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At least 415 records · Page 23
All sourcesNASA NTRSDOE OSTI
NASA NTRS2004

Advanced Oxide Material Systems for 1650 C Thermal/Environmental Barrier Coating Applications

Advanced thermal and environmental barrier coatings (TEBCs) are being developed for low-emission SiC/SiC ceramic matrix composite (CMC) combustor and vane applications to extend the CMC liner and vane temperature capability to 1650 C (3000 F) in oxidizing and water-vapor-containing combustion environments. The advanced 1650 C TEBC system is required to have a better high-temperature stability, lower thermal conductivity, and more resistance to sintering and thermal stress than current coating systems under engine high-heat-flux and severe thermal cycling conditions. In this report, the thermal conductivity and water vapor stability of selected candidate hafnia-, pyrochlore- and magnetoplumbite-based TEBC materials are evaluated. The effects of dopants on the materials properties are also discussed. The test results have been used to downselect the TEBC materials and help demonstrate the feasibility of advanced 1650 C coatings with long-term thermal cycling durability.

Zhu, Dong-Ming↗
NASA NTRS2005

Low NO(x) Combustor Development

The goal of these efforts was the development of an ultra-low emissions, lean-burn combustor for the High Speed Civil Transport. The HSCT Mach 2.4 FLADE C1 Cycle was selected as the baseline engine cycle. A preliminary compilation of performance requirements for the HSCT combustor system was developed. The emissions goals of the program, baseline engine cycle, and standard combustor performance requirements were considered in developing the compilation of performance requirements. Seven combustor system designs were developed. The development of these system designs was facilitated by the use of spreadsheet-type models which predicted performance of the combustor systems over the entire flight envelope of the HSCT. A chemical kinetic model was developed for an LPP combustor and employed to study NO(x) formation kinetics, and CO burnout. These predictions helped to define the combustor residence time. Five fuel-air mixer concepts were analyzed for use in the combustor system designs. One of the seven system designs, one using the Swirl-Jet and Cyclone Swirler fuel-air mixers, was selected for a preliminary mechanical design study.

Kastl, J. A.↗
NASA NTRS2005

Channels and Fan-like Features on Titan Surface Imaged by the Cassini RADAR

During two close flybys of Titan on October 26, 2004, and February 15, 2005, the Cassini s radar instrument acquired synthetic-aperture radar (SAR) data revealing Titan s complex surface and intriguing geological features. Fan-like and apparently flow-related features are connected to sinuous and linear features which resemble channels. The fan-like features and channels appear to be relatively SAR-bright and suggest surface roughness properties at the scale and bigger than the Ku-band, and possible volume scattering. A strong correlation between the SAR-bright and radiometric cold regions has been observed. The correlation is consistent with radiometric cold areas being caused by volume scattering at Ku as due to broken low-loss ice and resulting low emissivity as with the surfaces of Europa and Ganymede.

Paganelli, F.↗
NASA NTRS1998

Mixing of Multiple Jets With a Confined Subsonic Crossflow

Results from a recently completed enhanced mixing program are summarized in the two technical papers. These studies were parts of a High Speed Research (HSR)-supported joint Government/industry/university program that involved, in addition to the NASA Lewis Research Center, researchers at United Technologies Research Center, Allison Engine Company, CFD Research Corporation, and the University of California, Irvine. The studies investigated the mixing of jets injected normal to a confined subsonic mainsteam in both rectangular and cylindrical ducts. Experimental and computational studies were performed in both nonreacting and reacting flows. The orifice geometries and flow conditions were selected as typical of the complex three-dimensional flows in the combustion chambers in low-emission gas turbine engines. The principal conclusion from both the experiments and modeling was that the momentum-flux ratio J and orifice spacing S/H were the most significant flow and geometry variables, respectively. Conserved scalar distributions were similar-independent of reaction, orifice diameter H/d, and shape-when the orifice spacing and the square root of the momentum-flux ratio were inversely proportional. Jet penetration was critical, and penetration decreased as either momentum-flux ratio or orifice spacing decreased. We found that planar averages must be considered in context with the distributions. The mass-flow ratios and the orifices investigated were often very large. The jet-to-mainstream mass-flow ratio was varied from significantly less than 1 to greater than 1. The orifice-area to mainstream-cross-sectional-area was varied from approx. 0 to 0.5, and the axial planes of interest were often just downstream of the orifice trailing edge. Three-dimensional flow was a key part of efficient mixing and was observed for all configurations. As an example of the results, the accompanying figure shows the effects of different rates of mass addition on the opposite walls of a rectangular duct.

Holdeman, James D.↗
NASA NTRS1998

Advanced Subsonic Combustion Rig

Researchers from the NASA Lewis Research Center have obtained the first combustion/emissions data under extreme future engine operating conditions. In Lewis' new world-class 60-atm combustor research facility--the Advanced Subsonic Combustion Rig (ASCR)--a flametube was used to conduct combustion experiments in environments as extreme as 900 psia and 3400 F. The greatest challenge for combustion researchers is the uncertainty of the effects of pressure on the formation of nitrogen oxides (NOx). Consequently, U.S. engine manufacturers are using these data to guide their future combustor designs. The flametube's metal housing has an inside diameter of 12 in. and a length of 10.5 in. The flametube can be used with a variety of different flow paths. Each flow path is lined with a high-temperature, castable refractory material (alumina) to minimize heat loss. Upstream of the flametube is the injector section, which has an inside diameter of 13 in. and a length of 0.5-in. It was designed to provide for quick changeovers. This flametube is being used to provide all U.S. engine manufacturers early assessments of advanced combustion concepts at full power conditions prior to engine production. To date, seven concepts from engine manufacturers have been evaluated and improved. This collaborated development can potentially give U.S. engine manufacturers the competitive advantage of being first in the market with advanced low-emission technologies.

Lee, Chi-Ming↗
NASA NTRS1999

Combustion Instabilities Modeled

NASA Lewis Research Center's Advanced Controls and Dynamics Technology Branch is investigating active control strategies to mitigate or eliminate the combustion instabilities prevalent in lean-burning, low-emission combustors. These instabilities result from coupling between the heat-release mechanisms of the burning process and the acoustic flow field of the combustor. Control design and implementation require a simulation capability that is both fast and accurate. It must capture the essential physics of the system, yet be as simple as possible. A quasi-one-dimensional, computational fluid dynamics (CFD) based simulation has been developed which may meet these requirements. The Euler equations of mass, momentum, and energy have been used, along with a single reactive species transport equation to simulate coupled thermoacoustic oscillations. A very simple numerical integration scheme was chosen to reduce computing time. Robust boundary condition procedures were incorporated to simulate various flow conditions (e.g., valves, open ends, and choked inflow) as well as to accommodate flow reversals that may arise during large flow-field oscillations. The accompanying figure shows a sample simulation result. A combustor with an open inlet, a choked outlet, and a large constriction approximately two thirds of the way down the length is shown. The middle plot shows normalized, time-averaged distributions of the relevant flow quantities, and the bottom plot illustrates the acoustic mode shape of the resulting thermoacoustic oscillation. For this simulation, the limit cycle peak-to-peak pressure fluctuations were 13 percent of the mean. The simulation used 100 numerical cells. The total normalized simulation time was 50 units (approximately 15 oscillations), which took 26 sec on a Sun Ultra2.

Paxson, Daniel E.↗
NASA NTRS2000

Engine-Scale Combustor Rig Designed, Fabricated, and Tested for Combustion Instability Control Research

Low-emission combustor designs are prone to combustor instabilities. Because active control of these instabilities may allow future combustors to meet both stringent emissions and performance requirements, an experimental combustor rig was developed for investigating methods of actively suppressing combustion instabilities. The experimental rig has features similar to a real engine combustor and exhibits instabilities representative of those in aircraft gas turbine engines. Experimental testing in the spring of 1999 demonstrated that the rig can be tuned to closely represent an instability observed in engine tests. Future plans are to develop and demonstrate combustion instability control using this experimental combustor rig. The NASA Glenn Research Center at Lewis Field is leading the Combustion Instability Control program to investigate methods for actively suppressing combustion instabilities. Under this program, a single-nozzle, liquid-fueled research combustor rig was designed, fabricated, and tested. The rig has many of the complexities of a real engine combustor, including an actual fuel nozzle and swirler, dilution cooling, and an effusion-cooled liner. Prior to designing the experimental rig, a survey of aircraft engine combustion instability experience identified an instability observed in a prototype engine as a suitable candidate for replication. The frequency of the instability was 525 Hz, with an amplitude of approximately 1.5-psi peak-to-peak at a burner pressure of 200 psia. The single-nozzle experimental combustor rig was designed to preserve subcomponent lengths, cross sectional area distribution, flow distribution, pressure-drop distribution, temperature distribution, and other factors previously found to be determinants of burner acoustic frequencies, mode shapes, gain, and damping. Analytical models were used to predict the acoustic resonances of both the engine combustor and proposed experiment. The analysis confirmed that the test rig configuration and engine configuration had similar longitudinal acoustic characteristics, increasing the likelihood that the engine instability would be replicated in the rig. Parametric analytical studies were performed to understand the influence of geometry and condition variations and to establish a combustion test plan. Cold-flow experiments verified that the design values of area and flow distributions were obtained. Combustion test results established the existence of a longitudinal combustion instability in the 500-Hz range with a measured amplitude approximating that observed in the engine. Modifications to the rig configuration during testing also showed the potential for injector independence. The research combustor rig was developed in partnership with Pratt & Whitney of West Palm Beach, Florida, and United Technologies Research Center of East Hartford, Connecticut. Experimental testing of the combustor rig took place at United Technologies Research Center.

DeLaat, John C.↗
NASA NTRS2001

One-Dimensional Spontaneous Raman Measurements Made in a Gas Turbine Combustor

The NASA Glenn Research Center and the aerospace industry are designing and testing low-emission combustor concepts to build the next generation of cleaner, more fuel efficient aircraft powerplants. These combustors will operate at much higher inlet temperatures and at pressures that are up to 3 to 5 times greater than combustors in the current fleet. From a test and analysis viewpoint, there is an increasing need for measurements from these combustors that are nonintrusive, simultaneous, multipoint, and more quantitative. Glenn researchers have developed several unique test facilities (refs. 1 and 2) that allow, for the first time, optical interrogation of combustor flow fields, including subcomponent performance, at pressures ranging from 1 to 60 bar (1 to 60 atm). Experiments conducted at Glenn are the first application of a visible laser-pumped, one-dimensional, spontaneous Raman-scattering technique to analyze the flow in a high-pressure, advanced-concept fuel injector at pressures thus far reaching 12 bar (12 atm). This technique offers a complementary method to the existing two- and three-dimensional imaging methods used, such as planar laser-induced fluorescence. Raman measurements benefit from the fact that the signal from each species is a linear function of its density, and the relative densities of all major species can be acquired simultaneously with good precision. The Raman method has the added potential to calibrate multidimensional measurements by providing an independent measurement of species number-densities at known points within the planar laser-induced fluorescence images. The visible Raman method is similar to an ultraviolet-Raman technique first tried in the same test facility (ref. 3). However, the visible method did not suffer from the ultraviolet technique's fuel-born polycyclic aromatic hydrocarbon fluorescence interferences.

DeGroot, Wilhelmus A.↗
NASA NTRS2004

Advanced Oxide Material Systems For 1650 C Thermal/Environmental Barrier Coating Applications

Advanced thermal/environmental barrier coatings (T/EBCs) are being developed for low emission SiC/SiC ceramic matrix composite (CMC) combustor and vane applications to extend the CMC liner and vane temperature capability to 1650 C (3000 F) in oxidizing and water-vapor containing combustion environments. The 1650 C T/EBC system is required to have better thermal stability, lower thermal conductivity, and improved sintering and thermal stress resistance than current coating systems. In this paper, the thermal conductivity, water vapor stability and cyclic durability of selected candidate zirconia-/hafnia-, pyrochlore- and magnetoplumbite-based T/EBC materials are evaluated. The test results have been used to downselect the T/EBC coating materials, and help demonstrate advanced 1650OC coatings feasibility with long-term cyclic durability.

Zhu, Dongming↗
NASA NTRS2003

National Combustion Code Validated Against Lean Direct Injection Flow Field Data

Most combustion processes have, in some way or another, a recirculating flow field. This recirculation stabilizes the reaction zone, or flame, but an unnecessarily large recirculation zone can result in high nitrogen oxide (NOx) values for combustion systems. The size of this recirculation zone is crucial to the performance of state-of-the-art, low-emissions hardware. If this is a large-scale combustion process, the flow field will probably be turbulent and, therefore, three-dimensional. This research dealt primarily with flow fields resulting from lean direct injection (LDI) concepts, as described in Research & Technology 2001. LDI is a concept that depends heavily on the design of the swirler. The LDI concept has the potential to reduce NOx values from 50 to 70 percent of current values, with good flame stability characteristics. It is cost effective and (hopefully) beneficial to do most of the design work for an LDI swirler using computer-aided design (CAD) and computer-aided engineering (CAE) tools. Computational fluid dynamics (CFD) codes are CAE tools that can calculate three-dimensional flows in complex geometries. However, CFD codes are only beginning to correctly calculate the flow fields for complex devices, and the related combustion models usually remove a large portion of the flow physics.

Iannetti, Anthony C.↗
NASA NTRS2006

A Two-dimensional Cartesian and Axisymmetric Study of Combustion-acoustic Interaction

This paper describes a study of a lean premixed (LP) methane-air combustion wave in a two-dimensional Cartesian and axisymmetric coordinate system. Lean premixed combustors provide low emission and high efficiency; however, they are susceptible to combustion instabilities. The present study focuses on the behavior of the flame as it interacts with an external acoustic disturbance. It was found that the flame oscillations increase as the disturbance amplitude is increased. Furthermore, when the frequency of the disturbance is at resonance with a chamber frequency, the instabilities increase. For the axisymmetric geometry, the flame is found to be more unstable compared to the Cartesian case. In some cases, these instabilities were severe and led to flame extinction. In the axisymmetric case, several passive control devices were tested to assess their effectiveness. It is found that an acoustic cavity is better able at controlling the pressure fluctuations in the chamber.

Hood, Caroline↗
NASA NTRS2007

An Overview of Aerospace Propulsion Research at NASA Glenn Research Center

NASA Glenn Research center is the recognized leader in aerospace propulsion research, advanced technology development and revolutionary system concepts committed to meeting the increasing demand for low noise, low emission, high performance, and light weight propulsion systems for affordable and safe aviation and space transportation needs. The technologies span a broad range of areas including air breathing, as well as rocket propulsion systems, for commercial and military aerospace applications and for space launch, as well as in-space propulsion applications. The scope of work includes fundamentals, components, processes, and system interactions. Technologies developed use both experimental and analytical approaches. The presentation provides an overview of the current research and technology development activities at NASA Glenn Research Center .

Reddy, D. R.↗
NASA NTRS2006

HSCT Anticipated Seal Needs Turbomachinery Seals Combustor Seals

The High Speed Civil Transport (HSCT) engine concept is a large mixed flow turbofan similar in construction to current military fighter engines. The mission, however, is quite different. The engine will operate for long periods of time at very high Mach numbers and high altitudes. The engine is required to have very low emissions and noise levels to be acceptable in commercial service. Current thrust levels are in the 55000 lb range. At the current supercruise speed requirement of Mach 2.4, the engine inlet temperature will be at least 380 F. This is the lowest cycle temperature expected anywhere in the propulsion system.Seals will be exposed to operate at this temperature and higher for thousands of hours without failure. Durability, cost, and weight will all be very important in determining the type of seals selected for a successful HSCT engine.

Henry, John↗
NASA NTRS2007

Solar Airplanes and Regenerative Fuel Cells

A solar electric aircraft with the potential to "fly forever" has captured NASA's interest, and the concept for such an aircraft was pursued under Aeronautics Environmental Research Aircraft and Sensor Technology (ERAST) project. Feasibility of this aircraft happens to depend on the successful development of solar power technologies critical to NASA's Exploration Initiatives; hence, there was widespread interest throughout NASA to bring these technologies to a flight demonstration. The most critical is an energy storage system to sustain mission power during night periods. For the solar airplane, whose flight capability is already limited by the diffuse nature of solar flux and subject to latitude and time of year constraints, the feasibility of long endurance flight depends on a storage density figure of merit better than 400-600 watt-hr per kilogram. This figure of merit is beyond the capability of present day storage technologies (other than nuclear) but may be achievable in the hydrogen-oxygen regenerative fuel cell (RFC). This potential has led NASA to undertake the practical development of a hydrogen-oxygen regenerative fuel cell, initially as solar energy storage for a high altitude UAV science platform but eventually to serve as the primary power source for NASAs lunar base and other planet surface installations. Potentially the highest storage capacity and lowest weight of any non-nuclear device, a flight-weight RFC aboard a solar-electric aircraft that is flown continuously through several successive day-night cycles will provide the most convincing demonstration that this technology's widespread potential has been realized. In 1998 NASA began development of a closed cycle hydrogen oxygen PEM RFC under the Aeronautics Environmental Research Aircraft and Sensor Technology (ERAST) project and continued its development, originally for a solar electric airplane flight, through FY2005 under the Low Emissions Alternative Power (LEAP) project. Construction of the closed loop system began in 2002 at the NASA Glenn Research Center in Cleveland, Ohio. System checkout was completed, and testing began, in July of 2003. The initial test sequences were done with only a fuel cell or electrolyzer in the test rig. Those tests were used to verify the test apparatus, procedures, and software. The first complete cycles of the fully closed loop, regenerative fuel cell system were successfully completed in the following September. Following some hardware upgrades to increase reactant recirculation flow, the test rig was operated at full power in December 2003 and again in January 2004. In March 2004 a newer generation of fuel cell and electrolyzer stacks was substituted for the original hardware and these stacks were successfully tested at full power under cyclic operation in June of 2004.

Bents, David J.↗
NASA NTRS2007

High-Heat-Flux Cyclic Durability of Thermal and Environmental Barrier Coatings

Advanced ceramic thermal and environmental barrier coatings will play an increasingly important role in future gas turbine engines because of their ability to protect the engine components and further raise engine temperatures. For the supersonic vehicles currently envisioned in the NASA fundamental aeronautics program, advanced gas turbine engines will be used to provide high power density thrust during the extended supersonic flight of the aircraft, while meeting stringent low emission requirements. Advanced ceramic coating systems are critical to the performance, life and durability of the hot-section components of the engine systems. In this work, the laser and burner rig based high-heat-flux testing approaches were developed to investigate the coating cyclic response and failure mechanisms under simulated supersonic long-duration cruise mission. The accelerated coating cracking and delamination mechanism under the engine high-heat-flux, and extended supersonic cruise time conditions will be addressed. A coating life prediction framework may be realized by examining the crack initiation and propagation in conjunction with environmental degradation under high-heat-flux test conditions.

Zhu, Dongming↗
NASA NTRS2008

Thermal and Environmental Barrier Coating Development for Advanced Propulsion Engine Systems

Ceramic thermal and environmental barrier coatings (TEBCs) are used in gas turbine engines to protect engine hot-section components in the harsh combustion environments, and extend component lifetimes. Advanced TEBCs that have significantly lower thermal conductivity, better thermal stability and higher toughness than current coatings will be beneficial for future low emission and high performance propulsion engine systems. In this paper, ceramic coating design and testing considerations will be described for turbine engine high temperature and high-heat-flux applications. Thermal barrier coatings for metallic turbine airfoils and thermal/environmental barrier coatings for SiC/SiC ceramic matrix composite (CMC) components for future supersonic aircraft propulsion engines will be emphasized. Further coating capability and durability improvements for the engine hot-section component applications can be expected by utilizing advanced modeling and design tools.

Zhu, Dongming↗
NASA NTRS2009

An Overview of NASA's Contributions to Energy Technology

The National Aeronautics and Space Administration (NASA) is well known for its many contributions to advancing technology for the aviation and space industries. It may be surprising to some that it has also made a major impact in advancing energy technologies. This paper presents a historic overview of some of the energy programs that NASA was involved in, as well as presenting some current energy-related work that is relevant to both aerospace and non-aerospace needs. In the past, NASA developed prototype electric cars, low-emission gas turbines, wind turbines, and solar-powered villages, to name a few of the major energy projects. The fundamental expertise in fluid mechanics, heat transfer, thermodynamics, mechanical and electrical engineering, and other related fields, found in NASA s workforce, can easily be applied to develop creative solutions to energy problems in space, aviation, or terrestrial systems.

Lyons, Valerie J.↗
NASA NTRS2009

NEXT Long-Duration Test Neutralizer Performance and Erosion Characteristics

The NASA's Evolutionary Xenon Thruster (NEXT) program is developing the next-generation ion propulsion system with significant enhancements beyond the state-of-the-art to provide future NASA science missions with enhanced capabilities at a low total development cost. A Long-Duration Test (LDT) was initiated in June 2005, to verify the NEXT propellant throughput capability to a qualification-level of 450 kg, 1.5 times the anticipated throughput requirement of 300 kg per thruster based on mission analyses. As of September 2, 2009, the thruster has accumulated 24,400 hr of operation with extensive durations at the following input powers: 6.9, 4.7, 1.1, and 0.5 kW. The thruster has processed 434 kg of xenon, surpassing the NASA Solar Technology Application Readiness (NSTAR) program thruster propellant throughput demonstrated during the extended life testing of the Deep Space 1 flight spare ion thruster and approaching the NEXT development qualification throughput goal of 450 kg. The NEXT LDT has demonstrated a total impulse of 16.1 10(exp 6zzz0 N s; the highest total impulse ever demonstrated by an ion thruster. A reduction in neutralizer flow margin has been the only appreciable source of thruster performance degradation. The behavior of the neutralizer is not easily predicted due to both erosion and deposition observed in previous wear tests. Spot-to-plume mode transition flow data and in-situ erosion results for the LDT neutralizer are discussed. This loss of flow margin has been addressed through a combination of a design change in the prototype-model neutralizer to increase flow margin at low emission current and to update the NEXT throttle table to ensure adequate flow margin as a function of propellant throughput processed. The new throttle table will be used for future LDT operations. The performance of the NEXT LDT neutralizer is consistent with that observed for long-life hollow cathodes. The neutralizer life-limiting failure modes are progressing as expected and the neutralizer data indicate none of the neutralizer failures are imminent.

Herman, Daniel A.↗