SCIENTIFIC AND TECHNICAL AEROSPACE REPORTS
A Biweekly Publication of the National Aeronautics and Space Administration
VOLUME 43, ISSUE 16 - AUGUST 12, 2005
07 AIRCRAFT PROPULSION AND POWER
Includes primary propulsion systems and related systems and components, e.g., gas turbine engines, compressors, and fuel systems; and onboard auxiliary power plants for aircraft.
For related information see also 20 Spacecraft Propulsion and Power; 28 Propellants and Fuels; and 44 Energy Production and Conversion.
20050192615 Cleveland State Univ., Cleveland, OH, USA, NASA Glenn Research Center, Cleveland, OH, USA
Vibration-Based Method Developed to Detect Cracks in Rotors During Acceleration Through Resonance
Sawicki, Jerzy T.; Baaklini, George Y.; Gyekenyesi, Andrew L.; Research and Technology 2003; May 2004; 4 pp.; In English; No Copyright; Avail: CASI; A01, Hardcopy
In recent years, there has been an increasing interest in developing rotating machinery shaft crack-detection methodologies and online techniques. Shaft crack problems present a significant safety and loss hazard in nearly every application of modern turbomachinery. In many cases, the rotors of modern machines are rapidly accelerated from rest to operating speed, to reduce the excessive vibrations at the critical speeds. The vibration monitoring during startup or shutdown has been receiving growing attention (ref. 1), especially for machines such as aircraft engines, which are subjected to frequent starts and stops, as well as high speeds and acceleration rates. It has been recognized that the presence of angular acceleration strongly affects the rotor’s maximum response to unbalance and the speed at which it occurs. Unfortunately, conventional nondestructive evaluation (NDE) methods have unacceptable limits in terms of their application for online crack detection. Some of these techniques are time consuming and inconvenient for turbomachinery service testing. Almost all of these techniques require that the vicinity of the damage be known in advance, and they can provide only local information, with no indication of the structural strength at a component or system level. In addition, the effectiveness of these experimental techniques is affected by the high measurement noise levels existing in complex turbomachine structures. Therefore, the use of vibration monitoring along with vibration analysis has been receiving increasing attention. Derived from text
Rotors; Vibration
20050194591 NASA Glenn Research Center, Cleveland, OH, USA
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Acoustics and Thrust of Separate Flow Exhaust Nozzles With Mixing Devices Investigated for High Bypass Ratio Engines
Saiyed, Naseem H.; Research and Technology 1999; March 2000; 2 pp.; In English; Original contains black and white illustrations; No Copyright; Avail: CASI; A01, Hardcopy
Typical installed separate-flow exhaust nozzle system. The jet noise from modern turbofan engines is a major contributor to the overall noise from commercial aircraft. Many of these engines use separate nozzles for exhausting core and fan streams. As a part of NASA s Advanced Subsonic Technology (AST) program, the NASA Glenn Research Center at Lewis Field led an experimental investigation using model-scale nozzles in Glenn s Aero-Acoustic Propulsion Laboratory. The goal of the investigation was to develop technology for reducing the jet noise by 3 EPNdB. Teams of engineers from Glenn, the NASA Langley Research Center, Pratt & Whitney, United Technologies Research Corporation, the Boeing Company, GE Aircraft Engines, Allison Engine Company, and Aero Systems Engineering contributed to the planning and implementation of the test. Derived from text
Aeroacoustics; Thrust; Turbine Exhaust Nozzles; Exhaust Nozzles
20050195862 NASA Glenn Research Center, Cleveland, OH, USA
Probabilistic Risk-Based Approach to Aeropropulsion System Assessment Developed
Tong, Michael T.; Research and Technology 2000; March 2001; 3 pp.; In English; Original contains black and white illustrations; No Copyright; Avail: CASI; A01, Hardcopy
In an era of shrinking development budgets and resources, where there is also an emphasis on reducing the product development cycle, the role of system assessment, performed in the early stages of an engine development program, becomes very critical to the successful development of new aeropropulsion systems. A reliable system assessment not only helps to identify the best propulsion system concept among several candidates, it can also identify which technologies are worth pursuing. This is particularly important for advanced aeropropulsion technology development programs, which require an enormous amount of resources. In the current practice of deterministic, or point-design, approaches, the uncertainties of design variables are either unaccounted for or accounted for by safety factors. This could often result in an assessment with unknown and unquantifiable reliability. Consequently, it would fail to provide additional insight into the risks associated with the new technologies, which are often needed by decisionmakers to determine the feasibility and return-on-investment of a new aircraft engine. Derived from text
Aircraft Engines; Engine Design; Propulsion System Configurations
20050195868 Army Research Lab., Cleveland, OH, USA
Autonomous Propulsion System Technology Being Developed to Optimize Engine Performance Throughout the Lifecycle
Litt, Jonathan S.; Research and Technology 2003; May 2004; 6 pp.; In English; Original contains color illustrations; No Copyright; Avail: CASI; A02, Hardcopy
The goal of the Autonomous Propulsion System Technology (APST) project is to reduce pilot workload under both normal and anomalous conditions. Ongoing work under APST develops and leverages technologies that provide autonomous engine monitoring, diagnosing, and controller adaptation functions, resulting in an integrated suite of algorithms that maintain the propulsion system’s performance and safety throughout its life.
Engine-to-engine performance variation occurs among new engines because of manufacturing tolerances and assembly practices. As an engine wears, the performance changes as operability limits are reached. In addition to these normal phenomena, other unanticipated events such as sensor failures, bird ingestion, or component faults may occur, affecting pilot workload as well as compromising safety.
APST will adapt the controller as necessary to achieve optimal performance for a normal aging engine, and the safety net of APST algorithms will examine and interpret data from a variety of onboard sources to detect, isolate, and if possible, accommodate faults.
Situations that cannot be accommodated within the faulted engine itself will be referred to a higher level vehicle management system. This system will have the authority to redistribute the faulted engine’s functionality among other engines, or to replan the mission based on this new engine health information.
Work is currently underway in the areas of adaptive control to compensate for engine degradation due to aging, data fusion for diagnostics and prognostics of specific sensor and component faults, and foreign object ingestion detection. In addition, a framework is being defined for integrating all the components of APST into a unified system. A multivariable, adaptive, multimode control algorithm has been developed that accommodates degradation-induced thrust disturbances during throttle transients.
The baseline controller of the engine model currently being investigated has multiple control modes that are selected according to some performance or operational criteria. As the engine degrades, parameters shift from their nominal values. Thus, when a new control mode is swapped in, a variable that is being brought under control might have an excessive initial error. The new adaptive algorithm adjusts the controller gains on the basis of the level of degradation to minimize the disruptive influence of the large error on other variables and to recover the desired thrust response. Author
Adaptive Control; Propulsion System Configurations; Autonomy; Multivariable Control; Multisensor Fusion; Management Systems
20050195869 NASA Glenn Research Center, Cleveland, OH, USA
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Starting Vortex Identified as Key to Unsteady Ejector Performance
Paxson, Daniel E.; Research and Technology 2003; May 2004; 5 pp.; In English; Original contains color illustrations; No Copyright; Avail: CASI; A01, Hardcopy
Unsteady ejectors are currently under investigation for use in some pulse-detonation-engine-based propulsion systems. Experimental measurements made in the past, and recently at the NASA Glenn Research Center, have demonstrated that thrust augmentation can be enhanced considerably when the driver is unsteady.
In ejector systems, thrust augmentation is defined as = T(sup Total)/T(sup j), where T(sup Total) is the total thrust of the combined ejector and driving jet and T(sup j) is the thrust due to the driving jet alone. There are three images in this figure, one for each of the named thrust sources. The images are color contours of measured instantaneous vorticity. Each image is an ensemble average of at least 150 phase-locked measurements. The flow is from right to left, and the shape and location of each driver is shown on the far right of each image. The emitted vortex is a clearly defined ‘doughnut’ of highly vortical (spinning) flow. In these planar images, the vortex appears as two distorted circles, one above, and one below the axis of symmetry. Because they are spinning in the opposite direction, the two circles have vorticity of opposite sign and thus are different colors.
There is also a rectangle shown in each image. Its width represents the ejector diameter that was found experimentally to yield the highest thrust augmentation. It is apparent that the optimal ejector diameter is that which just ‘captures’ the vortex: that is, the diameter bounding the outermost edge of the vortex structure. The exact mechanism behind the enhanced performance is unclear; however, it is believed to be related to the powerful vortex emitted with each pulse of the unsteady driver. As such, particle imaging velocimetry (PIV) measurements were obtained for three unsteady drivers: a pulsejet, a resonance tube, and a speaker-driven jet.
All the drivers were tested with ejectors, and all exhibited performance enhancement over similarly sized steady drivers. The characteristic starting vortices of each driver are shown in these images. The images are color contours of measured instantaneous vorticity. Each image is an ensemble average of at least 150 phase-locked measurements. The flow is from right to left. The shape and location of each driver is shown on the far right of each image. The rectangle shown in each image represents the ejector diameter that was found experimentally to yield the highest thrust augmentation. It is apparent that the optimal ejector diameter is that which just ‘captures’ the vortex: that is, the diameter bounding the outermost edge of the vortex structure. Although not shown, it was observed that the emitted vortex spread as it traveled downstream.
The spreading rate for the pulsejet is shown as the dashed lines in the top image. A tapered ejector was fabricated that matched this shape. When tested, the ejector demonstrated superior performance to all those previously tested at Glenn (which were essentially of straight, cylindrical form), achieving a remarkable thrust augmentation of 2. The measured thrust augmentation is shown as a function of ejector length.
Also shown are the thrust augmentation values achieved with the straight, cylindrical ejectors of varying diameters. Here, thrust augmentation is plotted as a function of ejector length for several families of ejector diameters. It can be seen that large thrust augmentation values are indeed obtained and that they are sensitive to both ejector length and diameter, particularly the latter.
Five curves are shown. Four correspond to straight ejector diameters of 2.2, 3.0, 4.0, and 6.0 in. The fifth curve corresponds to the tapered ejector contoured to bound the emitted vortex. For each curve, there are several data points corresponding to different lengths. The largest value of thrust augmentation is 2.0 for the tapered ejector and 1.81 for the straight ejectors. Regardless of their diameters, all the ejectors trend toward peak performance at a particular leng. That the cross-sectional dimensions of optimal ejectors scaled precisely with the vortex dimensions on three separate pulsed thrust sources demonstrates that the action of the vortex is responsible for the enhanced ejector performance. The result also suggests that, in the absence of a complete understanding of the entrainment and augmentation mechanisms, methods of characterizing starting vortices may be useful for correlating and predicting unsteady ejector performance. Author
Ejectors; Unsteady Flow; Vortices; Fabrication; Propulsion System Performance
20050195873 NASA Glenn Research Center, Cleveland, OH, USA
Active Closed-Loop Stator Vane Flow Control Demonstrated in a Low-Speed Multistage Compressor
Bright, Michelle M.; Culley, Dennis E.; Strazisar, Anthony J.; Research and Technology 2003; May 2004; 4 pp.; In English; Original contains color and black and white illustrations; No Copyright; Avail: CASI; A01, Hardcopy
Closed-loop flow control was successfully demonstrated on the surface of stator vanes in NASA Glenn Research Center’s Low-Speed Axial Compressor (LSAC) facility. This facility provides a flow field that accurately duplicates the aerodynamics of modern highly loaded compressors. Closed-loop active flow control uses sensors and actuators embedded within engine components to dynamically alter the internal flow path during off-nominal operation in order to optimize engine performance and maintain stable operation. Derived from text
Feedback Control; Flow Regulators; Vanes; Turbocompressors; Stator Blades
20050195882 NASA Glenn Research Center, Cleveland, OH, USA
Ultra-Efficient Engine Technology (UEET) Program
Manthey, Lori A.; Research and Technology 2000; March 2001; 4 pp.; In English; Original contains color illustrations; No Copyright; Avail: CASI; A01, Hardcopy
The Ultra-Efficient Engine Technology (UEET) Program includes seven key projects that work with industry to develop and hand off revolutionary propulsion technologies that will enable future-generation vehicles over a wide range of flight speeds. A new program office, the Ultra-Efficient Engine Technology (UEET) Program Office, was formed at the NASA Glenn Research Center to manage an important National propulsion program for NASA. The Glenn-managed UEET Program, which began on October 1, 1999, includes participation from three other NASA centers (Ames, Goddard, and Langley), as well as five engine companies (GE Aircraft Engines, Pratt & Whitney, Honeywell, Allison/Rolls Royce, and Williams International) and two airplane manufacturers (the Boeing Company and Lockheed Martin Corporation). This 6-year, nearly $300 million program will address local air-quality concerns by developing technologies to significantly reduce nitrogen oxide (NOx) emissions. In addition, it will provide critical propulsion technologies to dramatically increase performance as measured in fuel burn reduction that will enable reductions of carbon dioxide (CO2) emissions. This is necessary to address the potential climate impact of long-term aviation growth. Derived from text
Technology Assessment; Propulsion System Performance; Air Quality
20050195883 NASA Glenn Research Center, Cleveland, OH, USA
Simplified Dynamic Model of Turbine Clearance Developed for Active Clearance Control Studies
Melcher, Kevin J.; Research and Technology 2003; May 2004; 4 pp.; In English; Original contains color illustrations; No Copyright; Avail: CASI; A01, Hardcopy
A simplified analytical model was developed and implemented to simulate changes in turbine tip clearance during the operation of a commercial gas turbine engine. The clearance model is an enabling technology for the fast-response active turbine tip-clearance control being developed by the NASA Glenn Research Center under the Ultra-Efficient Engine Technology Project. Derived from text
Dynamic Models; Gas Turbine Engines; Mathematical Models
20050196623 DYNACS Engineering Co., Inc., USA, NASA Glenn Research Center, Cleveland, OH, USA
One-Dimensional Spontaneous Raman Measurements Made in a Gas Turbine Combustor
DeGroot, Wilhelmus A.; Hicks, Yolanda R.; Locke, Randy J.; Anderson, Robert C.; Research and Technology 2000; March 2001; 3 pp.; In English; No Copyright; Avail: CASI; A01, Hardcopy
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. Author
Raman Spectra; Calibrating; Gas Turbines; Combustion Chambers; Flow Distribution; Polycyclic Aromatic Hydrocarbons; Laser Induced Fluorescence
20050196667 NASA Glenn Research Center, Cleveland, OH, USA
The GE-NASA RTA Hyperburner Design and Development
Lee, Jin-Ho; Winslow, Ralph; Buehrle, Robert J.; June 2005; 22 pp.; In English; 40th Combustion, 28th Airbreathing Propulsion, 22nd Propulsion Systems Hazards and 4th Modeling and Simulation Joint Subcommittee Meetings, 13-17 Jun. 2005, Charleston, SC, USA Contract(s)/Grant(s): WBS 22-065-92-43 Report No.(s): NASA/TM-2005-213803; E-15160; No Copyright; Avail: CASI; A03, Hardcopy
The Revolutionary Turbine Accelerator (RTA) project is a ground demonstration of a Mach 4 Turbine Based Combined Cycle engine. This new combined cycle engine developed for the ground-based demonstration will use a new type of augmentor called the hyperburner. The technical features of this new augmenter are introduced in this work. Some of the salient features include a new variable area bypass injector system and a new flame holder configuration. A summary of the hyperburner configuration and the supporting evidence obtained during the hyperburner rig experiments show that hyperburner is a viable burner concept capable of meeting the goals of the RTA ground engine demonstration project. Author
Turbine Engines; Supersonic Speed; Cycles; Burners; Injectors
20050196728 NASA Glenn Research Center, Cleveland, OH, USA
Hydrogen-powered Flight
Smith, Timothy D.; [2005]; 5 pp.; In English Contract(s)/Grant(s): 22-066-10-01 Report No.(s): E-15195; No Copyright; Avail: CASI; A01, Hardcopy
As the Nation moves towards a hydrogen economy the shape of aviation will change dramatically. To accommodate a switch to hydrogen the aircraft designs, propulsion, and power systems will look much different than the systems of today. Hydrogen will enable a number of new aircraft capabilities from high altitude long endurance remotely operated aircraft (HALE ROA) that will fly weeks to months without refueling to clean, zero emissions transport aircraft. Design and development of new hydrogen powered aircraft have a number of challenges which must be addressed before an operational system can become a reality. While the switch to hydrogen will be most outwardly noticeable in the aircraft designs of the future, other significant changes will be occurring in the environment. A switch to hydrogen for aircraft will completely eliminate harmful greenhouse gases such as carbon monoxide (CO), carbon dioxide (CO2), sulfur oxides (SOx), unburnt hydrocarbons and smoke. While these aircraft emissions are a small percentage of the amount produced on a daily basis, their placement in the upper atmosphere make them particularly harmful. Another troublesome gaseous emission from aircraft is nitrogen oxides (NOx) which contribute to ozone depletion in the upper atmosphere. Nitrogen oxide emissions are produced during the combustion process and are primarily a function of combustion temperature and residence time. The introduction of hydrogen to a gas turbine propulsion system will not eliminate NOx emissions; however the wide flammability range will make low NOx producing, lean burning systems feasible. A revolutionary approach to completely eliminating NOx would be to fly all electric aircraft powered by hydrogen air fuel cells. The fuel cells systems would only produce water, which could be captured on board or released in the lower altitudes. Currently fuel cell systems do not have sufficient energy densities for use in large aircraft, but the long term potential of eliminating greenhouse gas emissions makes it an intriguing and important field of research. Author (revised)
Hydrogen Fuels; Clean Fuels; Aircraft Fuels; Fuel Cells; Aircraft Engines
20050198876 NASA Glenn Research Center, Cleveland, OH, USA
Jet Engine Noise Generation, Prediction and Control, Chapter 86
Huff, Dennis L.; Envia, Edmane; September 20, 2004; 24 pp.; In English Report No.(s): E-14866; No Copyright; Avail: CASI; A03, Hardcopy
Aircraft noise has been a problem near airports for many years. It is a quality of life issue that impacts millions of people around the world. Solving this problem has been the principal goal of noise reduction research that began when commercial jet travel became a reality. While progress has been made in reducing both airframe and engine noise, historically, most of the aircraft noise reduction efforts have concentrated on the engines. This was most evident during the 1950 s and 1960 s when turbojet engines were in wide use. This type of engine produces high velocity hot exhaust jets during takeoff generating a great deal of noise. While there are fewer commercial aircraft flying today with turbojet engines, supersonic aircraft including high performance military aircraft use engines with similar exhaust flow characteristics. The Pratt & Whitney F100-PW-229, pictured in Figure la, is an example of an engine that powers the F-15 and F-16 fighter jets. The turbofan engine was developed for subsonic transports, which in addition to better fuel efficiency also helped mitigate engine noise by reducing the jet exhaust velocity. These engines were introduced in the late 1960 s and power most of the commercial fleet today. Over the years, the bypass ratio (that is the ratio of the mass flow through the fan bypass duct to the mass flow through the engine core) has increased to values approaching 9 for modern turbofans such as the General Electric s GE-90 engine (Figure lb). The benefits to noise reduction for high bypass ratio (HPBR) engines are derived from lowering the core jet velocity and temperature, and lowering the tip speed and pressure ratio of the fan, both of which are the consequences of the increase in bypass ratio. The HBPR engines are typically very large in diameter and can produce over 100,000 pounds of thrust for the largest engines. A third type of engine flying today is the turbo-shaft which is mainly used to power turboprop aircraft and helicopters. An example of this type of engine is shown in Figure IC, which is a schematic of the Honeywell T55 engine that powers the CH-47 Chinook helicopter. Since the noise from the propellers or helicopter rotors is usually dominant for turbo-shaft engines, less attention has been paid to these engines in so far as community noise considerations are concerned. This chapter will concentrate mostly on turbofan engine noise and will highlight common methods for their noise prediction and reduction. Derived from text
Aerodynamic Noise; Aircraft Noise; Engine Noise; Handbooks; Noise Reduction
20050198946 NASA Glenn Research Center, Cleveland, OH, USA
The Challenges Facing Future Conversion Systems for Space Power Applications
Schreiber, Jeffrey; [2004]; 11 pp.; In English; International Energy Conversion Engineering Conference, 16-19 Aug. 2004, Providence, RI, USA Contract(s)/Grant(s): WBS 22-972-20-01 Report No.(s): E-14842; No Copyright; Avail: CASI; A03, Hardcopy
High-efficiency, Stirling power convertors have been proposed for space power applications, ranging from relatively low-power radioisotope generators such as the 110 watt SRG110 to the higher-power 100 kWe SP-100. The NASA Glenn Research Center (GRC) has been involved in the supporting technology and development for both of these systems. Although the power levels are quite different, many of the challenges faced by both of these dynamic power conversion systems similar. A major challenge is fund in demonstration of the capability for high reliability and long-life of the power system when the wear mechanisms have been eliminated. A review is presented of the past efforts, including the status of current flight development efforts, and a projection of what the future might bring. Author
Space Power Reactors; Stirling Engines; Power Converters
20050199071 Propane Vehicle Demonstration Grant Program
Kerr, G.; Aug. 27, 2004; 14 pp.; In English Report No.(s): DE2005-832986; No Copyright; Avail: Department of Energy Information Bridge
The Propane Vehicle Demonstration Grants was established to demonstrate the benefits of new propane equipment. The US Department of Energy, the Propane Education & Research Council (PERC) and the Propane Vehicle Council (PVC) partnered in this program. The project impacted ten different states, 179 vehicles, and 15 new propane fueling facilities. Based on estimates provided, this project generated a minimum of 1,441,000 new gallons of propane sold for the vehicle market annually. Additionally, two new off-road engines were brought to the market. Projects originally funded under this project were the City of Portland, Colorado, Kansas City, Impco Technologies, Jasper Engines, Maricopa County, New Jersey State, Port of Houston, Salt Lake City Newspaper, Suburban Propane, Mutual Liquid Propane and Ted Johnson. NTIS
Motor Vehicles; Propane
20050199427 NASA Glenn Research Center, Cleveland, OH, USA, Ohio Aerospace Inst., OH, USA
Engine With Regression and Neural Network Approximators Designed
Patnaik, Surya N.; Hopkins, Dale A.; Research and Technology 2000; March 2001; 2 pp.; In English; No Copyright; Avail: CASI; A01, Hardcopy
At the NASA Glenn Research Center, the NASA engine performance program (NEPP, ref. 1) and the design optimization testbed COMETBOARDS (ref. 2) with regression and neural network analysis-approximators have been coupled to obtain a preliminary engine design methodology. The solution to a high-bypass-ratio subsonic waverotor-topped turbofan engine, which is shown in the preceding figure, was obtained by the simulation depicted in the following figure. This engine is made of 16 components mounted on two shafts with 21 flow stations. The engine is designed for a flight envelope with 47 operating points. The design optimization utilized both neural network and regression approximations, along with the cascade strategy (ref. 3). The cascade used three algorithms in sequence: the method of feasible directions, the sequence of unconstrained minimizations technique, and sequential quadratic programming. The normalized optimum thrusts obtained by the three methods are shown in the following figure: the cascade algorithm with regression approximation is represented by a triangle, a circle is shown for the neural network solution, and a solid line indicates original NEPP results. The solutions obtained from both approximate methods lie within one standard deviation of the benchmark solution for each operating point. The simulation improved the maximum thrust by 5 percent. The performance of the linear regression and neural network methods as alternate engine analyzers was found to be satisfactory for the analysis and operation optimization of air-breathing propulsion engines (ref. 4). Author
Engine Design; Regression Analysis; Neural Nets; Design Optimization; Turbofan Engines; Approximation
20050199433 NASA Glenn Research Center, Cleveland, OH, USA
Oil-Free Turbomachinery Being Developed
DellaCorte, Christopher; Valco, Mark J.; Research and Technology 2000; March 2001; 2 pp.; In English; No Copyright;Avail: CASI; A01, Hardcopy
NASA and the Army Research Laboratory (ARL) along with industry and university researchers, are developing Oil-Free technology that will have a revolutionary impact on turbomachinery systems used in commercial and military applications. System studies have shown that eliminating an engine’s oil system can yield significant savings in weight, maintenance, and operational costs. The Oil-Free technology (foil air bearings, high-temperature coatings, and advanced modeling) is being developed to eliminate the need for oil lubrication systems on high-speed turbomachinery such as turbochargers and gas turbine engines that are used in aircraft propulsion systems. The Oil-Free technology is enabled by recent breakthroughs in foil bearing load capacity, solid lubricant coatings, and computer-based analytical modeling. During the past fiscal year, a U.S. patent was awarded for the NASA PS300 solid lubricant coating, which was developed at the NASA Glenn Research Center. PS300 has enabled the successful operation of foil air bearings to temperatures over 650 C and has resulted in wear lives in excess of 100,000 start/stop cycles. This leapfrog improvement in performance over conventional solid lubricants (limited to 300 C) creates new application opportunities for high-speed, high-temperature Oil-Free gas turbine engines. On the basis of this break-through coating technology and the world’s first successful demonstration of an Oil-Free turbocharger in fiscal year 1999, industry is partnering with NASA on a 3-year project to demonstrate a small, Oil-Free turbofan engine for aeropropulsion. Author
Foil Bearings; Coating; High Temperature; Turbomachinery; Solid Lubricants; Lubrication Systems
20050199434 NASA Glenn Research Center, Cleveland, OH, USA
Technology Being Developed at Lawrence Berkeley National Laboratory: Ultra-Low- Emission Combustion Technologies for Heat and Power Generation
Cheng, Robert K.; Research and Technology 2000; March 2001; 3 pp.; In English; No Copyright; Avail: CASI; A01, Hardcopy
The Combustion Technologies Group at Lawrence Berkeley National Laboratory has developed simple, low-cost, yet robust combustion technologies that may change the fundamental design concept of burners for boilers and furnaces, and injectors for gas turbine combustors. The new technologies utilize lean premixed combustion and could bring about significant pollution reductions from commercial and industrial combustion processes and may also improve efficiency. The technologies are spinoffs of two fundamental research projects: An inner-ring burner insert for lean flame stabilization developed for NASA- sponsored reduced-gravity combustion experiments. A low-swirl burner developed for Department of Energy Basic Energy Sciences research on turbulent combustion. Derived from text
Combustion; Combustion Physics; Emission; Heat Generation; Gas Turbines
Source: NASA.
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