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  • Aircraft Propulsion and Power
  • 1995-1999  (108)
  • 1955-1959  (13)
  • 1
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    Simulation of Crack Propagation in Engine Rotating Components Under Variable Amplitude Loading (1999)
    In:  CASI
    Details
    Publication Date: 2004-12-03
    Description: The crack propagation life of tested specimens has been repeatedly shown to strongly depend on the loading history. Overloads and extended stress holds at temperature can either retard or accelerate the crack growth rate. Therefore, to accurately predict the crack propagation life of an actual component, it is essential to approximate the true loading history. In military rotorcraft engine applications, the loading profile (stress amplitudes, temperature, and number of excursions) can vary significantly depending on the type of mission flown. To accurately assess the durability of a fleet of engines, the crack propagation life distribution of a specific component should account for the variability in the missions performed (proportion of missions flown and sequence). In this report, analytical and experimental studies are described that calibrate/validate the crack propagation prediction capability for a disk alloy under variable amplitude loading. A crack closure based model was adopted to analytically predict the load interaction effects. Furthermore, a methodology has been developed to realistically simulate the actual mission mix loading on a fleet of engines over their lifetime. A sequence of missions is randomly selected and the number of repeats of each mission in the sequence is determined assuming a Poisson distributed random variable with a given mean occurrence rate. Multiple realizations of random mission histories are generated in this manner and are used to produce stress, temperature, and time points for fracture mechanics calculations. The result is a cumulative distribution of crack propagation lives for a given, life limiting, component location. This information can be used to determine a safe retirement life or inspection interval for the given location.
    Keywords: Aircraft Propulsion and Power
    Type: Design Principles and Methods for Aircraft Gas Turbine Engines; 38-1 - 38-8; RTO-MP-8
    Format: text
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  • 2
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    Laser Anemometer Measurements of the Flow Field in a 4:1 Pressure Ratio Centrifugal Impeller (1997)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: A laser-doppler anemometer was used to obtain flow-field velocity measurements in a 4:1 pressure ratio, 4.54 kg/s (10 lbm/s), centrifugal impeller, with splitter blades and backsweep, which was configured with a vaneless diffuser. Measured through-flow velocities are reported for ten quasi-orthogonal survey planes at locations ranging from 1% to 99% of main blade chord. Measured through-flow velocities are compared to those predicted by a 3-D viscous steady flow analysis (Dawes) code. The measurements show the development and progression through the impeller and vaneless diffuser of a through-flow velocity deficit which results from the tip clearance flow and accumulation of low momentum fluid centrifuged from the blade and hub surfaces. Flow traces from the CFD analysis show the origin of this deficit which begins to grow in the inlet region of the impeller where it is first detected near the suction surface side of the passage. It then moves toward the pressure side of the channel, due to the movement of tip clearance flow across the impeller passage, where it is cut by the splitter blade leading edge. As blade loading increases toward the rear of the channel the deficit region is driven back toward the suction surface by the cross-passage pressure gradient. There is no evidence of a large wake region that might result from flow separation and the impeller efficiency is relatively high. The flow field in this impeller is quite similar to that documented previously by NASA Lewis in a large low-speed backswept impeller.
    Keywords: Aircraft Propulsion and Power
    Type: NASA-TM-107541 , NAS 1.15:107541 , ARL-TR-1448 , E-10864 , GRC-E-DAA-TN18827 , Turbo-Expo 1997; Jun 02, 1997 - Jun 05, 1997; Orlando, FL; United States
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  • 3
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    Multidisciplinary Approach to Linear Aerospike Nozzle Optimization (1997)
    In:  CASI
    Details
    Publication Date: 2019-07-10
    Description: A model of a linear aerospike rocket nozzle that consists of coupled aerodynamic and structural analyses has been developed. A nonlinear computational fluid dynamics code is used to calculate the aerodynamic thrust, and a three-dimensional fink-element model is used to determine the structural response and weight. The model will be used to demonstrate multidisciplinary design optimization (MDO) capabilities for relevant engine concepts, assess performance of various MDO approaches, and provide a guide for future application development. In this study, the MDO problem is formulated using the multidisciplinary feasible (MDF) strategy. The results for the MDF formulation are presented with comparisons against sequential aerodynamic and structural optimized designs. Significant improvements are demonstrated by using a multidisciplinary approach in comparison with the single- discipline design strategy.
    Keywords: Aircraft Propulsion and Power
    Type: AIAA Paper 97-3374
    Format: application/pdf
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  • 4
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    The Design of Future Airbreathing Engine Systems within an Intelligent Synthesis Environment (1999)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: This paper describes a new Initiative proposed by the National Aeronautics and Space Administration (NASA). The purpose of this initiative is to develop a future design environment for engineering and science mission synthesis for use by NASA scientists and engineers. This new initiative is called the Intelligent Synthesis Environment (ISE). The paper describes the mission of NASA, future aerospace system characteristics, the current engineering design process, the ISE concept, and concludes with a description of possible ISE applications for the decision of air-breathing propulsion systems.
    Keywords: Aircraft Propulsion and Power
    Type: NASA/TM-1999-209192 , NAS 1.15:209192 , E-11706 , 14th International on Air Breathing Engines; Sep 05, 1999 - Sep 10, 1999; Florence; Italy
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  • 5
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    Application of Pressure Sensitive Paint to Confined Flow at Mach Number 2.5 (1998)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: Pressure sensitive paint (PSP) is a novel technology that is being used frequently in external aerodynamics. For internal flows in narrow channels, and applications at elevated nonuniform temperatures, however, there are still unresolved problems that complicate the procedures for calibrating PSP signals. To address some of these problems, investigations were carried out in a narrow channel with supersonic flows of Mach 2.5. The first set of tests focused on the distribution of the wall pressure in the diverging section of the test channel downstream of the nozzle throat. The second set dealt with the distribution of wall static pressure due to the shock/wall interaction caused by a 25 deg. wedge in the constant Mach number part of the test section. In addition, the total temperature of the flow was varied to assess the effects of temperature on the PSP signal. Finally, contamination of the pressure field data, caused by internal reflection of the PSP signal in a narrow channel, was demonstrated. The local wall pressures were measured with static taps, and the wall pressure distributions were acquired by using PSP. The PSP results gave excellent qualitative impressions of the pressure field investigated. However, the quantitative results, specifically the accuracy of the PSP data in narrow channels, show that improvements need to be made in the calibration procedures, particularly for heated flows. In the cases investigated, the experimental error had a standard deviation of +/- 8.0% for the unheated flow, and +/- 16.0% for the heated flow, at an average pressure of 11 kpa.
    Keywords: Aircraft Propulsion and Power
    Type: NASA/TM-1998-107527 , E-10842 , NAS 1.15:107527 , AIAA Paper 97-3214 , Joint Propulsion Conference; Jul 06, 1997 - Jul 09, 1997; Seattle, WA; United States
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  • 6
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    Modern Experimental Techniques in Turbine Engine Testing (1996)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: The paper describes application of two modern experimental techniques, thin-film thermocouples and pressure sensitive paint, to measurement in turbine engine components. A growing trend of using computational codes in turbomachinery design and development requires experimental techniques to refocus from overall performance testing to acquisition of detailed data on flow and heat transfer physics to validate these codes for design applications. The discussed experimental techniques satisfy this shift in focus. Both techniques are nonintrusive in practical terms. The thin-film thermocouple technique improves accuracy of surface temperature and heat transfer measurements. The pressure sensitive paint technique supplies areal surface pressure data rather than discrete point values only. The paper summarizes our experience with these techniques and suggests improvements to ease the application of these techniques for future turbomachinery research and code verifications.
    Keywords: Aircraft Propulsion and Power
    Type: NASA-TM-107383 , E-10572 , NAS 1.15:107383 , International Congress on Fluid Dynamics and Propulsion; Dec 29, 1996 - Dec 31, 1996; Cairo; Egypt
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  • 7
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    Dynamic Inlet Distortion Prediction with a Combined Computational Fluid Dynamics and Distortion Synthesis Approach (1996)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: A procedure has been developed for predicting peak dynamic inlet distortion. This procedure combines Computational Fluid Dynamics (CFD) and distortion synthesis analysis to obtain a prediction of peak dynamic distortion intensity and the associated instantaneous total pressure pattern. A prediction of the steady state total pressure pattern at the Aerodynamic Interface Plane is first obtained using an appropriate CFD flow solver. A corresponding inlet turbulence pattern is obtained from the CFD solution via a correlation linking root mean square (RMS) inlet turbulence to a formulation of several CFD parameters representative of flow turbulence intensity. This correlation was derived using flight data obtained from the NASA High Alpha Research Vehicle flight test program and several CFD solutions at conditions matching the flight test data. A distortion synthesis analysis is then performed on the predicted steady state total pressure and RMS turbulence patterns to yield a predicted value of dynamic distortion intensity and the associated instantaneous total pressure pattern.
    Keywords: Aircraft Propulsion and Power
    Type: NASA-CR-198053 , NAS 1.26:198053 , H-2129
    Format: application/pdf
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  • 8
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    The Effect of Vane-Blade Spacing on Transonic Turbine Stage Performance (1996)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: This paper presents the results of a computational study on the effect of axial spacing between the vane and blade rows of a transonic turbine stage. The study was performed on the mid-span section of a high-pressure turbine stage using a quasi-3D, unsteady Navier-Stokes solver that provides a fully interactive vane-blade unsteady flow solution. Three different cases were considered, corresponding to axial spacings of 20%, 40%, and 60% of the vane axial chord. The calculated vane and blade pressure distributions for the 40 percent case were found to compare favorably with experimental measurements acquired in a short-duration shock tunnel. In addition, the analysis shows a marked increase in the amplitude of the unsteady pressure fluctuations on the vane and blade surfaces as the spacing decreases. Time-averaged stage adiabatic efficiency predictions for each case are presented to show the effect of spacing on aerodynamic performance.
    Keywords: Aircraft Propulsion and Power
    Type: Loss Mechanisms and Unsteady Flows in Turbomachines; AGARD-CP-571
    Format: text
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  • 9
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    Jet Injection Used to Control Rotating Stall in a High-Speed Compressor (1998)
    In:  CASI
    Details
    Publication Date: 2018-06-05
    Description: 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.
    Keywords: Aircraft Propulsion and Power
    Type: Research and Technology 1997; NASA/TM-1998-206312
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  • 10
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    Three-Component Force Measurements on a Scramjet in a Reflected-Shock Tunnel (1998)
    In:  Other Sources
    Details
    Publication Date: 2019-07-13
    Description: A three-component stress-wave force-balance for a large scramjet has been designed, calibrated and tested in the HYPULSE reflected shock tunnel at GASL Inc., New York. The scramjet model is over 3-foot long and weighs in excess of 90 Ibm. The stress-wave force-balance is comprised of three stress bars which are attached to the model. Calibration results indicate that the force balance responds well within about 1 ms and that the sensitivity of the balance to the distribution of load is not large. Results with and without fuel injection were obtained in the tunnel operated for Mach 7 and Mach 10 flight simulation. These tests showed the force-balance can resolve axial force increments due to combustion of about 40 lb in the presence of model lift forces of 500-700 lb.
    Keywords: Aircraft Propulsion and Power
    Type: JANNAF Airbreathing Propulsion Subcommittee and 35th Combustion Subcommittee Meeting; 1; 35-52; CPIA-Publ-682-Vol-1
    Format: text
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