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Search indexed NASA NTRS and DOE OSTI research on propulsion, heat transfer, battery materials and energy systems. Follow report and document links to the original sources.

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At least 253 records · Page 14
All sourcesNASA NTRSDOE OSTI
NASA NTRS1948

Effect of Inlet Air Distortion on the Steady-State and Surge Characteristics of an Axial-Flow Turbojet Compressor

An investigation was conducted in an altitude test chamber to determine the effects of inlet airflow distortion on the compressor steady-state and surge characteristics of a high-pressure ratio, axial-flow turbojet engine. Circumferential-type inlet flow distortions were investigated, which covered a range of distortion sector angles from 20 deg to 168 deg and distortion levels up to 22 percent. The presence of inlet airflow distortions at the compressor face resulted in a substantial increase in the local pressure ratio in the distorted region, primarily for the inlet stages. The local pressure ratio in the distorted region for the inlet stages increased as either the distortion sector angle decreased or the percent distortion increased. The average compressor-surge pressure ratio was much more sensitive to inlet airflow distortions at lower engine speeds than at engine speeds near rated. Hence, compressor-surge margin reduction due to inlet airflow distortion was quite severe at the lower engine speeds. Although the average compressor-surge pressure ratio was generally reduced with inlet flow distortion, local pressure ratios across the distorted sector of the compressor were obtained during surge and were significantly greater than the normal compressor-surge pressure ratio. This was a result of increased loading of the inlet stages in the distorted region.

Ciepluch, Carl C.↗
NASA NTRS1982

The WISGSK: A computer code for the prediction of a multistage axial compressor performance with water ingestion

A computer code is presented for the prediction of off-design axial flow compressor performance with water ingestion. Four processes were considered to account for the aero-thermo-mechanical interactions during operation with air-water droplet mixture flow: (1) blade performance change, (2) centrifuging of water droplets, (3) heat and mass transfer process between the gaseous and the liquid phases and (4) droplet size redistribution due to break-up. Stage and compressor performance are obtained by a stage stacking procedure using representative veocity diagrams at a rotor inlet and outlet mean radii. The Code has options for performance estimation with (1) mixtures of gas and (2) gas-water droplet mixtures, and therefore can take into account the humidity present in ambient conditions. A test case illustrates the method of using the Code. The Code follows closely the methodology and architecture of the NASA-STGSTK Code for the estimation of axial-flow compressor performance with air flow.

Tsuchiya, T.↗
NASA NTRS1965

Compressor design system

Thermodynamic equations, solution, and experimental data for aerodynamic design of axial flow compressors

AERODYNAMICS↗
NASA NTRS1984

Four New Capabilities in NASTRAN for Dynamic and Aeroelastic Analyses of Rotating Cyclic Structures

Static aerothermoelastic design/analysis of axial-flow compressors, modal flutter analysis of axial-flow turbomachines, forced vibration analysis of rotating cyclic structures and modal flutter analysis of advanced turbopropellers with highly swept blades are four new capabilities developed and implemented in NASTRAN Level 17.7. The contents, applicability and usefulness of these capabilities which were developed and documented under the sponsorship of NASA's Lewis Research Center are discussed. Overall flowcharts and selected examples are presented.

Elchuri, V.↗
NASA NTRS1965

Compressor design requirements

Cycle analysis of turbojet and turboprop engines - axial flow compressor requirements for gas turbine engines

AXIAL FLOW COMPRESSOR↗
NASA NTRS1999

Multistage Turbomachinery Flows Simulated Numerically

At the NASA Lewis Research Center, a comprehensive assessment was made of the predictive capability of the average passage flow model as applied to multistage axial-flow compressors. This model, which describes the time-averaged flow field within a typical passage of a blade row embedded in a multistage configuration, is being widely used throughout U.S. aircraft industry as an integral part of their design systems. Rotor flow-angle deviation. In this work, detailed data taken within a four and one-half stage large low-speed compressor were used to assess the weaknesses and strengths of the predictive capabilities of the average passage flow model. The low-speed compressor blading is of modern design and employs stator end-bends. Measurements were made with slow- and high response instrumentation. The high-response measurements revealed the velocity components of both the rotor and stator wakes. From the measured wake profiles, we found that the flow exiting the rotors deviated from the rotor exit metal angle to a lesser degree than was predicted by the average passage flow model. This was found to be due to blade boundary layer transition, which recently has been shown to exist on multistage axial compressor rotor and stator blades, but was not accounted for in the average passage model. Consequently, a model that mimics the effects of blade boundary layer transition, Shih k-epsilon model, was incorporated into the average passage model. Simulations that incorporated this transition model showed a dramatic improvement in agreement with data. The altered model thus improved predictive capability for multistage axial-flow compressors, and this was verified by detailed experimental measurement.

Hathaway, Michael D.↗
NASA NTRS1974

Prediction of compressor stall for distorted and undistorted flow by use of a multistage compressor simulation on the digital computer

A simulation technique is presented for the prediction of compressor stall for axial-flow compressors for clean and distorted inlet flow. The simulation is implemented on the digital computer and uses stage stacking and lumped-volume gas dynamics. The resulting nonlinear differential equations are linearized about a steady-state operating point, and a Routh-Hurwitz stability test is performed on the linear system matrix. Parallel compressor theory is utilized to extend the technique to the distorted inlet flow problem. The method is applied to the eight-stage J85-13 compressor.

Daniele, C. J.↗
NASA NTRS1975

Prediction of compressor stall for distorted and undistorted flow by use of a multistage compressor simulation on the digital computer

A simulation technique is presented for the prediction of compressor stall for axial-flow compressors for clean and distorted inlet flow. The simulation is implemented on the digital computer and uses stage stacking and lumped-volume gas dynamics. The resulting nonlinear differential equations are linearized about a steady-state operating point, and a Routh-Hurwitz stability test is performed on the linear system matrix. Parallel compressor theory is utilized to extend the technique to the distorted inlet flow problem. The method is applied to the eight-stage J85-13 compressor. Analytical stall prediction for the undistorted stall line shows good agreement with experimental results. The predicted stall line for distorted inlet flow is in agreement with experimental results only for large distortion extents and/or low distortion levels. Results for low distortion extents and high distortion level do not agree with experimental results.

Daniele, C. J.↗
NASA NTRS1999

Inlet Flow Distortion and Unsteady Blade Response in a Transonic Axial-Compressor Rotor

This paper describes the unsteady blade surface pressures on the first-stage rotor blades of a two-stage transonic axial flow compressor experiencing inlet flow distortion. This study was conducted to demonstrate the ability of a full annulus unsteady Reynolds-averaged Navier-Stokes numerical technique to predict unsteady pressures on the rotor blades operating in a distorted inflow. A total pressure distortion produced by a variable mesh screen mounted near the inlet was used to excite the unsteady blade loading on the rotor. On-blade pressure transducers were used to measure the unsteady blade surface pressure. These pressures and the resulting transient load on the rotor blades were compared to the numerical prediction. It is important to develop numerical techniques to predict these transient loads to better understand the response of compressor blades to forcing functions. With this enhanced understanding and ability to predict these transient forces, more robust compressors can be developed. In the study, a high definition of the inlet flow distortion was achieved by rotating the distortion screens. In this manner the inlet flow distortion and the distortion at the first stage stator leading edge were measured at approximately every 0.7 degrees. This full annulus high definition of the inlet flow distortion was used as the inlet boundary condition for the numerical technique. The experimental measurements and numerical analyses are highly complementary in this study. Detailed comparisons between the measurements and the numerical analyses indicate that the current numerical procedure calculates the unsteady aerodynamic pressure on the blade surfaces reasonably well. Further, the agreement of the measured and predicted rotor exit flow distortion at the first stage stator leading edge provides verification of the numerical technique.

Rabe, D. C.↗
NASA NTRS1972

Performance of a single-stage axial-flow transonic compressor stage with a blade tip solidity of 1.7

The overall and blade-element performance of a transonic compressor stage is presented over the stable operating range at rotative speeds from 50 to 100 percent of design speed. Stage peak efficiency of 0.784 was obtained at a weight flow of 28.6 kilograms per second and a pressure ratio of 1.706. Stall margin at design speed was 11.4 percent. The peak efficiency being significantly less than design efficiency was attributed to: (1) the stator loss and the radial gradient of losses being much higher than design, (2) the losses and blockages associated with the rotor part-span dampers not being incorporated into the design, and (3) mismatch of the rotor and stator badle elements.

Moore, R. D.↗
NASA NTRS1977

Preliminary design study of advanced multistage axial flow core compressors

A preliminary design study was conducted to identify an advanced core compressor for use in new high-bypass-ratio turbofan engines to be introduced into commercial service in the 1980's. An evaluation of anticipated compressor and related component 1985 state-of-the-art technology was conducted. A parametric screening study covering a large number of compressor designs was conducted to determine the influence of the major compressor design features on efficiency, weight, cost, blade life, aircraft direct operating cost, and fuel usage. The trends observed in the parametric screening study were used to develop three high-efficiency, high-economic-payoff compressor designs. These three compressors were studied in greater detail to better evaluate their aerodynamic and mechanical feasibility.

Wisler, D. C.↗
NASA NTRS1978

Performance of single-stage axial-flow transonic compressor with rotor and stator aspect ratios of 1.19 and 1.26, respectively, and with design pressure ratio of 1.82

The overall and blade-element performance of a low-aspect-ratio transonic compressor stage is presented over the stable operating flow range at 70, 90, and 100 percent design speeds. At design speed the rotor and stage achieved peak efficiencies of 0.872 and 0.845 at pressure ratios of 1.875 and 1.842, respectively. The stage stall margin at design speed was 21.8 percent.

Reid, L.↗
NASA NTRS1980

Performance of single-stage axial-flow transonic compressor with rotor and stator aspect ratios of 1.19 and 1.26 respectively, and with design pressure ratio of 2.05

The overall and blade-element performances of a low-aspect-ratio transonic compressor stage are presented over the stable operating flow range for speeds from 50 to 100 percent of design. At design speed the rotor and stage achieved peak efficiencies of 0.876 and 0.840 at pressure ratios of 2.056 and 2.000, respectively. The stage stall margin at design speed was 10 percent.

Moore, R. D.↗
NASA NTRS1982

Performance of single-stage axial-flow transonic compressor with rotor and stator aspect ratios of 1.63 and 1.78, respectively, and with design pressure ratio of 1.82

The overall and blade-element performance of a transonic compressor stage is presented over the stable operating flow range for speeds from 50 to 100 percent of design. The stage was designed for a pressure ratio of 1.82 at a flow 20.2 kg/sec and a tip speed of 455 m/sec. At design speed the stage achieved a peak efficiency of 0.821 at a pressure ratio of 1.817. The stage stall margin at design speed based on conditions at stall and peak efficiency was about 11 percent.

Moore, R. D.↗
NASA NTRS1982

Performance of single-stage axial-flow transonic compressor with rotor and stator aspect ratios of 1.63 and 1.77, respectively, and with design pressure ratio of 2.05

The overall and blade-element performance of a transonic compressor stage is presented over the stable operating range for speeds from 50 to 100 percent of design. The stage was designed for a pressure ratio of 2.05 at a flow of 20.2 kg/sec and a tip speed of 455 m/sec. At design speed the rotor and stage achieved peak efficiencies of 0.849 and 0.831, respectively, at the minimum flow condition. The stage stall point occurred at a flow higher than the design flow.

Moore, R. D.↗