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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 361 records · Page 20
All sourcesNASA NTRSDOE OSTI
NASA NTRS1982

Combustion characteristics of hydrogen-carbon monoxide based gaseous fuels

The results of trials with a staged combustor designed to use coal-derived gaseous fuels and reduce the NO(x) emissions from nitrogen-bound fuels to 75 ppm and 37 ppm without bound nitrogen in 15% O2 are reported. The combustor was outfitted with primary zone regenerative cooling, wherein the air cooling the primary zone was passed into the combustor at 900 F and mixed with the fuel. The increase in the primary air inlet temperature eliminated flashback and autoignition, lowered the levels of CO, unburned hydrocarbons, and smoke, and kept combustion efficiencies to the 99% level. The combustor was also equipped with dual fuel injection to test various combinations of liquid/gas fuel mixtures. Low NO(x) emissions were produced burning both Lurgi and Winkler gases, regardless of the inlet pressure and temperature conditions. Evaluation of methanation of medium energy gases is recommended for providing a fuel with low NO(x) characteristics.

White, D. J.↗
NASA NTRS1984

Ceramic composite liner material for gas turbine combustors

Advanced commercial and military gas turbine engines may operate at combustor outlet temperatures in excess of 1920 K (3000 F). At these temperatures combustors liners experience extreme convective and radiative heat fluxes. The ability of a plasma sprayed ceramic coating to reduce liner metal temperature has been recognized. However, the brittleness of the ceramic layer and the difference in thermal expansion with the metal substrate has caused cracking, spalling and some separation of the ceramic coating. Research directed at turbine tip seals (or shrouds) has shown the advantage of applying the ceramic to a compliant metal pad. This paper discusses recent studies of applying ceramics to combustor liners in which yttria stabilized zirconia plasma sprayed on compliant metal substrates which were exposed to near stoichiometric combustion, presents performance and durability results, and describes a conceptual design for an advanced, small gas turbine combustor. Test specimens were convectively cooled or convective-transpiration cooled and were evaluated in a 10 cm square flame tube combustor at inlet air temperatures of 533 K (500 F) and at a pressure of 0.5 MPa (75 psia). The ceramics were exposed to flame temperatures in excess of 2000 K (3320 F). Results appear very promising with all 30 specimens surviving a screening test and one of two specimens surviving a cyclic durability test.

Ercegovic, D. B.↗
NASA NTRS1984

Atomization of liquid sheets in high pressure airflow

An investigation of liquid sheet atomization is made with combustor simulated inlet air pressures varied from 0.10 to 2.1 MPa. Mean drop diameters are measured with an improved scanning radiometer and correlated with the liquid and air stream Reynolds numbers, RE(1) and RE(A) and the airstream pressure sensitive group GC(2). These data are used in the modeling of the combustion process.

Ingebo, R. D.↗
NASA NTRS1984

Ceramic composite liner material for gas turbine combustors

The application of ceramics to gas turbine combustor liners to reduce liner metal temperature was studied in an experiment in which yttria-stabilized zirconia plasma was sprayed on compliant metal substrates exposed to near stoichiometric combustion. The strain isolation pad materials chosen were Hoskins Alloy 875 and BRUNSLLOY 534 Fiber Metal of 0.25 and 0.38 cm thicknesses and 35 and 45 percent density levels. Combustor screening tests of all specimens showed no evidence of deterioration or failure. Specimens exposed to flame temperatures in excess of 2100 K were convectively or convective-transpiration cooled and were evaluated in a 10 cm sq flame tube at inlet air temperature of 533 K and pressure of 0.5 MPa. The results suggest the superiority of a system composed of the Hoskins Alloy 875 compliant pad with 0.25 cm thickness and 35 percent density coupled with a NiCrAlY bond coat and a 8 percent Y2O3-ZrO2 ceramic top coat of 0.19 cm thickness.

Ercegovic, D. B.↗
NASA NTRS1983

Advanced automotive diesel engine system study

A conceptual study of an advanced automotive diesel engine is discussed. The engine concept selected for vehicle installation was a supercharged 1.4 liter, 4 cylinder spark assisted diesel of 14:1 compression ratio. A compounding unit consisting of a Lysholm compressor and expander is connected to the engine crankshaft by a belt drive. The inlet air charge is heated by the expander exhaust gas via a heat exchanger. Four levels of technology achievement on the selected engine concept were evaluated, from state-of-the-art to the ideal case. This resulted in the fuel economy increasing from 53.2 mpg to 81.7 mpg, and the 0-60 mph time decreasing from 17.6 seconds to 10.9 seconds.

Source record↗
NASA NTRS1985

Combustor flame flashback

A stainless steel, two-dimensional (rectangular), center-dump, premixed-prevaporized combustor with quartz window sidewalls for visual access was designed, built, and used to study flashback. A parametric study revealed that the flashback equivalence ratio decreased slightly as the inlet air temperature increased. It also indicated that the average premixer velocity and premixer wall temperature were not governing parameters of flashback. The steady-state velocity balance concept as the flashback mechanism was not supported. From visual observation several stages of burning were identified. High speed photography verified upstream flame propagation with the leading edge of the flame front near the premixer wall. Combustion instabilities (spontaneous pressure oscillations) were discovered during combustion at the dump plane and during flashback. The pressure oscillation frequency ranged from 40 to 80 Hz. The peak-to-peak amplitude (up to 1.4 psi) increased as the fuel/air equivalence ratio was increased attaining a maximum value just before flashback. The amplitude suddenly decreased when the flame stabilized in the premixer. The pressure oscillations were large enough to cause a local flow reversal. A simple test using ceramic fiber tufts indicated flow reversals existed at the premixer exit during flickering. It is suspected that flashback occurs through the premixer wall boundary layer flow reversal caused by combustion instability. A theoretical analysis of periodic flow in the premixing channel has been made. The theory supports the flow reversal mechanism.

Proctor, M. P.↗
NASA NTRS1987

Automated design of controlled diffusion blades

A numerical automation procedure has been developed to be used in conjunction with an inverse hodograph method for the design of controlled diffusion blades. With this procedure a cascade of airfoils with a prescribed solidity, inlet Mach number, inlet air flow angle, and air flow turning can be produced automatically. The trailing edge thickness of the airfoil, an important quantity in inverse methods, is also prescribed. The automation procedure consists of a multidimensional Newton iteration in which the objective design conditions are achieved by acting on the hodograph input parameters of the underlying inverse code. The method, although more general in scope, is applied in this paper to the design of axial flow compressor blade sections, and a wide range of examples is presented.

Sanz, Jose M.↗
NASA NTRS1988

Numerical simulation of hypersonic inlet flows with equilibrium or finite rate chemistry

An efficient numerical program incorporated with comprehensive high temperature gas property models has been developed to simulate hypersonic inlet flows. The computer program employs an implicit lower-upper time marching scheme to solve the two-dimensional Navier-Stokes equations with variable thermodynamic and transport properties. Both finite-rate and local-equilibrium approaches are adopted in the chemical reaction model for dissociation and ionization of the inlet air. In the finite rate approach, eleven species equations coupled with fluid dynamic equations are solved simultaneously. In the local-equilibrium approach, instead of solving species equations, an efficient chemical equilibrium package has been developed and incorporated into the flow code to obtain chemical compositions directly. Gas properties for the reaction products species are calculated by methods of statistical mechanics and fit to a polynomial form for C(p). In the present study, since the chemical reaction time is comparable to the flow residence time, the local-equilibrium model underpredicts the temperature in the shock layer. Significant differences of predicted chemical compositions in shock layer between finite rate and local-equilibrium approaches have been observed.

Yu, Sheng-Tao↗
NASA NTRS1986

LDV measurements in an annular combustor model

The design and setup of a Laser Doppler Velocimeter (LDV) system used to take velocity measurements in an annular combustor model are covered. The annular combustor model is of contemporary design using 60 degree flat vane swirlers, producing a strong recirculation zone. Detailed measurements are taken of the swirler inlet air flow and of the downstream enclosed swirling flow. The laser system used is a two color, two component system set up in forward scatter. Detailed are some of the special considerations needed for LDV use in the confined turbulent flow of the combustor model. The LDV measurements in a single swirler rig indicated that the flow changes radically in the first duct height. After this, a flow profile is set up and remains constant in shape. The magnitude of the velocities gradually decays due to viscous damping.

Barron, Dean A.↗
NASA NTRS1988

Automated design of controlled-diffusion blades

A numerical automation procedure has been developed to be used in conjunction with an inverse hodograph method for the design of controlled diffusion blades. With this procedure a cascade of airfoils with a prescribed solidity, inlet Mach number, inlet air flow angle, and air flow turning can be produced automatically. The trailing edge thickness of the airfoil, an important quantity in inverse methods, is also prescribed. The automation procedure consists of a multidimensional Newton iteration in which the objective design conditions are achieved by acting on the hodograph input parameters of the underlying inverse code. The method, although more general in scope, is applied in this paper to the design of axial flow compressor blade sections, and a wide range of examples is presented.

Sanz, J. M.↗
NASA NTRS1989

Automated design of controlled diffusion blades

A numerical automation procedure was developed to be used in conjunction with an inverse hodograph method for the design of controlled diffusion blades. With this procedure a cascade of airfoils with a prescribed solidity, inlet Mach No., inlet air flow angle and air flow turning can be produced automatically. The trailing edge thickness of the airfoil, an important quantity in inverse methods, is also prescribed. The automation procedure consists of a multi-dimensional Newton iteration in which the objective design conditions are achieved by acting on the hodograph input parameters of the underlying inverse code. The method, although more general in scope, is applied to the design of axial flow turbomachinery blade sections, both compressors and turbines. A collaborative effort with U.S. Engine Companies to identify designs of interest to the industry will be described.

Sanz, Jose M.↗
NASA NTRS1991

Static performance tests of a flight-type STOVL ejector

The design and development of thrust augmenting STOVL ejectors has typically been based on experimental iteration (i.e., trial and error). Static performance tests of a full scale vertical lift ejector were performed at primary flow temperatures up to 1560 R (1100 F). Flow visualization (smoke generators and yarn tufts) were used to view the inlet air flow, especially around the primary nozzle and end plates. Performance calculations are presented for ambient temperatures close to 480 R (20 F) and 535 R (75 F) which simulate seasonal aircraft operating conditions. Resulting thrust augmentation ratios are presented as functions of nozzle pressure ratio and temperature.

Barankiewicz, Wendy S.↗
NASA NTRS1991

Static performance tests of a flight-type STOVL ejector

The design and development of thrust augmenting STOVL ejectors has typically been based on experimental iteration (i.e., trial and error). Static performance tests of a full scale vertical lift ejector were performed at primary flow temperatures up to 1560 R (1100 F). Flow visualization (smoke generators and yarn tufts) were used to view the inlet air flow, especially around the primary nozzle and end plates. Performance calculations are presented for ambient temperatures close to 480 R (20 F) and 535 R (75 F) which simulate seasonal aircraft operating conditions. Resulting thrust augmentation ratios are presented as functions of nozzle pressure ratio and temperature.

Barankiewicz, Wendy S.↗
NASA NTRS1992

Nitric oxide formation in a lean, premixed-prevaporized jet A/air flame tube: An experimental and analytical study

An experimental and analytical study was performed on a lean, premixed-prevaporized Jet A/air flame tube. The NO(x) emissions were measured in a flame tube apparatus at inlet temperatures ranging from 755 to 866 K (900 to 1100 F), pressures from 10 to 15 atm, and equivalence ratios from 0.37 to 0.62. The data were then used in regressing an equation to predict the NO(x) production levels in combustors of similar design. Through an evaluation of parameters it was found that NO(x) is dependent on adiabatic flame temperature and combustion residence time, yet independent of pressure and inlet air temperature for the range of conditions studied. This equation was then applied to experimental data that were obtained from the literature, and a good correlation was achieved.

Lee, Chi-Ming↗
NASA NTRS1950

Combustion efficiency and altitude operational limits of three liquid hydrocarbon fuels having high volumetric energy content in a J33 single combustor

Combustion efficiency and altitude operational limits were determined in a J33 single combustor for AN-F-58 fuel and three liquid hydrocarbon fuels having high volumetric energy content (decalin, tetralin, and monomethylnaphthalene) at simulated altitude and combustor inlet-air conditions. At the conditions investigated, the combustion efficiency for the four fuels generally decreased with an increase in volumetric energy content. The altitude operational limits for decalin and tetralin fuels were higher than for AN-F-58 fuel; monomethylnaphthalene fuel gave the lowest altitude operational limit.

Stricker, Edward G↗
NASA NTRS1950

NACA research on combustors for aircraft gas turbines I : effects of operating variables on steady-state performance

Some of the systematic research conducted by the NACA on aircraft gas-turbine combustors is reviewed. Trends depicting the effect of inlet-air pressure, temperature, and velocity and fuel-air ratio on performance characteristics, such as combustion efficiency, maximum temperature rise attainable, pressure loss, and combustor-outlet temperature distribution are described for a variety of turbojet combustors of the liquid-fuel type. These trends are further discussed as effects significant to the turbojet engine, such as altitude operational limits, specific fuel consumption, thrust, acceleration, and turbine life.

Olson, Walter T↗
NASA NTRS1951

Preliminary results of turbojet-engine altitude-starting investigation

A spark energy of 2.13 joules per spark at 1 spark per second produced ignition to an altitude of 50,000 feet at a flight Mach number of 0.6. The minimum power requirements for ignition were obtained from a combination of low spark repetition rates and high spark energy. The altitude-ignition limit was also increased by increasing spark-gap immersion, fuel temperature, inlet-air temperature, and fuel volatility, and by decreasing flight Mach number. The maximum altitude at which flame propagation was accomplished from combusters with spark plugs to combusters without spark plugs to combustors without spark plugs was increased about 5000 feet by increasing fuel volatility.

COMBUSTION - IGNITION OF GASES↗
NASA NTRS1953

A comparison of typical national gas turbine establishment and NACA axial-flow compressor blade sections in cascade at low speed

Comparative cascade tests of the NGTE (National Gas Turbine Establishment of Great Britain) 10C4/30C50 and NACA 65-(12)10 axial flow compressor blade sections were conducted in a 5-inch low-speed cascade tunnel at the Langley Laboratory at air-inlet angles of 30 degrees, 45 degrees, and 60 degrees and a solidity of 1.0 by using the porous-wall technique. These NACA data for the NGTE 10C4/30C50 section were also compared with data from NGTE design charts for the same section. British and NACA incompressible cascade force-analysis equations are included.

CASCADES, EXPERIMENT↗