SCIENTIFIC AND TECHNICAL AEROSPACE REPORTS
A Biweekly Publication of the National Aeronautics and Space Administration
VOLUME 43, ISSUE 19 - SEPTEMBER 23, 2005
07 AIRCRAFT PROPULSION AND POWER - PART II
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.
20050214052 Pratt and Whitney Aircraft, USA
Pratt and Whitney UEET System Study and Demonstration Plan Nordeen, Craig; NASA Glenn Research Center UEET (Ultra-Efficient Engine Technology) Program: Agenda and Abstracts
[2001], pp. 45; In English; See also 20050214031; No Copyright; Abstract Only; Available from CASI only as part of the entire parent document
Pratt & Whitney has conducted an engine system and component study and concluded that a P&W geared turbofan will exceed all UEET requirements. A number of novel concepts have been identified that may exceed the current requirements by an even greater margin. The two phase technology demonstration plan is designed to prove component technologies that will enable future novel configurations. Author
Propulsion System Configurations; Turbofans; Turbofan Engines; Engine Parts; Propulsion System Performance
20050214054 General Electric Aircraft Engines, Cincinnati, OH, USA
Aspirating Seal Demonstration
Tseng, Tom; NASA Glenn Research Center UEET (Ultra-Efficient Engine Technology) Program: Agenda and Abstracts; [2001], pp. 46; In English; See also 20050214031; No Copyright; Abstract Only; Available from CASI only as part of the entire parent document
The performance benefits of using an aspirating seal compared to an existing labyrinth seal was verified through sub- and full-scale rig testing in the NASA Advanced Subsonic Technology (AST) Project, Contract No. NAS3-27720. This seal was designed for GE90 engine application at the low pressure turbine aft outer cavity seal location. The last phase of this program is to demonstrate the seal in the GE90 engine. This presentation is to brief the seal performance, engine test plan and status. Author
Labyrinth Seals; Engine Tests; Performance Tests; Cavities; Low Pressure
20050214060 Pratt and Whitney Aircraft, USA
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Variable Area Nozzle Development
Larkin, Mike; NASA Glenn Research Center UEET (Ultra-Efficient Engine Technology) Program: Agenda and Abstracts; [2001], pp. 40; In English; See also 20050214031; No Copyright; Abstract Only; Available from CASI only as part of the entire parent document
The ability to control fan nozzle exit area is an enabling technology for next generation high-bypass-ratio turbofan engines. Performance benefits for such designs are estimated at up to 8 percent in thrust specific fuel consumption (TSFC) and 5 to 6 percent in range (or mission fuel burn) relative to current fixed-geometry engines. Conventionally actuated variable area fan nozzle (VAN) concepts tend to be heavy and complicated, with significant aircraft integration, reliability and packaging issues. The goal of this effort was to eliminate these undesirable features and formulate a design that meets or exceeds leakage, durability, reliability, maintenance and manufacturing cost goals.AShape Memory Alloy (SMA) bundled cable actuator acting to move an array of flaps around the fan nozzle annulus is a concept that meets these requirements. In 1999, a full-scale sector model of this VAN system was built and then tested at the Jet Exit Test (JET) Facility at NASA Langley to demonstrate the system s ability to achieve 20 percent area variation of the nozzle under full-scale aerodynamic loads. In 2000, a closed loop position controller was developed and validated against full-scale loads at the JET Facility. Currently, flexible seal concepts to seal the nozzle flap-to-flap and other VAN-to-fixed structural elements are being examined. Potential seal materials are also being evaluated against chemical and temperature environment requirements. Finally, a durability test is in progress on the SMA bundled cable in the model at simulated full-scale loads. Author
Turbofan Engines; Turbine Exhaust Nozzles; Aerodynamic Loads; Feedback Control; Actuators
20050214063 Rolls-Royce Allison, Indianapolis, IN, USA
Rolls-Royce Low Emissions Combustor Development
Rizk, Nader; NASA Glenn Research Center UEET (Ultra-Efficient Engine Technology) Program: Agenda and Abstracts; [2001], pp. 6; In English; See also 20050214031; No Copyright; Abstract Only;Available from CASI only as part of the entire parent document
The objective for this program is to develop low emissions combustion systems, and integrate those systems into turbofan engines for subsonic defense systems and commercial regional applications. The progress made so far involved the development of two combustor concepts through testing in flame-tube, sector, and full combustor configurations. The testing of the full combustor under advanced regional engine cycles demon- stratedmore than50 percentN0,reductionfrom the 1996ICAO. TomeettheUEET70 percentN0,reduction goal, a concentric fuel-staged injector was tested in a flametube. This concept will be used in the combustor sector to be developed during the next 18 months. The effort made under this program supports Level I milestones 13,14,15, 16,18, and 22. Author
Combustion Chambers; Turbofan Engines; Fuel Injection
20050214064 Stanford Univ., CA, USA
Stanford Activities In Concepts for Advanced Gas Turbine Combustors
Edwards, Chris; NASA Glenn Research Center UEET (Ultra-Efficient Engine Technology) Program: Agenda and Abstracts; [2001], pp. 9; In English; See also 20050214031; No Copyright; Abstract Only;Available from CASI only as part of the entire parent document
This presentation reports the activities of a coordinated program of research to improve the efficiency and reduce emissions of aircraft gas turbine engines through innovative concepts for gas turbine combustion. Activities are in progress in four areas:
(1) combustor measurements for development and validation of LES methods for gas turbine combustor simulations,
(2) LES simulations of gas-phase combustion and spray combustion,
(3) construction and testing of a temporally modulated spray injector for use in the study and control of combustion instability, and
(4) development and testing of mesoscale injector arrays for use in gas turbine combustors.
Combustor measurements using gaseous fuel have been completed and liquid-fuel tests are now in progress. Measurements include mean temperature, major species, and two-component velocity mapping, and spray characterization in both combusting and cold-flow cases. Large Eddy Simulations (LES) of the combustor flow field are reported in a companion paper at this forum. Results are reported on development and characterization of a temporally modulated liquid fuel injector that has a size distribution function and cone angle that are independent of modulation frequency. This injector will provide a well-defined actuator for use in experimental and computational studies of combustion control. The mesoscale injector array is a spatial analog to the temporally modulated injector. It produces compact/ distributed, nearly premixed flames that have the potential to replace conventional approaches to combustion in both main-burner and turbine-reheat applications with significant performance improvements and emissions reductions. Results of prototype testing are reported.
Combustible Flow; Combustion Chambers; Combustion Control; Fuel Injection
20050214065 Stanford Univ., CA, USA
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Large Eddy Simulation (LES) of Gas Turbine Combustion
Moin, Parviz; NASA Glenn Research Center UEET (Ultra-Efficient Engine Technology) Program: Agenda and Abstracts; [2001], pp. 10; In English; See also 20050214031; No Copyright; Abstract Only; Available from CASI only as part of the entire parent document
Two new approaches to chemistry modeling for large eddy simulation (LES) of turbulent reacting flows are developed. Both approaches use the flamelet library and therefore include full complex chemistry. The progress variable approach, which maps the detailed chemical processes to a reduced set of tracking scalars, has been applied to a methane jet combustor for which experimental data is available. The progress variable approach is able to capture the unsteady, lifted flame dynamics observed in the experiment, to obtain good agreement with the experimental data, and to significantly out-perform the fast chemistry and steady flamelet models, which both predict an attached flame. The unsteady flamelet approach is the alternative method that is being developed in parallel at CTR for LES of turbulent combustion. The unsteady flamelet equations are solved simultaneously with the flow and scalar equations to account for transient chemical effects. These transient effects have been shown to be very important for prediction of pollutants such as nitrous oxides and smoke. This technique has been applied to the so-called Sandia Flame D, and the agreement of temperature and species concentrations with the data is excellent. The CTR program also includes a large effort in the development of numerical methods and codes for LES of turbulent combustion in complex geometry. An unstructured mesh code has been developed for parallel computer platforms. The numerical method is non-dissipative, which has been shown to be a critical property for LES. This code has been applied to a combustor sector of PW6000. In addition, a new numerical method for low Mach number flows capable of accurately resolving long wavelength acoustic waves responsible for combustion instabilities has been developed. Author
Gas Turbines; Combustion Chambers; Turbulent Combustion; Large Eddy SIMulation
20050214066 NASA Glenn Research Center, Cleveland, OH, USA
NASA National Combustion Code Simulations
Iannetti, Anthony; Davoudzadeh, Farhad; NASA Glenn Research Center UEET (Ultra-Efficient Engine Technology) Program: Agenda and Abstracts; [2001], pp. 11; In English; See also 20050214031; No Copyright; Abstract Only;Available from CASI only as part of the entire parent document
A systematic effort is in progress to further validate the National Combustion Code (NCC) that has been developed at NASA Glenn Research Center (GRC) for comprehensive modeling and simulation of aerospace combustion systems. The validation efforts include numerical simulation of the gas-phase combustor experiments conducted at the Center for Turbulence Research (CTR), Stanford University, followed by comparison and evaluation of the computed results with the experimental data. Presently, at GRC, a numerical model of the experimental gaseous combustor is built to simulate the experimental model. The constructed numerical geometry includes the flow development sections for air annulus and fuel pipe, 24 channel air and fuel swirlers, hub, combustor, and tail pipe. Furthermore, a three-dimensional multi-block, multi-grid grid (1.6 million grid points, 3-levels of multi-grid) is generated. Computational simulation of the gaseous combustor flow field operating on methane fuel has started. The computational domain includes the whole flow regime starting from the fuel pipe and the air annulus, through the 12 air and 12 fuel channels, in the combustion region and through the tail pipe. Author
Combustion Chambers; Computational Grids; Computer Programs; Fuel Combustion; Flow Distribution
20050214073 NASA Glenn Research Center, Cleveland, OH, USA
Turbine Chemistry Modeling
Liu, Nan-Suey; Wey, Thomas; NASA Glenn Research Center UEET (Ultra-Efficient Engine Technology) Program: Agenda and Abstracts; [2001], pp. 50; In English; See also 20050214031; No Copyright; Abstract Only; Available from CASI only as part of the entire parent document
Many of the engine exhaust species resulting in significant environmental impact exist in trace amounts. Recent research, e.g., conducted at MIT-AM, has pointed to the intra-engine environment as a possible site for important trace chemistry activity. In addition, the key processes affecting the trace species activity occurring downstream in the air passages of the turbine and exhaust nozzle are not well understood. Most recently, an effort has been initiated at NASA Glenn Research Center under the UEET Program to evaluate and further develop CFD-based technology for modeling and simulation of intra-engine trace chemical changes relevant to atmospheric effects of pollutant emissions from aircraft engines. This presentation will describe the current effort conducted at Glenn; some preliminary results relevant to the trace species chemistry in a turbine passage will also be presented to indicate the progress to date. Author
Computational Fluid Dynamics; Trace Elements; Environmental Surveys; Exhaust Emission; Contaminants
20050214076 Army Research Lab., Cleveland, OH, USA
Compressor Flow Control Concepts, 1, Advanced Casing Treatment
Hathaway, Michael D.; NASA Glenn Research Center UEET (Ultra-Efficient Engine Technology) Program: Agenda and Abstracts; [2001], pp. 27; In English; See also 20050214031; No Copyright; Abstract Only; Available from CASI only as part of the entire parent document
The goal of this research program was to develop an improved casing treatment concept for application to high-speed compressors which improves on existing capabilities. As such, a state-of-the-art CFD code (APNASA) was employed in a computationally based parametric investigation of the impact of casing suction and injection on the stability and performance of a low-speed fan rotor wherein the stalling mass flow was controlled by tip flow field breakdown. The parametric investigation was guided by observed trends in end wall flow characteristics as stall is approached, and based on the hypothesis that application of suction or blowing can mitigate these trends. The best suction and injection configurations were then combined to yield a self-recirculating casing treatment concept which improves on existing recirculating casing treatment designs. The results of this parametric investigation yielded (1) identification of the fluid mechanisms which precipitate stall of tip critical blade rows, and (2) an approach to recirculated casing treatment design which produces benefits in both compressor stall range and efficiency. Subsequent application of this approach to a high speed transonic rotor successfully yielded similar improvements in stall range with no loss in compressor efficiency. Author
Computational Fluid Dynamics; Compressor Effýciency; Compressors; Flow Characteristics; Flow Distribution; Suction; Stability; Injection
20050214080 NASA Glenn Research Center, Cleveland, OH, USA
Status of Glenn-HT Code Applications to Complex Cooling Geometries
Heidmann, James; NASA Glenn Research Center UEET (Ultra-Efficient Engine Technology) Program: Agenda and Abstracts; [2001], pp. 30; In English; See also 20050214031; No Copyright; Abstract Only; Available from CASI only as part of the entire parent document
The Glenn-HT code is a 3D Navier-Stokes solver that has been used and validated for a variety of convective heat transfer problems associated with turbine flows. These flows have included tip clearance, simplified internal cooling, and film cooling. The multi-block capability of the code makes it particularly useful for the complex geometries of such flows. One of the goals of the UEET program is to reduce turbine cooling flow while increasing turbine inlet temperature. The Glenn-HT code gives researchers a tool to analyze the flow within the very complicated geometries associated with actual cooled turbine designs. Through these analyses and their comparison with experimental data, it is hoped to extend the applicability of the Glenn-HT code for use as a tool to improve turbine cooling designs to meet UEET goals. Author
Navier-Stokes Equation; Convective Heat Transfer; Turbines; Film Cooling; Inlet Temperature
20050214081 Mississippi State Univ., MS, USA
MSU-TURBO Enhancements, 1, Enhancements and Validation Overview
Chen, Jen-Ping; NASA Glenn Research Center UEET (Ultra-Efficient Engine Technology) Program: Agenda and Abstracts; [2001], pp. 32; In English; See also 20050214031; No Copyright; Abstract Only; Available from CASI only as part of the entire parent document
Numerical simulation of the unsteady flows between multiple stages in turbomachinery has become a desired method to help in the understanding of the complex flows found in modem compressor designs. The computation of complete flow fields involving multiple blades typically requires a tremendous amount of computing resources, which soon exceeds that which can be provided by most single-processor supercomputers. A high-fidelity computer simulation of the complete four-stage UEET compressor will greatly aid in the development of this advanced technology. Parallel computing provides a practical way for such large-scale simulations. A parallel code based on a serial version of the TURBO code was developed. This code inherited most of the features found in the serial TURBO code. It was made to support general blocking domain decomposition, which will greatly aid in resolving details of the flow field. Thus far, validation has been by comparing results from the serial and parallel codes using NASA Stage 37. Computation of a 1.5-stage UEET initial design was also compared to the APNASA results. Estimated parallel efficiency of this code is about 85 percent. Author
Unsteady Flow; Turbomachinery; Computerized SIMulation; Flow Distribution; Compressors; Estimating
20050214420 NASA Glenn Research Center, Cleveland, OH, USA
Computer Code for Gas Turbine Engine Weight and Disk Life Estimation Improved
Tong, Michael T.; Research and Technology 2002; March 2003; 2 pp.; In English; No Copyright;Avail: CASI; A01, Hardcopy
Engine weight is a key design parameter for any new aircraft engine. It affects aircraft range and is a key element in fuel burn. Weight is also considered to be an indicator of engine cost. Reliable engine-weight estimation at the conceptual design stage is critical to the development of new aircraft engines. It helps to identify the best engine concept from among several candidates. At the NASA Glenn Research Center, the Weight Analysis of Turbine Engines (WATE) computer code, originally developed by Boeing Aircraft, is being used to estimate the engine weight of various conceptual engine designs. The code was originally developed for NASA in 1979, but since then substantial improvements were made to the code to improve the weight calculations for most of the engine components. Recently, a series of efforts were performed at Glenn to enhance the capability of the WATE code. The major enhancements include the incorporation of improved weight-calculation routines for the compressor and turbine disks using the finite-difference technique. In addition, the stress distribution for various disk geometries was incorporated for a life-prediction module to calculate disk life. A material database, consisting of the material data of most of the commonly used aerospace materials was also incorporated into WATE. Collectively, these enhancements provide a more realistic and systematic way to calculate engine weight. The current effort paves the way for an integrated engine design tool, which would easily allow engine developers to perform design tradeoffs between engine weight, durability, and cost. To demonstrate the new capabilities, Glenn researchers used the improved WATE code to perform an engine weight/life tradeoff assessment on a 90 000-lb-thrust-class commercial turbofan engine. Derived from text
Engine Design; Design Analysis; Turbine Engines; Gas Turbine Engines
20050214604 United Technologies Corp., East Hartford, CT, USA
Alternative Liquid Fuel Effects on Cooled Silicon Nitride Marine Gas Turbine Airfoils. Final Report
Holowczak, J.; Feb. 2002; 64 pp.; In English Report No.(s): DE2005-836627; No Copyright; Avail: Department of Energy Information Bridge
With prior support from the Office of Naval Research, DARPA, and U.S. Department of Energy, United Technologies is developing and engine environment testing what we believe to be the first internally cooled silicon nitride ceramic turbine vane in the USA. The vanes are being developed for the FT8, an aeroderivative stationary/marine gas turbine. The current effort resulted in further manufacturing and development and prototyping by two U.S. based gas turbine grade silicon nitride component manufacturers, preliminary development of both alumina, and YTRIA based environmental barrier coatings (EBC’s) and testing or ceramic vanes with an EBC coating. NTIS
Airfoils; Gas Turbine Engines; Gas Turbines; Liquid Fuels; Marine Propulsion; Nitrides; Silicon Nitrides
Source: NASA.
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