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At least 91 records · Page 5
All sourcesNASA NTRSDOE OSTI
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

Effect of Mach number on over-all performance of single-stage axial-flow compressor designed for high pressure ratio

A complete stage of an axial-flow compressor designed to produce a high pressure ratio was investigated at speeds from 110 to 130 percent of design speed (836 ft/sec). The data obtained and the results of a previous investigation of the same compressor at speeds from 50 to 100 percent of design speed were used to determine the effect of Mach number on over-all performance. The peak total-pressure ratio increased from 1.095 to 1.515 and the peak adiabatic efficiency decreased from 0.93 to 0.89 as the relative inlet Mach number increased from 0.34 to 0.77. At a relative inlet Mach number of 0.91, a total-pressure ratio of 1.635 was obtained at an adiabatic efficiency of 0.74.

COMPRESSORS - AXIAL FLOW↗
NASA NTRS1943

Performance of NACA Eight-stage Axial-flow Compressor Designed on the Basis of Airfoil Theory

The NACA has conducted an investigation to determine the performance that can be obtained from a multistage axial-flow compressor based on airfoil research. A theory was developed; an eight-stage axial-flow compressor was designed, constructed, and tested. The performance of the compressor was determined for speeds from 5000 to 14,000 r.p.m with varying air flow at each speed. Most of the tests were made with air at room temperature. The performance was determined in accordance with the Committee's recommended procedure for testing superchargers. The expected performance was obtained, showing that a multistage compressor of high efficiency can be designed by the application of airfoil theory.

Sinnette, John T↗
NASA NTRS1947

Investigation of X24C-2 10-Stage Axial-Flow Compressor: Effect of Inlet-Air Pressure and Temperature of Performance - 2

Effect of inlet-air pressure and temperature on the performance of the X24-2 10-Stage Axial-Flow Compressor from the X24C-2 turbojet engine was evaluated. Speeds of 80, 89, and 100 percent of equivalent design speed with inlet-air pressures of 6 and 12 inches of mercury absolute and inlet-air temperaures of approximately 538 degrees, 459 degrees,and 419 degrees R ( 79 degrees, 0 degrees, and minus 40 degrees F). Results were compared with prior investigations.

Finger, Harold B.↗
NASA NTRS1947

Investigation of 10-Stage Axial-Flow X24C-2 Compressor: Performance at Inlet Pressure of 21 Inches Mercury Absolute and Inlet Temperature of 538 R - 1

The performance at inlet pressure of 21 inches mercury absolute and inlet temperature of 538 R for the 10-stage axial-flow X24C-2 compressor from the X24C-2 turbojet engine was investigated. the peak adiabatic temperature-rise efficiency for a given speed generally occurred at values of pressure coefficient fairly close to 0.35.For this compressor, the efficiency data at various speeds could be correlated on two converging curves by the use of a polytropic loss factor derived.

Schum, Harold J.↗
NASA NTRS1974

Calculation of flow distribution in large radius ratio stages of axial flow turbines and comparison of theory and experiment

A method of calculating stage parameters and flow distribution of axial turbines is described. The governing equations apply to space between the blade rows and are based on the assumption of rotationally symmetrical, compressible, adiabatic flow conditions. Results are presented for stage design and flow analysis calculations. Theoretical results from the calculation system are compared with experimental data from low pressure steam turbine tests.

Herzog, J.↗
NASA NTRS1948

Investigation of X24C-2 10-Stage Axial-Flow Compressor. III - Surge Characteristics

Compressor operation at low air flows for a given speed is limited by unstable flow conditions, commonly called surge. An investigation of surge in centrifugal compressors (reference 1) showed that the pulsation of pressures and velocities occurred when the slope of the compressor characteristic curve was positive and that the magnitude and frequency, as well as the incidence of surge, depended on the capacity and resistance of the total system. Although the theory presented in reference 1 is applicable to axial-floe compressors, little experimental information is available on the surge characteristics of the individual stages of axial-flow compressors, or on the variation of the surge characteristics with operating conditions. During the investigation to determine the performance of the X24C-2 compressor (references 2 and 3), instrumentation was added to study the surge characteristics and to determine the effect of speed and inlet pressure on the frequency, amplitude, and phase relation of the pressure pulsations behind each stage.

Buckner, Howard A., Jr.↗
NASA NTRS1951

Some effects of blade trailing-edge thickness on performance of a single-stage axial-flow compressor

A set of modified NACA 65-series blower blades designed for axial inlet velocity, high inlet Mach number, and high blade loading was investigated for trailing-edge thickness of 0.015, 0.030, and 0.045 inch to determine the effect of trailing-edge thickness on single-stage axial-flow-compressor performance. Trailing-edge thickness effects were small except at the highest tip speed investigated (915 ft/sec). Trailing-edge thickness up to 30 percent of maximum blade thickness were used without sacrifice of performance of NACA 65-series blades.

AIRFOILS - THICKNESS↗
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 NTRS2009

Compressor Study to Meet Large Civil Tilt Rotor Engine Requirements

A vehicle concept study has been made to meet the requirements of the Large Civil Tilt Rotorcraft vehicle mission. A vehicle concept was determined, and a notional turboshaft engine system study was conducted. The engine study defined requirements for the major engine components, including the compressor. The compressor design-point goal was to deliver a pressure ratio of 31:1 at an inlet weight flow of 28.4 lbm/sec. To perform a conceptual design of two potential compressor configurations to meet the design requirement, a mean-line compressor flow analysis and design code were used. The first configuration is an eight-stage axial compressor. Some challenges of the all-axial compressor are the small blade spans of the rear-block stages being 0.28 in., resulting in the last-stage blade tip clearance-to-span ratio of 2.4%. The second configuration is a seven-stage axial compressor, with a centrifugal stage having a 0.28-in. impeller-exit blade span. The compressors conceptual designs helped estimate the flow path dimensions, rotor leading and trailing edge blade angles, flow conditions, and velocity triangles for each stage.

Veres, Joseph P.↗
NASA NTRS2009

Compressor Study to Meet Large Civil Tilt Rotor Engine Requirements

A vehicle concept study has been made to meet the requirements of the Large Civil Tilt Rotorcraft vehicle mission. A vehicle concept was determined, and a notional turboshaft engine system study was conducted. The engine study defined requirements for the major engine components, including the compressor. The compressor design-point goal was to deliver a pressure ratio of 31:1 at an inlet weight flow of 28.4 lbm/sec. To perform a conceptual design of two potential compressor configurations to meet the design requirement, a mean-line compressor flow analysis and design code were used. The first configuration is an eight-stage axial compressor. Some challenges of the all-axial compressor are the small blade spans of the rear-block stages being 0.28 in., resulting in the last-stage blade tip clearance-to-span ratio of 2.4 percent. The second configuration is a seven-stage axial compressor, with a centrifugal stage having a 0.28-in. impeller-exit blade span. The compressors conceptual designs helped estimate the flow path dimensions, rotor leading and trailing edge blade angles, flow conditions, and velocity triangles for each stage.

Veres, Joseph P.↗
NASA NTRS1957

Analytical and Experimental Investigation of the Effects of Compressor Interstage Air Bleed on Performance Characteristics of a 13-stage Axial-flow Compressor

Air was bled over the fifth-and tenth-stage rotor-blade rows through ports designed to pass 11 and 9 percent of the inlet flow, respectively, at 80 percent speed. Along the rated operating line the maximum speed at which rotating stall was encountered was lowered by either of these bleeds, and the stall patterns below these speeds were altered so that no dangerous resonant rotor-blade bending vibrations were excited. The combination of the two bleeds completely eliminated rotating stall to at least 50 percent speed. The compressor-discharge weight flow was decreased only at intermediate speeds, and the overall pressure ratio was affected only at intermediate speeds, and the overall pressure ratio was affected only by the combination bleed at intermediate speeds. Fifth-stage bleed increased compressor efficiency at low speeds, and tenth-stage bleed decreased efficiency at intermediate speeds.

Lucas, James G↗
NASA NTRS2007

The Attenuation of a Detonation Wave by an Aircraft Engine Axial Turbine Stage

A Constant Volume Combustion Cycle Engine concept consisting of a Pulse Detonation Combustor (PDC) followed by a conventional axial turbine was simulated numerically to determine the attenuation and reflection of a notional PDC pulse by the turbine. The multi-stage, time-accurate, turbomachinery solver TURBO was used to perform the calculation. The solution domain consisted of one notional detonation tube coupled to 5 vane passages and 8 rotor passages representing 1/8th of the annulus. The detonation tube was implemented as an initial value problem with the thermodynamic state of the tube contents, when the detonation wave is about to exit, provided by a 1D code. Pressure time history data from the numerical simulation was compared to experimental data from a similar configuration to verify that the simulation is giving reasonable results. Analysis of the pressure data showed a spectrally averaged attenuation of about 15 dB across the turbine stage. An evaluation of turbine performance is also presented.

VanZante, Dale↗
NASA NTRS1959

Performance of Typical Rear-Stage Axial-Flow Compressor Rotor Blade Row at Three Different Blade Setting Angles

A comparison of the performance of a single-stage rotor run at three different blade setting angles is presented. The rotor was of a design typical for a last stage of a multistage compressor. At each setting angle, the rotor blade row was operated from 53 to 100 percent of equivalent maximum speed (850 ft/sec tip speed) at constant inlet pressure. Hot-wire anemometry was used to observe rotating-stall and surge patterns in time unsteady flow. Results indicated that an increase in peak pressure ratio and an increase in maximum equivalent weight flow were obtained at each speed investigated when the blade setting angle was decreased. An increase in peak efficiency was achieved with decrease in blade setting angle for part of the range of speeds investigated. However, the peak efficiencies for the three blade setting angles were approximately the same at the maximum speed investigated. The flow ranges for all three configurations were about the same at minimum speed and decreased at almost the same rate when the rotative speed was increased through part of the range of speeds investigated. At maximum speed, the flow range for the smallest setting angle was considerably less than the flow range for the other two configurations. A decrease in efficiency and flow range for the smallest blade setting angle at maximum speed can be attributed primarily to a Mach number effect. In addition, because of the difference in projected axial chord lengths at the casing wall, some effect on performance could be expected from the change in three-dimensional flow occurring at the tip. Rotating-stall characteristics for the two smaller blade setting angles were essentially the same. Only surge could be detected for the largest blade setting angle in the unstable-flow region of operation.

Kussoy, Marvin I.↗
NASA NTRS1959

Investigation of a 4.5-Inch-Mean-Diameter Two-Stage Axial-Flow Turbine Suitable for Auxiliary Power Drives

The design and experimental investigation of a 4.5-inch-mean-diameter two-stage turbine are presented herein and used to study the effect of size on the efficiency of turbines in the auxiliary power drive class. The results of the experimental investigation indicated that design specific work was obtained at design speed at a total-to-static efficiency of 0.639. At design pressure ratio, design static-pressure distribution through the turbine was obtained with an equivalent specific work output of 33.2 Btu per pound and an efficiency of 0.656. It was found that, in the design of turbines in the auxiliary power drive class, Reynolds number plays an important part in the selection of the design efficiency. Comparison with theoretical efficiencies based on a loss coefficient and velocity diagrams are presented. Close agreement was obtained between theory and experiment when the loss coefficient was adjusted for changes in Reynolds number to the -1/5 power.

Wong, Robert Y.↗