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At least 127 records · Page 7
All sourcesNASA NTRSDOE OSTI
NASA NTRS1991

The numerical simulation of a high-speed axial flow compressor

The advancement of high-speed axial-flow multistage compressors is impeded by a lack of detailed flow-field information. Recent development in compressor flow modeling and numerical simulation have the potential to provide needed information in a timely manner. The development of a computer program is described to solve the viscous form of the average-passage equation system for multistage turbomachinery. Programming issues such as in-core versus out-of-core data storage and CPU utilization (parallelization, vectorization, and chaining) are addressed. Code performance is evaluated through the simulation of the first four stages of a five-stage, high-speed, axial-flow compressor. The second part addresses the flow physics which can be obtained from the numerical simulation. In particular, an examination of the endwall flow structure is made, and its impact on blockage distribution assessed.

Mulac, Richard A.↗
NASA NTRS1953

Analysis of Off-design Performance of a 16-stage Axial-flow Compressor with Various Blade Modifications

The over-all performance of a 16-stage axial-flow compressor was determined with various stator-blade resettings and a reduction in solidity of the rotor blades in the last three stages. It was shown that little control over the sudden change in slope of the surge-limit line at intermediate speeds was obtained with the blade modifications attempted, except that some change in speed at which the change in slope occurred could be effected by stator-blade resettings. Interstage data indicated that the severe surge limit at intermediate speeds was caused by stall of the inlet stage, which, because of stage interaction effects, resulted in a simultaneous decrease in performance of the following five or six stages. Stage data are presented which indicate the flow and pressure-ratio range over which each stage is required to operate at compressor speeds from 50 to 100 percent of design speed.

COMPRESSORS - AXIAL-FLOW↗
NASA NTRS1950

The Supersonic Axial-Flow Compressor

An investigation has been made to explore the possibilities of axial-flow compressors operating with supersonic velocities into the blade rows. Preliminary calculations showed that very high pressure ratios across a stage, together with somewhat increased mass flows, were apparently possible with compressors which decelerated air through the speed of sound in their blading. The first phase of the investigation was the development of efficient supersonic diffusers to decelerate air through the speed of sound. The present report is largely a general discussion of some of the essential aerodynamics of single-stage supersonic axial-flow compressors. As an approach to the study of supersonic compressors, three possible velocity diagrams are discussed briefly. Because of the encouraging results of this study, an experimental single-stage supersonic compressor has been constructed and tested in Freon-12. In this compressor, air decelerates through the speed of sound in the rotor blading and enters the stators at subsonic speeds. A pressure ratio of about 1.8 at an efficiency of about 80 percent has been obtained.

Kantrowitz, Arthur↗
NASA NTRS1951

Discussion of boundary-layer characteristics near the casing of an axial-flow compressor

Boundary-layer velocity profiles on the casing of an axial-flow compressor behind the guide vanes and rotor were measured and resolved into two components: along the streamline of the flow and perpendicular to it. Boundary-layer thickness and the deflection of the boundary layer at the wall were the generalizing parameters. By use of these results and the momentum-integral equations, the characteristics of boundary on the walls of axial-flow compressor are qualitatively discussed. Important parameters concerning secondary flow in the boundary layer appear to be turning of the flow and the product of boundary-layer thickness and streamline curvature outside the boundary layer. Two types of separation are shown to be possible in three dimensional boundary layer.

BOUNDARY LAYER CHARACTERISTICS - INTERNAL AERODYNA↗
NASA NTRS1986

Performance studies on an axial flow compressor stage

A low-speed, medium loaded axial flow compressor stage is studied experimentally and theoretically. The flow compressor facility, composed of an inlet guide vane row, a rotor blade row, and a stator blade row, and the principles of the streamline curvature method (SCM) and the Douglas-Neumann cascade program are described. The radial distribution of the flow properties, the rotor blade static pressure distribution, and the lift coefficient and relative flow angle derived experimentally and theoretically are compared. It is determined that there is good correlation between the experimental flow properties and the SCM data, the Douglas-Neumann cascade program and experimental rotor blade static pressure data, and the experimental and theoretical lift coefficients only in the midspan region. Modifications to the SCM and the Douglas-Neumann cascade program in order to improve their accuracy are discussed.

Sitaram, N.↗
NASA NTRS1988

Effect of stage loading on endwall flows in an axial flow compressor rotor

This paper reports results from investigations conducted to determine the effect of stage loading on endwall flows in a low speed axial flow compressor. These investigations consisted of two sets of measurements. The first set consisted of radial transverse of flow properties at the rotor inlet and exit, at five flow coefficients. These measurements are used to determine the boundary layer integral parameters. The displacement thicknesses at the rotor hub and tip agree reasonably well with Smith's (1970) correlation for multistage axial compressors. The second set consisted of measurements of static pressures on the rotor blade at four flow coefficients. From these measurements, lift coefficient is determined. Also, loss of lift coefficient near the tip is calculated, and is attributed mainly to the tip leakage flows.

Sitaram, N.↗
NASA NTRS1996

Experimental Investigation of the Flow Field in a Transonic, Axial Flow Compressor with Respect to the Development of Blockage and Loss

A detailed experimental investigation to understand and quantify the development of loss and blockage in the flow field of a transonic, axial flow compressor rotor has been undertaken. Detailed laser anemometer measurements were acquired upstream, within, and downstream of a transonic, axial compressor rotor operating at design and off-design conditions. The rotor was operated at 100%, 85%, 80%, and 60% of design speed which provided inlet relative Mach numbers at the blade tip of 1.48, 1.26, 1.18, and 0.89 respectively. At design speed the blockage is evaluated ahead of the rotor passage shock, downstream of the rotor passage shock, and near the trailing edge of the blade row. The blockage is evaluated in the core flow area as well as in the casing endwall region. Similarly at pm speed conditions for the cases of (1) where the rotor passage shock is much weaker than that at design speed and (2) where there is no rotor passage shock, the blockage and loss are evaluated and compared to the results at design speed. Specifically, the impact of the rotor passage shock on the blockage and loss development, pertaining to both the shock/boundary layer interactions and the shock/tip clearance flow interactions, is discussed. In addition, the blockage evaluated from the experimental data is compared to (1) an existing correlation of blockage development which was based on computational results, and (2) computational results on a limited basis. The results indicate that for this rotor the blockage in the endwall region is 2-3 times that of the core flow region and the blockage in the core flow region more than doubles when the shock strength is sufficient to separate the suction surface boundary layer. The distribution of losses in the care flow region indicate that the total loss is primarily comprised of the shock loss when the shock strength is not sufficient to separate the suction surface boundary layer. However, when the shock strength is sufficient to separate the suction surface boundary layer, the profile loss is comparable to the shock loss and can exceed the shock loss.

Suder, Kenneth L.↗
NASA NTRS1952

Preliminary Investigation of Flow Fluctuations During Surge and Blade Row Stall in Axial-flow Compressors / c Merle C. Huppert

A preliminary investigation of the flow fluctuations of surge and blade row stall was conducted with three single-stage axial-flow compressors with hub-tip ratios of 0.9,0.8, and 0.5 and with a multistage axial-flow compressor. Flow fluctuations of large amplitude associated with stall were detected in all compressors investigated. The fluctuations were caused by low flow regions affecting 25 to 40 percent of the annulus area and propagating in the direction of compressor rotation, but at a lower speed. Mild audible surge was obtained with the single-stage compressors with hub-tip ratios of 0.9 and 0.8. No audible surge was detected with the 0.5 hub-tip ratio stage. A severe audible surge was obtained with the multistage axial-flow compressor. The surge point was coincident with the sharp drop in pressure ratio due to stall.

ENGINES, TURBO-JET↗
NASA NTRS1952

Discussion of Boundary-Layer Characteristics Near the Wall of an Axial-Flow Compressor

The boundary-layer velocity profiles in the tip region of an axial-flow compressor downstream of the guide vanes and downstream of the rotor were measured by use of total-pressure and claw-type yaw probes. These velocities were resolved into two components: one along the streamline of the flow outside the boundary layer, and the other perpendicular to it. The affinity among all profiles was thus demonstrated with the boundary-layer thickness and the deflection of the boundary layer at the wall as the generalizing parameters. By use of these results and the momentum-integral equations, boundary-layer characteristics on the walls of an axial-flow compressor were qualitatively evaluated.

Mager, Artur↗
NASA NTRS1974

Blade selection for a modern axial-flow compressor

The procedures leading to successful design of an axial flow compressor are discussed. The three related approaches to cascade selection are: (1) experimental approach which relies on the use of experimental results from identical cascades to satisfy the velocity diagrams calculated, (2) a purely analytical procedure whereby blade shapes are calculated from the theoretical cascade and viscous flow equations, and (3) a semiempirical procedure which used experimental data together with the theoretically derived functional relations to relate the cascade parameters. Diagrams of typical transonic blade sections with uncambered leading edges are presented.

Wright, L. C.↗
NASA NTRS1948

Analysis of effect of basic design variables on subsonic axial-flow-compressor performance

A blade-element theory for axial-flow compressors has been developed and applied to the analysis of the effects of basic design variables such as Mach number, blade loading, and velocity distribution on compressor performance. A graphical method that is useful for approximate design calculations is presented. The relations among several efficiencies useful in compressor design are derived and discussed. The possible gains in useful operating range obtainable by the use of adjustable stator blades are discussed and a rapid approximate method of calculating blade-angle resettings is shown by an example. The relative Mach number is shown to be a dominant factor in determining the pressure ratio.

Sinnette, John T , Jr↗
NASA NTRS1951

Effect of Blade-surface Finish on Performance of a Single-stage Axial-flow Compressor

A set of modified NACA 5509-34 rotor and stator blades was investigated with rough-machine, hand-filed, and highly polished surface finishes over a range of weight flows at six equivalent tip speeds from 672 to 1092 feet per second to determine the effect of blade-surface finish on the performance of a single-stage axial-flow compressor. Surface-finish effects decreased with increasing compressor speed and with decreasing flow at a given speed. In general, finishing blade surfaces below the roughness that may be considered aerodynamically smooth on the basis of an admissible-roughness formula will have no effect on compressor performance.

COMPRESSORS - AXIAL FLOW↗
NASA NTRS1974

Computer program for definition of transonic axial-flow compressor blade rows

A method is presented for designing axial-flow compressor blading from blade elements defined on cones which pass through the blade-edge streamline locations. Each blade-element centerline is composed of two segments which are tangent to each other. The centerline and surfaces of each segment have constant change of angle with path distance. The stacking line for the blade elements can be leaned in both the axial and tangential directions. The output of the computer program gives coordinates for fabrication and properties for aeroelastic analysis for planar blade sections. These coordinates and properties are obtained by interpolation across conical blade elements. The program is structured to be coupled with an aerodynamic design program.

Crouse, J. E.↗