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Search indexed NASA NTRS and DOE OSTI research on propulsion, heat transfer, battery materials and energy systems. Follow report and document links to the original sources.

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At least 127 records · Page 7
All sourcesNASA NTRSDOE OSTI
NASA NTRS1991

New estimates of nitrous oxide emissions from biomass burning

The recent discovery of an artifact producing increased levels of N2O in combustion gas samples collected and stored in grab bottles before chemical analysis has resulted in the downgrading of fossil-fuel combustion and the questioning of biomass burning as important sources of N2O. As almost all reported analyses of N2O produced from biomass burning have involved essentially the same collection and analysis protocols as used in the fossil-fuel studies, this source of N2O must also be reexamined. Here, measurements of N2O made over a large prescribed fire using a near real-time in situ measurement technique are reported and compared with measurements of N2O from simultaneously collected grab-bottle samples. The results from 27 small laboratory biomass test fires are also used to help clarify the validity of earlier assessments. It is concluded that biomass burning contributes about seven percent of atmospheric N2O, as opposed to earlier estimates of several times this value.

Cofer, W. R., III↗
NASA NTRS1981

Investigations of coherent anti-Stokes Raman spectroscopy /CARS/ for practical combustion diagnostics

Coherent anti-Stokes Raman spectroscopy (CARS) is a coherent wave-mixing process in which the signal emerges as a laser-like beam in a precise direction. The spectroscopic technique is appropriate to spatially and temporally resolved measurements of temperature and major species concentrations in combustion. It has been generated from all of the dominant constituents in air-fed, hydrogen, and hydrocarbon fueled combustion, and found applicable to practical combustion systems. High pressure effects on CARS spectra have also been examined.

Eckbreth, A. C.↗
NASA NTRSDate not supplied

Lean Blowout (LBO) Simulations in a Rich-Burn Quick-Quench Lean-Burn (RQL) Gas Turbine Combustor

The transient processes involving lean blowout (LBO) in a RQL combustor such as the single cup combustor in the National Jet Fuels Combustion Program (NJFCP) referee rig are simulated with the reduced mechanisms based on Hybrid Chemistry for two different fuels, Cat-A2 and Cat-C1. Cat-A2 fuel is Jet-A with average or nominal properties based on a survey of the Petroleum Quality Information System database. The Cat-C1, an alternative jet fuel, is a GEVO alcohol-to-jet (ATJ) fuel composed almost entirely of two highly-branched C12 and C16 iso-paraffins. Experiments conducted in the NJFCP referee rig established a stable flame at an equivalence ratio of 0.096 (the near LBO condition) as the starting condition for LBO tests performed with all Category A and C fuels. Simulations of the approach to LBO verify that a stable flame is established at the near-LBO condition for each fuel before reducing fuel flow rate (and thus equivalence ratio) in a step-wise manner. At each new equivalence ratio, the global heat release rate and total fuel evaporation time histories are monitored to verify a new stable flame condition has been established before performing the next reduction in equivalence ratio until the flame is approaching the blowout. Main conclusions are threefold: (1) For Cat-C1 fuel the equivalence ratio for LBO is around 0.087. (2) For Cat-A2 fuel the equivalence ratio for LBO is either 0.082 or 0.078 depended on the time span of equivalence ratio of 0.085. (3) The averaged difference from LES on LBO for Cat-C1 relative to Cat-A2 is 8.75% versus 7.8% from the experimental data.

LES, Lean blow out↗
NASA NTRSDate not supplied

Assessment of the Fluid Network for the Effusion Holes in a RQL (Rich-burn, Quick-quench, Lean-burn) Gas Turbine Combustor

The transient processes involving lean blowout (LBO) in a RQL combustor such as the single cup combustor in the National Jet Fuels Combustion Program (NJFCP) referee rig are simulated with the reduced mechanisms based on Hybrid Chemistry for three different fuels, A2, C1 and C4. Experiments conducted in the NJFCP referee rig established a stable flame at a global equivalence ratio of 0.096 (the near LBO condition) as the starting condition for LBO tests performed with all A and C fuels. Simulations of the approach to LBO verify that a stable flame is established at the near-LBO condition for each fuel before reducing fuel flow rate (and thus equivalence ratio) in a stepwise manner. Main conclusions are: (1) For A2 fuel the equivalence ratio for LBO is 0.081 (2) For C1 fuel the equivalence ratio for LBO is 0.087 or 0.085depended on different step-down transition paths. (3) For C4 fuel the equivalence ratio for LBO is 0.087 or 0.084 depended on types of mesh or boundary conditions. (4) The fuel relative difference from LES on LBO for C1 fuel relative to A2 fuel is 7.40% or 4.94% versus 7.83% from the experimental data. (5) The fuel relative difference from LES on LBO for C4 fuel relative to A2 is 6.17% or 7.69% versus 5.21% from the experimental data

Lean blow out↗
NASA NTRSDate not supplied

Assessment of the Fluid Network for the Effusion Holes in a RQL (Rich-burn, Quick-quench, Lean-burn) Gas Turbine Combustor

The transient processes involving lean blowout (LBO) in a RQL combustor such as the single cup combustor in the National Jet Fuels Combustion Program (NJFCP) referee rig are simulated with the reduced mechanisms based on Hybrid Chemistry for three different fuels, A2, C1 and C4. Experiments conducted in the NJFCP referee rig established a stable flame at a global equivalence ratio of 0.096 (the near LBO condition) as the starting condition for LBO tests performed with all A and C fuels. Simulations of the approach to LBO verify that a stable flame is established at the near-LBO condition for each fuel before reducing fuel flow rate (and thus equivalence ratio) in a stepwise manner. Main conclusions are: (1) For A2 fuel the equivalence ratio for LBO is 0.081 (2) For C1 fuel the equivalence ratio for LBO is 0.087 or 0.085 depended on different step-down transition paths. (3) For C4 fuel the equivalence ratio for LBO is 0.087 or 0.084 depended on types of mesh or boundary conditions. (4) The fuel relative difference from LES on LBO for C1 fuel relative to A2 fuel is 7.40% or 4.94% versus 7.83% from the experimental data. (5) The fuel relative difference from LES on LBO for C4 fuel relative to A2 is 6.17% or 7.69% versus 5.21% from the experimental data.

Lean Blow out↗
NASA NTRS1993

Combustion of solid fuel in very low speed oxygen streams

In reduced gravity, the combustion of solid fuel in low-speed flow can be studied. The flame behavior in this low-speed regime will fill a void in our understanding of the flow effect on combustion. In addition, it is important for spacecraft fire safety considerations. In this work, modeling and experimental work on low-speed forced-concurrent-flow flame spread are carried out. In addition, experiments on reduced-gravity buoyant-flow flame spread are performed.

Tien, James S.↗
NASA NTRS2008

Stagnation point reverse flow combustor

A method for combusting a combustible fuel includes providing a vessel having an opening near a proximate end and a closed distal end defining a combustion chamber. A combustible reactants mixture is presented into the combustion chamber. The combustible reactants mixture is ignited creating a flame and combustion products. The closed end of the combustion chamber is utilized for directing combustion products toward the opening of the combustion chamber creating a reverse flow of combustion products within the combustion chamber. The reverse flow of combustion products is intermixed with combustible reactants mixture to maintain the flame.

Zinn, Ben T.↗
NASA NTRSDate not supplied

Numerical Simulation of Lean Blowout of Alternative Fuels in 7-element Lean Direct Injector

This research presents the result of numerically simulating 7-element swirl-venturi Lean Direct Injector (SV-LDI) lean blowout (LBO) experiments conducted at NASA Glenn Research Center in May of 2019. After simulating a cold flow case to confirm the pressure drop agrees well with the experiment, additional cases with two different fuels (an average jet fuel and a Gevo alcohol-to-jet fuel) from the National Jet Fuels Combustion Program (NJFCP) were computed to numerically determine the LBO condition. The procedure to approach the LBO follows the method used in the experiment where the air mass flow rate is gradually increased while the fuel supply is maintained. Transient history of global heat release rate as a function of air flow rate is presented, as well as temperature contours at different conditions to give a visual representation of the flame state. The Open National Combustion Code (OpenNCC) used in this research adopted reduced HyChem (Hybrid Chemistry) models along with k-LES turbulence model and a Lagrangian spray model that takes into account droplet internal temperature distribution affected by the shear force on the droplet surface. The transport equations and chemical reaction terms are integrated together to enhance conservation of chemical species that are especially important in the near LBO conditions. After showing the computed range of LBO agrees well with the experimental measurements, time averaged solutions of both fuels at their initial condition and at their limiting condition just before LBO are compared in detail to facilitate the understanding of LBO mechanism.

combustion↗
NASA NTRS1991

Scramjet mixing establishment times for a pulse facility

A numerical simulation of the temporally developing flow through a generic scramjet combustor duct is presented for stagnation conditions typical of flight at Mach 13 as produced by a shock tunnel pulse facility. The particular focus is to examine the start up transients and to determine the time required for certain flow parameters to become established. The calculations were made with a Navier-Stokes solver SPARK with temporally relaxing inflow conditions derived from operation of the T4 shock tunnel at the University of Queensland in Australia. Calculations at nominal steady inflow conditions were made for comparison. The generic combustor geometry includes the injection of hydrogen fuel from the base of a centrally located strut. In both cases, the flow was assumed laminar and fuel combustion was not included. The establishment process is presented for viscous parameters in the boundary layer and for parameters related to the fuel mixing.

Rogers, R. Clayton↗
NASA NTRS2011

HIFiRE Direct-Connect Rig (HDCR) Phase I Scramjet Test Results from the NASA Langley Arc-Heated Scramjet Test Facility

A series of hydrocarbon-fueled direct-connect scramjet ground tests has been completed in the NASA Langley Arc-Heated Scramjet Test Facility (AHSTF) at simulated Mach 8 flight conditions. These experiments were part of an initial test phase to support Flight 2 of the Hypersonic International Flight Research Experimentation (HIFiRE) Program. In this flight experiment, a hydrocarbon-fueled scramjet is intended to demonstrate transition from dual-mode to scramjet-mode operation and verify the scramjet performance prediction and design tools A performance goal is the achievement of a combusted fuel equivalence ratio greater than 0.7 while in scramjet mode. The ground test rig, designated the HIFiRE Direct Connect Rig (HDCR), is a full-scale, heat sink test article that duplicates both the flowpath lines and a majority of the instrumentation layout of the isolator and combustor portion of the flight test hardware. The primary objectives of the HDCR Phase I tests were to verify the operability of the HIFiRE isolator/combustor across the simulated Mach 6-8 flight regime and to establish a fuel distribution schedule to ensure a successful mode transition. Both of these objectives were achieved prior to the HiFIRE Flight 2 payload Critical Design Review. Mach 8 ground test results are presented in this report, including flowpath surface pressure distributions that demonstrate the operation of the flowpath in scramjet-mode over a small range of test conditions around the nominal Mach 8 simulation, as well as over a range of fuel equivalence ratios. Flowpath analysis using ground test data is presented elsewhere; however, limited comparisons with analytical predictions suggest that both scramjet-mode operation and the combustion performance objective are achieved at Mach 8 conditions.

Cabell, Karen↗
NASA NTRS1975

International Symposium on Combustion, 15th, Tokyo, Japan, August 25-31, 1974, Proceedings

Recent theoretical and experimental studies concerned with detonation and pressure wave combustion interaction, fire and explosion research and safety, heterogeneous combustion, flame-flow interactions, kinetics of elementary reactions, pollution control in and by combustion systems, and ignition are presented. Some of the topics covered include critical power density for direct initiation of unconfined gaseous detonations, extinction of laminar diffusion flames for liquid fuels, combustion of bulk titanium in oxygen, flame propagation in small spheres of unconfined and slightly confined flammable mixtures, kinetics of the reaction of nitric oxide with hydrogen, production of chemi-ions and formation of CH and CH2 radicals in methane-oxygen and ethylene-oxygen flames, NOx emission characteristics in two-stage combustion, and spherical ignition of oxyhydrogen behind a reflected shock wave. Individual items are announced in this issue.

Source record↗
NASA NTRS1995

High-pressure combustion of binary fuel sprays

The ultimate objective of this study is to obtain fundamental information relevant to combustion processes that occur in fuel sprays of practical interest at high pressures in internal combustion engines. Since practical fuels are multicomponent and derived from petroleum, the present work involves the model alkane mixture of n-heptane and n-hexadecane. Since burning droplets in sprays can interact with each other, the present work involves investigation of the effects of this interaction on flame shapes and droplet burning times. The small droplets in practical combustion chambers are not significantly influenced by buoyancy. Since such small droplets are difficult to study experimentally, the present work takes advantage of microgravity to lessen buoyancy and enable information about droplet interactions to be obtained by studying larger droplets. The results are intended to provide fundamental understanding that can be used in improving descriptions of practical spray combustion.

Mikami, Masato↗
NASA NTRS1975

The effect of initial flow nonuniformity on second-stage fuel injection and combustion in a supersonic duct

The effects of flow nonuniformity on second-stage hydrogen fuel injection and combustion in supersonic flow were evaluated. The first case, second-stage fuel injection into a uniform duct flow, produced data indicating that fuel mixing is considerably slower than estimates based on an empirical mixing correlation. The second-case, two-stage fuel injection (or second-stage fuel injection into a nonuniform duct flow), produced a large interaction between stages with extensive flow separation. For this case the measured wall pressure, heat transfer, and amount of reaction at the duct exit were significantly greater than estimates based on the mixing correlation. Substantially more second-stage fuel burned in the second case than in the first case. Overall effects of unmixedness/chemical kinetics were found not to be significant at the exit for stoichiometric fuel injection.

Russin, W. R.↗
NASA NTRS1995

In situ observations in aircraft exhaust plumes in the lower stratosphere at midlatitudes

Instrumentation on the NASA ER-2 high-altitude aircraft has been used to observe engine exhaust from the same aircraft while operating in the lower stratosphere. Encounters with the exhaust plume occurred approximately 10 min after emission with spatial scales near 2 km and durations of up to 10 s. Measurements include total reactive nitrogen, NO(y), the component species NO and NO2, CO2, H2O, CO, N2O, condensation nuclei, and meteorological parameters. The integrated amounts of CO2 and H2O during the encounters are consistent with the stoichiometry of fuel combustion (1:1 molar). Emission indices (EI) for NO(x) (= NO + NO2), CO, and N2O are calculated using simultaneous measurements of CO2. EI values for NO(x) near 4 g/(kg fuel) are in good agreement with values scaled from limited ground-based tests of the ER-2 engine. Non-NO(x) species comprise less than about 20% of emitted reactive nitrogen, consistent with model evaluations. In addition to demonstrating the feasibility of aircraft plume detection, these results increase confidence in the projection of emissions from current and proposed supersonic aircraft fleets and hence in the assessment of potential long-term changes in the atmosphere.

Fahey, D. W.↗
NASA NTRSDate not supplied

Effects of Confinement on Flame Spread in Microgravity

Solid fuel combustion experiments aboard the ISS examine the effects of confinement on a concurrent, purely-forced-flow flame spread in microgravity environment. The results for a thin, cotton-fiberglass-blended textile fabric fuel are presented. Flat baffles of differing materials are used to alter the radiative boundary conditions with transparent polycarbonate, black anodized aluminum (reflectance ~ 0), and highly polished aluminum (reflectance ~ 1). The baffles are parallel to the fuel sheet and placed symmetrically on each side. The inter-baffle distance is varied to change the boundary conditions for the flow. In all tests, samples are ignited at the upstream leading edge and allowed to burn to completion. Results show that the flame reaches a steady length and spread rate at low flow speeds (< 15 cm/s) for all tested inter-baffle distances. As the distance decreases, the flame length and spread rate first increase then decrease showing an optimal inter-baffle distance. For all baffle types, the flame either fails to ignite or extinguishes before reaching the end of the sample when the inter-baffle distance is too small (~ 1 cm). This is attributed to the reduction of oxygen supply to the flame zone and heat loss to the baffles. The results also show at the same inter-baffle distance, flame length and spread rate are highest for polished aluminum baffles, and lowest for transparent polycarbonate baffles. The differences are most prominent at intermediate tested baffle distances. While the radiative heat feedback from the baffles is expected to increase when the baffle distance decreases, the combustion is limited by the reduced oxygen supply. Near this limit, flame lengths and spread rates are similar for all baffle types.

Microgravity↗
NASA NTRS2010

HIFiRE Direct-Connect Rig (HDCR) Phase I Ground Test Results from the NASA Langley Arc-Heated Scramjet Test Facility

The initial phase of hydrocarbon-fueled ground tests supporting Flight 2 of the Hypersonic International Flight Research Experiment (HIFiRE) Program has been conducted in the NASA Langley Arc-Heated Scramjet Test Facility (AHSTF). The HIFiRE Program, an Air Force-lead international cooperative program includes eight different flight test experiments designed to target specific challenges of hypersonic flight. The second of the eight planned flight experiments is a hydrocarbon-fueled scramjet flight test intended to demonstrate dual-mode to scramjet-mode operation and verify the scramjet performance prediction and design tools. A performance goal is the achievement of a combusted fuel equivalence ratio greater than 0.7 while in scramjet mode. The ground test rig, designated the HIFiRE Direct Connect Rig (HDCR), is a full-scale, heat sink, direct-connect ground test article that duplicates both the flowpath lines and the instrumentation layout of the isolator and combustor portion of the flight test hardware. The primary objectives of the HDCR Phase I tests are to verify the operability of the HIFiRE isolator/combustor across the Mach 6.0-8.0 flight regime and to establish a fuel distribution schedule to ensure a successful mode transition prior to the HiFIRE payload Critical Design Review. Although the phase I test plans include testing over the Mach 6 to 8 flight simulation range, only Mach 6 testing will be reported in this paper. Experimental results presented here include flowpath surface pressure, temperature, and heat flux distributions that demonstrate the operation of the flowpath over a small range of test conditions around the nominal Mach 6 simulation, as well as a range of fuel equivalence ratios and fuel injection distributions. Both ethylene and a mixture of ethylene and methane (planned for flight) were tested. Maximum back pressure and flameholding limits, as well as a baseline fuel schedule, that covers the Mach 5.84-6.5 test space have been identified.

Hass, Neal E.↗
NASA NTRS2001

Effect of Slow External Flow on Flame Spreading over Solid Material: Opposed Spreading over Polyethylene Wire Insulation

The effect of slow external flow on solid combustion is very important from the view of fire safety in space because the solid material in spacecraft is generally exposed to the low air flow for ventilation. Further, the effect of low external flow on fuel combustion is generally fundamental information for industrial combustion system, such as gas turbine, boiler incinerator and so on. However, it is difficult to study the effect of low external flow on solid combustion in normal gravity, because the buoyancy-induced flow strongly disturbs the flow field, especially for low flow velocity. In this research therefore, the effect of slow external flow on opposed flame spreading over polyethylene (PE) wire insulation have been investigated in microgravity. The microgravity environment was provided by Japan Microgravity Center (JAMIC) in Japan and KC-135 at NASA GRC. The tested flow velocity range is 0-30cm/s with different oxygen concentration and inert gas component.

Fujita, O.↗