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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 37 records · Page 2
All sourcesNASA NTRSDOE OSTI
NASA NTRS1985

Study for the optimization of a transport aircraft wing for maximum fuel efficiency. Volume 1: Methodology, criteria, aeroelastic model definition and results

Work performed in the design of a transport aircraft wing for maximum fuel efficiency is documented with emphasis on design criteria, design methodology, and three design configurations. The design database includes complete finite element model description, sizing data, geometry data, loads data, and inertial data. A design process which satisfies the economics and practical aspects of a real design is illustrated. The cooperative study relationship between the contractor and NASA during the course of the contract is also discussed.

Radovcich, N. A.↗
NASA NTRS1983

Design study of a dual-cycle turbofan-ramjet engine for a hypersonic aircraft

Computer modelling was used with two different designs of an advanced turbofan-ramjet in order to derive performance predictions. The engine would enable an aircraft to take-off, accelerate to Mach 5.0, and climb to 90,000 ft. The two concepts included a turbofan with a ramjet annularly wrapped around it and a side-by-side configuration with the ramjet having a rectangular shape and mounted alongside the turbofan. The studies were performed to model weight, length, fuel efficiency, and the requirements of the thrust/drag ratio to exceed unity over the entire flight path. LH2 would be used for fuel and to regeneratively cool the combustion chamber. Turbofan operation with and without afterburner and with and without the ramjet inlet open were examined, as were variable areas for the burners. A side-by-side configuration displayed the best performance predictions, with a ramjet mass flow being 75 percent that of the turbofan and maximum temperatures being equal.

Zumwalt, G. W.↗
NASA NTRS1982

Additional experiments on flowability improvements of aviation fuels at low temperatures, volume 2

An investigation was performed to study flow improver additives and scale-model fuel heating systems for use with aviation hydrocarbon fuel at low temperatures. Test were performed in a facility that simulated the heat transfer and temperature profiles anticipated in wing fuel tanks during flight of long-range commercial aircraft. The results are presented of experiments conducted in a test tank simulating a section of an outer wing integral fuel tank approximately full-scale in height, chilled through heat exchange panels bonded to the upper and lower horizontal surfaces. A separate system heated lubricating oil externally by a controllable electric heater, to transfer heat to fuel pumped from the test tank through an oil-to-fuel heat exchanger, and to recirculate the heated fuel back to the test tank.

Stockemer, F. J.↗
NASA NTRS2018

Vehicle-Level System Impact of Boundary Layer Ingestion for the NASA D8 Concept Aircraft

The purpose of this study was to evaluate the vehicle-level impact of a boundary layer ingestion (BLI) propulsion system on a commercial transport aircraft concept. The NASA D8 (ND8) aircraft was chosen as the BLI concept aircraft to be studied. A power balance methodology developed by the Massachusetts Institute of Technology was adapted for use with the existing NASA sizing and performance tools to model the fuel consumption impact of BLI on the ND8. A key assumption for the BLI impact assessment was a 3.5% efficiency penalty associated with designing a fan for and operating in the distorted flow caused by BLI. The ND8 was compared to several other ND8-like aircraft that did not utilize BLI in order to determine the fuel consumption benefit attributable to BLI. Analytically “turning off” BLI on the ND8 without accounting for the physical requirements of redirecting the boundary layer or resizing the aircraft to meet the performance constraints resulted in a 2.8% increase in block fuel consumption to fly the design mission. When this non-physical aircraft was resized to meet the performance constraints, the block fuel consumption was 4.0% greater than the baseline ND8. The ND8 was also compared to an ND8-like aircraft with conventionally podded engines under the wing. This configuration had a 5.6% increase in block fuel consumption compared to the baseline ND8. This result is more reflective of the real world impact if BLI is not an available technology for the ND8 design. The BLI benefit results presented for this study should not be applied to other aircraft that have a propulsion-airframe integration design or BLI implementation different from the ND8.

Marien, Ty V.↗
NASA NTRS1994

Proton exchange membrane fuel cells for space and electric vehicle applications: From basic research to technology development

The proton exchange membrane fuel cell (PEMFC) is one of the most promising electrochemical power sources for space and electric vehicle applications. The wide spectrum of R&D activities on PEMFC's, carried out in our Center from 1988 to date, is as follows (1) Electrode Kinetic and Electrocatalysis of Oxygen Reduction; (2) Optimization of Structures of Electrodes and of Membrane and Electrode Assemblies; (3) Selection and Evaluation of Advanced Proton Conducting Membranes and of Operating Conditions to Attain High Energy Efficiency; (4) Modeling Analysis of Fuel Cell Performance and of Thermal and Water Management; and (5) Engineering Design and Development of Multicell Stacks. The accomplishments on these tasks may be summarized as follows: (1) A microelectrode technique was developed to determine the electrode kinetic parameters for the fuel cell reactions and mass transport parameters for the H2 and O2 reactants in the proton conducting membrane. (2) High energy efficiencies and high power densities were demonstrated in PEMFCs with low platinum loading electrodes (0.4 mg/cm(exp 2) or less), advanced membranes and optimized structures of membrane and electrode assemblies, as well as operating conditions. (3) The modeling analyses revealed methods to minimize mass transport limitations, particularly with air as the cathodic reactant; and for efficient thermal and water management. (4) Work is in progress to develop multi-kilowatt stacks with the electrodes containing low platinum loadings.

Srinivasan, Supramaniam↗
NASA NTRS2017

Thermal System Modeling for Lunar and Martian Surface Regenerative Fuel Cell Systems

The Advanced Exploration Systems (AES) Advanced Modular Power Systems (AMPS) Project is investigating different power systems for various lunar and Martian mission concepts. The AMPS Fuel Cell (FC) team has created two system-level models to evaluate the performance of regenerative fuel cell (RFC) systems employing different fuel cell chemistries. Proton Exchange Membrane fuel cells PEMFCs contain a polymer electrolyte membrane that separates the hydrogen and oxygen cavities and conducts hydrogen cations (protons) across the cell. Solid Oxide fuel cells (SOFCs) operate at high temperatures, using a zirconia-based solid ceramic electrolyte to conduct oxygen anions across the cell. The purpose of the modeling effort is to down select one fuel cell chemistry for a more detailed design effort. Figures of merit include the system mass, volume, round trip efficiency, and electrolyzer charge power required. PEMFCs operate at around 60 degrees Celsius versus SOFCs which operate at temperatures greater than 700 degrees Celsius. Due to the drastically different operating temperatures of the two chemistries the thermal control systems (TCS) differ. The PEM TCS is less complex and is characterized by a single pump cooling loop that uses deionized water coolant and rejects heat generated by the system to the environment via a radiator. The solid oxide TCS has its own unique challenges including the requirement to reject high quality heat and to condense the steam produced in the reaction. This paper discusses the modeling of thermal control systems for an extraterrestrial RFC that utilizes either a PEM or solid oxide fuel cell.

thermal modeling↗
NASA NTRS2017

Thermal Design for Extra-Terrestrial Regenerative Fuel Cell System

The Advanced Exploration Systems (AES) Advanced Modular Power Systems (AMPS) Project is investigating different power systems for various lunar and Martian mission concepts. The AMPS Fuel Cell (FC) team has created two system-level models to evaluate the performance of regenerative fuel cell (RFC) systems employing different fuel cell chemistries. Proton Exchange Membrane fuel cells PEMFCs contain a polymer electrolyte membrane that separates the hydrogen and oxygen cavities and conducts hydrogen cations (protons) across the cell. Solid Oxide fuel cells (SOFCs) operate at high temperatures, using a zirconia-based solid ceramic electrolyte to conduct oxygen anions across the cell. The purpose of the modeling effort is to down select one fuel cell chemistry for a more detailed design effort. Figures of merit include the system mass, volume, round trip efficiency, and electrolyzer charge power required. PEMFCs operate at around 60 C versus SOFCs which operate at temperatures greater than 700 C. Due to the drastically different operating temperatures of the two chemistries the thermal control systems (TCS) differ. The PEM TCS is less complex and is characterized by a single pump cooling loop that uses deionized water coolant and rejects heat generated by the system to the environment via a radiator. The solid oxide TCS has its own unique challenges including the requirement to reject high quality heat and to condense the steam produced in the reaction. This paper discusses the modeling of thermal control systems for an extraterrestrial RFC that utilizes either a PEM or solid oxide fuel cell.

thermal modeling↗
NASA NTRS1976

Langley facility for tests at Mach 7 of subscale, hydrogen-burning, airframe-integratable, scramjet models

Modifications to a 20-megawatt arc-heated facility for testing a hydrogen-burning, airframe-integratable, subscale, scramjet model are described. Arc-heated flow is mixed with unheated air to furnish a test flow duplicating Mach 7 flight. (Stagnation temperature is 2220 K.) Modifications to the commercially available heater to improve survivability and smoothness are described. Pitot profiles show uniform flow and a slightly thinner nozzle boundary layer than predicted. Comparison of the tunnel boundary layer, which will be ingested by the engine model, with the boundary layer that a flight engine might ingest from its vehicle forebody shows a difference in the density distribution through the boundary layer. Calculations of wall heating and transient wall temperatures of the engine model show that for a 30-sec burn, the heat sink model requires cooling at selected locations to avoid thermal-stress, cycle-life problems. Model performance predictions show that fuel equivalence ratio and nozzle exit area both have large effects on thrust. Average inlet entrance Mach number (as affected by boundary-layer ingestion) has little effect on thrust.

Boatright, W. B.↗
NASA NTRS1980

Study of research and development requirements of small gas-turbine combustors

A survey is presented of the major small-engine manufacturers and governmental users. A consensus was undertaken regarding small-combustor requirements. The results presented are based on an evaluation of the information obtained in the course of the study. The current status of small-combustor technology is reviewed. The principal problems lie in liner cooling, fuel injection, part-power performance, and ignition. Projections of future engine requirements and their effect on the combustor are discussed. The major changes anticipated are significant increases in operating pressure and temperature levels and greater capability of using heavier alternative fuels. All aspects of combustor design are affected, but the principal impact is on liner durability. An R&D plan which addresses the critical combustor needs is described. The plan consists of 15 recommended programs for achieving necessary advances in the areas of liner thermal design, primary-zone performance, fuel injection, dilution, analytical modeling, and alternative-fuel utilization.

Demetri, E. P.↗
NASA NTRS1990

Hybrid Propulsion Technology Program

Future launch systems of the United States will require improvements in booster safety, reliability, and cost. In order to increase payload capabilities, performance improvements are also desirable. The hybrid rocket motor (HRM) offers the potential for improvements in all of these areas. The designs are presented for two sizes of hybrid boosters, a large 4.57 m (180 in.) diameter booster duplicating the Advanced Solid Rocket Motor (ASRM) vacuum thrust-time profile and smaller 2.44 m (96 in.), one-quater thrust level booster. The large booster would be used in tandem, while eight small boosters would be used to achieve the same total thrust. These preliminary designs were generated as part of the NASA Hybrid Propulsion Technology Program. This program is the first phase of an eventual three-phaes program culminating in the demonstration of a large subscale engine. The initial trade and sizing studies resulted in preferred motor diameters, operating pressures, nozzle geometry, and fuel grain systems for both the large and small boosters. The data were then used for specific performance predictions in terms of payload and the definition and selection of the requirements for the major components: the oxidizer feed system, nozzle, and thrust vector system. All of the parametric studies were performed using realistic fuel regression models based upon specific experimental data.

Jensen, G. E.↗
NASA NTRS1982

Fuel optimal maneuvers for spacecraft with fixed thrusters

Several mathematical models, including a minimum integral square criterion problem, were used for the qualitative investigation of fuel optimal maneuvers for spacecraft with fixed thrusters. The solutions consist of intervals of "full thrust" and "coast" indicating that thrusters do not need to be designed as "throttleable" for fuel optimal performance. For the primary model considered, singular solutions occur only if the optimal solution is "pure translation". "Time optimal" singular solutions can be found which consist of intervals of "coast" and "full thrust". The shape of the optimal fuel consumption curve as a function of flight time was found to depend on whether or not the initial state is in the region admitting singular solutions. Comparisons of fuel optimal maneuvers in deep space with those relative to a point in circular orbit indicate that qualitative differences in the solutions can occur. Computation of fuel consumption for certain "pure translation" cases indicates that considerable savings in fuel can result from the fuel optimal maneuvers.

Carter, T. C.↗
NASA NTRS1992

Performance comparison of axisymmetric and three-dimensional hydrogen film coolant injection in a 110 N hydrogen/oxygen rocket

An experimental performance comparison of two geometrically different fuel film coolant injection sleeves was conducted on a 110 N gaseous hydrogen/oxygen rocket. One sleeve had slots milled axially down the walls and the other had a smooth surface to give axisymmetric flow. The comparison was made to investigate a conclusion in an earlier study that attributed a performance underprediciton to a simplifying modeling assumption of axisymmetric fuel film flow. The smooth sleeve had higher overall performance at one film coolant percentage and approximately the same or slightly better at another. The study showed that the lack of modeling of three-dimensional effects was not the cause of the performance underprediciton as speculated in earlier analytical studies.

Arrington, Lynn A.↗
NASA NTRS1992

Performance Comparison of Axisymmetric and Three-dimensional Hydrogen Film Coolant Injection in a 110N Hydrogen/oxygen Rocket

An experimental performance comparison of two geometrically different fuel film coolant injection sleeves was conducted on a 110 N gaseous hydrogen/oxygen rocket. One sleeve had slots milled axially down the walls and the other had a smooth surface to give axisymmetric flow. The comparison was made to investigate a conclusion in an earlier study that attributed a performance underprediction to a symplifying modeling assumption of axisymmetric fuel film flow. The smooth sleeve had higher overall performance at one film coolant percentage and approximately the same or slightly better at another. The study showed that the lack of modeling of three-dimensional effects was not the cause of the performance underprediction as speculated in earlier analytical studies.

Arrington, Lynn A.↗
NASA NTRS1993

Fuels and Combustion Technology for Advanced Aircraft Engines [Les Propergols et les Systèmes de Combustion pour les Moteurs d'Aéronefs]

The Conference Proceedings contains the 38 papers presented at the Propulsion and Energetics Panel 81st Symposium on "Fuels and Combustion Technology for Advanced Aircraft Engines" which was held from 10th-14th May 1993, in Fiuggi, Italy. The Technical Evaluation Report and the Keynote Address are included at the beginning, and discussions follow most papers. The Symposium was arranged in the following Sessions: Technology Overview Papers (2); Modelling: Pollutant Formation (4); Modelling: Combustor Design (5); High Temperature Fuels and Fuel Systems (6); Combustion Research: Performance (6); Combustion Research: Emissions (5); Fuel Atomization: Diagnostics and Modelling (5); and Combustion Research: Flowfleld and Mixing (4). The last paper is a contribution from Russia not allocated to a session. New technologies for low NOx combustors and advanced high-pressure/high temperature cycle engines result in unique problems in design and performance. There have been significant advances in modelling and diagnostics to aid the development of these technologies. The purpose of the Symposium was to bring together experts from industry, research establishments and universities to discuss fundamental and applied research in these areas as relevant to the development of advanced gas turbine engines, to exchange practical experience and to discuss the state of the art.

Fuels for aircraft↗
NASA NTRS1994

Space power systems technology

Reported here is a series of studies which examine several potential catalysts and electrodes for some fuel cell systems, some materials for space applications, and mathematical modeling and performance predictions for some solid oxide fuel cells and electrolyzers. The fuel cell systems have a potential for terrestrial applications in addition to solar energy conversion in space applications. Catalysts and electrodes for phosphoric acid fuel cell systems and for polymer electrolyte membrane (PEM) fuel cell and electrolyzer systems were examined.

Coulman, George A.↗
NASA NTRS2003

A Nonlinear Model for Fuel Atomization in Spray Combustion

Most gas turbine combustion codes rely on ad-hoc statistical assumptions regarding the outcome of fuel atomization processes. The modeling effort proposed in this project is aimed at developing a realistic model to produce accurate predictions of fuel atomization parameters. The model involves application of the nonlinear stability theory to analyze the instability and subsequent disintegration of the liquid fuel sheet that is produced by fuel injection nozzles in gas turbine combustors. The fuel sheet is atomized into a multiplicity of small drops of large surface area to volume ratio to enhance the evaporation rate and combustion performance. The proposed model will effect predictions of fuel sheet atomization parameters such as drop size, velocity, and orientation as well as sheet penetration depth, breakup time and thickness. These parameters are essential for combustion simulation codes to perform a controlled and optimized design of gas turbine fuel injectors. Optimizing fuel injection processes is crucial to improving combustion efficiency and hence reducing fuel consumption and pollutants emissions.

Liu, Nan-Suey↗
NASA NTRS1984

Fuel property effects on USAF gas turbine engine combustors and afterburners

Since the early 1970s, the cost and availability of aircraft fuel have changed drastically. These problems prompted a program to evaluate the effects of broadened specification fuels on current and future aircraft engine combustors employed by the USAF. Phase 1 of this program was to test a set of fuels having a broad range of chemical and physical properties in a select group of gas turbine engine combustors currently in use by the USAF. The fuels ranged from JP4 to Diesel Fuel number two (DF2) with hydrogen content ranging from 14.5 percent down to 12 percent by weight, density ranging from 752 kg/sq m to 837 kg/sq m, and viscosity ranging from 0.830 sq mm/s to 3.245 sq mm/s. In addition, there was a broad range of aromatic content and physical properties attained by using Gulf Mineral Seal Oil, Xylene Bottoms, and 2040 Solvent as blending agents in JP4, JP5, JP8, and DF2. The objective of Phase 2 was to develop simple correlations and models of fuel effects on combustor performance and durability. The major variables of concern were fuel chemical and physical properties, combustor design factors, and combustor operating conditions.

Reeves, C. M.↗
NASA NTRS2019

The Effect of Turbulence Modeling on the Mixing Characteristics of Several Fuel Injectors at Hypervelocity Flow Conditions

CFD analysis is presented on the effects of turbulence modeling choices on the mixing characteristics and performance of three fuel injectors at hypervelocity flow conditions. The analyses were carried out with the VULCAN-CFD solver using Reynolds-Averaged Simulations (RAS). The hypervelocity flow conditions match the high Mach number flow of the experiments conducted as a part of the Enhanced Injection and Mixing Project (EIMP) at the NASA Langley Research Center. The three injectors are the baseline configurations used in the experiments and represent three categories of injectors typically considered individually or in combination for fueling high-speed propulsive devices. The current work discusses the impact of the turbulence model and the turbulent Schmidt number on the mixing flow field behavior and the mixing performance as described by the one-dimensional values of the Mach number, total pressure recovery, and the mixing efficiency. Because planar laser induced fluorescence (PLIF) images are available from the EIMP experiments, the sensitivity of the synthetic LIF signal to turbulence modeling choices is also examined to determine whether PLIF can be extended beyond its intended qualitative visualization purpose and used to guide CFD turbulence model and parameter selections. It is found that the mixing performance, as quantified using mixing efficiency, exhibits a strong sensitivity to both turbulence model choice and turbulent Schmidt number value. However, the synthetic LIF signal only demonstrates a modest level of sensitivity, which suggests that PLIF is of limited use for guiding CFD turbulence model and parameter selections.

Drozda, Tomasz G.↗