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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 127 records · Page 7
All sourcesNASA NTRSDOE OSTI
NASA NTRS1987

High-surface-area, dual-function oxygen electrocatalysts for space power applications

The processes of hydration/dehydration and carbonation/decarbonation are investigated as an approach to provide higher surface area mixed metal oxides that are more active electrochemically. These materials are candidates for use as electrocatalysts and electrocatalyst supports for alkaline electrolyzers and fuel cells. For the case of the perovskite, LaCoO3 , higher surface areas were achieved with no change in structure and a more active oxygen electrocatalyst.

Ham, David O.↗
NASA NTRS1988

Evaluation and Improvement of Liquid Propellant Rocket Chugging Analysis Techniques. Part 2: a Study of Low Frequency Combustion Instability in Rocket Engine Preburners Using a Heterogeneous Stirred Tank Reactor Model

During the shutdown of the space shuttle main engine, oxygen flow is shut off from the fuel preburner and helium is used to push the residual oxygen into the combustion chamber. During this process a low frequency combustion instability, or chug, occurs. This chug has resulted in damage to the engine's augmented spark igniter due to backflow of the contents of the preburner combustion chamber into the oxidizer feed system. To determine possible causes and fixes for the chug, the fuel preburner was modeled as a heterogeneous stirred tank combustion chamber, a variable mass flow rate oxidizer feed system, a constant mass flow rate fuel feed system and an exit turbine. Within the combustion chamber gases were assumed perfectly mixed. To account for liquid in the combustion chamber, a uniform droplet distribution was assumed to exist in the chamber, with mean droplet diameter determined from an empirical relation. A computer program was written to integrate the resulting differential equations. Because chamber contents were assumed perfectly mixed, the fuel preburner model erroneously predicted that combustion would not take place during shutdown. The combustion rate model was modified to assume that all liquid oxygen that vaporized instantaneously combusted with fuel. Using this combustion model, the effect of engine parameters on chamber pressure oscillations during the SSME shutdown was calculated.

Bartrand, Timothy A.↗
NASA NTRS2016

A High-Heritage Blunt-Body Entry, Descent, and Landing Concept for Human Mars Exploration

Human-scale landers require the delivery of much heavier payloads to the surface of Mars than is possible with entry, descent, and landing (EDL) approaches used to date. A conceptual design was developed for a 10 m diameter crewed Mars lander with an entry mass of approx.75 t that could deliver approx.28 t of useful landed mass (ULM) to a zero Mars areoid, or lower, elevation. The EDL design centers upon use of a high ballistic coefficient blunt-body entry vehicle and throttled supersonic retro-propulsion (SRP). The design concept includes a 26 t Mars Ascent Vehicle (MAV) that could support a crew of 2 for approx.24 days, a crew of 3 for approx.16 days, or a crew of 4 for approx.12 days. The MAV concept is for a fully-fueled single-stage vehicle that utilizes a single pump-fed 250 kN engine using Mono-Methyl Hydrazine (MMH) and Mixed Oxides of Nitrogen (MON-25) propellants that would deliver the crew to a low Mars orbit (LMO) at the end of the surface mission. The MAV concept could potentially provide abort-to-orbit capability during much of the EDL profile in response to fault conditions and could accommodate return to orbit for cases where the MAV had no access to other Mars surface infrastructure. The design concept for the descent stage utilizes six 250 kN MMH/MON-25 engines that would have very high commonality with the MAV engine. Analysis indicates that the MAV would require approx.20 t of propellant (including residuals) and the descent stage would require approx.21 t of propellant. The addition of a 12 m diameter supersonic inflatable aerodynamic decelerator (SIAD), based on a proven flight design, was studied as an optional method to improve the ULM fraction, reducing the required descent propellant by approx.4 t.

Mars lander↗
NASA NTRS2016

A High-Heritage Blunt-Body Entry, Descent, and Landing Concept for Human Mars Exploration

Human-scale landers require the delivery of much heavier payloads to the surface of Mars than is possible with entry, descent, and landing (EDL) approaches used to date. A conceptual design was developed for a 10 m diameter crewed Mars lander with an entry mass of approx. 75 t that could deliver approx. 28 t of useful landed mass (ULM) to a zero Mars areoid, or lower, elevation. The EDL design centers upon use of a high ballistic coefficient blunt-body entry vehicle and throttled supersonic retro-propulsion (SRP). The design concept includes a 26 t Mars Ascent Vehicle (MAV) that could support a crew of 2 for approx. 24 days, a crew of 3 for approx.16 days, or a crew of 4 for approx.12 days. The MAV concept is for a fully-fueled single-stage vehicle that utilizes a single pump-fed 250 kN engine using Mono-Methyl Hydrazine (MMH) and Mixed Oxides of Nitrogen (MON-25) propellants that would deliver the crew to a low Mars orbit (LMO) at the end of the surface mission. The MAV concept could potentially provide abort-to-orbit capability during much of the EDL profile in response to fault conditions and could accommodate return to orbit for cases where the MAV had no access to other Mars surface infrastructure. The design concept for the descent stage utilizes six 250 kN MMH/MON-25 engines that would have very high commonality with the MAV engine. Analysis indicates that the MAV would require approx. 20 t of propellant (including residuals) and the descent stage would require approx. 21 t of propellant. The addition of a 12 m diameter supersonic inflatable aerodynamic decelerator (SIAD), based on a proven flight design, was studied as an optional method to improve the ULM fraction, reducing the required descent propellant by approx.4 t.

Mars lander↗
NASA NTRSDate not supplied

Mixing and Recovery Performance of Triplet Impinging Injectors with Varying Flow Areas for Detonative Combustion

Detonation combustion devices such as rotating detonation rocket engines (RDREs) have been proposed as an alternative to traditional deflagration engines as detonative combustion has a higher theoretical thermodynamic efficiency. Improved performance of an RDRE depends heavily on the propellant injection and mixing. Impingement type injectors have been shown to achieve higher characteristic velocity or combustion efficiency for both detonative and deflagrative engines. This work is exploratory in nature and focuses on unlike triplet impinging injectors to study how the flow area of the injector affects the mixing of propellants and recovery under the transient, high-peak pressure conditions unique to RDREs. A ratio of fuel flow area to oxidizer flow area is set to achieve a given mass flow rate and thus propellant mixture ratio. Four different fluid flow areas are studies while keeping the fuel/oxidizer ratio constant. This work has found that, while maintaining a fixed ratio fuel to oxidizer flow area, the injectors with smaller orifice areas show better resistance to backflow in steady state tests and have shorter recovery times in backflow experiments. Even with the shorter recovery times, it is found that the shortest recovery time is still an order of magnitude slower than the frequency of a single detonation wave traversing a 3” RDRE.

Michaela R. Hemming↗
NASA NTRS1993

Computational analysis of Variable Thrust Engine (VTE) performance

The Variable Thrust Engine (VTE) of the Orbital Maneuvering Vehicle (OMV) uses a hypergolic propellant combination of Monomethyl Hydrazine (MMH) and Nitrogen Tetroxide (NTO) as fuel and oxidizer, respectively. The performance of the VTE depends on a number of complex interacting phenomena such as atomization, spray dynamics, vaporization, turbulent mixing, convective/radiative heat transfer, and hypergolic combustion. This study involved the development of a comprehensive numerical methodology to facilitate detailed analysis of the VTE. An existing Computational Fluid Dynamics (CFD) code was extensively modified to include the following models: a two-liquid, two-phase Eulerian-Lagrangian spray model; a chemical equilibrium model; and a discrete ordinate radiation heat transfer model. The modified code was used to conduct a series of simulations to assess the effects of various physical phenomena and boundary conditions on the VTE performance. The details of the models and the results of the simulations are presented.

Giridharan, M. G.↗
NASA NTRS1999

The Effect of Air Preheat at Atmospheric Pressure on the Formation of NO(x) in the Quick-Mix Sections of an Axially Staged Combustor

The Rich-burn/Quick-mix/Lean-burn (RQL) combustor concept has been proposed to minimize the formation of nitrogen oxides (NO(x)) in gas turbine systems. The success of this combustor strategy is dependent upon the efficiency of the mixing section bridging the fuel-rich and fuel-lean stages. Note that although these results were obtained from an experiment designed to study an RQL mixer, the link between mixing and NOx signatures is considerably broader than this application, in that the need to understand this link exists in most advanced combustors. The experiment reported herein was designed to study the effects of inlet air temperature on NO(x) formation in a mixing section. The results indicate that NO(x) emission is increased for all preheated cases compared to non-preheated cases. When comparing the various mixing modules, the affect of jet penetration is important, as this determines where NO(x) concentrations peak, and affects overall NO(x) production. Although jet air comprises 70 percent of the total airflow, the impact that jet air preheat has on overall NO(x) emissions is small compared to preheating both main and jet air flow.

Vardakas, M. A.↗
NASA NTRS2006

NASA Tech Briefs, November 2006

Topics include: Simulator for Testing Spacecraft Separation Devices; Apparatus for Hot Impact Testing of Material Specimens; Instrument for Aircraft-Icing and Cloud-Physics Measurements; Advances in Measurement of Skin Friction in Airflow; Improved Apparatus for Testing Monoball Bearings; High-Speed Laser Scanner Maps a Surface in Three Dimensions; Electro-Optical Imaging Fourier-Transform Spectrometer; Infrared Instrument for Detecting Hydrogen Fires; Modified Coaxial Probe Feeds for Layered Antennas; Detecting Negative Obstacles by Use of Radar; Cryogenic Pound Circuits for Cryogenic Sapphire Oscillators; PixelLearn; New Software for Predicting Charging of Spacecraft; Conversion Between Osculating and Mean Orbital Elements; Generating a 2D Representation of a Complex Data Structure; Making Activated Carbon by Wet Pressurized Pyrolysis; Composite Solid Electrolyte Containing Li+- Conducting Fibers; Electrically Conductive Anodized Aluminum Surfaces; Rapid-Chill Cryogenic Coaxial Direct-Acting Solenoid Valve; Variable-Tension-Cord Suspension/Vibration- Isolation System; Techniques for Connecting Superconducting Thin Films; Versatile Friction Stir Welding/Friction Plug Welding System; Thermal Spore Exposure Vessels; Enumerating Spore-Forming Bacteria Airborne with Particles; Miniature Oxidizer Ionizer for a Fuel Cell; Miniature Ion-Array Spectrometer; Promoted-Combustion Chamber with Induction Heating Coil; Miniature Ion-Mobility Spectrometer; Mixed-Salt/Ester Electrolytes for Low-Temperature Li+ Cells; Miniature Free-Space Electrostatic Ion Thrusters; Miniature Bipolar Electrostatic Ion Thruster; Holographic Plossl Retroreflectors; Miniature Electrostatic Ion Thruster With Magnet; Using Apex To Construct CPM-GOMS Models; Sequence Detection for PPM Optical Communication With ISI; Algorithm for Rapid Searching Among Star-Catalog Entries; Expectation-Based Control of Noise and Chaos; Radio Heating of Lunar Soil to Release Gases; Using Electrostriction to Manipulate Ullage in Microgravity; Equations for Scoring Rules When Data Are Missing; Insulating Material for Next-Generation Spacecraft; and Pseudorandom Switching for Adding Radar to the AFF Sensor.

Source record↗
NASA NTRS1986

The flame structure and vorticity generated by a chemically reacting transverse jet

An analytical model describing the behavior of a turbulent fuel jet injected normally into a cross flow is developed. The model places particular emphasis on the contrarotating vortex pair associated with the jet, and predicts the flame length and shape based on entrainment of the oxidizer by the fuel jet. Effects of buoyancy and density variations in the flame are neglected in order to isolate the effects of large-scale mixing. The results are compared with a simulation of the transverse reacting jet in a liquid (acid-base) system. For a wide range of ratios of the cross flow to jet velocity, the model predicts flame length quite well. In particular, the observed transitional behavior in the flame length between cross-flow velocity to jet velocity of orifice ratios of 0.0 to 0.1, yielding an approximate minimum at the ratio 0.05, is reproduced very clearly by the present model. The transformation in flow structure that accounts for this minimum arises from the differing components of vorticity dominant in the near-field and far-field regions of the jet.

Karagozian, A. R.↗
NASA NTRS1977

Start up system for hydrogen generator used with an internal combustion engine

A hydrogen generator provides hydrogen rich product gases which are mixed with the fuel being supplied to an internal combustion engine for the purpose of enabling a very lean mixture of that fuel to be used, whereby nitrous oxides emitted by the engine are minimized. The hydrogen generator contains a catalyst which must be heated to a pre-determined temperature before it can react properly. To simplify the process of heating up the catalyst at start-up time, either some of the energy produced by the engine such as engine exhaust gas, or electrical energy produced by the engine, or the engine exhaust gas may be used to heat up air which is then used to heat the catalyst.

Houseman, J.↗
NASA NTRS2014

The Enhanced Injection and Mixing Project at NASA Langley

An overview of a research project aimed at enhancing fuel injection and mixing for scramjet engines at flight Mach numbers greater than 8 is presented. The specific objectives of this project are to increase knowledge and understanding of the fundamental physics governing scramjet fuel-air mixing, to develop strategies for improving injector performance and to develop functional relationships between mixing efficiency, losses due to mixing (such as total pressure loss, drag) and combustor/mixing length. The research approach consists of a closely coupled experimental and computational effort with experiments intended for exploring injection concepts and for anchoring computational fluid dynamics (CFD) simulations. The experiments are being conducted in the NASA Langley Arc-Heated Scramjet Test Facility and consist of helium injection into a Mach 6 air stream. Measurements include in-stream helium mole fraction, Pitot pressure, and total temperature as well as wall pressures and temperatures. Additionally, nitric oxide planar laser-induced fluorescence (NO PLIF) imaging will be used for flow visualization of the fuel plume. The numerical simulations, performed with the VULCAN CFD code, provide flowfield details unobtainable from the experiment. The current paper describes the experimental apparatus and diagnostic systems which have been developed and/or tested in preparation for the mixing tests and discusses numerical simulations performed to reduce the risk associated with the experimental design.

K F Cabell↗
NASA NTRS1992

A second-generation high speed civil transport: Stingray

The Stingray is the second-generation High Speed Civil Transport (HSCT) designed for the 21st Century. This aircraft is designed to be economically viable and environmentally sound transportation competitive in markets currently dominated by subsonic aircraft such as the Boeing 747 and upcoming McDonnell Douglas MD-12. With the Stringray coming into service in 2005, a ticket price of 21 percent over current subsonic airlines will cover operational costs with a 10 percent return on investment. The cost per aircraft will be $202 million with the Direct Operating Cost equal to $0.072 per mile per seat. This aircraft has been designed to be a realistic aircraft that can be built within the next ten to fifteen years. There was only one main technological improvement factor used in the design, that being for the engine specific fuel consumption. The Stingray, therefore, does not rely on technology that does not exist. The Stingray will be powered by four mixed flow turbofans that meet both nitrous oxide emissions and FAR 36 Stage 3 noise regulations. It will carry 250 passengers a distance of 5200 nautical miles at a speed of Mach 2.4. The shape of the Stingray, while optimized for supersonic flight, is compatible with all current airline facilities in airports around the world. As the demand for economical, high-speed flight increases, the Stingray will be ready and able to meet those demands.

Engdahl, Sean↗
NASA NTRS1993

In-stream measurements of combustion during Mach 5 to 7 tests of the Hypersonic Research Engine (HRE)

Results of in-stream combustion measurements taken during Mach 5 to 7 true simulation testing of the Hypersonic Research Engine/Aerothermodynamic Integration Model (HRE/AIM) are presented. These results, the instrumentation techniques, and configuration changes to the engine installation that were required to test this model are described. In test runs at facility Mach numbers of 5 to 7, an exhaust instrumentation ring which formed an extension of the engine exhaust nozzle shroud provided diagnostic measurements at 10 circumferential locations in the HRE combustor exit plane. The measurements included static and pitot pressures using conventional conical probes, combustion gas temperatures from cooled-gas pyrometer probes, and species concentration from analysis of combustion gas samples. Results showed considerable circumferential variation, indicating that efficiency losses were due to nonuniform fuel distribution or incomplete mixing. Results using the Mach 7 facility nozzle but with Mach 6 temperature simulation, 1590 to 1670 K, showed indications of incomplete combustion. Nitric oxide measurements at the combustor exit peaked at 2000 ppmv for stoichiometric combustion at Mach 6.

Lezberg, Erwin A.↗
NASA NTRS2007

High-Temperature Desulfurization of Heavy Fuel-Derived Reformate Gas Streams for SOFC Applications

Desulfurization of the hot reformate gas produced by catalytic partial oxidation or autothermal reforming of heavy fuels, such as JP-8 and jet fuels, is required prior to using the gas in a solid oxide fuel cell (SOFC). Development of suitable sorbent materials involves the identification of sorbents with favorable sulfidation equilibria, good kinetics, and high structural stability and regenerability at the SOFC operating temperatures (650 to 800 C). Over the last two decades, a major barrier to the development of regenerable desulfurization sorbents has been the gradual loss of sorbent performance in cyclic sulfidation and regeneration at such high temperatures. Mixed oxide compositions based on ceria were examined in this work as regenerable sorbents in simulated reformate gas mixtures and temperatures greater than 650 C. Regeneration was carried out with dilute oxygen streams. We have shown that under oxidative regeneration conditions, high regeneration space velocities (greater than 80,000 h(sup -1)) can be used to suppress sulfate formation and shorten the total time required for sorbent regeneration. A major finding of this work is that the surface of ceria and lanthanan sorbents can be sulfided and regenerated completely, independent of the underlying bulk sorbent. This is due to reversible adsorption of H2S on the surface of these sorbents even at temperatures as high as 800 C. La-rich cerium oxide formulations are excellent for application to regenerative H2S removal from reformate gas streams at 650 to 800 C. These results create new opportunities for compact sorber/regenerator reactor designs to meet the requirements of solid oxide fuel cell systems at any scale.

Flytzani-Stephanopoulos, Maria↗
NASA NTRS1977

Experimental clean combustor program, phase 3

A two-stage vortex burning and mixing combustor and associated fuel system components were successfully tested at steady state and transient operating conditions. The combustor exceeded the program goals for all three emissions species, with oxides of nitrogen 10 percent below the goal, carbon monoxide 26 percent below the goal, and total unburned hydrocarbons 75 percent below the goal. Relative to the JT9D-7 combustor, the oxides of nitrogen were reduced by 58 percent, carbon monoxide emissions were reduced by 69 percent, and total unburned hydrocarbons were reduced by 9 percent. The combustor efficiency and exit temperature profiles were comparable to those of production combustor. Acceleration and starting characteristics were deficient relative to the production engine.

Roberts, R.↗
NASA NTRS1982

A coalescence/dispersion model for turbulent flame stability

A stochastic mixing model is employed that couples finite rate mixing with finite rate combustion chemistry to examine the effects of combustor inlet conditions on the ignition limits of lean premixed turbulent flames. The kinetics of the fuel-oxidation process is described by means of a two-step reaction mechanism. The model predictions of the lean ignition limits of premixed turbulent propane-air flames are verified through comparisons with data in the literature. The variations of the lean limits are found to compare favorably with experimental magnitudes and trends.

Radhakrishnan, K.↗
NASA NTRS2000

Quantitative Measurements of Nitric Oxide Concentration in High-Pressure, Swirl-Stabilized Spray Flames

Lean direct-injection (LDI) spray flames offer the possibility of reducing NO(sub x) emissions from gas turbines by rapid mixing of the liquid fuel and air so as to drive the flame structure toward partially-premixed conditions. We consider the technical approaches required to utilize laser-induced fluorescence methods for quantitatively measuring NO concentrations in high-pressure LDI spray flames. In the progression from atmospheric to high-pressure measurements, the LIF method requires a shift from the saturated to the linear regime of fluorescence measurements. As such, we discuss quantitative, spatially resolved laser-saturated fluorescence (LSF), linear laser-induced fluorescence (LIF), and planar laser-induced fluorescence (PLIF) measurements of NO concentration in LDI spray flames. Spatially-resolved LIF measurements of NO concentration (ppm) are reported for preheated, LDI spray flames at pressures of two to five atmospheres. The spray is produced by a hollow-cone, pressure-atomized nozzle supplied with liquid heptane. NO is excited via the Q(sub 2)(26.5) transition of the gamma(0,0) band. Detection is performed in a two nanometer region centered on the gamma(0,1) band. A complete scheme is developed by which quantitative NO concentrations in high-pressure LDI spray flames can be measured by applying linear LIF. NO is doped into the reactants and convected through the flame with no apparent destruction, thus allowing a NO fluorescence calibration to be taken inside the flame environment. The in-situ calibration scheme is validated by comparisons to a reference flame. Quantitative NO profiles are presented and analyzed so as to better understand the operation of lean-direct injectors for gas turbine combustors. Moreover, parametric studies are provided for variations in pressure, air-preheat temperature, and equivalence ratio. Similar parametric studies are performed for lean, premixed-prevaporized flames to permit comparisons to those for LDI flames. Finally, PLIF is expanded to high pressure in an effort to quantify the detected fluorescence image for LDI flames. Success is achieved by correcting the PLIF calibration via a single-point LIF measurement. This procedure removes the influence of any preferential background that occurs in the PLIF detection window. In general, both the LIF and PLIF measurements verify that the LDI strategy could be used to reduce NO(sub x) emissions in future gas turbine combustors.

Cooper, Clayton S.↗
NASA NTRS1993

A High Performance H2-Cl2 Fuel Cell for Space Power Applications

NASA has numerous airborne/spaceborne applications for which high power and energy density power sources are needed. The proton exchange membrane fuel cell (PEMFC) is an attractive candidate for such a power source. PEMFC's offer many advantages for airborne/spaceborne applications. They have high power and energy densities, convert fuel to electrical power with high efficiency at both part and full load, and can rapidly startup and shutdown. In addition, PEMFC's are lightweight and operate silently. A significant impediment to the attainment of very high power and energy densities by PEMFC's is their current exclusive reliance on oxygen as the oxidant. Conventional PEMFC's oxidize hydrogen at the anode and reduce oxygen at the cathode. The electrode kinetics of oxygen reduction are known to be highly irreversible, incurring large overpotential losses. In addition, the modest open circuit potential of 1.2V for the H2-O2 fuel cell is unattainable due to mixed potential effects at the oxygen electrode. Because of the high overpotential losses, cells using H2 and O2 are capable of achieving high current densities only at very low cell voltages, greatly curtailing their power output. Based on experimental work on chlorine reduction in a gas diffusion electrode, we believe significant increases in both the energy and power densities of PEMFC systems can be achieved by employing chlorine as an alternative oxidant.

Anderson, Everett B.↗