; The 2004 NASA Faculty Fellowship Program Research Reports; January 2005, pp. II-1 - II-5; In English; See also 20050215300; Original contains color and black and white illustrations; No Copyright; Avail.: CASI: A01, Hardcopy
Aerocapture is a type of aero-assist technology in which an interplanetary spacecraft is equipped with a device shaped to optimize atmospheric drag during a pass through a planet s upper atmosphere. An aerocapture maneuver alters the spacecraft s trajectory so that it is captured as a satellite of that planet. Aerocapture is related to aerobraking. In aerobraking, a spacecraft lowers its orbit around a planet by repeated passes through that planet s upper atmosphere. As part of the public outreach effort of the NASA MSFC In-Space Propulsion research effort, a poster is being constructed relating to the new technology of aerocapture. Entitled ‘The Technology & Art of Aerocapture,' this poster will speak to three groups of people. The first and broadest audience is the general public who may or may not be familiar with aerocapture as a concept. The second group is the scientifically educated and informed public that is familiar with the concept of aerocapture but may or may not know the finer points about it. Finally, the third audience consists of the group that has worked with and is familiar with aerocapture. The poster is designed to educate and inspire all these audiences. Author
Aerocapture; Interplanetary Spacecraft; Aerobraking; Aeroassist; Drag
20050215314 Rose-Hulman Inst. of Tech., Terre Haute, IN, USA
Enhancement and Analysis of Real-Time Radiography Images
Doering, Edward R.; The 2004 NASA Faculty Fellowship Program Research Reports; January 2005, pp. XII-1 - XII-5; In English; See also 20050215300; Original contains color and black and white illustrations; No Copyright; Avail.: CASI: A01, Hardcopy
Shuttle Redesigned Solid Rocket Motor (RSRM) nozzle interiors fabricated from carbon phenolic composite exhibit ‘ply lift' when hot fired. The composite surface is smooth when fabricated, but the individual plies separate and lift away from the surface when exposed to high temperature and high-pressure exhaust gas. It shows a cross section of a post-fired composite in which ply lift is evident as dark fissures. Surface charring is also visible as a darker band about 0.2 inches thick. Charring is normal, but ply lift is not desirable since the fissures could possibly initiate an abnormal exhaust path from the RSRM. The underlying mechanisms of ply lift are under investigation as part of the Shuttle Return-To-Flight Program. Derived from text
Solid Propellant Rocket Engines; Fabrication; Carbon-Phenolic Composites; Nozzle Design
20050215320 Alabama Univ., Huntsville, AL, USA
Control of Combustion-Instabilities Through Various Passive Devices
Frendi, Kader; The 2004 NASA Faculty Fellowship Program Research Reports; January 2005, pp. XV-1 - XV-24; In English; See also 20050215300; Original contains color and black and white illustrations; No Copyright; Avail.: CASI: A03, Hardcopy
It is well known that under some operating conditions, rocket engines (using solid or liquid fuels) exhibit unstable modes of operation that can lead to engine malfunction and shutdown. The sources of these instabilities are diverse and are dependent on fuel, chamber geometry and various upstream sources such as pumps, valves and injection mechanism.
It is believed that combustion-acoustic instabilities occur when the acoustic energy increase due to the unsteady heat release of the flame is greater than the losses of acoustic energy from the system [1, 2]. Giammar and Putnam [3] performed a comprehensive study of noise generated by gasfired industrial burners and made several key observations; flow noise was sometimes more intense than combustion roar, which tended to have a characteristic frequency spectrum. Turbulence was amplified by the flame.
The noise power varied directly with combustion intensity and also with the product of pressure drop and heat release rate. Karchmer [4] correlated the noise emitted from a turbofan jet engine with that in the combustion chamber. This is important, since it quantified how much of the noise from an engine originates in the combustor. A physical interpretation of the interchange of energy between sound waves and unsteady heat release rates was given by Rayleigh [5] for inviscid, linear perturbations.
Bloxidge et al [6] extended Rayleigh's criterion to describe the interaction of unsteady combustion with one-dimensional acoustic waves in a duct. Solutions to the mass, momentum and energy conservation equations in the preand post-flame zones were matched by making several assumptions about the combustion process. They concluded that changes in boundary conditions affect the energy balance of acoustic waves in the combustor.
Abouseif et al [7] also solved the one-dimensional flow equations, but they used a onestep reaction to evaluate the unsteady heat release rate by relating it to temperature and velocity perturbations. Their analysis showed that oscillations arise from coupling between entropy waves produced at the flame and pressure waves originating from the nozzle.
Yang and Culick [8] assumed a thin flame sheet, which is distorted by velocity and pressure oscillations. Conservation equations were expressed in integral form and solutions for the acoustic wave equations and complex frequencies were obtained. The imaginary part of the frequency indicated stability regions of the flame. Activation energy asymptotics together with a one-step reaction were used by McIntosh [9] to study the effects of acoustic forcing and feedback on unsteady, one-dimensional flames. He found that the flame stability was altered by the upstream acoustic feedback.
Shyy et al [10] used a high-accuracy TVD scheme to simulate unsteady, one-dimensional longitudinal, combustion instabilities. However, numerical diffusion was not completely eliminated.
Recently, Prasad [11] investigated numerically the interactions of pressure perturbations with premixed flames. He used complex chemistry to study responses of pressure perturbations in one-dimensional combustors. His results indicated that reflected and transmitted waves differed significantly from incident waves. Author
Rocket Engines; Combustion; Heat Transfer; Pressure Oscillations; Frequency Stability; Acoustic Instability; Flame Stability; Malfunctions
20050215406 NASA Glenn Research Center, Cleveland, OH, USA
Physics of Colloids in Space--Plus (PCS+) Experiment Completed Flight Acceptance Testing
Doherty, Michael P.; Research and Technology 2003; May 2004; 3 pp.; In English; No Copyright; Avail.: CASI: A01, Hardcopy
The Physics of Colloids in Space--Plus (PCS+) experiment successfully completed system-level flight acceptance testing in the fall of 2003. This testing included electromagnetic interference (EMI) testing, vibration testing, and thermal testing. PCS+, an Expedite the Process of Experiments to Space Station (EXPRESS) Rack payload will deploy a second set of colloid samples within the PCS flight hardware system that flew on the International Space Station (ISS) from April 2001 to June 2002. PCS+ is slated to return to the ISS in late 2004 or early 2005. Derived from text
Colloids; Electromagnetic Interference; Vibration; Thermal Analysis
20050215408 NASA Glenn Research Center, Cleveland, OH, USA
Solid Fuel Delivery System Developed for Combustion Testing on the International Space Station
Frate, David T.; Research and Technology 2003; May 2004; 3 pp.; In English; No Copyright; Avail.: CASI: A01, Hardcopy
NASA initiated Bioastronautics and Human Research Initiatives in 2001 and 2003, respectively, to enhance the safety and performance of humans in space. The Flow Enclosure Accommodating Novel Investigations in Combustion of Solids (FEANICS) is a multiuser facility being built at the NASA Glenn Research Center to advance these initiatives by studying fire safety and the combustion of solid fuels in the microgravity environment of the International Space Station (ISS). One of the challenges for the FEANICS team was to build a system that allowed for several consecutive combustion tests to be performed with minimal astronaut crew interaction. FEANICS developed a fuel carousel that contains a various number of fuel samples, depending on the fuel width, and introduces them one at a time into a flow tunnel in which the combustion testing takes place. This approach will allow the science team to run the experiments from the ground, while only requiring the crew to change out carousels after several tests have been completed. Derived from text
Fuel Systems; Solid Propellants; Combustion; Performance Tests
20050215409 NASA Glenn Research Center, Cleveland, OH, USA
Combustion Module-2 Achieved Scientific Success on Shuttle Mission STS-107
Over, Ann P.; Research and Technology 2003; May 2004; 5 pp.; In English; No Copyright; Avail.: CASI: A01, Hardcopy
The familiar teardrop shape of a candle is caused by hot, spent air rising and cool fresh air flowing behind it. This type of airflow obscures many of the fundamental processes of combustion and is an impediment to our understanding and modeling of key combustion controls used for manufacturing, transportation, fire safety, and pollution. Conducting experiments in the microgravity environment onboard the space shuttles eliminates these impediments. NASA Glenn Research Center's Combustion Module-2 (CM-2) and its three experiments successfully flew on STS-107/Columbia in the SPACEHAB module and provided the answers for many research questions. However, this research also opened up new questions. The CM-2 facility was the largest and most complex pressurized system ever flown by NASA and was a precursor to the Glenn Fluids and Combustion Facility planned to fly on the International Space Station. CM-2 operated three combustion experiments: Laminar Soot Processes (LSP), Structure of Flame Balls at Low Lewis-Number (SOFBALL), and Water Mist Fire Suppression Experiment (Mist). Although Columbia's mission ended in tragedy with the loss of her crew and much data, most of the CM-2 results were sent to the ground team during the mission. Author
Combustion Chambers; Combustion Chemistry; Fire Prevention; Spacecraft Modules
20050215568 NASA Marshall Space Flight Center, Huntsville, AL, USA
Preliminary Assessment of Thrust Augmentation of NEP Based Missions
Chew, Gilbert; Pelaccio, Dennis G.; Chiroux, Robert; Pervan, Sherry; Rauwolf, Gerald A.; White, Charles; [2005]; 2 pp.; In English; AIAA 2005 Space Conference, 30 Aug. - 1 Sep. 2005, Long Beach, CA, USA; Copyright; Avail.: CASI: A01, Hardcopy
Science Applications International Corporation (SAIC), with support from NASA Marshall Space Flight Center, has conducted a preliminary study to compare options for augmenting the thrust of a conventional nuclear electric propulsion (NEP) system. These options include a novel nuclear propulsion system concept known as Hybrid Indirect Nuclear Propulsion (HINP) and conventional chemical propulsion. The utility and technical feasibility of the HINP concept are assessed, and features and potential of this new in-space propulsion system concept are identified. As part of the study, SAIC developed top-level design tools to model the size and performance of an HINP system, as well as for several chemical propulsion options, including liquid and gelled propellants. A mission trade study was performed to compare a representative HINP system with chemical propulsion options for thrust augmentation of NEP systems for a mission to Saturn's moon Titan. Details pertaining to the approach, features, initial demonstration results for HINP model development, and the mission trade study are presented. Key technology and design issues associated with the HINP concept and future work recommendations are also identified. Derived from text
Nuclear Electric Propulsion; Thrust Augmentation; Space Missions; NASA Space Programs; Technology Assessment
20050215598 NASA Glenn Research Center, Cleveland, OH, USA
Bimodal Nuclear Thermal Rocket Propulsion Investigated for Power-Rich, Artificial-Gravity Human Exploration Missions to Mars
Borowski, Stanley K.; McGuire Melissa L.; Dudzinski, Leonard A.; Research and Technology 2003; May 2004; 3 pp.; In English; No Copyright; Avail.: CASI: A01, Hardcopy
The NASA Glenn Research Center is involved in systems and mission analysis studies for Prometheus and the President s Vision for Space Exploration. In support of this effort, engineers at Glenn have been studying the application of nuclear thermal rocket (NTR) engines to human missions to Mars. The NTR is one of the leading propulsion options for future NASA human exploration missions because of its high specific impulse capability (I(sub sp) approx. 875 to 950 sec) and attractive engine thrust-to-weight ratio (\g3). Because only a relatively small amount of enriched uranium-235 fuel is consumed in an NTR during the primary propulsion maneuvers of a typical Mars mission, engines configured for both propulsive thrust and modest power generation (referred to as bimodal operation) provide the basis for a robust, power-rich stage with efficient propulsive capture capability. A family of modular bimodal NTR (BNTR) space-transfer-vehicle concepts has been developed that utilize a common core stage powered by three approx.15-klb(sub f) engines that produce 50 kWe of total electrical power for crew life support, high-data-rate communications with Earth, and an active refrigeration system for long-term, zero-boiloff liquid hydrogen (LH2) storage. Candidate nuclear fuels for BNTR engines include uranium carbide (UC2) particles with a chemical-vapor-deposited coating in graphite and uranium-carbide zirconium carbide (UC-ZrC) in graphite, which were developed during the Nuclear Engine for Rocket Vehicle Application (NERVA) program, as well as uranium oxide (UO2) in tungsten (W) metal cermet. These fuels, which are listed in order of increasing temperature capability, can produce hot hydrogen exhaust ranging from approx. 2550 to 2900 K. Each engine has its own closed-cycle Brayton rotating unit, capable of generating up to 25 kWe, that provides an engine-out capability. Under nominal conditions, each Brayton rotating unit would operate at two-thirds of the rated power (approx. 17 kWe). Author
Artificial Gravity; Systems Analysis; Nuclear Rocket Engines; Nuclear Propulsion; Uranium Carbides; Uranium Oxides; Nuclear Fuels
20050215635 NASA Marshall Space Flight Center, Huntsville, AL, USA
Microinstabilities in the Gasdynamic Mirror Propulsion System
Emrich, William; [2005]; 1 pp.; In English; 41st AIAA/ASME/SAE/ASEE Joint Propulsion Conference and Exhibit, 10-13 Jul. 2005, Tucson, AZ, USA; No Copyright; Avail.: Other Sources; Abstract Only
The gasdynamic mirror has been proposed as a concept which could form the basis of a highly efficient fusion rocket engine. Gasdynamic mirrors differ from most other mirror type plasma confinement schemes in that they have much larger aspect ratios and operate at somewhat higher plasma densities. There are several types of instabilities which are known to plague mirror type confinement schemes. These instabilities fall into two general classes. One class of instability is the Magnetohdrodynamic or MHD instability which induces gross distortions in the plasma geometry. The other class of instability is the ‘loss cone' microinstability which leads to general plasma turbulence. The ‘loss cone' microinstability is caused by velocity space asymmetries resulting from the loss of plasma having constituent particle velocities within the angle of the magnetic mirror ‘loss cone.' These instabilities generally manifest themselves in high temperature, moderately dense plasmas. The present study indicates that a GDM configured as a rocket engine might operate in a plasma regine where microinstabilities could potentially be significant. Author
Magnetohydrodynamic Stability; Gas Dynamics; Magnetic Mirrors; Propulsion System Configurations; Rocket Engines
20050215641 NASA Glenn Research Center, Cleveland, OH, USA
Options Studied for Managing Space Station Solar Array Electrical Hazards for Sequential Shunt Unit Replacement
Delleur, Ann M.; Kerslake, Thomas W.; Levy, Robert K.; Research and Technology 2003; May 2004; 3 pp.; In English; No Copyright; Avail.: CASI: A01, Hardcopy
The U.S. solar array strings on the International Space Station are connected to a sequential shunt unit (SSU). The job of the SSU is to shunt, or short, the excess current from the solar array, such that just enough current is provided downstream to maintain the 160-V bus voltage while meeting the power load demand and recharging the batteries. Should an SSU fail on-orbit, it would be removed and replaced with the on-orbit spare during an astronaut space walk or extravehicular activity (EVA) (see the photograph). However, removing an SSU during an orbit Sun period with input solar array power connectors fully energized could result in substantial hardware damage and/or safety risk to the EVA astronaut. The open-circuit voltage of cold solar-array strings can exceed 320 V, and warm solar-array strings could feed a short circuit with a total current level exceeding 240 A. Derived from text
Solar Arrays; Space Stations; Electric Batteries; Hazards; Failure; Management
20050215642
Probabilistic Analysis Techniques Applied to Complex Spacecraft Power System Modeling Research and Technology 2004
June 2005; 2 pp.; In English; No Copyright; Avail.: CASI: A01, Hardcopy
Electric power system performance predictions are critical to spacecraft, such as the International Space Station (ISS), to ensure that sufficient power is available to support all the spacecraft s power needs. In the case of the ISS power system, analyses to date have been deterministic, meaning that each analysis produces a single-valued result for power capability because of the complexity and large size of the model. As a result, the deterministic ISS analyses did not account for the sensitivity of the power capability to uncertainties in model input variables. Over the last 10 years, the NASA Glenn Research Center has developed advanced, computationally fast, probabilistic analysis techniques and successfully applied them to large (thousands of nodes) complex structural analysis models. These same techniques were recently applied to large, complex ISS power system models. This new application enables probabilistic power analyses that account for input uncertainties and produce results that include variations caused by these uncertainties. Specifically, N&R Engineering, under contract to NASA, integrated these advanced probabilistic techniques with Glenn s internationally recognized ISS power system model, System Power Analysis for Capability Evaluation (SPACE). Derived from text
Complex Systems; Performance Prediction; Spacecraft Power Supplies; Structural Analysis
20050216518 Air Force Research Lab., Edwards AFB, CA USA
Development of a Thrust Stand Micro-Balance to Assess Micropropulsion Performance
Pancotti, Anthony P.; Lilly, Taylor; Ketsdever, Andrew D.; Aguero, Victor; Schwoebel, Paul R.; Jun. 1, 2005; 6 pp.; In English Contract(s)/Grant(s): Proj-5026 Report No.(s): AD-A437017; No Copyright; Avail.: CASI: A02, Hardcopy
As proposed spacecraft and their associated thrusters have become smaller, technology has been developed to meet the demand for performance measurements for the extremely low force levels produced. For such thrusters, it is also desirable to measure the mass changes resulting from the utilization of propellant associated with either the steady state thrust or transient impulses. A thrust stand and novel data analysis method is presented to make both thrust (or impulse) and mass change measurements concurrently. It is shown that very accurate and repeatable measurements of mass can be made using an existing thrust stand system. Furthermore, it is shown that impulse and mass measurements can be resolved at the same time from a single thrust stand trace. DTIC
Balance; Low Thrust Propulsion; Mass; Mass Distribution; Measurement; Microbalances; Microthrust; Spacecraft Propulsion
20050217095 NASA Glenn Research Center, Cleveland, OH, USA
Forward Technology Solar Cell Experiment (FTSCE) for MISSE-5 Verified and Readied for Flight on STS-114
Jenkins, Phillip P.; Krasowski, Michael J.; Greer, Lawrence C.; Flatico, Joseph M.; Research and Technology 2004; June 2005; 3 pp.; In English; No Copyright; Avail.: CASI: A01, Hardcopy
The Forward Technology Solar Cell Experiment (FTSCE) is a space solar cell experiment built as part of the Fifth Materials on the International Space Station Experiment (MISSE-5): Data Acquisition and Control Hardware and Software. It represents a collaborative effort between the NASA Glenn Research Center, the Naval Research Laboratory, and the U.S. Naval Academy. The purpose of this experiment is to place current and future solar cell technologies on orbit where they will be characterized and validated. This is in response to recent on-orbit and ground test results that raised concerns about the in-space survivability of new solar cell technologies and about current ground test methodology. The various components of the FTSCE are assembled into a passive experiment container--a 2- by 2- by 4-in. folding metal container that will be attached by an astronaut to the outer structure of the International Space Station. Data collected by the FTSCE will be relayed to the ground through a transmitter assembled by the U.S. Naval Academy. Data-acquisition electronics and software were designed to be tolerant of the thermal and radiation effects expected on orbit. The experiment has been verified and readied for flight on STS-114. Derived from text
Solar Cells; Thermal Radiation; Computer Programs; Data Acquisition
20050217110 NASA Marshall Space Flight Center, Huntsville, AL, USA
Beamed Energy and Other Concepts for Aerospace Propulsion Applications
Cole, John W.; [2005]; 1 pp.; In English; No Copyright; Avail.: Other Sources; Abstract Only
Propulsion for aerospace applications is limited by two basic parameters: specific energy (MJ/kg) and specific power (KW/kg). Specific energy can perhaps be improved by increasing the energy content of propellants, increasing energy storage of other on-board devices, and by the use of intense off-board energy sources such as beamed energy. Several beamed energy concepts for space access have been investigated using Lasers and Microwave beams. Several preliminary concepts have been examined for high altitude platforms for commercial or military applications. Some of these results are described. Additionally, two concepts are briefly described for potentially improving on-board specific energy: Metallic Hydrogen and Magnetic Energy Storage. Author
Aerospace Engineering; Spacecraft Propulsion; Laser Power Beaming; Microwave Transmission
20050217123 NASA Marshall Space Flight Center, Huntsville, AL, USA
Fifth International Symposium on Liquid Space Propulsion
Garcia, R., Compiler; January 2005; 1167 pp.; In English; Fifth International Symposium on Liquid Space Propulsion, 27-30 Oct. 2003, Chattanooga, TN, USA; See also 20050217124 - 20050217152 Report No.(s): NASA/CP-2005-213607; M-1129; No Copyright; Avail.: CASI: A99, Hardcopy
Contents include the fiollowing: Theme: Life-life Combustion Devices Technology. Technical Sessions: International Perspectives. System Level Effects. Component Level Processes. Material Considerations. Design Environments --Predictions. Injector Design Technology. Design Environments -- Measurements. Panel Discussion: Views on future research and development needs and Symposium observations. AquariumWelcome and Southern Belle Riverboat Recognition Banquet evening events. Derived from text
Liquid Propellant Rocket Engines; Combustion; Life (Durability); Propulsion
20050217124 Boeing Co., USA
History of Rocket Propulsion at Rocketdyne
Fisher, Steve; Fifth International Symposium on Liquid Space Propulsion; January 2005; 40 pp.; In English; See also 20050217123; Original contains color and black and white illustrations; No Copyright; Avail.: CASI: A03, Hardcopy
From 0.01 LB thrust space engines to 1.8 million LB thrust F-1A booster. Nearly all propellant combinations: from exotic storables, tri-dyne, LOX/kerosene, LOX/hydrogen, peroxide/hydrocarbons, to LOX/alcohols and everything in between. Nearly all engine cycles: from pressure fed, gas generator, fuel rich staged, ‘oxidizer rich' staged, various expanders, self pressurizing, hot gas tap-off, to full flow (combined fuel and oxidizer rich) staged. Nearly all thrust chamber configurations: from conventional, dual bell, Aerospike, linear Aerospike, concentric booster/sustainer, to expansion-deflection. Derived from text
Liquid Propellant Rocket Engines; Thrust Chambers; Aerospike Engines; Propulsion; Gas Generators
20050217125 NASA Marshall Space Flight Center, Huntsville, AL, USA
Experimental Results for an Annular Aerospike with Differential Throttling
Ruf, Joseph H.; McDaniels, David M.; Fifth International Symposium on Liquid Space Propulsion; January 2005; 20 pp.; In English; See also 20050217123; Original contains color illustrations; No Copyright; Avail.: CASI: A03, Hardcopy
A) MSFC funded an internal study on Altitude Compensating Nozzles: 1) Develop an ACN design and performance prediction tool. 2) Design, build and test cold flow ACN nozzles. 3) An annular aerospike nozzle was designed and tested. 4) Incorporated differential throttling to assess Thrust Vector Control. B) Objective of the test hardware: 1) Provide design tool verification. 2) Provide benchmark data for CFD calculations. 3) Experimentally measure side force, or TVC, for a differentially throttled annular aerospike. Derived from text
Thrust Vector Control; Performance Prediction; Computational Fluid Dynamics; Annular Nozzles
20050217127 Centre National de la Recherche Scientifique, Paris, France
Investigation of Subcritical and Transcritical Cryogenic Combustion Using Imaging and Laser Techniques
Candel, S.; Juniper, M.; Scouflair, P.; Rolon, C.; Clauss, W.; Klimenko, D.; Oschwald, M.; Grisch, F.; Bouchardy, P.; Vingert, L.; Fifth International Symposium on Liquid Space Propulsion; January 2005; 26 pp.; In English; See also 20050217123; Original contains color and black and white illustrations; No Copyright; Avail.: CASI: A03, Hardcopy
The detailed understanding of liquid propellant combustion is necessary to the development of improved and more reliable propulsion systems. This article describes experimental investigations aimed at providing a fundamental basis for design and engineering of combustion components. It reports recent applications of imaging techniques and laser diagnostics to cryogenic combustion at high pressure. The flame structure is investigated in the trascritical range where the pressure exceeds the critical pressure of oxygen but the temperature of the injected liquid oxygen is below its critical value. Data are obtained from imaging of OH* radicals emission, backlighting and Coherent Anti-Stokes Raman Scattering of H2O molecules. Results from these different diagnostics provide a detailed view of the flame structure and some quantitative temperature profile measurements at various pressure. The data may be used to guide numerical modeling of transcritical flames and the theoretical and numerical analysis of the stabilization process. Results obtained may also be employed to devise engineering methodologies for component development aimed at service life extension. Author
Combustion; Cryogenics; Liquid Rocket Propellants; Propulsion System Performance; Propellant Combustion; Imaging Techniques; Mathematical Models
20050217128 Air Force Research Lab., USA
Understanding Injection Into High Pressure Supercritical Environments< p>
Chehroudi, B.; Talley, D.; Mayer, W.; Branam, R.; Smith, J. J.; Schik, A.; Oschwald, M.; Fifth International Symposium on Liquid Space Propulsion; January 2005; 27 pp.; In English; See also 20050217123; Original contains color and black and white illustrations; No Copyright; Avail.: CASI: A03, Hardcopy
The objectives of this paper are: Consolidate and compare the findings of two independent laboratories related to the injection of cryogenic liquids into high pressure supercritical environments. Derived from text
Cryogenics; Cold Flow Tests; Fracture Strength; Sound Waves
20050217130 NASA Marshall Space Flight Center, Huntsville, AL, USA
Operational Issues in the Development of a Cost-Effective Reusable LOX/LH2 Engine
Ballard, Richard O.; Fifth International Symposium on Liquid Space Propulsion; January 2005; 20 pp.; In English; See also 20050217123; Original contains color illustrations; No Copyright; Avail.: CASI: A03, Hardcopy
Contents include the following: SLI initiated under NASA Research Announcement (NRA) 8-30. Strategic Objectives. Make spaceflight safer (1 in 10000 mission LOV). Make spaceflight cheaper ($1000/lb payload). Two prototype LOX/LH2 engine systems funded under Cycle-1 of NRA8-30. COBRA (Pratt & Whitney / Aerojet). RS-83 (Rocketdyne). Derived from text
Reusable Launch Vehicles; Cost Effectiveness; Space Flight; Propulsion System Performance
20050217131 Japan Aerospace Exploration Agency, Japan
Research Achievement on Long-Life Thrust Chambers at NAL/KPL
Kumakawa, A.; Moriya, S.; Sato, M.; Fifth International Symposium on Liquid Space Propulsion; January 2005; 51 pp.; In English; See also 20050217123; Original contains color and black and white illustrations; No Copyright; Avail.: CASI: A04, Hardcopy
Since its establishment of in 1965, NAL ( National Aerospace Laboratory) / KPL ( Kakuda Space Propulsion Laboratory) has contributed to the R&D of Japanese launch vehicles, especially with respect to propulsion. It has been involved in the R&D of rocket engines such as the LE-5 engine and the LE-7 engine. These engines have been respectively used for the H-1 and H-2 launch vehicles. NAL/KPL has been merged and renamed as Space Propulsion Research Center of JAXA ( Japan Aerospace Exploration Agency) this October. Derived from text
Thrust Chambers; Rocket Engines; Propulsion
20050217132 NASA Marshall Space Flight Center, Huntsville, AL, USA
Combustion Device Failures During Space Shuttle Main Engine Development
Goetz, Otto K.; Monk, Jan C.; Fifth International Symposium on Liquid Space Propulsion; January 2005; 59 pp.; In English; See also 20050217123; Original contains color and black and white illustrations; No Copyright; Avail.: CASI: A04, Hardcopy
Major Causes: Limited Initial Materials Properties. Limited Structural Models - especially fatigue. Limited Thermal Models. Limited Aerodynamic Models. Human Errors. Limited Component Test. High Pressure. Complicated Control. Derived from text
Combustion; Engine Design; Failure; Aerodynamic Characteristics
20050217133 NASA Marshall Space Flight Center, Huntsville, AL, USA
Transient Mathematical Modeling for Liquid Rocket Engine Systems: Methods, Capabilities, and Experience
Seymour, David C.; Martin, Michael A.; Nguyen, Huy H.; Greene, William D.; Fifth International Symposium on Liquid Space Propulsion; January 2005; 36 pp.; In English; See also 20050217123; Original contains color illustrations; No Copyright; Avail.: CASI: A03, Hardcopy
The subject of mathematical modeling of the transient operation of liquid rocket engines is presented in overview form from the perspective of engineers working at the NASA Marshall Space Flight Center. The necessity of creating and utilizing accurate mathematical models as part of liquid rocket engine development process has become well established and is likely to increase in importance in the future. The issues of design considerations for transient operation, development testing, and failure scenario simulation are discussed. An overview of the derivation of the basic governing equations is presented along with a discussion of computational and numerical issues associated with the implementation of these equations in computer codes. Also, work in the field of generating usable fluid property tables is presented along with an overview of efforts to be undertaken in the future to improve the tools use for the mathematical modeling process. Author
Mathematical Models; Liquid Propellant Rocket Engines; Computer Programs
20050217134 Office National d'Etudes et de Recherches Aerospatiales, Paris, France
CFD Code Validation for Space Propulsion Applications
Vuillot, F.; Scherrer, D.; Habiballah, M.; Fifth International Symposium on Liquid Space Propulsion; January 2005; 55 pp.; In English; See also 20050217123; Original contains color and black and white illustrations; No Copyright;Avail.: CASI: A04, Hardcopy
The presentation covers: 1) SPACE PROPULSION (launcher applications): a. Present: Liquid and solid rocket propulsion, b. Future: Hypersonic airbreathing propulsion. 2) COMPETITIVE MARKET: a. Costs and delays b. Usefulness of numerical simulations: c. Selecting appropriate designs, d. Validating specific solutions. Helping design test benches and instrumentation/measurement plans f. Analyzing static firings, hot testings and flight data. 3) IMPORTANCE of CODE VALIDATION: a. Accuracy b. Reliability. Derived from text
Computational Fluid Dynamics; Reliability; Launchers; Propulsion; Air Breathing Engines
20050217137 Army Research Lab., Cleveland, OH, USA
Damage Assessment of Combustion Devices
Bonacuse, Peter J.; Fifth International Symposium on Liquid Space Propulsion; January 2005; 26 pp.; In English; See also 20050217123; Original contains color illustrations; No Copyright; Avail.: CASI: A03, Hardcopy
Complex interactions among damage mechanisms are possible depending on material, temperature range, loading conditions, and environment. Extrapolations from limited data can lead to unconservative and inaccurate life prediction of the combustor devices. Conditions representative of the service environment in the durability testing of materials and components are highly desirable. The accuracy of all life prediction models is highly dependent on the accuracy of the model inputs (temperatures, stresses/strains, material properties). It is important to identify the mechanisms causing damage in the combustor device and utilize the appropriate models to estimate the durability of the component in a reliable manner. There is always a distribution in the number of cycles or time to failure in materials subjected to repeated loading. Multiple experiments are necessary to quantify this variability, especially at longer expected lives (typically result in higher variability in life). To minimize damage in combustion devices: reduce thermal gradients, lower peak temperatures, reduce the effect of environment, and reduce the severity and number of thermal excursions. Author
Combustion Chambers; Damage Assessment; Temperature Gradients; Prediction Analysis Techniques; Performance Tests; Life (Durability); Accuracy; Degradation; Failure
20050217139 Seoul National Univ., Korea, Republic of
Subscale Testing and Design Issues of Injectors
Yoon, Youngbin; Jeung, In-Seuck; Fifth International Symposium on Liquid Space Propulsion; January 2005; 11 pp.; In English; See also 20050217123; Original contains color illustrations Contract(s)/Grant(s): M1-0104-00-0058; No Copyright; Avail.: CASI: A03, Hardcopy
The breakup characteristics of liquid sheets formed by like-doublet injector were investigated in the cold-flow and atmospheric ambient pressure condition. The sheet breakup wavelength, which induces the sheet to be broken into ligaments, as well as the sheet breakup length, which is important for the flame location, was measured using a stroboscopic light. Since these spray characteristics are affected by the flow characteristics of two liquid jets before they impinge on each other, we focused on the effects of orifice internal flow such as the cavitation phenomenon that occurs inside the sharp-edged orifice. From the experimental results, we found that the liquid jet turbulence delays the sheet breakup and makes shorter wavelengths of sheet. Since the turbulent strength of sharp-edged orifice is stronger than that of round-edged orifice, the shape of orifice entrance results in large differences in the spray characteristics. Using these results, we proposed empirical models on the spray characteristics of the like-doublet injector, and these models are believed to provide some useful and actual data for designing liquid rocket combustors. Author
Flow Characteristics; Fluid Jets; Injectors; Internal Flow; Orifice Flow; Spray Characteristics; Cavitation Flow; Combustion Chambers
20050217141 Korea Aerospace Research Inst., Daejeon, Korea, Republic of
Development of a Liquid Propellant Rocket, Korea Sounding Rocket (KSR)-III
Chae,Yeon Seok; Lee, Soo Yong; Seo, Seonghyeon; Fifth International Symposium on Liquid Space Propulsion; January 2005; 35 pp.; In English; See also 20050217123; Original contains color illustrations; No Copyright; Avail.: CASI: A03, Hardcopy
Korea Sounding Rocket-III has established itself as the first-ever bi-propellant liquid rocket designed and manufactured in Korea, which is fueled by pressure-fed, liquid oxygen and environment-friendly kerosene. The project on KSR-III took its development period of five years to achieve its successful launch campaign on November 28, 2002. The main objectives of the development of KSR-III were to acquire prerequisite experiences in the development of satellite launchers and to further strengthen capability of domestic industries related to space technology. The program has been accomplished by going through many development stages such as subscale model and full-scale engine combustion tests, propellant feeding system tests, integrated power plant tests, and eventually a flight test. Each development stage underwent various technical difficulties and challenges that had to be resolved and verified with reliable processes to minimize trials and errors. The subscale model combustion test provided useful data about the combustion efficiency of a split-triplet (F-OO-F) injector and suggested the best injector arrangement. Two types of subscale thrust chambers allowed us to determine the design of an injector and a faceplate for the full scale thrust chamber. The on-ground combustion tests of a full-scale thrust chamber had verified its performance and the endurance of hardware including ablative material. The thrust and chamber pressure of the engine are 13tonf @ sea level and 13.6 atm, respectively, which lasts for 59 seconds. A composite baffle successfully suppressed troublesome high frequency combustion instability at the beginning of the development. Throughout the program, three different versions of full-scale chambers had been designed and manufactured for the optimization of performance and weight. The propellant feeding system was tested and verified through water flow tests and each component showed acceptable performance for flight use. Eventually, the thrust chamber assembly and the propellant feeding system were vertically integrated to check the functionality of subsystems and verify the performance of the whole propulsion system and thus, a stage qualification test had been conducted before a final flight test. At long last, the series of all these tests resulted in the successful flight test of KSR-III. Author
Liquid Rocket Propellants; Combustion Effýciency; Combustion Stability; Engine Tests; Performance Tests; Propellant Tests
20050217142 Northrop Grumman Space Technology, United States
Lifetime Issues for In-Space Propulsion Systems
Dressler, Gordon; Fifth International Symposium on Liquid Space Propulsion; January 2005; 30 pp.; In English; See also 20050217123; Original contains color illustrations; No Copyright; Avail.: CASI: A03, Hardcopy
The Orbital Express program will be a key flight demonstration that spacecraft can attain many additional years of operational life by in-space refueling of propellants. A major improvement in the mission life capability of current radiation-cooled ACS/RCS thrusters can be obtained by elimination of sensitive anti-oxidation coatings through use of: regeneratively-cooled combustion chambers, or noble refractory metal combustion chambers. Derived from text
Combustion Chambers; Propellants; Propulsion System Performance; Regenerative Cooling; Service Life
20050217143 Purdue Univ., USA
Injectors for Long Life Combustion Devices
Bazarov, Vladimir; Fifth International Symposium on Liquid Space Propulsion; January 2005; 34 pp.; In English; See also 20050217123; Original contains color illustrations; No Copyright; Avail.: CASI: A03, Hardcopy
Life-limiting factors in LRE combustion devices: 1) High local heat fluxes, caused by high pressure, gas density and high velocity of combustion products; 2) Concentration non-uniformity and local zones of oxidizer-rich gases; 3) High level and wide spectrum of mechanical vibrations; 4) Combustion instability; 5) Combustion unsteadiness, accompanied by high amplitude and wide frequency range of gas pressure, velocity and vorticity pulsation; and 6) Most of them depend on injector design and operation. Derived from text
Injectors; Heat Flux; Gas Pressure; Combustion Stability; Combustion Products; Compressed Gas
20050217144 Air Force Research Lab., USA
Gas-Centered Swirl Coaxial Liquid Injector Evaluations
Cohn, A. K.; Strakey, P. A.; Talley, D. G.; Fifth International Symposium on Liquid Space Propulsion; January 2005; 39 pp.;
In English; See also 20050217123; Original contains color illustrations; No Copyright; Avail.: CASI: A03, Hardcopy Development of Liquid Rocket Engines is expensive. Extensive testing at large scales usually required. In order to verify engine lifetime, large number of tests required. Limited Resources available for development. Sub-scale cold-flow and hot-fire testing is extremely cost effective. Could be a necessary (but not sufficient) condition for long engine lifetime. Reduces overall costs and risk of large scale testing. Goal: Determine knowledge that can be gained from sub-scale cold-flow and hot-fire evaluations of LRE injectors. Determine relationships between cold-flow and hot-fire data. Derived from text
Liquid Propellant Rocket Engines; Performance Tests; Life (Durability)
20050217145 European Aeronautic Defence and Space Co., Munich, Germany
Injector Issues With Different Propellant Combinations
Haeseler, D.; Maedig, C.; Rubinski, V.; Kosmatcheva, V.; Berezhnoy, V.; Bratukhin, N.; Fifth International Symposium on Liquid Space Propulsion; January 2005; 51 pp.; In English; See also 20050217123; Original contains color illustrations Contract(s)/Grant(s): DLR Proj. 50JR9911; No Copyright; Avail.: CASI: A04, Hardcopy
Contents include the folowing: Experimental characterization of operation, efficiency, heat flux of: Injection element concepts for wide throttling range. Injection element concepts for large flowrate. Injection element concepts for hydrocarbon fuels (LOX-Methane and LOX-Kerosene). Activities in cooperation with KBKhA/Russia since 1993. Derived from text
Injectors; Propellants; Throttling; Flow Velocity
20050217146 Nauchno-Proizvodstvennoe Obedinenie Energomash, USSR
LOX Kerosene Oxygen-Rich Staged Combustion Rocket Engine Design and Life Issues
Chelkis, Feliks; Fifth International Symposium on Liquid Space Propulsion; January 2005; 33 pp.; In English; See also 20050217123; Original contains color illustrations; No Copyright; Avail.: CASI: A03, Hardcopy
The principal decisions with respect to the design of RD-170 (RD-171) were based on the experience in development of engines with oxygen rich staged combustion (ORSC). Development of engines of such type was initiated in the early sixties of the last century. For the past period of time newly developed engines have been made mainly with ORSC cycle. This period of time was the most important phase of the history of rocket engineering development in the Soviet Union. A major part of the production and test infrastructure that provided capabilities to develop engines with a high level of combustion chamber pressure was established then. In the same period of time, based on experience in engine development and operations, the modern scientific-engineering methods and standard documentation base now in use were generated. As of today, some ORSC engines have a very high performance both in Russia and in the whole world. Derived from text
Combustion Chambers; Engine Design; Liquid Oxygen; Rocket Engine Design
20050217147 Institut Superieur des Materiaux et de la Construction Mecanique, France
Material Requirements and Limitations for Liquid Rocket Engines
Inglebert, G.; Gras, R.; Fifth International Symposium on Liquid Space Propulsion; January 2005; 44 pp.; In English; See also 20050217123; Original contains color illustrations; No Copyright; Avail.: CASI: A03, Hardcopy
An outline includes: 1) Spatial systems work in severe environment Vacuum, cryogenic fluids, hot gases; 2) Key parameters for design: mechanical links, guiding, power transmission systems, sealing; 3) Engines for launchers: Turbopumps to compress LH2 and LO(x) \g 100 bars Combustion chamber and nozzle yields specific thrust; 4) Deep knowledge of materials Physical properties in the loading range (thermal and mechanical); 5) Accurate enough behavior models. Derived from text
Liquid Propellant Rocket Engines; Cryogenic Fluids; Combustion Chambers; Specific Impulse
20050217149 Pratt and Whitney Space Propulsion, West Palm Beach, FL, USA
Cooling and Life Issues for Long Life Combustion Chambers
Popp, Mike; Fifth International Symposium on Liquid Space Propulsion; January 2005; 26 pp.; In English; See also 20050217123; Original contains color illustrations; No Copyright; Avail.: CASI: A03, Hardcopy
Conclusions: 1) Long Combustion Chamber Life is enabled by reducing Chamber Liner Temperatures and Chemical Attack; 2) Injector-originated Film Cooling is widely used; 3) Chamber Wall-originated Film Cooling and Thermal Barrier Coatings are used successfully in RD-170/180 Flight Engines; 4) Transpiration Cooling has been successfully demonstrated; further testing and manufacturing process development needed 5) Thermal Barrier Coatings must demonstrate Fail/Safe Behavior; 6) Combination of Thermal Barrier Coating, Film Cooling, Transpiration Cooling (TBD) enables Long Life Chamber Construction. PERFORMANCE must be traded for LIFE: Long Life Design results in reduced Isp and increased mass, however, penalties are relatively small. Derived from text
Combustion Chambers; Cooling; Film Cooling; Injectors; Sweat Cooling; Thermal Control Coatings; Failure
20050217150 Office National d'Etudes et de Recherches Aerospatiales, Paris, France, Deutsche Forschungsanstalt fuer Luftund Raumfahrt, Lampoldshausen, Germany
Experimental Investigation and Modeling of the Ignition Transient of a Coaxial H2/O2-Injector
Oschwald, M.; Schmidt, V.; Klimenko, D.; Haidn, O.; Nicole, A.; Ordonneau, G.; Habiballah, M.; Fifth International Symposium on Liquid Space Propulsion; January 2005; 56 pp.; In English; See also 20050217123; Original contains color and black and white illustrations; No Copyright; Avail.: CASI: A04, Hardcopy
A test case have been defined to investigate experimentally and by numerical simulation the transient ignition phenomenology when igniting coaxial injected O2 and H2 by a laser. Using high-speed photography the temporal evolution of the flame and its anchoring at the injector could be visualized. From the analysis of the flame front movement flame velocities and convection velocities have been determined at specific phases during the ignition transient. Author
Experimentation; Models; Ignition; Oxygen; Hydrogen; Injectors; Numerical Analysis
20050217151 NASA Marshall Space Flight Center, Huntsville, AL, USA
Future Directions for Space Transportation and Propulsion at NASA
Sackheim, Robert L.; Fifth International Symposium on Liquid Space Propulsion; January 2005; 52 pp.; In English; See also 20050217123; Original contains black and white illustrations; No Copyright; Avail.: CASI: A04, Hardcopy
Contents include the following: Oxygen Compatible Materials. Manufacturing Technology Demonstrations. Turbopump Inducer Waterflow Test. Turbine Damping ‘Whirligig' Test. Single Element Preburner and Main Injector Test. 40K Multi-Element Preburner and MI. Full-Scale Battleship Preburner. Prototype Preburner Test Article. Full-Scale Prototype TCA. Turbopump Hot-Fire Test Article. Prototype Engine. Validated Analytical Models. Derived from text
Propulsion System Configurations; Technology Assessment; Preburners; Prototypes; Turbine Pumps
20050217196 NASA Glenn Research Center, Cleveland, OH, USA
Cooled Ceramic Matrix Composite Propulsion Structures Demonstrated
Jaskowiak, Martha H.; Dickens, Kevin W.; Research and Technology 2004; June 1, 2005; 4 pp.; In English; No Copyright; Avail.: CASI: A01, Hardcopy
NASA's Next Generation Launch Technology (NGLT) Program has successfully demonstrated cooled ceramic matrix composite (CMC) technology in a scramjet engine test. This demonstration represented the world s largest cooled nonmetallic matrix composite panel fabricated for a scramjet engine and the first cooled nonmetallic composite to be tested in a scramjet facility. Lightweight, high-temperature, actively cooled structures have been identified as a key technology for enabling reliable and low-cost space access. Tradeoff studies have shown this to be the case for a variety of launch platforms, including rockets and hypersonic cruise vehicles. Actively cooled carbon and CMC structures may meet high-performance goals at significantly lower weight, while improving safety by operating with a higher margin between the design temperature and material upper-use temperature. Studies have shown that using actively cooled CMCs can reduce the weight of the cooled flow-path component from 4.5 to 1.6 lb/sq ft and the weight of the propulsion system s cooled surface area by more than 50 percent. This weight savings enables advanced concepts, increased payload, and increased range. The ability of the cooled CMC flow-path components to operate over 1000 F hotter than the state-of-the-art metallic concept adds system design flexibility to space-access vehicle concepts. Other potential system-level benefits include smaller fuel pumps, lower part count, lower cost, and increased operating margin. Derived from text
Cooling Systems; Ceramic Matrix Composites; Propulsion; Composite Structures
20050217201 ZIN Technologies, Inc., Brook Park, OH, USA
Advanced Controller Developed for the Free-Piston Stirling Convertor
Gerber, Scott S.; Research and Technology 2004; June 2005; 3 pp.; In English; No Copyright; Avail.: CASI: A01, Hardcopy
A free-piston Stirling power convertor is being considered as an advanced power-conversion technology for future NASA deep-space missions requiring long-life radioisotope power systems. The NASA Glenn Research Center has identified key areas where advanced technologies can enhance the capability of Stirling energy-conversion systems. One of these is power electronic controls. Current power-conversion technology for Glenn-tested Stirling systems consists of an engine-driven linear alternator generating an alternating-current voltage controlled by a tuning-capacitor-based alternating-current peak voltage load controller. The tuning capacitor keeps the internal alternator electromotive force (EMF) in phase with its respective current (i.e., passive power factor correction). The alternator EMF is related to the piston velocity, which must be kept in phase with the alternator current in order to achieve stable operation. This tuning capacitor, which adds volume and mass to the overall Stirling convertor, can be eliminated if the controller can actively drive the magnitude and phase of the alternator current. Derived from text
Converters; Piston Engines; Stirling Cycle; Electromotive Forces; Energy Conversion
Source: NASA.
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