SCIENTIFIC AND TECHNICAL AEROSPACE REPORTS
A Biweekly Publication of the National Aeronautics and Space Administration
VOLUME 43, ISSUE 13 - JULY 1, 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.
20050177434 Florida Univ., Gainesville, FL USA
Semi-Closed Cycle Turbine Engines in U.S. Army Applications With Water Harvesting
Lear, William E.; Sherif, S. A.; Khan, J. R.; Meitner, Peter L.; Crittenden, John; Dec. 2004; 3 pp.; In English; Original contains color illustrations Report No.(s): AD-A433357; No Copyright; Avail: Defense Technical Information Center (DTIC)
Gas turbine power requirements for the U.S. Army vary tremendously depending upon the specific application, spanning the power range from a few Watts for individual soldier power supplies to Megawatt-scale engines for heavy land vehicles and air transport. Each application has its own set of key drivers which constrain the selection of an engine technology and in optimizing the design, the objective function dependencies vary widely. Hence, although the desire for compactness, efficiency, lightness, low observables, etc. is universal among the common Army applications, their relative importance is not. In addition, certain applications may require an engine attribute which is wholly unimportant in other cases, such as the need for heat and/or air conditioning in a distributed generation unit. Thus flexibility in the design of any type of engine is a significant advantage if it is to find a niche in several applications, an obvious benefit to minimizing Army inventories and costs. DTIC
Closed Cycles; Gas Turbines; Turbine Engines; Water
20050177898 NASA Lewis Research Center, Cleveland, OH, USA
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Steady-State Cycle Deck Launcher Developed for Numerical Propulsion System Simulation
VanDrei, Donald E.; Research and Technology 1996; March 1997; 2 pp.; In English; No Copyright; Avail: CASI; A01, Hardcopy
One of the objectives of NASA’s High Performance Computing and Communications Program’s (HPCCP) Numerical Propulsion System Simulation (NPSS) is to reduce the time and cost of generating aerothermal numerical representations of engines, called customer decks. These customer decks, which are delivered to airframe companies by various U.S. engine companies, numerically characterize an engine’s performance as defined by the particular U.S. airframe manufacturer.
Until recently, all numerical models were provided with a Fortran-compatible interface in compliance with the Society of Automotive Engineers (SAE) document AS681F, and data communication was performed via a standard, labeled common structure in compliance with AS681F. Recently, the SAE committee began to develop a new standard: AS681G. AS681G addresses multiple language requirements for customer decks along with alternative data communication techniques. Along with the SAE committee, the NPSS Steady-State Cycle Deck project team developed a standard Application Program Interface (API) supported by a graphical user interface. This work will result in Aerospace Recommended Practice 4868 (ARP4868).
The Steady-State Cycle Deck work was validated against the Energy Efficient Engine customer deck, which is publicly available. The Energy Efficient Engine wrapper was used not only to validate ARP4868 but also to demonstrate how to wrap an existing customer deck. The graphical user interface for the Steady-State Cycle Deck facilitates the use of the new standard and makes it easier to design and analyze a customer deck. This software was developed following I. Jacobson’s Object-Oriented Design methodology and is implemented in C++. The AS681G standard will establish a common generic interface for U.S. engine companies and airframe manufacturers. This will lead to more accurate cycle models, quicker model generation, and faster validation leading to specifications. The standard will facilitate cooperative work between industry and NASA.
The NPSS Steady-State Cycle Deck team released a batch version of the Steady-State Cycle Deck in March 1996. Version 1.1 was released in June 1996. During fiscal 1997, NPSS accepted enhancements and modifications to the Steady-State Cycle Deck launcher. Consistent with NPSS’ commercialization plan, these modifications will be done by a third party that can provide long-term software support. Author
Steady State; Propulsion System Performance; Launchers; Computerized Simulation; Numerical Analysis
20050177901 NASA Lewis Research Center, Cleveland, OH, USA
TURBO-AE: An Aeroelastic Code for Propulsion Applications
Bakhle, Milind A.; Research and Technology 1996; March 1997; 2 pp.; In English; No Copyright; Avail: CASI; A01, Hardcopy
NASA’s Advanced Subsonic Technology (AST) program is developing new technologies to increase the fuel efficiency of commercial aircraft engines, improve the safety of engine operation, and reduce engine emissions and noise. With the development of new designs for ducted fans, compressors, and turbines to achieve these goals, a basic aeroelastic requirement is that there should be no flutter or high resonant blade stresses in the operating regime.
To verify the aeroelastic soundness of these designs, we need an accurate prediction and analysis code. Such a two-dimensional viscous propulsion aeroelastic code, named TURBO-AE, is being developed at the NASA Lewis Research Center. The TURBO-AE aeroelastic code is based on a three-dimensional unsteady aerodynamic Euler/Navier-Stokes turbomachinery code TURBO, developed under a grant from NASA Lewis. TURBO-AE can model viscous flow effects that play an important role in certain aeroelastic problems, such as flutter with flow separation (or stall flutter) and flutter in the presence of shock and boundary-layer interaction. The structural dynamics representation of the blade in the TURBO-AE code is based on a normal mode representation.
A finite element analysis code, such as NASTRAN, is used to calculate in-vacuum vibration modes and the associated natural frequency. A work-per-cycle approach is used to determine aeroelastic (flutter) stability. With this approach, the motion of the blade is prescribed to be a harmonic vibration in a specified in vacuum normal mode. The aerodynamic forces acting on the vibrating blade and the work done by these forces on the vibrating blade during a cycle of vibration are calculated. If positive work is being done on the blade by the aerodynamic forces, the blade is dynamically unstable, since it will extract energy from the flow, leading to an increase in the amplitude of the blade’s oscillation.
Initial calculations have been done for a configuration representative of the Energy Efficient Engine fan rotor. The accompanying figure shows the work-per-cycle after each cycle of vibration. It can be seen that the work-per-cycle does not vary much after the fourth cycle. The negative sign of the converged work-per-cycle shows that the fan blade is dynamically stable and will not flutter.
TURBO-AE will provide a useful aeroelastic prediction/analysis capability for engine manufacturers. It will reduce design cycle times by allowing new blade designs to be verified for aeroelastic soundness before blades are built and tested. With this prediction capability, it will be possible to build thinner, lighter, and faster rotating blades without encountering aeroelastic problems like stall flutter and high-cycle fatigue due to forced vibrations. Author
Aeroelasticity; Turbomachinery; Commercial Aircraft; Subsonic Speed; Propulsion System Performance
20050177906 NASA Lewis Research Center, Cleveland, OH, USA
Rocket Engine Numerical Simulator (RENS)
Davidian, Kenneth O.; Research and Technology 1996; March 1997; 2 pp.; In English; No Copyright; Avail: Other Sources; Abstract Only
Work is being done at three universities to help today’s NASA engineers use the knowledge and experience of their Apolloera predecessors in designing liquid rocket engines.
Ground-breaking work is being done in important subject areas to create a prototype of the most important functions for the Rocket Engine Numerical Simulator (RENS). The goal of RENS is to develop an interactive, realtime application that engineers can utilize for comprehensive preliminary propulsion system design functions. RENS will employ computer science and artificial intelligence research in knowledge acquisition, computer code parallelization and objectification, expert system architecture design, and object-oriented programming.
In 1995, a 3-year grant from the NASA Lewis Research Center was awarded to Dr. Douglas Moreman and Dr. John Dyer of Southern University at Baton Rouge, Louisiana, to begin acquiring knowledge in liquid rocket propulsion systems. Resources of the University of West Florida in Pensacola were enlisted to begin the process of enlisting knowledge from senior NASA engineers who are recognized experts in liquid rocket engine propulsion systems. Dr. John Coffey of the University of West Florida is utilizing his expertise in interviewing and concept mapping techniques to encode, classify, and integrate information obtained through personal interviews.
The expertise extracted from the NASA engineers has been put into concept maps with supporting textual, audio, graphic, and video material. A fundamental concept map was delivered by the end of the first year of work and the development of maps containing increasing amounts of information is continuing. Find out more information about this work at the Southern University/University of West Florida.
In 1996, the Southern University/University of West Florida team conducted a 4-day group interview with a panel of five experts to discuss failures of the RL10 rocket engine in conjunction with the Centaur launch vehicle. The discussion was recorded on video and audio tape. Transcriptions of the entire proceedings and an abbreviated video presentation of the discussion highlights are under development.
Also in 1996, two additional 3-year grants were awarded to conduct parallel efforts that would complement the work being done by Southern University and the University ofWest Florida. Dr. Prem Bhalla of Jackson State University in Jackson, Mississippi, is developing the architectural framework for RENS. By employing the Rose Rational language and Booch Object Oriented Programming (OOP) technology, Dr. Bhalla is developing the basic structure of RENS by identifying and encoding propulsion system components, their individual characteristics, and cross-functionality and dependencies.
Dr. Ruknet Cezzar of Hampton University, located in Hampton, Virginia, began working on the parallelization and objectification of rocket engine analysis and design codes. Dr. Cezzar will use the Turbo C++ OOP language to translate important liquid rocket engine computer codes from FORTRAN and permit their inclusion into the RENS framework being developed at Jackson State University. The Southern University/University of West Florida grant was extended by 1 year to coordinate the conclusion of all three efforts in 1999. Author
Numerical Control; Liquid Propellant Rocket Engines; Design Analysis; Propulsion System Configurations; Rocket Engine Design; Computer Programs
20050177910 NASA Lewis Research Center, Cleveland, OH, USA
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COMETBOARDS Can Optimize the Performance of a Wave-Rotor-Topped Gas Turbine Engine
Patnaik, Surya N.; Research and Technology 1996; March 1997; 2 pp.; In English; No Copyright; Avail: CASI; A01, Hardcopy
A wave rotor, which acts as a high-technology topping spool in gas turbine engines, can increase the effective pressure ratio as well as the turbine inlet temperature in such engines. The wave rotor topping, in other words, may significantly enhance engine performance by increasing shaft horse power while reducing specific fuel consumption. This performance enhancement requires optimum selection of the wave rotor’s adjustable parameters for speed, surge margin, and temperature constraints specified on different engine components.
To examine the benefit of the wave rotor concept in engine design, researchers soft coupled NASA Lewis Research Center’s multidisciplinary optimization tool COMETBOARDS and the NASA Engine Performance Program (NEPP) analyzer. The COMETBOARDS-NEPP combined design tool has been successfully used to optimize wave-rotor-topped engines. For illustration, the design of a subsonic gas turbine wave-rotorenhanced engine with four ports for 47 mission points (which are specified by Mach number, altitude, and power-setting combinations) is considered.
The engine performance analysis, constraints, and objective formulations were carried out through NEPP, and COMETBOARDS was used for the design optimization. So that the benefits that accrue from wave rotor enhancement could be examined, most baseline variables and constraints were declared to be passive, whereas important parameters directly associated with the wave rotor were considered to be active for the design optimization. The engine thrust was considered as the merit function.
The wave rotor engine design, which became a sequence of 47 optimization subproblems, was solved successfully by using a cascade strategy available in COMETBOARDS. The graph depicts the optimum COMETBOARDS solutions for the 47 mission points, which were normalized with respect to standard results. As shown, the combined tool produced higher thrust for all mission points than did the other solution, with maximum benefits around mission points 11, 25, and 31. Such improvements can become critical, especially when engines are sized for these specific mission points. Author
Gas Turbine Engines; Wave Rotors; Spools; Engine Analyzers; Subsonic Flow; Fuel Consumption; Design Optimization
20050177915 NASA Lewis Research Center, Cleveland, OH, USA
Rotor-Stator Interaction Performance Effects
VanZante, Dale E.; Research and Technology 1996; March 1997; 1 pp.; In English; No Copyright; Avail: Other Sources; Abstract Only
Decreased axial spacing between blade rows in an axial compressor stage is thought to increase stage performance because of an unsteady process that occurs in the downstream blade row and acts on the upstream blade row wakes. This process results in the ‘recovery’ of part of the wake energy before all of this energy is irreversibly lost due to viscous diffusion. To study the wake-blade interaction mechanism, researchers at the NASA Lewis Research Center acquired two-component Laser Fringe Anemometer measurements of the rotor wake in the single-stage transonic compressor at two stage loading levels. The detailed measurements were acquired for one stator pitch in circumference at axial positions from the rotor trailing edge to 20 percent of the stator axial chord, at the exit of the stator passage, and downstream of the stator row including the stator wake. These data show that the changes in wake energy that occur inside the stator passage are not due to viscous dissipation alone, and thus the data provide evidence that ‘wake recovery’ is occurring. A time-accurate, three-dimensional Navier Stokes simulation of the compressor stator was done at the corresponding stage loading levels. The measurements and simulations are being used in combination to show the effects of stator blade loading, quantify the effects of viscosity, and quantify the stage efficiency gain due to the wake recovery process. The accuracy of simple models of the wake recovery process is being evaluated in an effort to include the effects of wake recovery in the NASA-developed Average Passage code for multistage turbomachinery simulations. Author
Rotors; Interactions; Performance Prediction; Stator Blades; Three Dimensional Models
20050177922 NASA Lewis Research Center, Cleveland, OH, USA
Stator Indexing in Multistage Compressors
Barankiewicz, Wendy S.; Research and Technology 1996; March 1997; 1 pp.; In English; No Copyright; Avail: CASI; A01, Hardcopy
The relative circumferential location of stator rows (stator indexing) is an aspect of multistage compressor design that has not yet been explored for its potential impact on compressor aerodynamic performance. Although the inlet stages of multistage compressors usually have differing stator blade counts, the aft stages of core compressors can often have stage blocks with equal stator blade counts in successive stages. The potential impact of stator indexing is likely greatest in these stages. To assess the performance impact of stator indexing, researchers at the NASA Lewis Research Center used the 4 ft diameter, four-stage NASA Low Speed Axial Compressor for detailed experiments. This compressor has geometrically identical stages that can circumferentially index stator rows relative to each other in a controlled manner; thus it is an ideal test rig for such investigations. AuthorTurbocompressors; Stator Blades; Performance Tests
20050179332 NASA Lewis Research Center, Cleveland, OH, USA, Pratt and Whitney Aircraft Group, USA
Fan Blade Deflection Measurement and Analyses Correlation
Mehmed, Oral; Janetzke, David C.; Research and Technology 1996; March 1997; 2 pp.; In English; No Copyright; Avail: CASI; A01, Hardcopy
Steady deflection measurements were taken of two identical NASA/Pratt & Whitney-designed fan blades while they were rotating in a vacuum in NASA Lewis Research Center’s Dynamic Spin Facility. The one-fifth-scale fan blades, which have a tip diameter of 22 in. and a pinroot retention, are of sparshell construction and were unducted for this test. The purpose of the test was to measure the change of the radial deflection of the blade tip and blade angle at selected radial stations along the blade span with respect to rotational speed. The procedure for radial deflection measurement had no precedent and was newly developed for this test. Radial deflection measurements were made to assure adequate tip clearance existed between the fan blades and the duct for a follow-on wind tunnel test. Also, blade angle deflection measurements were desired before pitchsetting parts for the wind tunnel test were finish machined. During the test, laser beams were aimed across the blade path into photodiodes to give signals that were used to determine blade angle change or tip radial deflection. These laser beams were set parallel to the spin axis at selected radial stations. Derived from text
Deflection; Fan Blades; Pitch (Inclination); Angles (Geometry)
20050179336 NASA Lewis Research Center, Cleveland, OH, USA, Pratt and Whitney Aircraft, USA
Fluidic Injection for Throat-Area Control and Thrust Vectoring
Lam, David W.; Research and Technology 1996; March 1997; 1 pp.; In English; No Copyright; Avail: CASI; A01, Hardcopy
The Fluidic Injection Nozzle Technology program is a national cooperative effort to develop fluidic area control and thrust vectoring concepts for advanced exhaust systems. Exhaust nozzles with vector flow capability will increase the maneuverability and survivability of future fighter aircraft. Currently used mechanical vectoring and area-control systems add weight and complexity to aircraft exhaust systems.With the use of fluidic injection, area-control and vectoring can be achieved without the added weight penalty. Under this program, the NASA Lewis Research Center entered into a cooperative test program with Pratt & Whitney to study the performance of the F119 nozzle with fluidic injection. Our area of interest was to measure flow and thrust coefficients, throat-area reduction, and vector angles. This experimental program was successfully completed in January 1996 in Lewis’ CE-22 facility. Various nozzle throat areas and expansion ratios were tested over a wide range of nozzle pressure ratios. Other model configurations included different injection locations at different injection angles. Results confirmed that fluidic injection is feasible in throat-area reduction and in vectoring. The data obtained from this test program were added to the current database, which can then be used to validate any future performance prediction methodology and computational fluid dynamics (CFD) analysis. Author
Aircraft Engines; Fluidics; Control Equipment; Exhaust Nozzles; Injection; Performance Prediction
20050179337 NASA Lewis Research Center, Cleveland, OH, USA, Allison Engine Co., Indianapolis, IN, USA
Forward Swept Compressor Testing
Miller, David P.; Research and Technology 1996; March 1997; 2 pp.; In English; No Copyright; Avail: CASI; A01, Hardcopy
A new forward-swept rotor designed by Allison Engine Company was tested in NASA Lewis Research Center’s CE-18 facility. This testing was a follow-on project sponsored by NASALewis to study range enhancements in small turbomachinery. The test was conducted against a baseline rotor design that was also tested in CE-18. The design point for the rotor was a rotor pressure ratio of 2.69, a mass flow of 10.52 lbm/sec, and an adiabatic efficiency of 89.1 percent. Test data indicate that the rotor met the pressure ratio of 2.69 with a 10.77 lbm/sec flow rate, a 87.5-percent adiabatic efficiency, and a 19.5-percent stall margin. The baseline rotor achieved a pressure ratio of 2.69 at a 10.77 lbm/sec flow rate with a stall margin of only 9.2 percent and an adiabatic efficiency of 87.0 percent. The major differences are the significant stall margin increase and the substantially higher off-design peak efficiencies of the forward-swept rotor. The substantially higher performance over the baseline rotor design makes the new design a viable technology candidate for future products. Author
Compressors; Pressure Ratio; Turbomachinery; Effýciency; Flow Velocity; Mass Flow
20050179350 NASA Lewis Research Center, Cleveland, OH, USA
Interactive Educational Tool for Turbofan and Afterburning Turbojet Engines
Benson, Thomas J.; Research and Technoloyg 1996; March 1997; 1 pp.; In English; No Copyright; Avail: Other Sources; Abstract Only
A workstation-based, interactive educational computer program has been developed at the NASA Lewis Research Center to aid in the teaching and understanding of turbine engine design and analysis. This tool has recently been extended to model the performance of two-spool turbofans and afterburning turbojets. The program solves for the flow conditions through the engine by using classical one-dimensional thermodynamic analysis found in various propulsion textbooks. Either an approximately thermally perfect or calorically perfect gas can be used in the thermodynamic analysis. Students can vary the design conditions through a graphical user interface; engine performance is calculated immediately. A variety of graphical formats are used to present results, including numerical results, moving bar charts, and student-generated temperature versus entropy (Ts), pressure versus specific volume (pv), and engine performance plots. The package includes user-controlled printed output, restart capability, online help screens, and a browser that displays teacher-prepared lessons in turbomachinery. The program runs on a variety of workstations or a personal computer using the UNIX operating system and X-based graphics. It is being tested at several universities in the midwestern USA; the source and executables are available free from the author. Author
Workstations; Computer Programs; Education; Turbine Engines; Engine Design; Spools; Turbofans; Afterburning; Turbojet Engines; Design Analysis
20050180555 Purdue Univ., West Lafayette, IN, USA, NASA Marshall Space Flight Center, Huntsville, AL, USA
High-Frequency, High-Temperature Fretting Experiments
Matlik, J. F.; Farris, T. N.; Haake, F. K.; Swanson, G. R.; Duke, G. C.; [2005]; 32 pp.; In English; Original contains black and white illustrations; No Copyright; Avail: CASI; A03, Hardcopy
Fretting is a structural damage mechanism observed when two nominally clamped surfaces are subjected to an oscillatory loading. A critical location for fretting induced damage has been identified at the blade/disk and blade/damper interfaces of gas turbine engine turbomachinery and space propulsion components. The high-temperature, high-frequency loading environment seen by these components lead to severe stress gradients at the edge-of-contact. These contact stresses drive crack nucleation and propagation in fretting and are very sensitive to the geometry of the contacting bodies, the contact loads, materials, temperature, and contact surface tribology (friction). To diagnose the threat that small and relatively undetectable fretting cracks pose to damage tolerance and structural integrity of in-service components, the objective of this work is to develop a well-characterized experimental fretting rig capable of investigating fretting behavior of advanced aerospace alloys subjected to load and temperature conditions representative of such turbomachinery components. Author
High Temperature; Fretting; Oscillations; Loads (Forces); Contact Loads; Structural Failure; Damage; Turbomachinery; Gas Turbine Engines
20050180656 NASA Lewis Research Center, Cleveland, OH, USA
Jet Injection Used to Control Rotating Stall in a High-Speed Compressor
Bright, Michelle M.; Strazisar, Anthony J.; Weigl, Harald J.; Spakovzsky, Zoltan; Paduano, James D.; Research and Technology 1997; April 1998; 3 pp.; In English; No Copyright; Avail: CASI; A01, Hardcopy
In a joint effort between the Massachusetts Institute of Technology (MIT) and the NASA Lewis Research Center, a new technology was demonstrated to identify and control rotating stall and surge in a single-stage, high-speed compressor. Through the use of highvelocity, high-frequency jet injectors, the instabilities of surge and stall were controlled in a high-speed compressor rig. Through the use of active stall control, modal instabilities that normally occur in the pressure measurements prior to stall were normalized and the range of the compressor was extended. Normally the events of rotating stall and surge instabilities limit the operation of the aeroengine compressor to a region below the surge line. To enhance the performance of the compressor, the Lewis/MIT team used active stall control methods to extend the normal operation of the compressor beyond the original stall point. Derived from text
Compressors; Rotating Stalls; Injection
20050180665 NASA Lewis Research Center, Cleveland, OH, USA, Allied-Signal Engines and Systems, Phoenix, AZ, USA
Active Pattern Factor Control for Gas Turbine Engines May, James E.; Research and Technology 1997; April 1998; 2 pp.; In English Contract(s)/Grant(s): NAS3-27752; No Copyright; Avail: CASI; A01, Hardcopy
Small variations in fuel/air mixture ratios within gas turbine combustors can result in measurable, and potentially detrimental, exit thermal gradients. Thermal gradients can increase emissions, as well as shorten the design life of downstream turbomachinery, particularly stator vanes. Uniform temperature profiles are usually sought through careful design and manufacturing of related combustor components. However, small componentto-component variations as well as numerous aging effects degrade system performance. To compensate for degraded thermal performance, researchers are investigating active, closed-loop control schemes. Derived from text
Combustion Chambers; Control Systems Design; Gas Turbines; Aircraft Engines
20050180832 NASA Lewis Research Center, Cleveland, OH, USA
Wave Rotor Research and Technology Development
Welch, Gerard E.; Research and Technology 1997; April 1998; 3 pp.; In English; No Copyright; Avail: CASI; A01, Hardcopy
Wave rotor technology offers the potential to increase the performance of gas turbine engines significantly, within the constraints imposed by current material temperature limits. The wave rotor research at the NASA Lewis Research Center is a three-element effort: 1) Development of design and analysis tools to accurately predict the performance of wave rotor components; 2) Experiments to characterize component performance; 3) System integration studies to evaluate the effect of wave rotor topping on the gas turbine engine system. Derived from text
Research and Development; Wave Rotors; Engine Parts; Performance Prediction
20050181421 NASA Lewis Research Center, Cleveland, OH, USA
Solar Electric Propulsion for Mars Exploration
Hack, Kurt J.; Research and Technology 1997; April 1998; 2 pp.; In English; No Copyright; Avail: CASI; A01, Hardcopy
Highly propellant-efficient electric propulsion is being combined with advanced solar power technology to provide a non-nuclear transportation option for the human exploration of Mars. By virtue of its high specific impulse, electric propulsion offers a greater change in spacecraft velocity for each pound of propellant than do conventional chemical rockets. As a result, a mission to Mars based on solar electric propulsion (SEP) would require fewer heavy-lift launches than a traditional all-chemical space propulsion scenario would. Performance, as measured by mass to orbit and trip time, would be comparable to the NASA design reference mission for human Mars exploration, which utilizes nuclear thermal propulsion; but it would avoid the issues surrounding the use of nuclear reactors in space. Derived from text
Solar Electric Propulsion; Mars Exploration; Propellants; Nuclear Propulsion; Electric Propulsion
20050181428 NASA Glenn Research Center, Cleveland, OH, USA
Closed Loop Active Flow Separation Detection and Control in a Multistage Compressor
Bright, Michelle M.; Culley, Dennis E.; Braunscheidel, Edward P.; Welch, Gerard E.; May 2005; 16 pp.; In English; 43rd Aerospace Sciences Meeting and Exhibit, 10-13 Jan. 2005, Reno, NV, USA; Original contains color illustrations Contract(s)/Grant(s): WBS 22-714-70-54 Report No.(s): NASA/TM-2005-213553; AIAA Paper 2005-0849; E-14993; No Copyright; Avail: CASI; A03, Hardcopy
Active closed loop flow control was successfully demonstrated on a full annulus of stator vanes in a low speed axial compressor. Two independent methods of detecting separated flow conditions on the vane suction surface were developed. The first technique detects changes in static pressure along the vane suction surface, while the second method monitors variation in the potential field of the downstream rotor. Both methods may feasibly be used in future engines employing embedded flow control technology. In response to the detection of separated conditions, injection along the suction surface of each vane was used. Injected mass flow on the suction surface of stator vanes is known to reduce separation and the resulting limitation on static pressure rise due to lowered diffusion in the vane passage. A control algorithm was developed which provided a proportional response of the injected mass flow to the degree of separation, thereby minimizing the performance penalty on the compressor system. Author
Detection; Feedback Control; Separated Flow; Turbocompressors
20050182032 NASA Lewis Research Center, Cleveland, OH, USA
Common Analysis Tool Being Developed for Aeropropulsion: The National Cycle Program Within the Numerical Propulsion System Simulation Environment
Follen, Gregory J.; Naiman, Cynthia G.; Research and Technology 1998; April 1999; 3 pp.; In English; No Copyright; Avail: CASI; A01, Hardcopy
The NASA Lewis Research Center is developing an environment for analyzing and designing aircraft engines-the Numerical Propulsion System Simulation (NPSS). NPSS will integrate multiple disciplines, such as aerodynamics, structure, and heat transfer, and will make use of numerical ‘zooming’ on component codes. Zooming is the coupling of analyses at various levels of detail. NPSS uses the latest computing and communication technologies to capture complex physical processes in a timely, cost-effective manner. The vision of NPSS is to create a ‘numerical test cell’ enabling full engine simulations overnight on cost-effective computing platforms. Through the NASA/Industry Cooperative Effort agreement, NASA Lewis and industry partners are developing a new engine simulation called the National Cycle Program (NCP). NCP, which is the first step toward NPSS and is its initial framework, supports the aerothermodynamic system simulation process for the full life cycle of an engine. U.S. aircraft and airframe companies recognize NCP as the future industry standard common analysis tool for aeropropulsion system modeling. The estimated potential payoff for NCP is a $50 million/yr savings to industry through improved engineering productivity. Derived from text
Propulsion; Aircraft Engines
20050182033 NASA Lewis Research Center, Cleveland, OH, USA
Flow of GE90 Turbofan Engine Simulated
Veres, Joseph P.; Research and Technology 1998; April 1999; 3 pp.; In English; No Copyright; Avail: CASI; A01, Hardcopy
The objective of this task was to create and validate a three-dimensional model of the GE90 turbofan engine (General Electric) using the APNASA (average passage) flow code. This was a joint effort between GE Aircraft Engines and the NASA Lewis Research Center. The goal was to perform an aerodynamic analysis of the engine primary flow path, in under 24 hours of CPU time, on a parallel distributed workstation system. Enhancements were made to the APNASA Navier-Stokes code to make it faster and more robust and to allow for the analysis of more arbitrary geometry. The resulting simulation exploited the use of parallel computations by using two levels of parallelism, with extremely high efficiency.The primary flow path of the GE90 turbofan consists of a nacelle and inlet, 49 blade rows of turbomachinery, and an exhaust nozzle. Secondary flows entering and exiting the primary flow path-such as bleed, purge, and cooling flows-were modeled macroscopically as source terms to accurately simulate the engine. The information on these source terms came from detailed descriptions of the cooling flow and from thermodynamic cycle system simulations. These provided boundary condition data to the three-dimensional analysis. A simplified combustor was used to feed boundary conditions to the turbomachinery. Flow simulations of the fan, high-pressure compressor, and high- and low-pressure turbines were completed with the APNASA code. Derived from text
Turbofan Engines; Turbocompressors; Thrust Augmentation
20050182036 NASA Lewis Research Center, Cleveland, OH, USA
Affordable Manufacturing Technologies Being Developed for Actively Cooled Ceramic Components
Bhatt, Ramakrishna T.; Research and Technology 1998; April 1999; 1 pp.; In English; No Copyright; Avail: CASI; A01, Hardcopy
Efforts to improve the performance of modern gas turbine engines have imposed increasing service temperature demands on structural materials. Through active cooling, the useful temperature range of nickel-base superalloys in current gas turbine engines has been extended, but the margin for further improvement appears modest. Because of their low density, high-temperature strength, and high thermal conductivity, in situ toughened silicon nitride ceramics have received a great deal of attention for cooled structures. However, high processing costs have proven to be a major obstacle to their widespread application. Advanced rapid prototyping technology, which is developing rapidly, offers the possibility of an affordable manufacturing approach. Derived from text
Gas Turbine Engines; Ceramics
Source: NASA.
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