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Results for “Combustion simulation”

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 55 records · Page 3
All sourcesNASA NTRSDOE OSTI
NASA NTRS1992

Multi-dimensional high order essentially non-oscillatory finite difference methods in generalized coordinates

The nonlinear stability of compact schemes for shock calculations is investigated. In recent years compact schemes were used in various numerical simulations including direct numerical simulation of turbulence. However to apply them to problems containing shocks, one has to resolve the problem of spurious numerical oscillation and nonlinear instability. A framework to apply nonlinear limiting to a local mean is introduced. The resulting scheme can be proven total variation (1D) or maximum norm (multi D) stable and produces nice numerical results in the test cases. The result is summarized in the preprint entitled 'Nonlinearly Stable Compact Schemes for Shock Calculations', which was submitted to SIAM Journal on Numerical Analysis. Research was continued on issues related to two and three dimensional essentially non-oscillatory (ENO) schemes. The main research topics include: parallel implementation of ENO schemes on Connection Machines; boundary conditions; shock interaction with hydrogen bubbles, a preparation for the full combustion simulation; and direct numerical simulation of compressible sheared turbulence.

Shu, Chi-Wang↗
NASA NTRS1972

The simulation of high pressure hydrogen/oxygen rocket engines.

Initial experiments have established the feasibility of using a backward running detonation wave in a shock tube to produce the correct enthalpy and chemistry of the simulated combustion chamber gases. This technique produces plume Reynolds numbers higher than those achievable by any other known technique. The practical pressure limitations are being explored and detailed surveys of both single and multiple engine plumes and impingement data on a cylindrical body are being obtained. Facility details to allow safe handling of the high pressure hydrogen and oxygen up to a design chamber pressure of 20,000 psi are described.

Hopkins, H.↗
NASA NTRS1992

Operating characteristics at Mach 4 of an inlet having forward-swept, sidewall-compression surfaces

A three-dimensional, forward-swept sidewall-compression inlet has been tested in the Mach 4 Blowdown Facility at the NASA Langley Research Center to examine parameters that affect inlet starting, operability, and performance. The inlet was designed to simultaneously provide good starting and mass capture characteristics through the combination of forward-swept sidewalls and the aft placement of the downstream cowl. Parametrics examined included top surface compression, geometric contraction ratio, sidewall compression angle, boundary-layer thickness on the vehicle undersurface, and cowl position. To simulate combustion effects on inlet operation, the inlet was mechanically back pressured to determine maximum achievable combustor-to-inlet pressure ratio prior to inlet unstart. Inlet operability and performance are shown in terms of inlet starting (pulsed and self-starting) and unstarting characteristics, inlet static pressure distributions, combustor back pressure limits, mass capture, and flow uniformity.

Hudgens, Julie A.↗
NASA NTRS1993

An experimental study of the effects of bodyside compression on forward swept sidewall compression inlets ingesting a turbulent boundary layer

Forward swept sidewall compression inlets have been tested in the Mach 4 Blowdown Facility at the NASA Langley Research Center to study the effects of bodyside compression surfaces on inlet performance in the presence of an incoming turbulent boundary layer. The measurements include mass flow capture and mean surface pressure distributions obtained during simulated combustion pressure increases downstream of the inlet. The kerosene-lampblack surface tracer technique has been used to obtain patterns of the local wall shear stress direction. Inlet performance is evaluated using starting and unstarting characteristics, mass capture, mean surface pressure distributions and permissible back pressure limits. The results indicate that inlet performance can be improved with selected bodyside compression surfaces placed between the inlet sidewalls.

Rodi, Patrick E.↗
NASA NTRS1959

Investigation on the use of a Freely Rotating Rotor at the Cowl Face of a Supersonic Conical Inlet to Reduce Inlet Flow Distortion

An investigation has been made on the use of a freely rotating rotor at the cowl face of a supersonic conical diffuser to determine its effectiveness in reducing inlet flow distortion and the penalty in terms of total-pressure loss imposed by such a device when distortions are negligible. Tests were made with a rotor having an inlet tip diameter of 2.18 inches and a ratio of hub radius to tip radius of 0.52, in conjunction with a conical inlet having a 25 deg semi-vertex cone angle, at a Mach number of 2.1 over an angle-of-attack range of 0 deg to 8 deg. A simplified analysis showing that a supersonic, freely rotating rotor with maximum solidity for noninterference between blades will operate in an undistorted flow with a total-pressure defect of 1 percent or less was experimentally verified. Overall total-pressure distortions of 0.1 to 0.4 and Mach number distortions of 0.4 to 1.4, obtained at 4 deg to 8 deg angle of attack, were reduced about 30 percent and 23 percent, respectively, because of the presence of the rotor, with no measurable total-pressure loss. The rotor increased the peak total-pressure recovery at the simulated combustion chamber 1 1/2 and 3 1/2 percent at 6 deg and 8 deg angles of attack, respectively. This increase is attributed to lower diffusion duct losses as a consequence of a more uniform flow created by the rotor.

Goldberg, Theodore J.↗
NASA NTRS2000

National Combustion Code: Parallel Implementation and Performance

The National Combustion Code (NCC) is being developed by an industry-government team for the design and analysis of combustion systems. CORSAIR-CCD is the current baseline reacting flow solver for NCC. This is a parallel, unstructured grid code which uses a distributed memory, message passing model for its parallel implementation. The focus of the present effort has been to improve the performance of the NCC flow solver to meet combustor designer requirements for model accuracy and analysis turnaround time. Improving the performance of this code contributes significantly to the overall reduction in time and cost of the combustor design cycle. This paper describes the parallel implementation of the NCC flow solver and summarizes its current parallel performance on an SGI Origin 2000. Earlier parallel performance results on an IBM SP-2 are also included. The performance improvements which have enabled a turnaround of less than 15 hours for a 1.3 million element fully reacting combustion simulation are described.

Quealy, A.↗
NASA NTRS2002

Hybrid Propulsion Testing at Marshall Space Flight Center

Hybrid propulsion testing involving eleven and twenty-four inch motors performed at the Marshall Space Flight Center (MSFC) from the early 1990's to the present are discussed. Topics covered include: Solid Propulsion Investigation Program, Joint NASA Industry Research and Development (JIRAD) program, Large Subscale Solid Rocket Combustion Simulator (LSSRCS), Hybrid Propulsion Demonstration Program (HPDP), Hybrid Propulsion for Launch Vehicle Booster (HPTLV), Peroxide Hybrid Upper Stage (PHUS) and Solid Fuel Torch (SFT).

Prince, Andrew S.↗
NASA NTRS2002

National Combustion Code Parallel Performance Enhancements

The National Combustion Code (NCC) is being developed by an industry-government team for the design and analysis of combustion systems. The unstructured grid, reacting flow code uses a distributed memory, message passing model for its parallel implementation. The focus of the present effort has been to improve the performance of the NCC code to meet combustor designer requirements for model accuracy and analysis turnaround time. Improving the performance of this code contributes significantly to the overall reduction in time and cost of the combustor design cycle. This report describes recent parallel processing modifications to NCC that have improved the parallel scalability of the code, enabling a two hour turnaround for a 1.3 million element fully reacting combustion simulation on an SGI Origin 2000.

Quealy, Angela↗
NASA NTRS1995

Affect of Brush Seals on Wave Rotor Performance Assessed

The NASA Lewis Research Center's experimental and theoretical research shows that wave rotor topping can significantly enhance gas turbine engine performance levels. Engine-specific fuel consumption and specific power are potentially enhanced by 15 and 20 percent, respectively, in small (e.g., 400 to 700 hp) and intermediate (e.g., 3000 to 5000 hp) turboshaft engines. Furthermore, there is potential for a 3- to 6-percent specific fuel consumption enhancement in large (e.g., 80,000 to 100,000 lbf) turbofan engines. This wave-rotor-enhanced engine performance is accomplished within current material-limited temperature constraints. The completed first phase of experimental testing involved a three-port wave rotor cycle in which medium total pressure inlet air was divided into two outlet streams, one of higher total pressure and one of lower total pressure. The experiment successfully provided the data needed to characterize viscous, partial admission, and leakage loss mechanisms. Statistical analysis indicated that wave rotor product efficiency decreases linearly with the rotor to end-wall gap, the square of the friction factor, and the square of the passage of nondimensional opening time. Brush seals were installed to further minimize rotor passage-to-cavity leakage. The graph shows the effect of brush seals on wave rotor product efficiency. For the second-phase experiment, which involves a four-port wave rotor cycle in which heat is added to the Brayton cycle in an external burner, a one-dimensional design/analysis code is used in conjunction with a wave rotor performance optimization scheme and a two-dimensional Navier-Stokes code. The purpose of the four-port experiment is to demonstrate and validate the numerically predicted four-port pressure ratio versus temperature ratio at pressures and temperatures lower than those that would be encountered in a future wave rotor/demonstrator engine test. Lewis and the Allison Engine Company are collaborating to investigate wave rotor integration in an existing turboshaft engine. Recent theoretical efforts include simulating wave rotor dynamics (e.g., startup and load-change transient analysis), modifying the one-dimensional wave rotor code to simulate combustion internal to the wave rotor, and developing an analytical wave rotor design/analysis tool based on macroscopic balances for parametric wave rotor/engine analysis.

Source record↗
NASA NTRS2012

Environmental Durability and Stress Rupture of EBC/CMCs

This research focuses on the strength and creep performance of SiC fiber-reinforced SiC ceramic matrix composite (CMC) environmental barrier coating (EBC) systems under complex simulated engine environments. Tensile-strength and stress-rupture testing was conducted to illustrate the material properties under isothermal and thermal gradient conditions. To determine material durability, further testing was conducted under exposure to thermal cycling, thermal gradients and simulated combustion environments. Emphasis is placed on experimental techniques as well as implementation of non-destructive evaluation, including modal acoustic emission and electrical resistivity monitoring, to characterize strength degradation and damage mechanisms. Currently, little is known about the behavior of EBC-CMCs under these conditions; consequently, this work will prove invaluable in the development of structural components for use in high temperature applications.

Appleby, Matthew↗
NASA NTRS2021

High Temperature Solid Particle Erosion in a Melt-Infiltrated SiC/SiC Ceramic Matrix Composite

Ceramic matrix composites (CMCs) are an enabling propulsion material system that offer weight benefits over current Ni-based superalloys, and have higher temperature capabilities that can reduce cooling requirements. Incorporating CMCs into the hot section of gas-turbine engines therefore leads to an increase in engine efficiency. While significant advancements have been made, challenges still remain for current and next-generation gas-turbines; particularly when operating in dust-laden or erosive environments. Solid particles entrained in the gas flow can impact engine hardware resulting in localized damage and material removal due to repeated, cumulative impacts. In this study, the erosion behavior of a melt-infiltrated (MI) silicon carbide fiber-reinforced silicon carbide (SiC/SiC) CMC is investigated at high temperature (1,200 °C) in a simulated combustion environment using 150 µm alumina particles as erodent. Particle impact velocities ranged from 100 to 200 m/s and the angle of impingement varied from 30° to 90°. Erosion testing was also performed on α-SiC to elucidate similarities and differences in the erosion response of the composite compared to that of a monolithic ceramic. Scanning electron microscopy (SEM) was used to study the post-erosion damage morphology and the governing mechanisms of material removal.

Ceramic matrix composites (CMCs), solid particle e↗
NASA NTRS2004

Carbon Dioxide Dispersion in the Combustion Integrated Rack Simulated Numerically

When discharged into an International Space Station (ISS) payload rack, a carbon dioxide (CO2) portable fire extinguisher (PFE) must extinguish a fire by decreasing the oxygen in the rack by 50 percent within 60 sec. The length of time needed for this oxygen reduction throughout the rack and the length of time that the CO2 concentration remains high enough to prevent the fire from reigniting is important when determining the effectiveness of the response and postfire procedures. Furthermore, in the absence of gravity, the local flow velocity can make the difference between a fire that spreads rapidly and one that self-extinguishes after ignition. A numerical simulation of the discharge of CO2 from PFE into the Combustion Integrated Rack (CIR) in microgravity was performed to obtain the local velocity and CO2 concentration. The complicated flow field around the PFE nozzle exits was modeled by sources of equivalent mass and momentum flux at a location downstream of the nozzle. The time for the concentration of CO2 to reach a level that would extinguish a fire anywhere in the rack was determined using the Fire Dynamics Simulator (FDS), a computational fluid dynamics code developed by the National Institute of Standards and Technology specifically to evaluate the development of a fire and smoke transport. The simulation shows that CO2, as well as any smoke and combustion gases produced by a fire, would be discharged into the ISS cabin through the resource utility panel at the bottom of the rack. These simulations will be validated by comparing the results with velocity and CO2 concentration measurements obtained during the fire suppression system verification tests conducted on the CIR in March 2003. Once these numerical simulations are validated, portions of the ISS labs and living areas will be modeled to determine the local flow conditions before, during, and after a fire event. These simulations can yield specific information about how long it takes for smoke and combustion gases produced by a fire to reach a detector location, how large the fire would be when the detector alarms, and the behavior of the fire until it has been extinguished. This new capability could then be used to optimize the location of fire detectors and fire-suppression ports as well as to evaluate the effectiveness of fire suppressants and response strategies. Numerical data collected from these simulations could also be used to develop a virtual reality fire event for crew training and fire safety awareness. This work is funded by NASA's Bioastronautics Initiative, which has the objective of ensuring and enhancing the health, safety, and performance of humans in space. As part of this initiative, the Microgravity Combustion Science Branch at the NASA Glenn Research Center is conducting spacecraft fire safety research to significantly improve fire safety on inhabited spacecraft.

Wu, Ming-Shin↗
NASA NTRS2002

Numerical Simulation of the Combustion of Fuel Droplets: Finite Rate Kinetics and Flame Zone Grid Adaptation (CEFD)

The NASA Nebraska Space Grant (NSGC) & EPSCoR programs have continued their effort to support outstanding research endeavors by funding the Numerical Simulation of the Combustion of Fuel Droplets study at the University of Nebraska at Lincoln (UNL). This team of researchers has developed a transient numerical model to study the combustion of suspended and moving droplets. The engines that propel missiles, jets, and many other devices are dependent upon combustion. Therefore, data concerning the combustion of fuel droplets is of immediate relevance to aviation and aeronautical personnel, especially those involved in flight operations. The experiments being conducted by Dr. Gogos and Dr. Nayagam s research teams, allow investigators to gather data for comparison with theoretical predictions of burning rates, flame structures, and extinction conditions. The consequent improved hndamental understanding droplet combustion may contribute to the clean and safe utilization of fossil hels (Williams, Dryer, Haggard & Nayagam, 1997, 72). The present state of knowledge on convective extinction of he1 droplets derives fiom experiments conducted under normal gravity conditions. However, any data obtained with suspended droplets under normal gravity are grossly affected by gravity. The need to obtain experimental data under microgravity conditions is therefore well justified and addresses one of the goals of NASA s Human Exploration and Development of Space (HEDS) microgravity combustion experiment.

Gogos, George↗
NASA NTRS2000

Modeling of Nonacoustic Combustion Instability in Simulations of Hybrid Motor Tests

A transient model of a hybrid motor was formulated to study the cause and elimination of nonacoustic combustion instability. The transient model was used to simulate four key tests out of a series of seventeen hybrid motor tests conducted by Thiokol, Rocketdyne, and Martin Marietta at NASA Marshall Space Flight Center (MSFC). These tests were performed under the Hybrid Propulsion Technology for Launch Vehicle Boosters (HPTLVB) program. The first test resulted in stable combustion. The second test resulted in large-amplitude, 6.5-Hz chamber pressure oscillations that gradually damped away by the end of the test. The third test resulted in large-amplitude, 7.5-Hz chamber pressure oscillations that were sustained throughout the test. The seventh test resulted in elimination of combustion instability with the installation of an orifice immediately upstream of the injector. Formulation and implementation of the model are the scope of this presentation. The current model is an independent continuation of modeling presented previously by joint Thiokol-Rocketdyne collaborators Boardman, Hawkins, Wassom. and Claflin. The previous model simulated an unstable independent research and development (IR&D) hybrid motor test performed by Thiokol. There was very good agreement between the model and test data. Like the previous model, the current model was developed using Matrix-x simulation software. However, tests performed at MSFC under the HPTLVB program were actually simulated. ln the current model, the hybrid motor, consisting of the liquid oxygen (lox) injector, the multiport solid fuel grain, and nozzle, was simulated. The lox feedsystem, consisting of the tank, venturi. valve, and feed lines, was also simulated in the model. All components of the hybrid motor and lox feedsystem are treated by a lumped-parameter approach. Agreement between the results of the transient model and actual test data was very good. This agreement between simulated and actual test data indicated that the combustion instability in the hybrid motor was due to two causes: 1. a lox feed system of insufficient stiffness, and 2. a lox injector with an impedance (it pressure drop that was too low to provide damping against the feed system oscillations. Also, it was discovered that testing with a new grain of solid fuel sustained the combustion instability. However, testing with a used grain of solid fuel caused the combustion instability to gradually decay.

Rocker, M.↗
NASA NTRS1990

Direct simulation of turbulent combustion

Understanding and modeling of turbulent combustion are key-problems in the computation of numerous practical systems. Because of the lack of analytical theories in this field and of the difficulty of performing precise experiments, direct simulation appears to be one of the most attractive tools to use in addressing this problem. The present work can be split into two parts: (1) Development and validation of a direct simulation method for turbulent combustion; (2) Applications of the method to premixed turbulent combustion problems. The goal of part 1 is to define and to test a numerical method for direct simulation of reacting flows. A high level of confidence should be attached to direct simulation results, and this can only be achieved through extensive validation tests. In part 2, direct simulation is used to address some of the many critical problems related to turbulent combustion. At the present time, I have limited this work to premixed combustion and considered only four basic issues: (1) The effect of pressure waves on flame propagation; (2) The interaction between flame fronts and vortices; (3) The influence of curvature on premixed flame fronts; and (4) The validation of flamelet models for premixed turbulent combustion.

Poinsot, T. J.↗
NASA NTRS1979

Simulation of liquid propellant rocket engine combustion instabilities

A simulation technique for studying the high frequency combustion instabilities of liquid propellant rocket engines has been developed and used to investigate various aspects of instability phenomena. Of importance was investigation of the significance of the method of coupling the combustion and the gas dynamics of the system. Two coupling processes were studied: linear response of the combustion process to pressure fluctuations, and the nature of the resulting instabilities; and nonlinear response of the combustion process to velocity fluctuations, and the nature of the resulting instabilities. For the combustion model studied, nonlinear (velocity) coupling was found to more closely characterize liquid propellant instabilities.

Ventrice, M. B.↗