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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 19 records
All sourcesNASA NTRSDOE OSTI
NASA NTRS2003

Characteristics of Gaseous Diffusion Flames with High Temperature Combustion Air in Microgravity

The characteristics of gaseous diffusion flames have been obtained using high temperature combustion air under microgravity conditions. The time resolved flame images under free fall microgravity conditions were obtained from the video images obtained. The tests results reported here were conducted using propane as the fuel and about 1000 C combustion air. The burner included a 0.686 mm diameter central fuel jet injected into the surrounding high temperature combustion air. The fuel jet exit Reynolds number was 63. Several measurements were taken at different air preheats and fuel jet exit Reynolds number. The resulting hybrid color flame was found to be blue at the base of the flame followed by a yellow color flame. The length and width of flame during the entire free fall conditions has been examined. Also the relative flame length and width for blue and yellow portion of the flame has been examined under microgravity conditions. The results show that the flame length decreases and width increases with high air preheats in microgravity condition. In microgravity conditions the flame length is larger with normal temperature combustion air than high temperature air.

Ghaderi, M.↗
NASA NTRS1986

Ignition and combustion temperatures determined by laser heating

A laser heating technique and facility have been developed to study metal ignition and combustion in high-pressure oxygen. The ignition and combustion temperatures, estimates of oxidation rates, and ignition and combustion morphology can be determined. This facility and the laser heating techniques are described. Examples of the type of data obtained are presented and discussed. The ignition temperature curves for an aluminum alloy - Unified Numbering System (UNS) A96061, a stainless steel - UNS S30200, and two nickel alloys - UNS N07718 and N04400 are given.

Bransford, J. W.↗
NASA NTRS1953

Analytical evaluation of effect of equivalence ratio inlet-air temperature and combustion pressure on performance of several possible ram-jet fuels

The results of an analytical investigation of the theoretical air specific impulse performance and adiabatic combustion temperatures of several possible ram-jet fuels over a range of equivalence ratios, inlet-air temperatures, and combustion pressures, is presented herein. The fuels include octane-1, 50-percent-magnesium slurry, boron, pentaborane, diborane, hydrogen, carbon, and aluminum. Thermal effects from high combustion temperatures were found to effect considerably the combustion performance of all the fuels. An increase in combustion pressure was beneficial to air specific impulse at high combustion temperatures. The use of these theoretical data in engine operation and in the evaluation of experimental data is described.

COMBUSTION - TURBINE ENGINES↗
NASA NTRS1988

Laser Diagnostics for combustion temperature and species measurements

Laser optical diagnostic techniques for the measurement of combustion gaseous-phase temperatures and, or species concentrations are discussed. The techniques fall into two classes: incoherent (Rayleigh scattering, spontaneous Raman scattering, laser induced fluorescence spectroscopy) and coherent (coherent anti-Stokes Raman spectroscopy). The advantages, disadvantages and applicability of each method are outlined.

Eckbreth, Alan C.↗
NASA NTRS1983

Preliminary tests of an advanced high-temperature combustion system

A combustion system has been developed to operate efficiently and with good durability at inlet pressures to 4.05 MPa (40 atm), inlet air temperatures to 900 K, and exhaust gas temperatures to 2480 K. A preliminary investigation of this system was conducted at inlet pressures to 0.94 MPa (9 atm), a nominal inlet air temperature of 560 K, and exhaust gas temperatures to 2135 K. A maximum combustion efficiency of 98.5 percent was attained at a fuel-air ratio of 0.033; the combustion efficiency decreased to about 90 percent as the fuel-air ratio was increased to 0.058. An average liner metal temperature of 915 K, 355 kelvins greater than the nominal inlet air temperature, was reached with an average exhaust gas temperature of 2090 K. The maximum local metal temperature at this condition was about 565 kelvins above the nominal inlet air temperature and decreased to 505 kelvins above with increasing combustor pressure. Tests to determine the isothermal total pressure loss of the combustor showed a liner loss of 1.1 percent and a system loss of 6.5 percent.

Wear, J. D.↗
NASA NTRS1983

Comparison of CARS combustion temperatures with standard techniques

This paper describes a comparative study of temperatures measured with Coherent Anti-Stokes Raman Spectroscopy (CARS) and those using standard measurement techniques. CARS temperatures from nitrogen vibrational-rotational spectra have been obtained in a premixed hydrogen-air laboratory flame supported on a flat flame burner over a temperature range from about 1000 K to 2100 K. The techniques used for comparison were sodium line reversal, thin wire thermocouples and heat balance calculations. These studies show a good agreement over the entire temperature range studied.

Antcliff, R. R.↗
NASA NTRS2002

Combustion Temperature Measurement by Spontaneous Raman Scattering in a Jet-A Fueled Gas Turbine Combustor Sector

Spontaneous vibrational Raman scattering was used to measure temperature in an aviation combustor sector burning jet fuel. The inlet temperature ranged from 670 K (750 F) to 756 K (900 F) and pressures from 13 to 55 bar. With the exception of a discrepancy that we attribute to soot, good agreement was seen between the Raman-derived temperatures and the theoretical temperatures calculated from the inlet conditions. The technique used to obtain the temperature uses the relationship between the N2 anti-Stokes and Stokes signals, within a given Raman spectrum. The test was performed using a NASA-concept fuel injector and Jet-A fuel over a range of fuel/air ratios. This work represents the first such measurements in a high-pressure, research aero-combustor facility.

Hicks, Yolanda R.↗
NASA NTRS1981

Ultra-lean combustion at high inlet temperatures

Combustion at inlet air temperatures of 1100 to 1250 K was studied for application to advanced automotive gas turbine engines. Combustion was initiated by the hot environment, and therefore no external ignition source was used. Combustion was stabilized without a flameholder. The tests were performed in a 12 cm diameter test section at a pressure of 2.5 x 10 to the 5th power Pa, with reference velocities of 32 to 60 m/sec and at maximum combustion temperatures of 1350 to 1850 K. Number 2 diesel fuel was injected by means of a multiple source fuel injector. Unburned hydrocarbons emissions were negligible for all test conditions. Nitrogen oxides emissions were less than 1.9 g NO2/kg fuel for combustion temperatures below 1680 K. Carbon monoxide emissions were less than 16 g CO/kg fuel for combustion temperatures greater than 1600 K, inlet air temperatures higher than 1150 K, and residence times greater than 4.3 microseconds.

Anderson, D. N.↗
NASA NTRS1981

Ultra-lean combustion at high inlet temperatures

Combustion at inlet-air temperatures of 1100 to 1250 K was studied for application to advanced automotive gas turbine engines. Combustion was initiated by the hot environment, and therefore no external ignition source was used. Combustion was stabilized without a flameholder. The tests were performed in a 12-cm-diameter test section at a pressure of 250,000 Pa, with reference velocities of 32 to 60 m/s and at maximum combustion temperatures of 1350 to 1850 K. Number 2 diesel fuel was injected by means of a multiple-source fuel injector. Unburned hydrocarbon emissions were negligible for all test conditions. Nitrogen oxide emissions were less than 1.9 g NO2/kg fuel for combustion temperatures below 1680 K. Carbon monoxide emissions were less than 16 g CO/kg fuel for combustion temperatures greater than 1600 K, inlet-air temperatures higher than 1150 K, and residence times greater than 4.3 ms.

Anderson, D. N.↗
NASA NTRS2000

Gas-Phase Influence on Quasisteady "Liquid Flames" in Gravitational Fields

We consider the SHS (self-propagating high-temperature synthesis) process for synthesizing materials. In this process a powder mixture of reactants is cold pressed into a sample, which is ignited at one end. A high temperature combustion wave then propagates through the sample converting reactants to the desired product material. In this process, melting of some or all the components is often observed. Therefore, we study combustion waves propagating through a high caloricity inorganic powder mixture whose combustion temperature exceeds the melting temperatures of many components. The solid matrix is thus destroyed by the propagating combustion wave due to melting ahead of the reaction zone, and a liquid bath is formed which contains gaseous bubbles. The waves propagate in the presence of a gravitational field. Due to the effect of gravity, there is relative motion between the rising bubbles and the descending bath, which affects the composition of the medium, its thermophysical properties, the 'liquid flame' structure, and the propagation velocity. To enhance our understanding of phenomena associated with the interaction of the relative motion with the propagating combustion wave we formulate and analyze a relatively simple mathematical model of liquid flames in a gravitational field. We describe the wave structure and combustion characteristics including the combustion velocity. We compare our results to existing experimental observations and suggest new experiments to be performed. We consider the effects of gravity and, in particular, examine both microgravity and large gravity conditions.

Shkadinsky, K. G.↗
NASA NTRS2005

Development of Ionic Liquid Monopropellants for In-Space Propulsion

A family of new, low toxicity, high energy monopropellants is currently being evaluated at NASA Marshall Space Flight Center for in-space rocket engine applications such as reaction control engines. These ionic liquid monopropellants, developed in recent years by the Air Force Research Laboratory, could offer system simplification, less in-flight thermal management, and reduced handling precautions, while increasing propellant energy density as compared to traditional storable in-space propellants such as hydrazine and nitrogen tetroxide. However, challenges exist in identifying ignition schemes for these ionic liquid monopropellants, which are known to burn at much hotter combustion temperatures compared to traditional monopropellants such as hydrazine. The high temperature combustion of these new monopropellants make the use of typical ignition catalyst beds prohibitive since the catalyst cannot withstand the elevated temperatures. Current research efforts are focused on monopropellant ignition and burn rate characterization, parameters that are important in the fundamental understanding of the monopropellant behavior and the eventual design of a thruster. Laboratory studies will be conducted using alternative ignition techniques such as laser-induced spark ignition and hot wire ignition. Ignition delay, defined as the time between the introduction of the ignition source and the first sign of light emission from a developing flame kernel, will be measured using Schlieren visualization. An optically-accessible liquid monopropellant burner, shown schematically in Figure 1 and similar in design to apparatuses used by other researchers to study solid and liquid monopropellants, will be used to determine propellant burn rate as a function of pressure and initial propellant temperature. The burn rate will be measured via high speed imaging through the chamber s windows.

Blevins, John A.↗
NASA NTRS2005

Development of Ionic Liquid Monopropellants for In-Space Propulsion

A family of new, low toxicity, high energy monopropellants is currently being evaluated at NASA Marshall Space Flight Center for in-space rocket engine applications such as reaction control engines. These ionic liquid monopropellants, developed in recent years by the Air Force Research Laboratory, could offer system simplification, less in-flight thermal management, and reduced handling precautions, while increasing propellant energy density as compared to traditional storable in-space propellants such as hydrazine and nitrogen tetroxide. However, challenges exist in identifying ignition schemes for these ionic liquid monopropellants, which are known to burn at much hotter combustion temperatures compared to traditional monopropellants such as hydrazine. The high temperature combustion of these new monopropellants make the use of typical ignition catalyst beds prohibitive since the catalyst cannot withstand the elevated temperatures. Current research efforts are focused on monopropellant ignition and burn rate characterization, parameters that are important in the fundamental understanding of the monopropellant behavior and the eventual design of a thruster. Laboratory studies will be conducted using alternative ignition techniques such as laser-induced spark ignition and hot wire ignition. Ignition delay, defined as the time between the introduction of the ignition source and the first sign of light emission from a developing flame kernel, will be measured using Schlieren visualization. An optically-accessible liquid monopropellant burner will be used to determine propellant burn rate as a function of pressure and initial propellant temperature. The burn rate will be measured via high speed imaging through the chamber s windows.

Blevins, John A.↗
NASA NTRS2001

Combustion of Gaseous Fuels with High Temperature Air in Normal- and Micro-gravity Conditions

The objective of this study is determine the effect of air preheat temperature on flame characteristics in normal and microgravity conditions. We have obtained qualitative (global flame features) and some quantitative information on the features of flames using high temperature combustion air under normal gravity conditions with propane and methane as the fuels. This data will be compared with the data under microgravity conditions. The specific focus under normal gravity conditions has been on determining the global flame features as well as the spatial distribution of OH, CH, and C2 from flames using high temperature combustion air at different equivalence ratio.

Wang, Y.↗
NASA NTRS1947

Determination of ram-jet combustion-chamber temperatures by means of total-pressure surveys

A method is described by which the total temperature of the gases at the combustion-chamber outlet of a ram-jet engine may be determined from the loss in total pressure measured across the combustion chamber. A working chart is presented by means of which the ratio of the total temperature of the gases at the combustion-chamber outlet to the total temperature of the gases at the combustion-chamber inlet may be determined from the measured loss of total pressure across the combustion chamber and the known values of air flow, total pressure, and total temperature at the combustion-chamber inlet. Values of total-temperature ratio across the combustion chamber of a 20-inch ram jet were obtained in the Cleveland altitude wind tunnel over a range of pressure altitudes from 6000 to 15,000 feet. The difference between the temperature ratio across the combustion chamber determined from the chart and that obtained from the thermocouple measurement was within 6.2 percent of the thermocouple-temperature ratio and was within the accuracy of the thermocouple measurements.

Pinkel, I Irving↗
NASA NTRS2003

Processing Methods Established To Fabricate Porous Oxide Ceramic Spheres for Thermal Barrier Coating Applications

As gas turbine technology advances, the demand for efficient engines and emission reduction requires a further increase in operating temperatures, but combustion temperatures are currently limited by the temperature capability of the engine components. The existing thermal barrier coating (TBC) technology does not provide sufficient thermal load reduction at a 3000 F (1649 C) operating condition. Advancement in thermal barrier coating technology is needed to meet this aggressive goal. One concept for improving thermal barrier coating effectiveness is to design coating systems that incorporate a layer that reflects or scatters photon radiation. This can be achieved by using porous structures. The refractive index mismatch between the solid and pore, the pore size, and the pore density can be engineered to efficiently scatter photon radiation. Under NASA s Ultra-Efficient Engine Technology (UEET) Program, processing methods to fabricate porous ceramic spheres suitable for scattering photon radiation at elevated temperatures have been established. A straightforward templating process was developed at the NASA Glenn Research Center that requires no special processing equipment. The template was used to define particle shape, particle size, and pore size. Spherical organic cation exchange resins were used as a structure-directing template. The cation exchange resins have dual template capabilities that can produce different pore architectures. This process can be used to fabricate both metal oxide and metal carbide spheres.

Dynys, Frederick W.↗
NASA NTRS1993

Fuel-rich catalytic combustion of a high density fuel

Fuel-rich catalytic combustion (ER is greater than 4) of the high density fuel exo-tetrahydrocyclopentadiene (JP-10) was studied over the equivalence ratio range 5.0 to 7.6, which yielded combustion temperatures of 1220 to 1120 K. The process produced soot-free gaseous products similar to those obtained with iso-octane and jet-A in previous studies. The measured combustion temperature agreed well with that calculated assuming soot was not a combustion product. The process raised the effective hydrogen/carbon (H/C) ratio from 1.6 to over 2.0, thus significantly improving the combustion properties of the fuel. At an equivalence ratio near 5.0, about 80 percent of the initial fuel carbon was in light gaseous products and about 20 percent in larger condensable molecules. Fuel-rich catalytic combustion has now been studied for three fuels with H/C ratios of 2.25 (iso-octane), 1.92 (jet-A), and 1.6 (JP-10). A comparison of the product distribution of these fuels shows that, in general, the measured concentrations of the combustion products were monotonic functions of the H/C ratio with the exception of hydrogen and ethylene. In these cases, data for JP-10 fell between iso-octane and jet-A rather than beyond jet-A. It is suggested that the ring cross-linking structure of JP-10 may be responsible for this behavior. All the fuels studied showed that the largest amounts of small hydrocarbon molecules and the smallest amounts of large condensable molecules occurred at the lower equivalence ratios. This corresponds to the highest combustion temperatures used in these studies. Although higher temperatures may improve this mix, the temperature is limited. First, the life of the present catalyst would be greatly shortened when operated at temperatures of 1300 K or greater. Second, fuel-rich catalytic combustion does not produce soot because the combustion temperatures used in the experiments were well below the threshold temperature (1350 K) for the formation of soot. Increasing the temperature above this value would remove the soot-free nature of the process. Since all the fuels studied show a similar breakdown of the primary fuel into smaller molecular combustion products, this technique can be applied to all hydrocarbon fuels.

Brabbs, Theodore A.↗
NASA NTRS1987

Bonding Lexan and sapphire to form high-pressure, flame-resistant window

Flammable materials have been studied in normal gravity and microgravity for many years. Photography plays a major role in the study of the combustion process giving a permanent visual record that can be analyzed. When these studies are extended to manned spacecraft, safety becomes a primary concern. The need for a high-pressure, flame-resistant, shatter-resistant window permitting photographic recording of combustion experiments in manned spacecraft prompted the development of a method for bonding Lexan and sapphire. Materials that resist shattering (e.g., Lexan) are not compatible with combustion experiments; the material loses strength at combustion temperatures. Sapphire is compatible with combustion temperatures in oxygen-enriched atmospheres but is subject to shattering. Combining the two materials results in a shatter-resistant, flame-resistant window. Combustion in microgravity produces a low-visibility flame; however, flame propagation and flame characteristics are readily visible as long as there is no deterioration of the image. Since an air gap between the Lexan and the sapphire would reduce transmission, a method was developed for bonding these unlike materials to minimize light loss.

Richardson, William R.↗
NASA NTRS1970

Composite propellant combustion

Temperature measurements in gaseous reaction zone near surface of burning composite solid propellant, using modified line reversal pyrometer

Derr, R. L.↗