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At least 73 records · Page 4
All sourcesNASA NTRSDOE OSTI
NASA NTRS2004

New Reduced Two-Time Step Method for Calculating Combustion and Emission Rates of Jet-A and Methane Fuel With and Without Water Injection

A simplified kinetic scheme for Jet-A, and methane fuels with water injection was developed to be used in numerical combustion codes, such as the National Combustor Code (NCC) or even simple FORTRAN codes that are being developed at Glenn. The two time step method is either an initial time averaged value (step one) or an instantaneous value (step two). The switch is based on the water concentration in moles/cc of 1x10(exp -20). The results presented here results in a correlation that gives the chemical kinetic time as two separate functions. This two step method is used as opposed to a one step time averaged method previously developed to determine the chemical kinetic time with increased accuracy. The first time averaged step is used at the initial times for smaller water concentrations. This gives the average chemical kinetic time as a function of initial overall fuel air ratio, initial water to fuel mass ratio, temperature, and pressure. The second instantaneous step, to be used with higher water concentrations, gives the chemical kinetic time as a function of instantaneous fuel and water mole concentration, pressure and temperature (T4). The simple correlations would then be compared to the turbulent mixing times to determine the limiting properties of the reaction. The NASA Glenn GLSENS kinetics code calculates the reaction rates and rate constants for each species in a kinetic scheme for finite kinetic rates. These reaction rates were then used to calculate the necessary chemical kinetic times. Chemical kinetic time equations for fuel, carbon monoxide and NOx were obtained for Jet-A fuel and methane with and without water injection to water mass loadings of 2/1 water to fuel. A similar correlation was also developed using data from NASA's Chemical Equilibrium Applications (CEA) code to determine the equilibrium concentrations of carbon monoxide and nitrogen oxide as functions of overall equivalence ratio, water to fuel mass ratio, pressure and temperature (T3). The temperature of the gas entering the turbine (T4) was also correlated as a function of the initial combustor temperature (T3), equivalence ratio, water to fuel mass ratio, and pressure.

Molnar, Melissa↗
NASA NTRS2000

Test Capabilities and Recent Experiences in the NASA Langley 8-Foot High Temperature Tunnel

The NASA Langley 8-Foot High Temperature Tunnel is a combustion-heated hypersonic blowdown-to-atmosphere wind tunnel that provides flight enthalpy simulation for Mach numbers of 4, 5, and 7 through an altitude range from 50,000 to 120,000 feet. The open-.jet test section is 8-ft. in diameter and 12-ft. long. The test section will accommodate large air-breathing hypersonic propulsion systems as well as structural and thermal protection system components. Stable wind tunnel test conditions can be provided for 60 seconds. Additional test capabilities are provided by a radiant heater system used to simulate ascent or entry heating profiles. The test medium is the combustion products of air and methane that are burned in a pressurized combustion chamber. Oxygen is added to the test medium for air-breathing propulsion tests so that the test gas contains 21 percent molar oxygen. The facility was modified extensively in the late 1980's to provide airbreathing propulsion testing capability. In this paper, a brief history and general description of the facility are presented along with a discussion of the types of supported testing. Recently completed tests are discussed to explain the capabilities this facility provides and to demonstrate the experience of the staff.

Hodge, Jeffrey S.↗
NASA NTRS1984

Modification of NASA Langley 8 foot high temperature tunnel to provide a unique national research facility for hypersonic air-breathing propulsion systems

A planned modification of the NASA Langley 8-Foot High Temperature Tunnel to make it a unique national research facility for hypersonic air-breathing propulsion systems is described, and some of the ongoing supporting research for that modification is discussed. The modification involves: (1) the addition of an oxygen-enrichment system which will allow the methane-air combustion-heated test stream to simulate air for propulsion testing; and (2) supplemental nozzles to expand the test simulation capability from the current nominal Mach number to 7.0 include Mach numbers 3.0, 4.5, and 5.0. Detailed design of the modifications is currently underway and the modified facility is scheduled to be available for tests of large scale propulsion systems by mid 1988.

Kelly, H. N.↗
NASA NTRS1984

Modification of NASA Langley 8 Foot High Temperature Tunnel to provide a unique national research facility for hypersonic air-breathing propulsion systems

A planned modification of the NASA Langley 8-Foot High Temperature Tunnel to make it a unique national research facility for hypersonic air-breathing propulsion systems is described, and some of the ongoing supporting research for that modification is discussed. The modification involves: (1) the addition of an oxygen-enrichment system which will allow the methane-air combustion-heated test stream to simulate air for propulsion testing; and (2) supplemental nozzles to expand the test simulation capability from the current nominal Mach number to 7.0 include Mach numbers 3.0, 4.5, and 5.0. Detailed design of the modifications is currently underway and the modified facility is scheduled to be available for tests of large scale propulsion systems by mid 1988.

Kelly, H. N.↗
NASA NTRS1991

Research capabilities of the NASA Langley 8-foot high temperature tunnel

The NASA Langley Research Center's 8-Foot High Temperature Tunnel (8' HTT) has been modified to facilitate the testing of hypersonic airbreathing propulsion systems in addition to aerothermal load definition and structural concept verification at Mach 4, 5, and 7. The 8' HTT simulates flight from 60 to 125 kft with run times of 1 to 2 min. The 8-ft diameter and 12-ft-long free jet length enables the testing of large engines or multiple subscale engines. Methane and air combustion products provide the true temperature environment; an oxygen system enriches the combustion products to the same volume fraction of oxygen as air to enable the testing of airbreathing engines. A hydrogen system provides model fuel for cooling and combustion. High use tunnel components have been upgraded or replaced.

Puster, Richard L.↗
NASA NTRS1992

Combustion-wave ignition for rocket engines

The combustion wave ignition concept was experimentally studied in order to verify its suitability for application in baffled sections of a large booster engine combustion chamber. Gaseous oxygen/gaseous methane (GOX/GH4) and gaseous oxygen/gaseous hydrogen (GOX/GH2) propellant combinations were evaluated in a subscale combustion wave ignition system. The system included four element tubes capable of carrying ignition energy simultaneously to four locations, simulating four baffled sections. Also, direct ignition of a simulated Main Combustion Chamber (MCC) was performed. Tests were conducted over a range of mixture ratios and tube geometries. Ignition was consistently attained over a wide range of mixture ratios. And at every ignition, the flame propagated through all four element tubes. For GOX/GH4, the ignition system ignited the MCC flow at mixture ratios from 2 to 10 and for GOX/GH2 the ratios is from 2 to 13. The ignition timing was found to be rapid and uniform. The total ignition delay when using the MCC was under 11 ms, with the tube-to-tube, as well as the run-to-run, variation under 1 ms. Tube geometries were found to have negligible effect on the ignition outcome and timing.

Liou, Larry C.↗
NASA NTRS2010

Combustion Stability Analyses of Coaxial Element Injectors with Liquid Oxygen/Liquid Methane Propellants

Liquid rocket engines using oxygen and methane propellants are being considered by the National Aeronautics and Space Administration (NASA) for in-space vehicles. This propellant combination has not been previously used in a flight-qualified engine system, so limited test data and analysis results are available at this stage of early development. NASA has funded several hardware-oriented activities with oxygen and methane propellants over the past several years with the Propulsion and Cryogenic Advanced Development (PCAD) project, under the Exploration Technology Development Program. As part of this effort, the NASA Marshall Space Flight Center has conducted combustion stability analyses of several of the configurations. This paper presents test data and analyses of combustion stability from the recent PCAD-funded test programs at the NASA MSFC. These test programs used swirl coaxial element injectors with liquid oxygen and liquid methane propellants. Oxygen was injected conventionally in the center of the coaxial element, and swirl was provided by tangential entry slots. Injectors with 28-element and 40-element patterns were tested with several configurations of combustion chambers, including ablative and calorimeter spool sections, and several configurations of fuel injection design. Low frequency combustion instability (chug) occurred with both injectors, and high-frequency combustion instability occurred at the first tangential (1T) transverse mode with the 40-element injector. In most tests, a transition between high-amplitude chug with gaseous methane flow and low-amplitude chug with liquid methane flow was readily observed. Chug analyses of both conditions were conducted using techniques from Wenzel and Szuch and from the Rocket Combustor Interactive Design and Analysis (ROCCID) code. The 1T mode instability occurred in several tests and was apparent by high-frequency pressure measurements as well as dramatic increases in calorimeter-measured heat flux throughout the chamber. Analyses of the transverse mode were conducted with ROCCID and empirical methods such as Hewitt d/V. This paper describes the test hardware configurations, test data, analysis methods, and presents results of the various analyses.

Hulka, J. R.↗
NASA NTRS1992

Computational method to predict thermodynamic, transport, and flow properties for the modified Langley 8-foot high-temperature tunnel

The Langley 8 foot high temperature tunnel (8 ft HTT) is used to test components of hypersonic vehicles for aerothermal loads definition and structural component verification. The test medium of the 8 ft HTT is obtained by burning a mixture of methane and air under high pressure; the combustion products are expanded through an axisymmetric conical contoured nozzle to simulate atmospheric flight at Mach 7. This facility was modified to raise the oxygen content of the test medium to match that of air and to include Mach 4 and Mach 5 capabilities. These modifications will facilitate the testing of hypersonic air breathing propulsion systems for a wide range of flight conditions. A computational method to predict the thermodynamic, transport, and flow properties of the equilibrium chemically reacting oxygen enriched methane-air combustion products was implemented in a computer code. This code calculates the fuel, air, and oxygen mass flow rates and test section flow properties for Mach 7, 5, and 4 nozzle configurations for given combustor and mixer conditions. Salient features of the 8 ft HTT are described, and some of the predicted tunnel operational characteristics are presented in the carpet plots to assist users in preparing test plans.

Venkateswaran, S.↗
NASA NTRS1981

Advanced oxygen-hydrocarbon Earth-to-orbit propulsion

Liquid oxygen/hydrocarbon (LO2/HC) rocket engine cycles for a surface to orbit transportation system were evaluated. A consistent engine system data base is established for defining advantages and disadvantages, system performance and operating limits, engine parametric data, and technology requirements for candidate engine systems. Preliminary comparisons of the engine cycles utilizing delivered specific impulse values are presented. Methane and propane staged combustion cycles are the highest LO2/HC performers. The hydrogen cooled LO2/methane dual throat engine was found to be the highest performing. Technology needs identified in the study include: high temperature turbines; oxidizer-rich preburners; LO2, methane, and propane cooling; methane and propane fuel-rich preburners; the HC fuel turbopump; and application of advanced composite materials to the engine system. Parametric sensitivity analysis data are displayed which show the effect of variations in engine thrust, mixture ratio, chamber pressure, area ratio, cycle life, and turbine inlet temperature on specific impulse and engine weight.

Obrien, C. J.↗
NASA NTRS2014

Combustion Stability Characteristics of the Project Morpheus Liquid Oxygen / Liquid Methane Main Engine

The project Morpheus liquid oxygen (LOX) / liquid methane (LCH4) main engine is a Johnson Space Center (JSC) designed ~5,000 lbf-thrust, 4:1 throttling, pressure-fed cryogenic engine using an impinging element injector design. The engine met or exceeded all performance requirements without experiencing any in- ight failures, but the engine exhibited acoustic-coupled combustion instabilities during sea-level ground-based testing. First tangential (1T), rst radial (1R), 1T1R, and higher order modes were triggered by conditions during the Morpheus vehicle derived low chamber pressure startup sequence. The instability was never observed to initiate during mainstage, even at low power levels. Ground-interaction acoustics aggravated the instability in vehicle tests. Analysis of more than 200 hot re tests on the Morpheus vehicle and Stennis Space Center (SSC) test stand showed a relationship between ignition stability and injector/chamber pressure. The instability had the distinct characteristic of initiating at high relative injection pressure drop at low chamber pressure during the start sequence. Data analysis suggests that the two-phase density during engine start results in a high injection velocity, possibly triggering the instabilities predicted by the Hewitt stability curves. Engine ignition instability was successfully mitigated via a higher-chamber pressure start sequence (e.g., ~50% power level vs ~30%) and operational propellant start temperature limits that maintained \cold LOX" and \warm methane" at the engine inlet. The main engine successfully demonstrated 4:1 throttling without chugging during mainstage, but chug instabilities were observed during some engine shutdown sequences at low injector pressure drop, especially during vehicle landing.

Melcher, John C.↗
NASA NTRS2010

Combustion and Performance Analyses of Coaxial Element Injectors with Liquid Oxygen/Liquid Methane Propellants

Liquid rocket engines using oxygen and methane propellants are being considered by the National Aeronautics and Space Administration (NASA) for in-space vehicles. This propellant combination has not been previously used in a flight-qualified engine system, so limited test data and analysis results are available at this stage of early development. NASA has funded several hardware-oriented activities with oxygen and methane propellants over the past several years with the Propulsion and Cryogenic Advanced Development (PCAD) project, under the Exploration Technology Development Program. As part of this effort, the NASA Marshall Space Flight Center has conducted combustion, performance, and combustion stability analyses of several of the configurations. This paper summarizes the analyses of combustion and performance as a follow-up to a paper published in the 2008 JANNAF/LPS meeting. Combustion stability analyses are presented in a separate paper. The current paper includes test and analysis results of coaxial element injectors using liquid oxygen and liquid methane or gaseous methane propellants. Several thrust chamber configurations have been modeled, including thrust chambers with multi-element swirl coax element injectors tested at the NASA MSFC, and a uni-element chamber with shear and swirl coax injectors tested at The Pennsylvania State University. Configurations were modeled with two one-dimensional liquid rocket combustion analysis codes, the Rocket Combustor Interaction Design and Analysis (ROCCID), and the Coaxial Injector Combustion Model (CICM). Significant effort was applied to show how these codes can be used to model combustion and performance with oxygen/methane propellants a priori, and what anchoring or calibrating features need to be applied or developed in the future. This paper describes the test hardware configurations, presents the results of all the analyses, and compares the results from the two analytical methods

Hulka, J. R.↗
NASA NTRS1991

Symposium (International) on Combustion, 23rd, Universite d'Orleans, France, July 22-27, 1990, Proceedings

The present symposium on combustion discusses reaction kinetics, NO(x) kinetics, premixed, diffusion, and nonsteady flames, turbulent combustion, hazardous waste, fluidized bed combustion, coal boilers and furnaces, engines, heterogeneous kinetics, heterogeneous, droplet, and microgravity combustion, and high-temperature synthesis. Attention is given to reactions of biphenyl, methylnaphthalenes, and phenanthrene with atomic oxygen in the gas phase, the oxidation of ortho-xylene, the effects of water on combustion kinetics at high pressure, and the formation and measurement of N2O in combustion systems. Topics addressed include large ions in premixed benzene-oxygen flames, the structure and kinetics of CH4/N2O flames, the propagation of unsteady hydrogen premixed flames near flammability limits, and the basic structure of lean propane flames. Also considered are OH measurements of piloted diffusion flames of nitrogen-diluted methane near extinction, waste combustion, preferential oxidation of carbon surfaces, and reburning mechanisms in a pulverized coal combustor.

Source record↗
NASA NTRS1978

Computer Program for the Design and Off-Design Performance of Turbojet and Turbofan Engine Cycles

The rapid computer program is designed to be run in a stand-alone mode or operated within a larger program. The computation is based on a simplified one-dimensional gas turbine cycle. Each component in the engine is modeled thermo-dynamically. The component efficiencies used in the thermodynamic modeling are scaled for the off-design conditions from input design point values using empirical trends which are included in the computer code. The engine cycle program is capable of producing reasonable engine performance prediction with a minimum of computer execute time. The current computer execute time on the IBM 360/67 for one Mach number, one altitude, and one power setting is about 0.1 seconds. about 0.1 seconds. The principal assumption used in the calculation is that the compressor is operated along a line of maximum adiabatic efficiency on the compressor map. The fluid properties are computed for the combustion mixture, but dissociation is not included. The procedure included in the program is only for the combustion of JP-4, methane, or hydrogen.

Morris, S. J.↗
NASA NTRS1982

Two-stage earth-to-orbit vehicles with dual-fuel propulsion in the Orbiter

Earth-to-orbit vehicle studies of future replacements for the Space Shuttle are needed to guide technology development. Previous studies that have examined single-stage vehicles have shown advantages for dual-fuel propulsion. Previous two-stage system studies have assumed all-hydrogen fuel for the Orbiters. The present study examined dual-fuel Orbiters and found that the system dry mass could be reduced with this concept. The possibility of staging the booster at a staging velocity low enough to allow coast-back to the launch site is shown to be beneficial, particularly in combination with a dual-fuel Orbiter. An engine evaluation indicated the same ranking of engines as did a previous single-stage study. Propane and RP-1 fuels result in lower vehicle dry mass than methane, and staged-combustion engines are preferred over gas-generator engines. The sensitivity to the engine selection is less for two-stage systems than for single-stage systems.

Martin, J. A.↗
NASA NTRS2000

Scale-Invariant Forms of Conservation Equations in Reactive Fields and a Modified Hydro-Thermo-Diffusive Theory of Laminar Flames

A scale-invariant model of statistical mechanics is applied to present invariant forms of mass, energy, linear, and angular momentum conservation equations in reactive fields. The resulting conservation equations at molecular-dynamic scale are solved by the method of large activation energy asymptotics to describe the hydro-thermo-diffusive structure of laminar premixed flames. The predicted temperature and velocity profiles are in agreement with the observations. Also, with realistic physico-chemical properties and chemical-kinetic parameters for a single-step overall combustion of stoichiometric methane-air premixed flame, the laminar flame propagation velocity of 42.1 cm/s is calculated in agreement with the experimental value.

Sohrab, Siavash H.↗
NASA NTRS1990

Carbon-isotopic analysis of dissolved acetate

Heating of dried, acetate-containing solids together with oxalic acid dihydrate conveniently releases acetic acid for purification by gas chromatography. For determination of the carbon-isotopic composition of total acetate, the acetate-containing zone of the chromatographic effluent can be routed directly to a combustion furnace coupled to a vacuum system allowing recovery, purification, and packaging of CO2 for mass-spectrometric analysis. For analysis of methyl carbon, acetic acid can be cryogenically trapped from the chromatographic effluent, then transferred to a tube containing excess NaOH. The tube is evacuated, sealed, and heated to 500 degrees C to produce methane by pyrolysis of sodium acetate. Subsequent combustion of the methane allows determination of the 13C content at the methyl position in the parent acetate. With typical blanks, the standard deviation of single analyses is less than 0.4% for acetate samples larger than 5 micromoles. A full treatment of uncertainties is outlined.

Non-NASA Center↗