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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 145 records · Page 8
All sourcesNASA NTRSDOE OSTI
NASA NTRS1994

Numerical Simulation of Shock-Induced Combustion Past Blunt Bodies Using Shock-Fitting Technique

Two-dimensional axisymmetric, reacting viscous flow over blunt projectiles is computed to study shock induced combustion at Mach 5.11 and Mach 6.46 in hydrogen-air mixture. A finite-difference, shock-fitting method is used to solve the complete set of Navier Stokes and species conservation equations. In this approach, the bow shock represents a boundary of the computational domain and is treated as a discontinuity across which Rankine-Hugoniot conditions are applied. All interior details of the flow such as compression waves, reaction front, and the wall boundary layer are captured automatically in the solution. Since shock-fitting approach reduces the amount of artificial dissipation, all the intricate details of the flow are captured much more clearly than has been possible with the shock-capturing approach. This has allowed an improved understanding of the physics of shock-induced combustion over blunt projectiles and the numerical results can now be explained more readily with one dimensional wave-interaction model than before.

Ahuja, J. K.↗
NASA NTRS1994

Numerical simulation of shock-induced combustion past blunt bodies using shock-fitting technique

Two-dimensional axisymmetric, reacting viscous flow over blunt projectiles is computed to study shock-induced combustion at Mach 5.11 and Mach 6.46 in hydrogen-air mixture. A finite-difference, shock-fitting method is used to solve the complete set of Navier-Stokes and species conservation equations. In this approach, the bow shock represents a boundary of the computational domain and is treated as a discontinuity across which Rankine-Hugoniot conditions are applied. All interior details of the flow such as compression waves, reaction front, and the wall boundary layer are captured automatically in the solution. Since shock-fitting approach reduces the amount of artificial dissipation, all the intricate details of the flow are captured much more clearly than has been possible with the shock-capturing approach. This has allowed an improved understanding of the physics of shock-induced combustion over blunt projectiles and the numerical results can now be explained more readily with one-dimensional wave-interaction model than before.

Ahuja, J. K.↗
NASA NTRS1990

Calculations of gaseous H2/O2 thruster

Calculations were made for the gaseous H2/O2 thruster for the Space Station by using the RPLUS code which employs an implicit finite volume, LUSSOR scheme to solve the Navier-Stokes equations and the species equations. The combustion processes of hydrogen and oxygen are modeled by a 9 species and 18 step reaction mechanism and the turbulence is simulated by the Baldwin-Lomax turbulence model for the thruster wall boundary layer and the modified Prandtl's mixing length model for the reacting shear layer. Results are presented for different mixture ratios and fuel film cooling percents and compared with the experimental data. The calculated performance predictions for the thruster agree well with the experimental data and the results demonstrate that the RPLUS code can be used for design and analysis of thrusters and rockets.

Kim, S. C.↗
NASA NTRS1992

Atmospheric chemistry in the Arctic and subarctic - Influence of natural fires, industrial emissions, and stratospheric inputs

Layers with enhanced concentrations of trace gases intercepted by the NASA Electra aircraft over Alaska during the Arctic Boundary Layer Expedition (ABLE 3A) in July-August 1988 are discussed. Haze layers apparently associated with boreal fires were enriched in hydrocarbons and NO(y), with emission factors corresponding closely to laboratory data for smoldering combustion. It is argued that atmospheric composition was strongly modified by wildfires during several periods of the ABLE 3A mission. The associated enhancement of NO(y) was smaller than observed for most other combustion processes but was nonetheless significant in the context of very low background concentrations. Ozone production in fire plumes was negligible. Ambient O3 was supplied by the stratosphere, with little direct input from midlatitude source during summer. It is argued that NO(y) was supplied about equally by the stratosphere and by wildfires. Hydrocarbons and CO appear to derive from biomass fires and from human activities.

Wofsy, S. C.↗
NASA NTRS2009

Integrated High Payoff Rocket Propulsion Technology (IHPRPT) SiC Recession Model

SiC stability and recession rates were modeled in hydrogen/oxygen combustion environments for the Integrated High Payoff Rocket Propulsion Technology (IHPRPT) program. The IHPRPT program is a government and industry program to improve U.S. rocket propulsion systems. Within this program SiC-based ceramic matrix composites are being considered for transpiration cooled injector faceplates or rocket engine thrust chamber liners. Material testing under conditions representative of these environments was conducted at the NASA Glenn Research Center, Cell 22. For the study described herein, SiC degradation was modeled under these Cell 22 test conditions for comparison to actual test results: molar mixture ratio, MR (O2:H2) = 6, material temperatures to 1700 C, combustion gas pressures between 0.34 and 2.10 atm, and gas velocities between 8,000 and 12,000 fps. Recession was calculated assuming rates were controlled by volatility of thermally grown silica limited by gas boundary layer transport. Assumptions for use of this model were explored, including the presence of silica on the SiC surface, laminar gas boundary layer limited volatility, and accuracy of thermochemical data for volatile Si-O-H species. Recession rates were calculated as a function of temperature. It was found that at 1700 C, the highest temperature considered, the calculated recession rates were negligible, about 200 m/h, relative to the expected lifetime of the material. Results compared favorably to testing observations. Other mechanisms contributing to SiC recession are briefly described including consumption of underlying carbon and pitting. A simple expression for liquid flow on the material surface was developed from a one-dimensional treatment of the Navier-Stokes Equation. This relationship is useful to determine under which conditions glassy coatings or thermally grown silica would flow on the material surface, removing protective layers by shear forces. The velocity of liquid flow was found to depend on the gas velocity, the viscosity of gas and liquid, as well as the thickness of the gas boundary layer and the liquid layer. Calculated flow rates of a borosilicate glass coating compared well to flow rates observed for this coating tested on a SiC panel in Cell 22.

Opila, E. J.↗
NASA NTRS1979

Theory of deposition of condensible impurities on surfaces immersed in combustion gases

The components resulting from the deposition of inorganic salts (e.g., Na2S04) and oxides present in the combustion products from gas turbine engines were investigated. Emphasis was placed on the effects of multicomponent vapor transport, thermophoretic transport of vapor and small particles to actively cooled surfaces, variable fluid properties within mass transfer boundary layers, and free stream turbulence.

Rosner, D. E.↗
NASA NTRSDate not supplied

Pretest Simulations of a Supersonic Mixing and Combustion Validation Experiment to Assess Sensitivities

The reliance on CFD simulations to develop, design, and optimize scramjet systems (or components) has become commonplace. This reliance inevitably hinges on the ability of the computational analyst to quantify the level of confidence in their computational results. Unfortunately, nearly all the measured data available for this assessment comes from antiquated experimental datasets, or from tests that focused on the extraction of scramjet system (or component) performance. The objective of a CFD validation experiment is to quantify the predictive accuracy of one or more of the CFD physics submodels, implying that other uncertainties related to replicating the facility flow environment (e.g., knowledge of boundary conditions) must be minimized to the extent possible. This inevitably places stringent requirements on the quality and quantity of measurements taken to accurately specify inflow, outflow, and surface conditions for the CFD simulations; in addition to the measurements taken for the validation of physics submodels. This places additional demands on the experimental process above and beyond those for test article performance assessment. A recent high speed code credibility workshop series sponsored by AFRL identified a gap in existing validation data for fundamental assessments of turbulent mixing and combustion CFD closure models at scramjet engine relevant conditions. To address this gap, engineers at AFRL have designed a coaxial jet flame configuration that will be tested at two facilities (Research Cell 19 at the Air Force Research Lab, and at Purdue University). The effort described here documents pretest simulations of this validation experiment with the goal of fleshing out the extent of the facility flowpath that must be included to adequately reproduce the facility test section flow environment. The findings indicate that the flow around the support structure for the fuel injection centerbody upstream of the facility nozzle generates disturbances that persist throughout the nozzle expansion process; corrupting the azimuthal symmetry that was desired in the fuel/air mixing region of the test section. Simulations without this support structure maintained a high degree of azimuthal symmetry up until the fuel injection plane. However, even in this scenario the azimuthal symmetry was not maintained once the centerbody boundary layer transitioned to a wake flow downstream of the fuel injection plane.

CFD↗
NASA NTRSDate not supplied

Pretest Simulations of a Supersonic Mixing and Combustion Validation Experiment to Assess Sensitivities

The reliance on CFD simulations to develop, design, and optimize scramjet systems (or components) has become commonplace. This reliance inevitably hinges on the ability of the computational analyst to quantify the level of confidence in their computational results. Unfortunately, nearly all the measured data available for this assessment comes from antiquated experimental datasets, or from tests that focused on the extraction of scramjet system (or component) performance. The objective of a CFD validation experiment is to quantify the predictive accuracy of one or more of the CFD physics submodels, implying that other uncertainties related to replicating the facility flow environment (e.g., knowledge of boundary conditions) must be minimized to the extent possible. This inevitably places stringent requirements on the quality and quantity of measurements taken to accurately specify inflow, outflow, and surface conditions for the CFD simulations; in addition to the measurements taken for the validation of physics submodels. This places additional demands on the experimental process above and beyond those for test article performance assessment. A recent high speed code credibility workshop series sponsored by AFRL identified a gap in existing validation data for fundamental assessments of turbulent mixing and combustion CFD closure models at scramjet engine relevant conditions. To address this gap, engineers at AFRL have designed a coaxial jet flame configuration that will be tested at two facilities (Research Cell 19 at the Air Force Research Lab, and at Purdue University). The effort described here documents pretest simulations of this validation experiment with the goal of fleshing out the extent of the facility flowpath that must be included to adequately reproduce the facility test section flow environment. The findings indicate that the flow around the support structure for the fuel injection centerbody upstream of the facility nozzle generates disturbances that persist throughout the nozzle expansion process; corrupting the azimuthal symmetry that was desired in the fuel/air mixing region of the test section. Simulations without this support structure maintained a high degree of azimuthal symmetry up until the fuel injection plane. However, even in this scenario the azimuthal symmetry was not maintained once the centerbody boundary layer transitioned to a wake flow downstream of the fuel injection plane.

CFD↗
NASA NTRS1986

Experimental verification of corrosive vapor deposition rate theory in high velocity burner rigs

The ability to predict deposition rates is required to facilitate modelling of high temperature corrosion by fused salt condensates in turbine engines. A corrosive salt vapor deposition theory based on multicomponent chemically frozen boundary layers (CFBL) has been successfully verified by high velocity burner rig experiments. The experiments involved internally air-impingement cooled, both rotating full and stationary segmented cylindrical collectors located in the crossflow of sodium-seeded combustion gases. Excellent agreement is found between the CFBL theory an the experimental measurements for both the absolute amounts of Na2SO4 deposition rates and the behavior of deposition rate with respect to collector temperature, mass flowrate (velocity) and Na concentration.

Gokoglu, S. A.↗
NASA NTRS1986

Experimental verification of corrosive vapor deposition rate theory in high velocity burner rigs

The ability to predict deposition rates is required to facilitate modelling of high temperature corrosion by fused salt condensates in turbine engines. A corrosive salt vapor deposition theory based on multicomponent chemically frozen boundary layers (CFBL) has been successfully verified by high velocity burner rig experiments. The experiments involved internally air-impingement cooled, both rotating full and stationary segmented cylindrical collectors located in the crossflow of sodium-seeded combustion gases. Excellent agreement is found between the CFBL theory and the experimental measurements for both the absolute amounts of Na2SO4 deposition rates and the behavior of deposition rate with respect to collector temperature, mass flowrate (velocity) and Na concentration.

Gokoglu, Suleyman A.↗
NASA NTRS1998

Annual Research Briefs, 1998

The topics contained in this progress report are direct numerical simulation of turbulent non-premixed combustion with realistic chemistry; LES of non-premixed turbulent reacting flows with conditional source term estimation; measurements of the three-dimensional scalar dissipation rate in gas-phase planar turbulent jets; direct simulation of a jet diffusion flame; on the use of interpolating wavelets in the direct numerical simulation of combustion; on the use of a dynamically adaptive wavelet collocation algorithm in DNS (direct numerical simulation) of non-premixed turbulent combustion; 2D simulations of Hall thrusters; computation of trailing-edge noise at low mach number using LES and acoustic analogy; weakly nonlinear modeling of the early stages of bypass transition; interactions between freestream turbulence and boundary layers; interfaces at the outer boundaries of turbulent motions; largest scales of turbulent wall flows; the instability of streaks in near-wall turbulence; an implementation of the v(sup 2) - f model with application to transonic flows; heat transfer predictions in cavities; a structure-based model with stropholysis effects; modeling a confined swirling coaxial jet; subgrid-scale models based on incremental unknowns for large eddy simulations; subgrid scale modeling taking the numerical error into consideration; towards a near-wall model for LES of a separated diffuser flow; on the feasibility of merging LES with RANS (Reynolds Averaging Numerical simulation) for the near-wall region of attached turbulent flows; large-eddy simulation of a separated boundary layer; numerical study of a channel flow with variable properties; on the construction of high order finite difference schemes on non-uniform meshes with good conservation properties; development of immersed boundary methods for complex geometries; and particle methods for micro and macroscale flow simulations.

Spinks, Debra↗
NASA NTRS1979

Autoignition of hydrogen injected transverse to a supersonic airstream

A parametric study of autoignition characteristics of transverse fuel injector patterns for a hydrogen-fueled, supersonic-combustion ramjet was conducted in direct-connect tests over a range of simulated hypersonic flight conditions. Autoignition limits, defined directly by visual observations and indirectly by measured wall pressure and temperature, are compared with a semi-empirical ignition limit model which accounts for first order effects on autoignition. It is shown that to initiate combustion in the immediate region of the fuel injector, ignition must occur in the separation region ahead of the fuel jet. Autoignition within this separation zone exhibits strong dependence on flow pressure and temperature, injector diameter, local boundary layer energy thickness, wall temperature, and on fluid dynamic interactions with adjacent downstream facing steps (such as inlet isolation steps). Secondary effects are also discussed. The empirical model is used to define limits of test temperature, pressure, and scale required for autoignition in a current subscale engine test program.

Mcclinton, C. R.↗
NASA NTRS2012

Status of Turbulence Modeling for Hypersonic Propulsion Flowpaths

This report provides an assessment of current turbulent flow calculation methods for hypersonic propulsion flowpaths, particularly the scramjet engine. Emphasis is placed on Reynolds-averaged Navier-Stokes (RANS) methods, but some discussion of newer meth- ods such as Large Eddy Simulation (LES) is also provided. The report is organized by considering technical issues throughout the scramjet-powered vehicle flowpath including laminar-to-turbulent boundary layer transition, shock wave / turbulent boundary layer interactions, scalar transport modeling (specifically the significance of turbulent Prandtl and Schmidt numbers) and compressible mixing. Unit problems are primarily used to conduct the assessment. In the combustor, results from calculations of a direct connect supersonic combustion experiment are also used to address the effects of turbulence model selection and in particular settings for the turbulent Prandtl and Schmidt numbers. It is concluded that RANS turbulence modeling shortfalls are still a major limitation to the accuracy of hypersonic propulsion simulations, whether considering individual components or an overall system. Newer methods such as LES-based techniques may be promising, but are not yet at a maturity to be used routinely by the hypersonic propulsion community. The need for fundamental experiments to provide data for turbulence model development and validation is discussed.

Georgiadis, Nicholas J.↗
NASA NTRS1984

Analytic modeling of a spray diffusion flame

A detailed model for a spray diffusion flame is described. The model is based on the boundary layer form of the equations of motion, with droplet transport accounted for using a discretized droplet size distribution function. Interphase transport of mass and energy are accounted for, with a flame-sheet model used to describe the combustion process on a droplet scale. Near dynamic equilibrium is assumed for the description of droplet transport; droplets can diffuse relative to the gas phase. Gas-phase mixing is accounted for using a two-equation turbulence model; buoyancy effects are included, with a temperature fluctuation equation used to account for buoyancy effects on turbulence structure. Thermal radiation from gas-phase CO2 and H2O is included. Gas-phase chemical kinetics are modeled using a 20-reaction, 10-species version of the advanced quasi-global chemical kinetics formulation. Results are compared with data for a vaporizing Freon spray and a pentane spray flame. It is shown that the computational approach provides a reasonably valid picture of the overall development of a spray diffusion flame, and, furthermore, provides a useful tool for the parametric examination of the spray combustion process.

Harsha, P. T.↗
NASA NTRS1985

Rates of elementary reactions - Measurement and applications

Techniques used for characterizing elementary chemical reaction kinetics are explored. Flash- or laser-photolysis (FP) involves producing reactive species on the psec time scale and monitoring the changes spectroscopically. In the discharge flow (DF) method, reactive species are produced continuously in a flow of an inert gas containing the reactants. FP avoids surface and transport effects, while DF allows several reactions to be studied in different regions of one flow. Transport and surface boundary layer models are defined for DF calculations and sample calculations are carried out to illustrate the difficulties inherent in theoretically defining the rate constants for elementary reactions. Applications of the models thus far derived in atmospheric science and combustion studies are discussed.

Kaufman, F.↗
NASA NTRS2004

A Resonant Pulse Detonation Actuator for High-Speed Boundary Layer Separation Control

A variety of different types of actuators have been previously investigated as flow control devices. Potential applications include the control of boundary layer separation in external flows, as well as jet engine inlet and diffuser flow control. The operating principles for such devices are typically based on either mechanical deflection of control surfaces (which include MEMS flap devices), mass injection (which includes combustion driven jet actuators), or through the use of synthetic jets (diaphragm devices which produce a pulsating jet with no net mass flow). This paper introduces some of the initial flow visualization work related to the development of a relatively new type of combustion-driven jet actuator that has been proposed based on a pulse detonation principle. The device is designed to utilize localized detonation of a premixed fuel (Hydrogen)-air mixture to periodically inject a jet of gas transversely into the primary flow. Initial testing with airflow successfully demonstrated resonant conditions within the range of acoustic frequencies expected for the design. Schlieren visualization of the pulsating air jet structure revealed axially symmetric vortex flow, along with the formation of shocks. Flow visualization of the first successful sustained oscillation condition is also demonstrated for one configuration of the current test section. Future testing will explore in more detail the onset of resonant combustion and the approach to conditions of sustained resonant detonation.

Beck, B. T.↗
NASA NTRS1982

Modeling parameter influences on MHD swirl combustion nozzle design

Attention is given to a research project which has the goal to develop a two-stage slagging gasifier-combustor in the form of a high-intensity combustor, taking into account a suitable aerodynamic design of the second stage nozzle which will prevent the separation of the boundary layer as the flow turns from axial to radial direction. The specific objectives of the present investigation are to test the effect of various second-stage nozzle geometries, flow rates, swirl number, and distribution in the first and second stages upon the corresponding flowfield in the second stage. Special emphasis is given to the avoidance of boundary layer separation as the flow turns from axial to radial direction into the MHD disk generator.

Lilley, D. G.↗
NASA NTRS1999

Measurements of Acidic Gases and Aerosol Species Aboard the NASA DC-8 Aircraft During the Pacific Exploratory Mission in the Tropics (PEM-Tropics A)

We received funding to provide measurements of nitric acid (HNO3), formic acid (HCOOH), acetic acid (CH3COOH), and the chemical composition of aerosols aboard the NASA Ames DC-8 research aircraft during the PEM-Tropics A mission. These measurements were successfully completed and the final data resides in the electronic archive (ftp-gte.larc.nasa.gov) at NASA Langley Research Center. For the PEM-Tropics A mission the University of New Hampshire group was first author of four different manuscripts. Three of these have now appeared in the Journal of Geophysical Research-Atmospheres, included in the two section sections on PEM-Tropics A. The fourth manuscript has just recently been submitted to this same journal as a stand alone paper. All four of these papers are included in this report. The first paper (Influence of biomass combustion emissions on the distribution of acidic trace gases over the Southern Pacific basin during austral springtime) describes the large-scale distributions of HNO3, HCOOH, and CH3COOH. Arguments were presented to show, particularly in the middle tropospheric region, that biomass burning emissions from South America and Africa were a major source of acidic gases over the South Pacific basin. The second paper (Aerosol chemical composition and distribution during the Pacific Exploratory Mission (PEM) Tropics) covers the aerosol aspects of our measurement package. Compared to acidic gases, O3, and selected hydrocarbons, the aerosol chemistry showed little influence from biomass burning emissions. The data collected in the marine boundary layer showed a possible marine source of NH3 to the troposphere in equatorial areas. This source had been speculated on previously, but our data was the first collected from an airborne platform to show its large-scale features. The third paper (Constraints on the age and dilution of Pacific Exploratory Mission-Tropics biomass burning plumes from the natural radionuclide tracer Pb-210) utilized the unexpectedly high concentrations of Pb-210 in the combustion plumes to estimate their ages and mixing along the transport route to the South Pacific basin. The final paper in the group (Tropospheric reactive-odd nitrogen over the South Pacific in austral springtime). This paper provides a summary of reactive nitrogen during PEM-Tropics A, with HNO3 and PAM showing the most impact from combustion emission.

Talbot, Robert W.↗