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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 1,117 records · Page 62
All sourcesNASA NTRSDOE OSTI
NASA NTRS1991

Constitutive and damage material modeling in a high pressure hydrogen environment

Numerous components in reusable space propulsion systems such as the SSME are exposed to high pressure gaseous hydrogen environments. Flow areas and passages in the fuel turbopump, fuel and oxidizer preburners, main combustion chamber, and injector assembly contain high pressure hydrogen either high in purity or as hydrogen rich steam. Accurate constitutive and damage material models applicable to high pressure hydrogen environments are therefore needed for engine design and analysis. Existing constitutive and cyclic crack initiation models were evaluated only for conditions of oxidizing environments. The main objective is to evaluate these models for applicability to high pressure hydrogen environments.

Russell, D. A.↗
NASA NTRSDate not supplied

Validated Reduced Computational Methods for Realistic RDC Injection Modeling

This paper outlines a simplified computational method without chemistry for single and multi-injector performance for rotating detonation combustion (RDC) environment as well as relevant performance metrics. For non-reactive cases, an unsteady pressure profile from a Method of Characteristics (MOC) simulation was used to simulate an RDC combustion wave travelling around the annulus. In reactive cases, a pre-mixed stoichiometric mixture of hydrogen and air is injected into a 2D unwrapped combustion chamber with a specified injector area. The performance is evaluated based on several performance metrics including unsteady discharge coefficient, average mass flux coming into the combustor, and percentage time obstructed. Finally, the computational method was validated through an experimental RDC test setup with a single element injector in a hydrogen-air detonation driver environment.

Propulsion↗
NASA NTRS1991

User's manual for rocket combustor interactive design (ROCCID) and analysis computer program. Volume 2: Appendixes A-K

The appendices A-K to the user's manual for the rocket combustor interactive design (ROCCID) computer program are presented. This includes installation instructions, flow charts, subroutine model documentation, and sample output files. The ROCCID program, written in Fortran 77, provides a standardized methodology using state of the art codes and procedures for the analysis of a liquid rocket engine combustor's steady state combustion performance and combustion stability. The ROCCID is currently capable of analyzing mixed element injector patterns containing impinging like doublet or unlike triplet, showerhead, shear coaxial and swirl coaxial elements as long as only one element type exists in each injector core, baffle, or barrier zone. Real propellant properties of oxygen, hydrogen, methane, propane, and RP-1 are included in ROCCID. The properties of other propellants can be easily added. The analysis models in ROCCID can account for the influences of acoustic cavities, helmholtz resonators, and radial thrust chamber baffles on combustion stability. ROCCID also contains the logic to interactively create a combustor design which meets input performance and stability goals. A preliminary design results from the application of historical correlations to the input design requirements. The steady state performance and combustion stability of this design is evaluated using the analysis models, and ROCCID guides the user as to the design changes required to satisfy the user's performance and stability goals, including the design of stability aids. Output from ROCCID includes a formatted input file for the standardized JANNAF engine performance prediction procedure.

Muss, J. A.↗
NASA NTRS2019

An Analysis of the Strayton Engine, a Brayton and Stirling Cycle Recuperating Engine

This paper explores the novel Strayton engine concept. This engine combines the cycles of a Brayton engine with that of a Stirling engine to create a highly efficient recuperating gas turbine engine. In the explored case, both Brayton cycle and Stirling cycle engines are used to generate electrical power. Additionally, the Stirling engine is used to draw heat out of the Brayton turbine (acting to cool the turbine blades), while also pumping heat into Brayton cycle just before combustion occurs (acting as the mechanism for recuperation). The purpose of this paper is to detail the system level modeling techniques used to generate the simulation, perform a cycle analysis of the combined cycle engine, identify key technologies and challenges associated with the concept, and compare potential performance gains with existing gas turbine engines and internal combustion engines. Topics such as controls, blade cooling effects, engine weight, and heat transfer using heat pipe are also explored. Results from this work show potential architectures that could provide the required heat transfer rates, potential control strategies, and performance benefits, including efficiency gains between 10% and 3% on engines ranging from 200HP to 670HP with the combined cycle engine.

Chapman, Jeffryes W.↗
NASA NTRS2019

An Analysis of the Strayton Engine, a Brayton and Stirling Cycle Recuperating Engine

This paper explores the novel Strayton engine concept. This engine combines the cycles of a Brayton engine with that of a Stirling engine to create a highly efficient recuperating gas turbine engine. In the explored case, both Brayton cycle and Stirling cycle engines are used to generate electrical power. Additionally, the Stirling engine is used to draw heat out of the Brayton turbine (acting to cool the turbine blades), while also pumping heat into Brayton cycle just before combustion occurs (acting as the mechanism for recuperation). The purpose of this paper is to detail the system level modeling techniques used to generate the simulation, perform a cycle analysis of the combined cycle engine, identify key technologies and challenges associated with the concept, and compare potential performance gains with existing gas turbine engines and internal combustion engines. Topics such as controls, blade cooling effects, engine weight, and heat transfer using heat pipe are also explored. Results from this work show potential architectures that could provide the required heat transfer rates, potential control strategies, and performance benefits, including efficiency gains between 10% and 3% on engines ranging from 200HP to 670HP with the combined cycle engine.

Gas turbine engine↗
NASA NTRS1977

Analytical study of laser supported combustion waves in hydrogen

A one-dimensional energy equation, with constant pressure and area, was used to model the LSC wave. This equation balances convection, conduction, laser energy absorption, radiation energy loss and radiation energy transport. Solutions of this energy equation were obtained to give profiles of temperature and other properties, as well as the relation between laser intensity and mass flux through the wave. The flow through the LSC wave was then conducted through a variable pressure, variable area streamtube to accelerate it to high speed, with the propulsion application in mind. A numerical method for coupling the LSC wave model to the streamtube flow was developed, and a sample calculation was performed. The result shows that 42% of the laser power has been radiated away by the time the gas reaches the throat. It was concluded that in the radially confined flows of interest for propulsion applications, transverse velocities would be less important than in the unconfined flows where air experiments have been conducted.

Kemp, N. H.↗
NASA NTRSDate not supplied

Application of Artificial Intelligence for Surface PM2.5 Estimations from Geostationary Satellite and Atmospheric Numerical Model Data

PM2.5, particulate matter (PM) with a diameter less than or equal to 2.5 μm, is emitted from anthropogenic fuel combustion and forest fires. Due to their small size, PM2.5 can penetrate into respiratory systems and cause or exacerbate serious illness. The US Environmental Protection Agency (EPA) regulates the levels of surface PM2.5 but surface monitoring has spatial and temporal limitations. The Aerosol Optical Depth (AOD) retrievals from the Geostationary Operational Environmental Satellite (GOES) missions and meteorological factors can be utilized as an alternative technique to estimate surface PM2.5 levels at a higher spatial and temporal resolution compared to surface monitors. Traditional estimation approaches rely on linear regression techniques and have limitations modeling the nonlinear relationship between the meteorological factors, AOD retrievals, and surface PM2.5. We compare different machine learning techniques and identify the best-suited model that can represent the nonlinearity between the factors affecting PM2.5 levels

Manisha Khatri↗
NASA NTRS1993

Intake flow modeling in a four stroke diesel using KIVA3

Intake flow for a dual intake valved diesel engine is modeled using moving valves and realistic geometries. The objectives are to obtain accurate initial conditions for combustion calculations and to provide a tool for studying intake processes. Global simulation parameters are compared with experimental results and show good agreement. The intake process shows a 30 percent difference in mass flows and average swirl in opposite directions across the two intake valves. The effect of the intake process on the flow field at the end of compression is examined. Modeling the intake flow results in swirl and turbulence characteristics that are quite different from those obtained by conventional methods in which compression stroke initial conditions are assumed.

Hessel, R. P.↗
NASA NTRS2010

Demonstration of a Fiber Optic Regression Probe

The capability to provide localized, real-time monitoring of material regression rates in various applications has the potential to provide a new stream of data for development testing of various components and systems, as well as serving as a monitoring tool in flight applications. These applications include, but are not limited to, the regression of a combusting solid fuel surface, the ablation of the throat in a chemical rocket or the heat shield of an aeroshell, and the monitoring of erosion in long-life plasma thrusters. The rate of regression in the first application is very fast, while the second and third are increasingly slower. A recent fundamental sensor development effort has led to a novel regression, erosion, and ablation sensor technology (REAST). The REAST sensor allows for measurement of real-time surface erosion rates at a discrete surface location. The sensor is optical, using two different, co-located fiber-optics to perform the regression measurement. The disparate optical transmission properties of the two fiber-optics makes it possible to measure the regression rate by monitoring the relative light attenuation through the fibers. As the fibers regress along with the parent material in which they are embedded, the relative light intensities through the two fibers changes, providing a measure of the regression rate. The optical nature of the system makes it relatively easy to use in a variety of harsh, high temperature environments, and it is also unaffected by the presence of electric and magnetic fields. In addition, the sensor could be used to perform optical spectroscopy on the light emitted by a process and collected by fibers, giving localized measurements of various properties. The capability to perform an in-situ measurement of material regression rates is useful in addressing a variety of physical issues in various applications. An in-situ measurement allows for real-time data regarding the erosion rates, providing a quick method for empirically anchoring any analysis geared towards lifetime qualification. Erosion rate data over an operating envelope could also be useful in the modeling detailed physical processes. The sensor has been embedded in many regressing media for the purposes of proof-of-concept testing. A gross demonstration of its capabilities was performed using a sanding wheel to remove layers of metal. A longer-term demonstration measurement involved the placement of the sensor in a brake pad, monitoring the removal of pad material associated with the normal wear-and-tear of driving. It was used to measure the regression rates of the combustable media in small model rocket motors and road flares. Finally, a test was performed using a sand blaster to remove small amounts of material at a time. This test was aimed at demonstrating the unit's present resolution, and is compared with laser profilometry data obtained simultaneously. At the lowest resolution levels, this unit should be useful in locally quantifying the erosion rates of the channel walls in plasma thrusters. .

Korman, Valentin↗
NASA NTRS1991

Analytical combustion/emissions research related to the NASA high-speed research program

Increasing the pressure and temperature of the engines of new generation supersonic airliners increases the emissions of nitrogen oxides to a level that would have an adverse impact on the Earth's protective ozone layer. In the process of implementing low emissions combustor technologies, NASA Lewis Research Center has pursued a combustion analysis program to guide combustor design processes, to identify potential concepts of greatest promise, and to optimize them at low cost, with short turn-around time. The approach is to upgrade and apply advanced computer programs for gas turbine applications. Efforts have been made to improve the code capabilities of modeling the physics. Test cases and experiments are used for code validation. To provide insight into the combustion process and combustor design, two-dimensional and three-dimensional codes such as KIVA-II and LeRC 3D have been used. These codes are operational and calculations have been performed to guide low emissions combustion experiments.

Nguyen, H. Lee↗
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 NTRS1972

Further analytical study of hybrid rocket combustion

Analytical studies of the transient and steady-state combustion processes in a hybrid rocket system are discussed. The particular system chosen consists of a gaseous oxidizer flowing within a tube of solid fuel, resulting in a heterogeneous combustion. Finite rate chemical kinetics with appropriate reaction mechanisms were incorporated in the model. A temperature dependent Arrhenius type fuel surface regression rate equation was chosen for the current study. The governing mathematical equations employed for the reacting gas phase and for the solid phase are the general, two-dimensional, time-dependent conservation equations in a cylindrical coordinate system. Keeping the simplifying assumptions to a minimum, these basic equations were programmed for numerical computation, using two implicit finite-difference schemes, the Lax-Wendroff scheme for the gas phase, and, the Crank-Nicolson scheme for the solid phase.

Hung, W. S. Y.↗
NASA NTRS1975

Predicted exhaust emissions from a methanol and jet fueled gas turbine combustor

A computer model of a gas turbine combustor has been used to predict the kinetic combustion and pollutant formation processes for methanol and simulated jet fuel. Use of the kinetic reaction mechanisms has also allowed a study of ignition delay and flammability limit of these two fuels. The NOX emissions for methanol were predicted to be from 69 to 92% lower than those for jet fuel at the same equivalence ratio which is in agreement with experimentally observed results. The high heat of vaporization of methanol lowers both the combustor inlet mixture temperatures and the final combustion temperatures. The lower combustion temperatures lead to low NOX emissions while the lower inlet mixture temperatures increase methanol's ignition delay. This increase in ignition delay dictates the lean flammability limit of methanol to be 0.8, while jet fuel is shown to combust at 0.4.

Adelman, H. G.↗
NASA NTRS1976

Numerical optimization of gasdynamic lasers

A sophisticated numerical optimization procedure has been coupled to a detailed theoretical model of the carbon dioxide gasdynamic laser (GDL) for the purpose of systematically evaluating the performance of these systems. The integrated model was used to predict the performance of a wide variety of combustion-driven gasdynamic lasers and was then used to maximize the theoretical available specific energy of these lasers. Up to six design variables and three nonlinear constraints were used in these optimizations, although the method is not limited to this number. Results are presented to show the performance of typical lasers before and after optimization. Improvements approaching a factor of two in the available specific energy are found, suggesting that numerical optimization techniques are valuable in improving the design of gasdynamic lasers.

Saunders, R. C., III↗
NASA NTRS1986

A comprehensive model to determine the effects of temperature and species fluctuations on reaction rates in turbulent reacting flows

A computationally-viable model describing the interaction between fluid-mechanical turbulence and finite-rate combustion reactions, principally in high-speed flows was developed. Chemical kinetic mechanisms, complete and global, were developed describing the finite rate reaction of fuels of interest to NASA. These fuels included principally hydrogen and silane, although a limited amount of work involved hydrocarbon fuels as well.

Chinitz, W.↗
NASA NTRS1989

Radiation Effects on Flow Characteristics in Combustion Chambers

A JANNAF sponsored workshop was held to discuss the importance and role of radiative heat transfer in rocket combustion chambers. The potential impact of radiative transfer on hardware design, reliability, and performance was discussed. The current state of radiative transfer prediction capability in CFD modeling was reviewed and concluded to be substantially lacking in both the physical models used and the radiative property data available. There is a clear need to begin to establish a data base for making radiation calculations in rocket combustion chambers. A natural starting point for this effort would be the NASA thermochemical equilibrium code (CEC).

Brewster, M. Q.↗
NASA NTRS1957

Free-jet Tests of a 1.1-inch-diameter Supersonic Ram-jet Engine

Results are presented of free-jet tests of a 1.1-inch-diameter hydrogen-burning ram-jet engine over a Mach number range from 1.42 to 2.28 and a Reynolds number range from 6.01 times 10 to the 6th power to 15.78 times 10 to the 6th power. Tests demonstrated the reliability and wide operating range of the engine and showed its suitability for installation on wind-tunnel models of airplane and missile configurations. A comparison of engine operation with combustion-chamber lengths of 3.33 and 1.51 engine diameters was made at a Mach number of 2.06. A maximum test thrust coefficient of 0.905 was obtained at fuel-air ratio of 0.034 and a Mach number of 2.06 for the engine with the 3.33-diameter combustion chamber.

ENGINES, RAM-JET↗
NASA NTRS1994

Computer program for calculation of complex chemical equilibrium compositions and applications. Part 1: Analysis

This report presents the latest in a number of versions of chemical equilibrium and applications programs developed at the NASA Lewis Research Center over more than 40 years. These programs have changed over the years to include additional features and improved calculation techniques and to take advantage of constantly improving computer capabilities. The minimization-of-free-energy approach to chemical equilibrium calculations has been used in all versions of the program since 1967. The two principal purposes of this report are presented in two parts. The first purpose, which is accomplished here in part 1, is to present in detail a number of topics of general interest in complex equilibrium calculations. These topics include mathematical analyses and techniques for obtaining chemical equilibrium; formulas for obtaining thermodynamic and transport mixture properties and thermodynamic derivatives; criteria for inclusion of condensed phases; calculations at a triple point; inclusion of ionized species; and various applications, such as constant-pressure or constant-volume combustion, rocket performance based on either a finite- or infinite-chamber-area model, shock wave calculations, and Chapman-Jouguet detonations. The second purpose of this report, to facilitate the use of the computer code, is accomplished in part 2, entitled 'Users Manual and Program Description'. Various aspects of the computer code are discussed, and a number of examples are given to illustrate its versatility.

Gordon, Sanford↗