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Search indexed NASA NTRS and DOE OSTI research on propulsion, heat transfer, battery materials and energy systems. Follow report and document links to the original sources.

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At least 505 records · Page 28
All sourcesNASA NTRSDOE OSTI
NASA NTRS1980

Chemical propulsion technology

An overview of NASA's low thrust liquid chemical propulsion program is presented with particular emphasis on thrust system technology in the ten to one thousand pound thrust range. Key technology issues include high performance of cooled low thrust engines; small cryogenic pumps; multiple starts-shutdowns (10) with slow ramps (approximately 10 seconds); thrust variation - 4/1 in flight and 20/1 between flights; long life (100 hours); improved system weight and size; and propellant selection.

Priem, R. J.↗
NASA NTRS2016

Foundational Methane Propulsion Related Technology Efforts, and Challenges for Applications to Human Exploration Beyond Earth Orbit

Current interest in human exploration beyond earth orbit is driving requirements for high performance, long duration space transportation capabilities. Continued advancement in photovoltaic power systems and investments in high performance electric propulsion promise to enable solar electric options for cargo delivery and pre-deployment of operational architecture elements. However, higher thrust options are required for human in-space transportation as well as planetary descent and ascent functions. While high thrust requirements for interplanetary transportation may be provided by chemical or nuclear thermal propulsion systems, planetary descent and ascent systems are limited to chemical solutions due to their higher thrust to weight and potential planetary protection concerns. Liquid hydrogen fueled systems provide high specific impulse, but pose challenges due to low propellant density and the thermal issues of long term propellant storage. Liquid methane fueled propulsion is a promising compromise with lower specific impulse, higher bulk propellant density and compatibility with proposed in-situ propellant production concepts. Additionally, some architecture studies have identified the potential for commonality between interplanetary and descent/ascent propulsion solutions using liquid methane (LCH4) and liquid oxygen (LOX) propellants. These commonalities may lead to reduced overall development costs and more affordable exploration architectures. With this increased interest, it is critical to understand the current state of LOX/LCH4 propulsion technology and the remaining challenges to its application to beyond earth orbit human exploration. This paper provides a survey of NASA's past and current methane propulsion related technology efforts, assesses the accomplishments to date, and examines the remaining risks associated with full scale development.

Brown, Thomas↗
NASA NTRS1997

Preliminary Design of Low-Thrust Interplanetary Missions

For interplanetary missions, highly efficient electric propulsion systems can be used to increase the mass delivered to the destination and/or reduce the trip time over typical chemical propulsion systems. This technology is being demonstrated on the Deep Space 1 mission - part of NASA's New Millennium Program validating technologies which can lower the cost and risk and enhance the performance of future missions. With the successful demonstration on Deep Space 1, future missions can consider electric propulsion as a viable propulsion option. Electric propulsion systems, while highly efficient, produce only a small amount of thrust. As a result, the engines operate during a significant fraction of the trajectory. This characteristic makes it much more difficult to find optimal trajectories. The methods for optimizing low-thrust trajectories are typically categorized as either indirect, or direct. Indirect methods are based on calculus of variations, resulting in a two-point boundary value problem that is solved by satisfying terminal constraints and targeting conditions. These methods are subject to extreme sensitivity to the initial guess of the variables - some of which are not physically intuitive. Adding a gravity assist to the trajectory compounds the sensitivity. Direct methods parameterize the problem and use nonlinear programming techniques to optimize an objective function by adjusting a set of variables. A variety of methods of this type have been examined with varying results. These methods are subject to the limitations of the nonlinear programming techniques. In this paper we present a direct method intended to be used primarily for preliminary design of low-thrust interplanetary trajectories, including those with multiple gravity assists. Preliminary design implies a willingness to accept limited accuracy to achieve an efficient algorithm that executes quickly.

Sims, Jon A.↗
NASA NTRS1978

Two dimensional wedge/translating shroud nozzle

A jet propulsion exhaust nozzle is reported for multi-engine installations which produces high internal/external, thrust-minus-drag, performance for transonic cruise or transonic acceleration as well as improved performance at subsonic and supersonic speeds. A two dimensional wedge/translating shroud provides the variable nozzle exit geometry needed to achieve high engine performance over a wide range of throttle power settings.

Maiden, D. L.↗
NASA NTRS2002

The FRC Acceleration Space Thruster (FAST) Experiment

The objective of the FRC (Field Reversed Configuration) Acceleration Space Thruster (FAST) Experiment is to investigate the use of a repetitive FRC source as a thruster, specifically for an NEP (nuclear electric propulsion) system. The Field Reversed Configuration is a plasmoid with a closed poloidal field line structure, and has been extensively studied as a fusion reactor core. An FRC thruster works by repetitively producing FRCs and accelerating them to high velocity. An FRC thruster should be capable of I(sub sp)'s in the range of 5,000 - 25,000 seconds and efficiencies in the range of 60 - 80 %. In addition, they can have thrust densities as high as 10(exp 6) N/m2, and as they are inductively formed, they do not suffer from electrode erosion. The jet-power should be scalable from the low to the high power regime. The FAST experiment consists of a theta-pinch formation chamber, followed by an acceleration stage. Initially, we will produce and accelerate single FRCs. The initial focus of the experiment will be on the ionization, formation and acceleration of a single plasmoid, so as to determine the likely efficiency and I(sub sp). Subsequently, we will modify the device for repetitive burst-mode operation (5-10 shots). A variety of diagnostics are or will be available for this work, including a HeNe interferometer, high-speed cameras, and a Thomson-scattering system. The status of the experiment will be described.

Martin, Adam↗
NASA NTRS1985

An investigation at static conditions of nonaxisymmetric nozzle thrust reverser port geometry including effects of rounding and rotating the port corner

An investigation has been conducted in the static test facility of the Langley 16-Foot Transonic Tunnel to determine the effects of rounding and rotating the port corner of a nonaxisymmetric thrust reversing port. High pressure air was used to simulate jet exhaust at nozzle pressure ratios up to 6.0. The investigation indicates that using a rounded corner in the reverser port passage increases the discharge coefficient significantly compared to that of a sharp cornered thrust reversing port. Incorporating a rotating cylinder into the rounded corner reverser port passage provides no additional increase at cylinder revolutions per minute up to 10,000. (Revolutions per minute of 390,000 were necessary to achieve a surface velocity equal to the duct velocity.) Highest levels of discharge coefficient are obtained by providing a large, round, cylindrical surface in the port upstream wall. Locating the reverser port in a constant area duct results in higher discharge coefficients than locating the port in a convergent duct.

Arbiter, D. G.↗
NASA NTRS2017

Near Earth Asteroid Scout: NASA's Solar Sail Mission to a NEA

NASA is developing a solar sail propulsion system for use on the Near Earth Asteroid (NEA) Scout reconnaissance mission and laying the groundwork for their use in future deep space science and exploration missions. Solar sails use sunlight to propel vehicles through space by reflecting solar photons from a large, mirror-like sail made of a lightweight, highly reflective material. This continuous photon pressure provides propellant-less thrust, allowing for very high delta V maneuvers on long-duration, deep space exploration. Since reflected light produces thrust, solar sails require no onboard propellant. The Near Earth Asteroid (NEA) Scout mission, funded by NASA’s Advanced Exploration Systems Program and managed by NASA MSFC, will use the sail as primary propulsion allowing it to survey and image Asteroid 1991VG and, potentially, other NEA’s of interest for possible future human exploration. The NEA Scout spacecraft is housed in a 6U CubeSat-form factor and utilizes an 86 square meter solar sail for a total mass less than 14 kilograms. The mission is in partnership with the Jet Propulsion Laboratory with support from Langley Research Center and science participants from various institutions. NEA Scout will be launched on the maiden flight of the Space Launch System in 2019. The solar sail for NEA Scout will be based on the technology developed and flown by the NASA NanoSail-D and flown on The Planetary Society’s Lightsail-A. Four approximately-7-meter stainless steel booms wrapped on two spools (two overlapping booms per spool) will be motor driven and pull the sail from its stowed volume. The sail material is an aluminized polyimide approximately 2.5 microns thick. As the technology matures, solar sails will increasingly be used to enable science and exploration missions that are currently impossible or prohibitively expensive using traditional chemical and electric propulsion systems. This paper will summarize the status of the NEA Scout mission and solar sail technology in general.

Near Earth Asteroid↗
NASA NTRS2001

Solar-Thermal Engine Testing

A solar-thermal engine serves as a high-temperature solar-radiation absorber, heat exchanger, and rocket nozzle. collecting concentrated solar radiation into an absorber cavity and transferring this energy to a propellant as heat. Propellant gas can be heated to temperatures approaching 4,500 F and expanded in a rocket nozzle, creating low thrust with a high specific impulse (I(sub sp)). The Shooting Star Experiment (SSE) solar-thermal engine is made of 100 percent chemical vapor deposited (CVD) rhenium. The engine 'module' consists of an engine assembly, propellant feedline, engine support structure, thermal insulation, and instrumentation. Engine thermal performance tests consist of a series of high-temperature thermal cycles intended to characterize the propulsive performance of the engines and the thermal effectiveness of the engine support structure and insulation system. A silicone-carbide electrical resistance heater, placed inside the inner shell, substitutes for solar radiation and heats the engine. Although the preferred propellant is hydrogen, the propellant used in these tests is gaseous nitrogen. Because rhenium oxidizes at elevated temperatures, the tests are performed in a vacuum chamber. Test data will include transient and steady state temperatures on selected engine surfaces, propellant pressures and flow rates, and engine thrust levels. The engine propellant-feed system is designed to Supply GN2 to the engine at a constant inlet pressure of 60 psia, producing a near-constant thrust of 1.0 lb. Gaseous hydrogen will be used in subsequent tests. The propellant flow rate decreases with increasing propellant temperature, while maintaining constant thrust, increasing engine I(sub sp). In conjunction with analytical models of the heat exchanger, the temperature data will provide insight into the effectiveness of the insulation system, the structural support system, and the overall engine performance. These tests also provide experience on operational aspects of the engine and associated subsystems, and will include independent variation of both steady slate heat-exchanger temperature prior to thrust operation and nitrogen inlet pressure (flow rate) during thrust operation. Although the Shooting Star engines were designed as thermal-storage engines to accommodate mission parameters, they are fully capable of operating as scalable, direct-gain engines. Tests are conducted in both operational modes. Engine thrust and propellant flow rate will be measured and thereby I(sub sp). The objective of these tests is to investigate the effectiveness of the solar engine as a heat exchanger and a rocket. Of particular interest is the effectiveness of the support structure as a thermal insulator, the integrity of both the insulation system and the insulation containment system, the overall temperature distribution throughout the engine module, and the thermal power required to sustain steady state fluid temperatures at various flow rates.

Tucker, Stephen↗
NASA NTRS1977

Static, noise, and transition tests of a combined-surface-blowing V/STOL lift/propulsion system

Efficient thrust vectoring and high levels of circulatory lift were obtained in tests of a half model V/STOL airplane by using a type of externally blown jet flap in which the jet exhaust from wing-mounted cruise fans is directed over both upper and lower surfaces of a flapped wing. Approximately 90% thrust recovery with 87 deg of thrust vectoring was achieved under static conditions using 89 deg of trailing edge flap deflection. The approximately 10% loss appears to be associated primarily with pressure losses due to the flap brackets or slot entries. The jet induced lift was shown to be 55% of the theoretical value for a fullspan jet-flapped wing, even though only 27.5% of the wingspan was immersed in the jet. Steady rate of descent capability in excess of 1,000 feet per minute is predicted. The possibility of significant aerodynamic-noise cancelling when blowing over both surfaces at high velocities is indicated.

Schoen, A. H.↗
NASA NTRS1990

Impact of emerging technologies on future combat aircraft agility

The foreseeable character of future within-visual-range air combat entails a degree of agility which calls for the integration of high-alpha aerodynamics, thrust vectoring, intimate pilot/vehicle interfaces, and advanced weapons/avionics suites, in prospective configurations. The primary technology-development programs currently contributing to these goals are presently discussed; they encompass the F-15 Short Takeoff and Landing/Maneuver Technology Demonstrator Program, the Enhanced Fighter Maneuverability Program, the High Angle-of-Attack Technology Program, and the X-29 Technology Demonstrator Program.

Nguyen, Luat T.↗
NASA NTRS2007

High-Heat-Flux Cyclic Durability of Thermal and Environmental Barrier Coatings

Advanced ceramic thermal and environmental barrier coatings will play an increasingly important role in future gas turbine engines because of their ability to protect the engine components and further raise engine temperatures. For the supersonic vehicles currently envisioned in the NASA fundamental aeronautics program, advanced gas turbine engines will be used to provide high power density thrust during the extended supersonic flight of the aircraft, while meeting stringent low emission requirements. Advanced ceramic coating systems are critical to the performance, life and durability of the hot-section components of the engine systems. In this work, the laser and burner rig based high-heat-flux testing approaches were developed to investigate the coating cyclic response and failure mechanisms under simulated supersonic long-duration cruise mission. The accelerated coating cracking and delamination mechanism under the engine high-heat-flux, and extended supersonic cruise time conditions will be addressed. A coating life prediction framework may be realized by examining the crack initiation and propagation in conjunction with environmental degradation under high-heat-flux test conditions.

Zhu, Dongming↗
NASA NTRS2011

High Voltage EEE Parts for EMA/EHA Applications on Manned Launch Vehicles

The objective of this paper is an assessment of high voltage electronic components required for high horsepower electric thrust vector control (TVC) systems for human spaceflight launch critical application. The scope consists of creating of a database of available Grade 1 electrical, electronic and electromechanical (EEE) parts suited to this application, a qualification path for potential non-Grade 1 EEE parts that could be used in these designs, and pathfinder testing to validate aspects of the proposed qualification plan. Advances in the state of the art in high power electric power systems enable high horsepower electric actuators, such as the electromechnical actuator (EMA) and the electro-hydrostatic actuator (EHA), to be used in launch vehicle TVC systems, dramaticly reducing weight, complexity and operating costs. Designs typically use high voltage insulated gate bipolar transistors (HV-IGBT). However, no Grade 1 HV-IGBT exists and it is unlikely that market factors alone will produce such high quality parts. Furthermore, the perception of risk, the lack of qualification methodoloy, the absence of manned space flight heritage and other barriers impede the adoption of commercial grade parts onto the critical path. The method of approach is to identify high voltage electronic component types and key parameters for parts currently used in high horsepower EMA/EHA applications, to search for higher quality substitutes and custom manufacturers, to create a database for these parts, and then to explore ways to qualify these parts for use in human spaceflight launch critical application, including grossly derating and possibly treating hybrid parts as modules. This effort is ongoing, but results thus far include identification of over 60 HV-IGBT from four manufacturers, including some with a high reliability process flow. Voltage ranges for HV-IGBT have been identified, as has screening tests used to characterize HV-IGBT. BSI BS ISO 21350 Space systems Off-the-shelf item utilization, developed from Marshall Work Instruction MWI8060.1 OFF-THE-SHELF HARDWARE UTILIZATION IN FLIGHT HARDWARE DEVELOPMENTwas found to provide guidance for including commercial off-the-shelf (COTS) hardware for use in critical applications.

Griffin, Trent↗
NASA NTRS1978

Mercurian global tectonics - A consequence of tidal despinning

The surface of Mercury exhibits a global tectonic system consisting of an ancient set of NE and NW tending lineaments and a younger set of planimetrically arcuate escarpments interpreted as thrust of high-angle reverse faults. The trends, distribution, and age relations of these tectonic features can be explained by a combination of tidal despinning and global contraction of the planet. In our model, early tidal despinning resulted in conjugate shear fractures trending roughly N60 deg E and N60 deg W which were subsequently modified by a variety of surface processes to produce the presently visible set of lineaments. Continued despinning plus global contraction produced thrust faults with roughly north-south trends. Final contraction may have postdated despinning and produced randomly oriented thrust faults. All of these events predated the formation of Caloris basin, because basin-associated deposits blanket both lineaments and arcuate thrust faults.

Melosh, H. J.↗
NASA NTRS2000

Magnetic Flux Compression Concept for Aerospace Propulsion and Power

The objective of this research is to investigate system level performance and design issues associated with magnetic flux compression devices for aerospace power generation and propulsion. The proposed concept incorporates the principles of magnetic flux compression for direct conversion of nuclear/chemical detonation energy into electrical power. Specifically a magnetic field is compressed between an expanding detonation driven diamagnetic plasma and a stator structure formed from a high temperature superconductor (HTSC). The expanding plasma cloud is entirely confined by the compressed magnetic field at the expense of internal kinetic energy. Electrical power is inductively extracted, and the detonation products are collimated and expelled through a magnetic nozzle. The long-term development of this highly integrated generator/propulsion system opens up revolutionary NASA Mission scenarios for future interplanetary and interstellar spacecraft. The unique features of this concept with respect to future space travel opportunities are as follows: ability to implement high energy density chemical detonations or ICF microfusion bursts as the impulsive diamagnetic plasma source; high power density system characteristics constrain the size, weight, and cost of the vehicle architecture; provides inductive storage pulse power with a very short pulse rise time; multimegajoule energy bursts/terawatt power bursts; compact pulse power driver for low-impedance dense plasma devices; utilization of low cost HTSC material and casting technology to increase magnetic flux conservation and inductive energy storage; improvement in chemical/nuclear-to-electric energy conversion efficiency and the ability to generate significant levels of thrust with very high specific impulse; potential for developing a small, lightweight, low cost, self-excited integrated propulsion and power system suitable for space stations, planetary bases, and interplanetary and interstellar space travel; potential for attaining specific impulses approaching 10 (exp 6) seconds, which would enable missions to the outer planets within ten years and missions at interstellar distances within fifty years.

Litchford, Ron J.↗
NASA NTRS1985

Stall recovery control strategy methodology and results

The nonrecoverable stall condition in high performance turbofan engines is characterized by thrust loss, rising turbine temperatures, high pressure compressor rotating stall, and loss of engine control. Stall recovery control is presently investigated by means of an engine system computer model capable of either surge or rotating stall postinstability operation. Several techniques are examined which can yield rapid recovery from stall; a composite strategy which involves reduction of engine speed to idle while simultaneously opening the 10th- and 14th-stage compressor bleed ports allowed recovery to speeds slightly higher than idle while combustor fuel flow continued.

Hopf, W. R.↗
NASA NTRS1990

Static investigation of a two-dimensional convergent-divergent exhaust nozzle with multiaxis thrust-vectoring capability

An investigation was conducted in the Static Test Facility of the NASA Langley 16-Foot Transonic Tunnel to determine the internal performance of two-dimensional convergent-divergent nozzles designed to have simultaneous pitch and yaw thrust vectoring capability. This concept utilized divergent flap rotation of thrust vectoring in the pitch plane and deflection of flat yaw flaps hinged at the end of the sidewalls for yaw thrust vectoring. The hinge location of the yaw flaps was varied at four positions from the nozzle exit plane to the throat plane. The yaw flaps were designed to contain the flow laterally independent of power setting. In order to eliminate any physical interference between the yaw flap deflected into the exhaust stream and the divergent flaps, the downstream corners of both upper and lower divergent flaps were cut off to allow for up to 30 deg of yaw flap deflection. The impact of varying the nozzle pitch vector angle, throat area, yaw flap hinge location, yaw flap length, and yaw flap deflection angle on nozzle internal performance characteristics, was studied. High-pressure air was used to simulate jet exhaust at nozzle pressure ratios up to 7.0. Static results indicate that configurations with the yaw flap hinge located upstream of the exit plane provide relatively high levels of thrust vectoring efficiency without causing large losses in resultant thrust ratio. Therefore, these configurations represent a viable concept for providing simultaneous pitch and yaw thrust vectoring.

Taylor, John G.↗
NASA NTRS1998

Elimination of High-Frequency Combustion Instability in the Fastrac Engine Thrust Chamber

NASA's Marshall Space Flight Center(MSFC) has been tasked with developing a 60,000 pound thrust, pump-fed, LOX/RP-1 engine under the Advanced Space Transportation Program(ASTP). This government-led design has been designated the Fastrac engine. The X-34 vehicle will use the Fastrac engine as the main propulsion system. The X-34 will be a suborbital vehicle developed by the Orbital Sciences Corporation. The X-34 vehicle will be launched from an L-1011 airliner. After launch, the X-34 vehicle will be able to climb to altitudes up to 250,000 feet and reach speeds up to Mach 8, over a mission range of 500 miles. The overall length, wingspan, and gross takeoff weight of the X-34 vehicle are 58.3 feet, 27.7 feet and 45,000 pounds, respectively. This report summarizes the plan of achieving a Fastrac thrust chamber assembly(TCA) stable bomb test that meets the JANNAF standards, the Fastrac TCA design, and the combustion instabilities exhibited by the Fastrac TCA during testing at MSFC's test stand 116 as determined from high-frequency fluctuating pressure measurements. This report also summarizes the characterization of the combustion instabilities from the pressure measurements and the steps taken to eliminate the instabilities.

Rocker, Marvin↗