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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 973 records · Page 54
All sourcesNASA NTRSDOE OSTI
NASA NTRS2014

Propulsion System Models for Rotorcraft Conceptual Design

The conceptual design code NDARC (NASA Design and Analysis of Rotorcraft) was initially implemented to model conventional rotorcraft propulsion systems, consisting of turboshaft engines burning jet fuel, connected to one or more rotors through a mechanical transmission. The NDARC propulsion system representation has been extended to cover additional propulsion concepts, including electric motors and generators, rotor reaction drive, turbojet and turbofan engines, fuel cells and solar cells, batteries, and fuel (energy) used without weight change. The paper describes these propulsion system components, the architecture of their implementation in NDARC, and the form of the models for performance and weight. Requirements are defined for improved performance and weight models of the new propulsion system components. With these new propulsion models, NDARC can be used to develop environmentally-friendly rotorcraft designs.

Propulsion System Models↗
NASA NTRSDate not supplied

Analysis of Additively Manufactured Inconel 718 Combustion Behavior in Promoted Oxygen Environments

Promoted combustion testing is a vital tool for engineers to establish the combustion and flammability characteristics of materials (metallic or otherwise) in oxygen enriched environments. Historically, much of the established data for metallic promoted combustion has been with regards to cast and wrought forms. However, with the emergence of additive manufacturing as a preferred method of fabrication, the need exists to evaluate how metals in that form behave. This paper will serve as a review of the work that has been done and an analysis of the nickel-based superalloy Inconel 718, a material popular for aerospace applications such as liquid fueled rocket components and turbine engines. Promoted combustion testing (per the ASTM G124 standard) was conducted on samples of both wrought and selective laser melted fabrication, to provide comparison of flammability response between materials produced by each manufacturing method. Additionally, post-build treatments were applied to test samples to identify any effects on performance provided by hot isostatic pressing, oxygen-getting wrap during HIP, stress relieving, and solutionizing/aging heat treatments. This study will utilize optical and scanning electron microscopy, energy dispersive spectroscopy, x-ray diffraction, and metallography to identify the differences in behavior of additively manufactured and wrought Inconel 718.

Additively Manufactured↗
NASA NTRSDate not supplied

Aeropropulsive Assessment of a BLI Tail-Cone Thruster Propulsion System for the NASA Common Research Model and a Transonic Truss-Braced Wing Configuration

This paper presents an investigation into the potential benefits of a boundary layer ingestion(BLI) aft fuselage propulsion system, or tail-cone thruster (TCT), when applied to two transport category aircraft: a conventional tube-and-wing configuration based on the NASA Common Research Model (CRM) and a transonic truss-braced wing (TTBW) configuration. Through the ingestion of low-momentum flow developed over the airframe of an aircraft, the TCT propulsion system technology can reduce overall power consumption and hence fuel burn and emissions. However, an accurate assessment of this potential requires the application of high-fidelity analysis tools that can capture the relevant propulsion-airframe integration effects, at least to first order. Toward this end, the present study employs aeropropulsive analysis tools based on the Reynolds-averaged Navier-Stokes equations to estimate the power savings coefficient, measured relative to equivalent non-BLI configurations of a given airframe. Results indicate shaft power savings of 4.1% and 5.9% can be obtained for the boundary-layer ingesting TCT configuration of the CRM and TTBW, respectively. The present study also examines the coupling between airframe aerodynamics and the inlet flow distortion experienced by the TCT of a given aircraft configuration through airframe component sensitivity studies.

AATT↗
NASA NTRS2016

Application of Deterministic and Probabilistic System Design Methods and Enhancements of Conceptual Design Tools for ERA Project

This report documents work done by the Aerospace Systems Design Lab (ASDL) at the Georgia Institute of Technology, Daniel Guggenheim School of Aerospace Engineering for the National Aeronautics and Space Administration, Aeronautics Research Mission Directorate, Integrated System Research Program, Environmentally Responsible Aviation (ERA) Project. This report was prepared under contract NNL12AA12C, "Application of Deterministic and Probabilistic System Design Methods and Enhancement of Conceptual Design Tools for ERA Project". The research within this report addressed the Environmentally Responsible Aviation (ERA) project goal stated in the NRA solicitation "to advance vehicle concepts and technologies that can simultaneously reduce fuel burn, noise, and emissions." To identify technology and vehicle solutions that simultaneously meet these three metrics requires the use of system-level analysis with the appropriate level of fidelity to quantify feasibility, benefits and degradations, and associated risk. In order to perform the system level analysis, the Environmental Design Space (EDS) [Kirby 2008, Schutte 2012a] environment developed by ASDL was used to model both conventional and unconventional configurations as well as to assess technologies from the ERA and N+2 timeframe portfolios. A well-established system design approach was used to perform aircraft conceptual design studies, including technology trade studies to identify technology portfolios capable of accomplishing the ERA project goal and to obtain accurate tradeoffs between performance, noise, and emissions. The ERA goal, shown in Figure 1, is to simultaneously achieve the N+2 benefits of a cumulative noise margin of 42 EPNdB relative to stage 4, a 75 percent reduction in LTO NOx emissions relative to CAEP 6 and a 50 percent reduction in fuel burn relative to the 2005 best in class aircraft. There were 5 research task associated with this research: 1) identify technology collectors, 2) model technology collectors in EDS, 3) model and assess ERA technologies, 4) LTO and cruise emission prediction, and 5) probabilistic analysis of technology collectors and portfolios.

Mavris, Dimitri N.↗
NASA NTRS1979

Perturbation solutions of combustion instability problems

A method involving approximate modal analysis using the Galerkin method followed by an approximate solution of the resulting modal-amplitude equations by the two-variable perturbation method (method of multiple scales) is applied to two problems of pressure-sensitive nonlinear combustion instability in liquid-fuel rocket motors. One problem exhibits self-coupled instability while the other exhibits mode-coupled instability. In both cases it is possible to carry out the entire linear stability analysis and significant portions of the nonlinear stability analysis in closed form. In the problem of self-coupled instability the nonlinear stability boundary and approximate forms of the limit-cycle amplitudes and growth and decay rates are determined in closed form while the exact limit-cycle amplitudes and growth and decay rates are found numerically. In the problem of mode-coupled instability the limit-cycle amplitudes are found in closed form while the growth and decay rates are found numerically. The behavior of the solutions found by the perturbation method are in agreement with solutions obtained using complex numerical methods.

Googerdy, A.↗
NASA NTRS2018

Study of a Tricarbide Grooved Ring Fuel Element for Nuclear Thermal Propulsion

Deep space exploration, especially that of Mars, is on the horizon as the next big challenge for space exploration. Nuclear propulsion, through which high thrust and efficiency can be achieved, is a promising option for decreasing the cost and logistics of such a mission. Work on nu- clear thermal engines goes back to the days of the NERVA program. Currently, nuclear thermal propulsion is under development again in various forms to provide a superior propulsion system for deep space exploration. The authors have been working to develop a concept nuclear thermal engine that uses a grooved ring fuel element as an alternative to the traditional hexagonal rod design. The authors are also studying the use of carbide fuels. The concept was developed in order to increase surface area and heat transfer to the propellant. The use of carbides would also raise the operating temperature of the reactor. It is hoped that this could lead to a higher thrust to weight nuclear thermal engine. This paper describes the modeling of neutronics, heat transfer, and fluid dynamics of this alternative nuclear fuel element geometry. Fabrication experiments of grooved rings from carbide refractory metals are also presented along with material characterization and interactions with a hot hydrogen environment. Results of experiments and associated analysis are desired densities with some success in material distribution and reaching a solid solution. Future work is needed to improve distribution of material, minimize oxidation during the milling process, and de ne a fabrication process that will serve for constructing grooved ring fuel rods for large system tests.

Taylor, Brian↗
NASA NTRS2012

Alternate-Fueled Combustor-Sector Performance

In order to realize alternative fueling for military and commercial use, the industry has set forth guidelines that must be met by each fuel. These aviation fueling requirements are outlined in MILDTL- 83133F(2008) or ASTM D 7566 Annex (2011) standards, and are classified as drop-in fuel replacements. This paper provides combustor performance data for synthetic-paraffinic-kerosene- (SPK-) type (Fisher-Tropsch (FT)) fuel and blends with JP-8+100, relative to JP-8+100 as baseline fueling. Data were taken at various nominal inlet conditions: 75 psia (0.52 MPa) at 500 F (533 K), 125 psia (0.86 MPa) at 625 F (603 K), 175 psia (1.21 MPa) at 725 F (658 K), and 225 psia (1.55 MPa) at 790 F (694 K). Combustor performance analysis assessments were made for the change in flame temperatures, combustor efficiency, wall temperatures, and exhaust plane temperatures at 3%, 4%, and 5% combustor pressure drop (% delta P) for fuel: air ratios (F/A) ranging from 0.010 to 0.025. Significant general trends show lower liner temperatures and higher flame and combustor outlet temperatures with increases in FT fueling relative to JP-8+100 fueling. The latter affects both turbine efficiency and blade/vane life.

Thomas, Anna E.↗
NASA NTRS2013

Alternate-Fueled Combustor-Sector Performance

In order to realize alternative fueling for military and commercial use, the industry has set forth guidelines that must be met by each fuel. These aviation fueling requirements are outlined in MIL-DTL-83133F(2008) or ASTM D 7566 Annex (2011) standards, and are classified as "drop-in" fuel replacements. This report provides combustor performance data for synthetic-paraffinic-kerosene- (SPK-) type (Fischer-Tropsch (FT)) fuel and blends with JP-8+100, relative to JP-8+100 as baseline fueling. Data were taken at various nominal inlet conditions: 75 psia (0.52 MPa) at 500 degF (533 K), 125 psia (0.86 MPa) at 625 degF (603 K), 175 psia (1.21 MPa) at 725 degF (658 K), and 225 psia (1.55 MPa) at 790 degF (694 K). Combustor performance analysis assessments were made for the change in flame temperatures, combustor efficiency, wall temperatures, and exhaust plane temperatures at 3, 4, and 5 percent combustor pressure drop (DP) for fuel:air ratios (F/A) ranging from 0.010 to 0.025. Significant general trends show lower liner temperatures and higher flame and combustor outlet temperatures with increases in FT fueling relative to JP-8+100 fueling. The latter affects both turbine efficiency and blade and vane lives.

Thomas, Anna E.↗
NASA NTRS2008

Turbine Design and Analysis for the J-2X Engine Turbopumps

Pratt and Whitney Rocketdyne and NASA Marshall Space Flight Center are developing the advanced upper stage J-2X engine based on the legacy design of the J-2/J-2S family of engines which powered the Apollo missions. The cryogenic propellant turbopumps have been denoted as Mark72-F and Mark72-0 for the fuel and oxidizer side, respectively. Special attention is focused on preserving the essential flight-proven design features while adapting the design to the new turbopump configuration. Advanced 3-D CFD analysis has been employed to verify turbine aero performance at current flow regime boundary conditions and to mitigate risks associated with stresses. A limited amount of redesign and overall configuration modifications allow for a robust design with performance level matching or exceeding requirement.

Marcu, Bogdan↗
NASA NTRS1990

Electrocatalysis for oxygen electrodes in fuel cells and water electrolyzers for space applications

The lead ruthenate pyrochlore Pb2Ru2O6.5, in both high- and low-area forms, has been characterized using thermogravimetric analysis, X-ray photoelectron spectroscopy, X-ray diffraction, cyclic voltammetry, and O2 reduction and generation kinetic-mechanistic studies. Mechanisms are proposed. Compounds in which part of the Ru is substituted with Ir have also been prepared. They exhibit somewhat better performance for O2 reduction in porous, gas-fed electrodes than the unsubstituted compound. The anodic corrosion resistance of pyrochlore-based porous electrodes was improved by using two different anionically conducting polymer overlayers, which slow down the diffusion of ruthenate and plumbate out of the electrode. The O2 generation performance was improved with both types of electrodes. With a hydrogel overlayer, the O2 reduction performance was also improved.

Prakash, Jai↗
NASA NTRS1991

Optimization of vehicle weight for Mars excursion missions

The excursion class mission to Mars is defined as a mission with one year durations coupled with a stay time of up to 30 days. The fuel budget for such a mission is investigated. The overall figure of merit in such an assessment is the vehicle weight ratio, the ratio between the wet vehicle weight to the dry vehicle weight. It is necessary to minimize the overall fuel budget for the mission in order to maximize the benefits that could be obtained by sending humans to Mars. Assumptions used in the analysis are: each mission will depart and terminate in low-earth-orbit (LEO) (500 km circular) and the maximum stay time at Mars is 30 days. The maximum mission duration is one year (355-375 days). The mass returned to LEO is 135,000 kg, the dropoff mass left at Mars is 168,000 kg. Three propulsive techniques for atmospheric interface are investigated: aerobraking, all-chemical propulsion, and nuclear thermal propulsion. By defining the fuel requirements, the space transfer vehicle's configuration is defined.

Ferebee, Melvin J., Jr.↗
NASA NTRS1994

Effect of power system technology and mission requirements on high altitude long endurance aircraft

An analysis was performed to determine how various power system components and mission requirements affect the sizing of a solar powered long endurance aircraft. The aircraft power system consists of photovoltaic cells and a regenerative fuel cell. Various characteristics of these components, such as PV cell type, PV cell mass, PV cell efficiency, fuel cell efficiency, and fuel cell specific mass, were varied to determine what effect they had on the aircraft sizing for a given mission. Mission parameters, such as time of year, flight altitude, flight latitude, and payload mass and power, were also altered to determine how mission constraints affect the aircraft sizing. An aircraft analysis method which determines the aircraft configuration, aspect ratio, wing area, and total mass, for maximum endurance or minimum required power based on the stated power system and mission parameters is presented. The results indicate that, for the power system, the greatest benefit can be gained by increasing the fuel cell specific energy. Mission requirements also substantially affect the aircraft size. By limiting the time of year the aircraft is required to fly at high northern or southern latitudes, a significant reduction in aircraft size or increase in payload capacity can be achieved.

Colozza, Anthony J.↗
NASA NTRS1996

Heuristic Modeling for TRMM Lifetime Predictions

Analysis time for computing the expected mission lifetimes of proposed frequently maneuvering, tightly altitude constrained, Earth orbiting spacecraft have been significantly reduced by means of a heuristic modeling method implemented in a commercial-off-the-shelf spreadsheet product (QuattroPro) running on a personal computer (PC). The method uses a look-up table to estimate the maneuver frequency per month as a function of the spacecraft ballistic coefficient and the solar flux index, then computes the associated fuel use by a simple engine model. Maneuver frequency data points are produced by means of a single 1-month run of traditional mission analysis software for each of the 12 to 25 data points required for the table. As the data point computations are required only a mission design start-up and on the occasion of significant mission redesigns, the dependence on time consuming traditional modeling methods is dramatically reduced. Results to date have agreed with traditional methods to within 1 to 1.5 percent. The spreadsheet approach is applicable to a wide variety of Earth orbiting spacecraft with tight altitude constraints. It will be particularly useful to such missions as the Tropical Rainfall Measurement Mission scheduled for launch in 1997, whose mission lifetime calculations are heavily dependent on frequently revised solar flux predictions.

Jordan, P. S.↗
NASA NTRS2004

Calculated Drag of an Aerial Refueling Assembly Through Airplane Performance Analysis

The aerodynamic drag of an aerial refueling assembly was calculated during the Automated Aerial Refueling project at the NASA Dryden Flight Research Center. An F/A-18A airplane was specially instrumented to obtain accurate fuel flow measurements and to determine engine thrust. A standard Navy air refueling store with a retractable refueling hose and paradrogue was mounted to the centerline pylon of the F/A-18A airplane. As the paradrogue assembly was deployed and stowed, changes in the calculated thrust of the airplane occurred and were equated to changes in vehicle drag. These drag changes were attributable to the drag of the paradrogue assembly. The drag of the paradrogue assembly was determined to range from 200 to 450 lbf at airspeeds from 170 to 250 KIAS. Analysis of the drag data resulted in a single drag coefficient of 0.0056 for the paradrogue assembly that adequately matched the calculated drag for all flight conditions. The drag relief provided to the tanker airplane when a receiver airplane engaged the paradrogue is also documented from 35 to 270 lbf at the various flight conditions tested. The results support the development of accurate aerodynamic models to be used in refueling simulations and control laws for fully autonomous refueling.

Vachon, Michael Jacob↗
NASA NTRS1981

Design for supercommuters

Action to be taken to prepare to implement efficient, modern commuter aircraft for the 1990s is outlined. The increase in the contribution of jet fuel costs to aircraft direct operating costs (DOC) is noted as the motivation for the introduction of turboprop-powered commuter aircraft, which use 15-20% less fuel per seat mile at short stage lengths, to replace larger jet transports. Designs proposed by various manufacturers which will make use of existing technology for 19-, 30- and 50-seat aircraft capable of carrying a full payload of passengers and baggage for 600 n mi and optimized for minimum DOC over a 100-n mi stage length are presented, and the improvements in fuel usage, DOC and passenger comfort to be obtained with the use of advanced technology are pointed out. The goals and considered technologies of the dedicated small-transport aircraft technology program recommended by a commuter air transport subcommittee of the NASA Advisory Council Aeronautics Advisory Committee to speed the development of commuter technology are then presented, with attention given to efforts of analysis, design and testing of propulsion systems, structures, aerodynamics and systems intended to result in 16-24% savings in DOC and up to 40% savings in fuel. The commuter development plans of various manufacturers are also indicated.

Williams, L. J.↗
NASA NTRS2005

Investigation of Ruthenium Dissolution in Advanced Membrane Electrode Assemblies for Direct Methanol Based Fuel Cell Stacks

Dissolution of ruthenium was observed in the 80-cell stack. Duration testing was performed in single cell MEAs to determine the pathway of cell degradation. EDAX analysis on each of the single cell MEAs has shown that the Johnson Matthey commercial catalyst is stable in DMFC operation for 250 hours, no ruthenium dissolution was observed. Changes in the hydrophobicity of the cathode backing papers was minimum. Electrode polarization analysis revealed that the MEA performance loss is attributed to changes in the cathode catalyst layer. Ruthenium migration does not seem to occur during cell operation but can occur when methanol is absent from the anode compartment, the cathode compartment has access to air, and the cells in the stack are electrically connected to a load (Shunt Currents). The open-to-air cathode stack design allowed for: a) The MEAs to have continual access to oxygen; and b) The stack to sustain shunt currents. Ruthenium dissolution in a DMFC stack can be prevented by: a) Developing an internally manifolded stacks that seal reactant compartments when not in operation; b) Bringing the cell voltages to zero quickly when not in operation; and c) Limiting the total number of cells to 25 in an effort to limit shunt currents.

alternative power sources↗
NASA NTRS1991

Integrated trajectory and control analysis for generic hypersonic vehicles

A tool which supports integrated assessment of air-breathing hypersonic vehicle trajectory management and control strategy has been developed by Hattis and Malchow (1991). Using a two-point boundary value problem solution technique, the tool can construct an entire near-minimum-fuel trajectory and desired control strategy from takeoff to orbit, while satisfying physically derived inequality constraints and while achieving efficient propulsive mode phasing. A strategy is also formulated to construct preliminary trajectory and control history representations with less computational burden than required for the overall flight profile assessment. A tabulated example hypersonic vehicle model has been used to demonstrate the integrated analysis methodology. To assure good numerical behavior when the algorithm uses tabulated model data, a multidimensional cubic spline curve smoothing routine has been developed. A dynamic-pressure-constrained, near-fuel-optimal trajectory was computed to provide examples of information that the methodology can provide. Previously unspecified propulsive discontinuities were located, flight regimes demanding rapid attitude changes were identified, available and closed-loop controller authority was ascertained, and inadequacies in vehicle and specific subsystem model representations were determined.

Hattis, Philip D.↗