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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 649 records · Page 36
All sourcesNASA NTRSDOE OSTI
NASA NTRS2002

Design Considerations of Istar Hydrocarbon Fueled Combustor Operating in Air Augmented Rocket, Ramjet and Scramjet Modes

The development and ground test of a rocket-based combined cycle (RBCC) propulsion system is being conducted as part of the NASA Marshall Space Flight Center (MSFC) Integrated System Test of an Airbreathing Rocket (ISTAR) program. The eventual flight vehicle (X-43B) is designed to support an air-launched self-powered Mach 0.7 to 7.0 demonstration of an RBCC engine through all of its airbreathing propulsion modes - air augmented rocket (AAR), ramjet (RJ), and scramjet (SJ). Through the use of analytical tools, numerical simulations, and experimental tests the ISTAR program is developing and validating a hydrocarbon-fueled RBCC combustor design methodology. This methodology will then be used to design an integrated RBCC propulsion system thai: produces robust ignition and combustion stability characteristics while maximizing combustion efficiency and minimizing drag losses. First order analytical and numerical methods used to design hydrocarbon-fueled combustors are discussed with emphasis on the methods and determination of requirements necessary to establish engine operability and performance characteristics.

Andreadis, Dean↗
NASA NTRS2004

Fuel-Flexible Gas Turbine Combustor Flametube Facility

Facility modifications have been completed to an existing combustor flametube facility to enable testing with gaseous hydrogen propellants at the NASA Glenn Research Center. The purpose of the facility is to test a variety of fuel nozzle and flameholder hardware configurations for use in aircraft combustors. Facility capabilities have been expanded to include testing with gaseous hydrogen, along with the existing hydrocarbon-based jet fuel. Modifications have also been made to the facility air supply to provide heated air up to 350 psig, 1100 F, and 3.0 lbm/s. The facility can accommodate a wide variety of flametube and fuel nozzle configurations. Emissions and performance data are obtained via a variety of gas sample probe configurations and emissions measurement equipment.

Little, James E.↗
NASA NTRS2004

Reliable and Affordable Control Systems Active Combustor Pattern Factor Control

Active, closed-loop control of combustor pattern factor is a cooperative effort between Honeywell (formerly AlliedSignal) Engines and Systems and the NASA Glenn Research Center to reduce emissions and turbine-stator vane temperature variations, thereby enhancing engine performance and life, and reducing direct operating costs. Total fuel flow supplied to the engine is established by the speed/power control, but the distribution to individual atomizers will be controlled by the Active Combustor Pattern Factor Control (ACPFC). This system consist of three major components: multiple, thin-film sensors located on the turbine-stator vanes; fuel-flow modulators for individual atomizers; and control logic and algorithms within the electronic control.

McCarty, Bob↗
NASA NTRS1995

Coupled Monte Carlo Probability Density Function/ SPRAY/CFD Code Developed for Modeling Gas-Turbine Combustor Flows

The success of any solution methodology for studying gas-turbine combustor flows depends a great deal on how well it can model various complex, rate-controlling processes associated with turbulent transport, mixing, chemical kinetics, evaporation and spreading rates of the spray, convective and radiative heat transfer, and other phenomena. These phenomena often strongly interact with each other at disparate time and length scales. In particular, turbulence plays an important role in determining the rates of mass and heat transfer, chemical reactions, and evaporation in many practical combustion devices. Turbulence manifests its influence in a diffusion flame in several forms depending on how turbulence interacts with various flame scales. These forms range from the so-called wrinkled, or stretched, flamelets regime, to the distributed combustion regime. Conventional turbulence closure models have difficulty in treating highly nonlinear reaction rates. A solution procedure based on the joint composition probability density function (PDF) approach holds the promise of modeling various important combustion phenomena relevant to practical combustion devices such as extinction, blowoff limits, and emissions predictions because it can handle the nonlinear chemical reaction rates without any approximation. In this approach, mean and turbulence gas-phase velocity fields are determined from a standard turbulence model; the joint composition field of species and enthalpy are determined from the solution of a modeled PDF transport equation; and a Lagrangian-based dilute spray model is used for the liquid-phase representation with appropriate consideration of the exchanges of mass, momentum, and energy between the two phases. The PDF transport equation is solved by a Monte Carlo method, and existing state-of-the-art numerical representations are used to solve the mean gasphase velocity and turbulence fields together with the liquid-phase equations. The joint composition PDF approach was extended in our previous work to the study of compressible reacting flows. The application of this method to several supersonic diffusion flames associated with scramjet combustor flow fields provided favorable comparisons with the available experimental data. A further extension of this approach to spray flames, three-dimensional computations, and parallel computing was reported in a recent paper. The recently developed PDF/SPRAY/computational fluid dynamics (CFD) module combines the novelty of the joint composition PDF approach with the ability to run on parallel architectures. This algorithm was implemented on the NASA Lewis Research Center's Cray T3D, a massively parallel computer with an aggregate of 64 processor elements. The calculation procedure was applied to predict the flow properties of both open and confined swirl-stabilized spray flames.

Source record↗
NASA NTRS1997

High-Pressure Combustion Testing Reveals Promise of Low-Emission Combustors for Advanced Subsonic Gas Turbines

NASA Lewis Research Center's new, world-class, 60-atm combustor research facility, the Advanced Subsonic Combustion Rig (ASCR), is in operation and producing highly unique research data. At operating pressures to 800 psia, emissions of nitrogen oxides were reduced by greater than 70 percent with an advanced fuel injector designed at NASA Lewis. Data, including exhaust emissions and pressure and temperature distributions, were acquired at high pressures and temperatures representative of future subsonic engines. Results to date represent an improved understanding of the formation of nitrogen oxides at these high pressures (twice the pressure of previous combustor tests) and temperatures.

Lyons, Valerie J.↗
NASA NTRS2005

Coherent Anti-Stokes Raman Spectroscopy (CARS) Measurements in Supersonic Combustors at NASA Langley Research Center

This paper describes the recent use of coherent anti-Stokes Raman spectroscopy (CARS) to study supersonic combustion at NASA Langley Research Center. CARS is a nonlinear optical measurement technique used to measure temperature and species mole fractions remotely in harsh environments. A CARS system has been applied to two different combustor geometries at NASA Langley. Both experiments used the same vitiated wind-tunnel facility to create an air flow that simulates flight at Mach numbers of 6 and 7 for the combustor inlet and both experiments used hydrogen fuel. In the first experiment, the hydrogen was injected supersonically at a 30-degree angle with respect to the incoming flow. In the second experiment, the hydrogen was injected sonically at normal incidence. While these injection schemes produced significantly different flow features, the CARS method provided mean temperature, N2, O2 and H2 maps at multiple downstream locations for both. The primary aim of these measurements was to provide detailed flowfield information for computational fluid dynamics (CFD) code validation.

Danehy, Paul M.↗
NASA NTRS2001

Laser High-Cycle Thermal Fatigue of Pulse Detonation Engine Combustor Materials Tested

Pulse detonation engines (PDE's) have received increasing attention for future aerospace propulsion applications. Because the PDE is designed for a high-frequency, intermittent detonation combustion process, extremely high gas temperatures and pressures can be realized under the nearly constant-volume combustion environment. The PDE's can potentially achieve higher thermodynamic cycle efficiency and thrust density in comparison to traditional constant-pressure combustion gas turbine engines (ref. 1). However, the development of these engines requires robust design of the engine components that must endure harsh detonation environments. In particular, the detonation combustor chamber, which is designed to sustain and confine the detonation combustion process, will experience high pressure and temperature pulses with very short durations (refs. 2 and 3). Therefore, it is of great importance to evaluate PDE combustor materials and components under simulated engine temperatures and stress conditions in the laboratory. In this study, a high-cycle thermal fatigue test rig was established at the NASA Glenn Research Center using a 1.5-kW CO2 laser. The high-power laser, operating in the pulsed mode, can be controlled at various pulse energy levels and waveform distributions. The enhanced laser pulses can be used to mimic the time-dependent temperature and pressure waves encountered in a pulsed detonation engine. Under the enhanced laser pulse condition, a maximum 7.5-kW peak power with a duration of approximately 0.1 to 0.2 msec (a spike) can be achieved, followed by a plateau region that has about one-fifth of the maximum power level with several milliseconds duration. The laser thermal fatigue rig has also been developed to adopt flat and rotating tubular specimen configurations for the simulated engine tests. More sophisticated laser optic systems can be used to simulate the spatial distributions of the temperature and shock waves in the engine. Pulse laser high-cycle thermal fatigue behavior has been investigated on a flat Haynes 188 alloy specimen, under the test condition of 30-Hz cycle frequency (33-msec pulse period and 10-msec pulse width including a 0.2-msec pulse spike; ref. 4). Temperature distributions were calculated with one-dimensional finite difference models. The calculations show that that the 0.2-msec pulse spike can cause an additional 40 C temperature fluctuation with an interaction depth of 0.08 mm near the specimen surface region. This temperature swing will be superimposed onto the temperature swing of 80 C that is induced by the 10-msec laser pulse near the 0.53-mm-deep surface interaction region.

Zhu, Dong-Ming↗
NASA NTRS2005

Ejector Enhanced Pulsejet Based Pressure Gain Combustors: An Old Idea With a New Twist

An experimental investigation of pressure-gain combustion for gas turbine application is described. The test article consists of an off-the-shelf valved pulsejet, and an optimized ejector, both housed within a shroud. The combination forms an effective can combustor across which there is a modest total pressure rise rather than the usual loss found in conventional combustors. Although the concept of using a pulsejet to affect semi-constant volume (i.e., pressure-gain) combustion is not new, that of combining it with a well designed ejector to efficiently mix the bypass flow is. The result is a device which to date has demonstrated an overall pressure rise of approximately 3.5 percent at an overall temperature ratio commensurate with modern gas turbines. This pressure ratio is substantially higher than what has been previously reported in pulsejet-based combustion experiments. Flow non-uniformities in the downstream portion of the device are also shown to be substantially reduced compared to those within the pulsejet itself. The standard deviation of total pressure fluctuations, measured just downstream of the ejector was only 5.0 percent of the mean. This smoothing aspect of the device is critical to turbomachinery applications since turbine performance is, in general, negatively affected by flow non-uniformities and unsteadiness. The experimental rig will be described and details of the performance measurements will be presented. Analyses showing the thermodynamic benefits from this level of pressure-gain performance in a gas turbine will also be assessed for several engine types. Issues regarding practical development of such a device are discussed, as are potential emissions reductions resulting from the rich burning nature of the pulsejet and the rapid mixing (quenching) associated with unsteady ejectors.

Paxson, Daniel E.↗
NASA NTRS2006

Restricted Acoustic Modal Analysis Applied to Internal Combustor Spectra and Cross-Spectra Measurements

A treatment of the modal decomposition of the pressure field in a combustor as determined by two Kulite pressure measurements is developed herein. It is applied to a Pratt & Whitney PW4098 engine combustor over a range of operating conditions. For modes other than the plane wave the new part of the treatment is the assumption that there are distinct frequency bands in which the individual modes, including the plane wave mode, overlap such that if circumferential mode m and circumferential mode m-1 are present than circumferential mode m 2 is not. Consequently, in the analysis used herein at frequencies above the first cut-off mode frequency, only pairs of circumferential modes are individually present at each frequency. Consequently, this is a restricted modal analysis. A new result is that the successful use of the same modal span frequencies over a range of operating conditions for this particular engine suggests that the temperature, T, and the velocity, v, of the flow at each operating condition are related by c(sup 2)-v(sup 2) = a constant where c is the speed of sound.

Miles, Jeffrey Hilton↗
NASA NTRS2007

Low-Emission Hydrogen Combustors for Gas Turbines Using Lean Direct Injection

One of the key technology challenges for the use of hydrogen in gas turbine engines is the performance of the combustion system, in particular the fuel injectors. To investigate the combustion performance of gaseous hydrogen fuel injectors flame tube combustor experiments were performed. Tests were conducted to measure the nitrogen oxide (NO(x)) emissions and combustion performance at inlet conditions of 588 to 811 K, 0.4 to 1.4 MPa, and equivalence ratios up to 0.48. All the injectors were based on Lean Direct Injection (LDI) technology with multiple injection points and quick mixing. One challenge to hydrogen-based premixing combustion systems is flashback since hydrogen has a reaction rate over 7 times that of Jet-A. To reduce the risk, design mixing times were kept short and velocities high to minimize flashback. Five fuel injector designs were tested in 6.35- and 8.9-cm-diameter flame tubes with non-vitiated heated air and gaseous hydrogen. Data is presented on measurements of NO(x) emissions and combustion efficiency for the hydrogen injectors at 2.540, 7.937, and 13.652 cm from the injector face. Results show that for some configurations, NO(x) emissions are comparable to that of state of the art Jet-A LDI combustor concepts.

Marek, C. John↗
NASA NTRS2005

Low Emission Hydrogen Combustors for Gas Turbines Using Lean Direct Injection

One of the key technology challenges for the use of hydrogen in gas turbine engines is the performance of the combustion system, in particular the fuel injectors. To investigate the combustion performance of gaseous hydrogen fuel injectors flame tube combustor experiments were performed. Tests were conducted to measure the nitrogen oxide (NOx) emissions and combustion performance at inlet conditions of 600 to 1000 deg F, 60 to 200 pounds per square inch absolute (psia), and equivalence ratios up to 0.48. All the injectors were based on Lean Direct Injection (LDI) technology with multiple injection points and quick mixing. One challenge to hydrogen based premixing combustion systems is flashback since hydrogen has a reaction rate over seven times that of Jet-A. To reduce the risk, design mixing times were kept short and velocities high to minimize flashback. Five fuel injector designs were tested in 2.5 and 3.5-in. diameter flame tubes with non-vitiated heated air and gaseous hydrogen. Data is presented on measurements of NOx emissions and combustion efficiency for the hydrogen injectors at 1.0, 3.125, and 5.375 in. from the injector face. Results show that for some configurations, NOx emissions are comparable to that of state of the art Jet-A LDI combustor concepts.

Marek, C. John↗
NASA NTRS1994

Combustor liner support assembly

A support assembly for a gas turbine engine combustor includes an annular frame having a plurality of circumferentially spaced apart tenons, and an annular combustor liner disposed coaxially with the frame and including a plurality of circumferentially spaced apart tenons circumferentially adjoining respective ones of the frame tenons for radially and tangentially supporting the liner to the frame while allowing unrestrained differential thermal radial movement therebetween.

Halila, Ely E.↗
NASA NTRS1980

Ring connection for porous combustor wall panels

A gas turbine engine combustor assembly of unique configuration has an outer wall made up of a plurality of axially extending multi-layered porous metal panels joined together at butt joints therebetween by a reinforcing and heat dissipation ring and a unique weld configuration to prevent thermal erosion of the ends of the porous metal panels at the butt joints; the combustor further including a unique inner wall made up of a plurality of like axially extending multi-layered porous metal panels joined at butt joints by a reinforcing and heat dissipation ring on the inner surface of the inner wall panels and an improved butt weld joint that prevents thermal erosion of the ends of the porous metal inner wall panels.

Verdouw, Albert J.↗
NASA NTRS1999

Turbofan engine with a low pressure turbine driven supercharger in a bypass duct operated by a fuel rich combustor and an afterburner

A multiple bypass turbofan engine includes a core Brayton Cycle gas generator with a fuel rich burning combustor and is provided with a variable supercharged bypass duct around the gas generator with a supercharging means in the supercharged bypass duct powered by a turbine not mechanically connected to the gas generator. The engine further includes a low pressure turbine driven forward fan upstream and forward of an aft fan and drivingly connected to a low pressure turbine by a low pressure shaft, the low pressure turbine being aft of and in serial flow communication with the core gas generator. A fan bypass duct is disposed radially outward of the core engine assembly and has first and second inlets disposed between the forward and aft fans. An inlet duct having an annular duct wall is disposed radially inward of the bypass duct and connects the second inlet to the bypass duct. A supercharger means for compressing air is drivingly connected to the low pressure turbine and is disposed in the inlet duct. A secondary combustor or augmentor is disposed in an exhaust duct downstream of and in fluid flow communication with the bypass duct and the gas generator.

Bartos, James W.↗
NASA NTRS2001

Modular combustor dome

A combustor dome module includes a mixer tube having a hollow heat shield sealingly joined around the outlet end thereof. The modules may then be assembled in an array for defining the combustor dome, with each module being individually removable therefrom.

Glynn, Christopher Charles↗
NASA NTRS2001

Combustor with fuel preparation chambers

An annular combustor having fuel preparation chambers mounted in the dome of the combustor. The fuel preparation chamber comprises an annular wall extending axially from an inlet to an exit that defines a mixing chamber. Mounted to the inlet are an air swirler and a fuel atomizer. The air swirler provides swirled air to the mixing chamber while the atomizer provides a fuel spray. On the downstream side of the exit, the fuel preparation chamber has an inwardly extending conical wall that compresses the swirling mixture of fuel and air exiting the mixing chamber.

Zelina, Joseph↗
NASA NTRS2002

Fuel injection assembly for gas turbine engine combustor

A fuel injection assembly for a gas turbine engine combustor, including at least one fuel stem, a plurality of concentrically disposed tubes positioned within each fuel stem, wherein a cooling supply flow passage, a cooling return flow passage, and a tip fuel flow passage are defined thereby, and at least one fuel tip assembly connected to each fuel stem so as to be in flow communication with the flow passages, wherein an active cooling circuit for each fuel stem and fuel tip assembly is maintained by providing all active fuel through the cooling supply flow passage and the cooling return flow passage during each stage of combustor operation. The fuel flowing through the active cooling circuit is then collected so that a predetermined portion thereof is provided to the tip fuel flow passage for injection by the fuel tip assembly.

Candy, Anthony J.↗
NASA NTRS2006

Fuel-air mixing apparatus for reducing gas turbine combustor exhaust emissions

A fuel-air mixer for use in a combustion chamber of a gas turbine engine is provided. The fuel air mixing apparatus comprises an annular fuel injector having a plurality of discrete plain jet orifices, a first swirler wherein the first swirler is located upstream from the fuel injector and a second swirler wherein the second swirler is located downstream from the fuel injector. The plurality of discrete plain jet orifices are situated between the highly swirling airstreams generated by the two radial swirlers. The distributed injection of the fuel between two highly swirling airstreams results in rapid and effective mixing to the desired fuel-air ratio and prevents the formation of local hot spots in the combustor primary zone. A combustor and a gas turbine engine comprising the fuel-air mixer of the present invention are also provided as well as a method using the fuel-air mixer of the present invention.

Zupanc, Frank J.↗