Experimental evaluation of outer case blowing or bleeding of single stage axial flow compressor. Part 4 - Performance of bleed insert configuration no. 3
Evaluation of bleed insert configuration of single stage axial flow compressor
Koch, C. C.↗
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Frequency response of an axial-flow compressor exposed to inlet pressure perturbations
Experimental results of a series of engine tests designed to obtain the stage dynamics of an eight-stage axial-flow compressor over the frequency range of 0.5 to 200 hertz are presented. The total pressure at the compressor face was varied by means of a secondary air jet system installed in the engine inlet and positioned to oppose the primary airflow. Total-pressure probes located at each compressor stage were used to obtain the frequency response of each compressor-stage total pressure to the average compressor-inlet total pressure. The engine operating conditions were chosen to illustrate the effects of changing the rotor speed, changing the exhaust nozzle area, and isolating the compressor discharge pressure perturbations from the fuel control and hence, the fuel flow.
Milner, E. J.↗
Analysis of supersonic stall bending flutter in axial-flow compressor by actuator disk theory
An analytical model was developed for predicting the onset of supersonic stall bending flutter in axial-flow compressors. The analysis is based on two-dimensional, compressible, unsteady actuator disk theory. It is applied to a rotor blade row by considering a cascade of airfoils. The effects of shock waves and flow separation are included in the model. Calculations show that the model predicts the onset, in an unshrouded rotor, of a bending flutter mode that exhibits many of the characteristics of supersonic stall bending flutter. The validity of the analysis for predicting this flutter mode is demonstrated.
Adamczyk, J. J.↗
Predicting the Performance of an Axial-Flow Compressor
Stage-stacking computer code (STGSTK) developed for predicting off-design performance of multi-stage axial-flow compressors. Code uses meanline stagestacking method. Stage and cumulative compressor performance calculated from representative meanline velocity diagrams located at rotor inlet and outlet meanline radii. Numerous options available within code. Code developed so user modify correlations to suit their needs.
Steinke, R. J.↗
End-wall and profile losses in a low-speed axial flow compressor rotor
The blade-to-blade variation of relative stagnation pressure losses in the tip region inside the rotor of a single-stage, axial-flow compressor, is presented and interpreted in this paper. The losses are measured at two flow coefficients (one at the design point and the other at the near peak pressure rise point) to discern the effect of blade loading on the end-wall losses. The tip clearance losses are found to increase with an increase in the pressure rise coefficient. The losses away from the tip region and near the hub regions are measured downstream. The losses are integrated and interpreted in this paper.
Lakshminarayana, B.↗
Performance of a 1380-foot-per-second-tip-speed axial-flow compressor rotor with a blade tip solidity of 1.3
Aerodynamic design parameters are presented along the overall and blade element performance, of an axial flow compressor rotor designed to study the effects of blade solidity on efficiency and stall margin. At design speed the peak efficiency was 0.844 and occurred at an equivalent weight flow of 63.5 lb/sec with a total pressure ratio of 1.801. Design efficiency, pressure ratio, and weight flow 0.814, 1.65, and 65.3(41.1 lb/sec/sq ft of annulus area), respectively. Stall margin for design speed was 6.4 percent based on the weight flow and pressure ratio values at peak efficiency and just prior to stall.
Hager, R. D.↗
Axial flow compressor computer program for calculating off design performance /Program 4/
Computer program for calculating off-design aerodynamic performance of axial flow compressor
Carmody, R. H.↗
Performance of 1380 foot per second tip-speed axial-flow compressor rotor blade tip solidity of 1.5
This presents the aerodynamic design parameters along with the overall and blade element performance of an axial-flow compressor rotor designed to study the effects of blade solidity on efficiency and stall margin. At design speed the peak efficiency was 0.892 and occurred at an equivalent weight flow of 65.0 lb/sec. The total pressure ratio was 1.83 and the total temperature ratio was 1.215. Design efficiency, weight flow, pressure ratio, and temperature ratio were 0.824, 65.3, 1.65, and 1.187, respectively. Stall margin for design speed was 10 percent based on the weight flow and pressure ratio values at peak efficiency and just prior to stall.
Ball, C. L.↗
STGSTK: A computer code for predicting multistage axial flow compressor performance by a meanline stage stacking method
A FORTRAN computer code is presented for off-design performance prediction of axial-flow compressors. Stage and compressor performance is obtained by a stage-stacking method that uses representative velocity diagrams at rotor inlet and outlet meanline radii. The code has options for: (1) direct user input or calculation of nondimensional stage characteristics; (2) adjustment of stage characteristics for off-design speed and blade setting angle; (3) adjustment of rotor deviation angle for off-design conditions; and (4) SI or U.S. customary units. Correlations from experimental data are used to model real flow conditions. Calculations are compared with experimental data.
Steinke, R. J.↗
Performance of 1380-foot-per-second tip-speed axial-flow compressor rotor with blade tip solidity of 1.1
The aerodynamic design parameters are presented along with the overall and blade element performance, of an axial-flow compressor rotor designed to study the effects of blade solidity on efficiency and stall margin. At design speed the peak efficiency was 0.853 and occured at an equivalent weight flow of 65.7lb/sec. The total pressure ratio was 1.68. Design efficiency, weight flow, pressure ratio, and temperature ratio were 0.822, 65.3, 1.65, and 1.187, respectively. Stall margin for design speed was 14 percent based on the weight flows and pressure ratios at peak efficiency and just prior to stall.
Janetzke, D. C.↗
Vibration Survey of Blades in 19XB Axial-Flow Compressor: Dynamic Investigation - 2
Strain-gage measurements were taken under operating conditions from blades of various stages of the 19XB axial-flow compressor in an effort to determine the reason for failures in the seventh and tenth stages. First bending-mode vibrations were detected in the first five stages of the compressor caused by each integral multiple of rotor speed from three through ten. Lead-wire failures in the last five stages resulted in incomplete data. The dynamic-vibration frequencies at various rotor speeds were compared with statically measured frequencies analytically corrected for the influence of centrifugal force. Large increases in vibration ani~litude with increased pressure ratio were observed. During surging operation, blade vibrations were not present. The effects of pressure ratio and surge indicate the existence of aerodynamic excitation as the cause of the blade vibrations.
Meyer, Andre J., Jr.↗
Effect of Various Blade Modifications in Performance of a 16-Stage Axial-flow Compressor. IV - Effect on Over-all Performance Characteristics of Decreasing Twelfth through Fifteenth Stage Stator-blade Angles 3 deg and Increasing Stator Angles in the Inlet Stages
The performance of a 16-stage axial-flow compressor, in which two modifications of unloaded inlet stages were combined with loaded exit stages, has been determined. In the first modification the exit stages were loaded by decreasing the twelfth through fifteenth stage stator angles 3 deg. as compared with the blade angles in the original compressor, and the inlet stages were unloaded by increasing the blade angles the following amounts: guide vanes and first-stage stator, 6 deg; second- and third-stage stators, 4 deg.; and fourth-stage stators, 3 deg. The over-all performance of this configuration was compared with that of the compressor with the original blade angles. The peak efficiency was increased at all speeds below design and the weight flow was higher at speeds below 80 percent of design, the same at 80 percent of design, and lower at speeds abovce 80 percent of design. The maximum reduction in weight flow occurred at design speed. The surge limit line was higher at speeds between 75 and 90 percent of design when presented on a pressure ratio against weight flow basis. The second configuration was the same as the first with the exception that the second-, third-, and fourth-stage stator blade angles were the same as in the compressor with the original blade angles. A comparison of the performance of this configuration with that of the compressor with the original blade angles showed the same general trends of changes in performance as the first configuration. Comparisons were made of compressor configurations to show the effects upon the performance of decreased loading in the inlet stages. Below 75 percent of design speed, decreased loading results in increased weight flow and peak efficiency; above 80 percent of design speed, decreased loading in the inlet stages results in decreased weight flow and small changes in peak efficiencies. Between 75 and 90 percent of design the changes in surge weight flow and pressure ratio were such that the surge limit line was raised with decreased loading in the inlet stages when presented as pressure ratio against weight flow.
Hatch, James E.↗
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.↗
Development of the seeding system used for laser velocimeter surveys of the NASA Low-Speed Centrifugal Compressor flow field
An atomizer-based system for distributing high-volume rates of seed material was developed to support laser velocimeter investigations of the NASA Low-Speed Centrifugal Compressor flow field. The seeding system and the major concerns that were addressed during its development are described. Of primary importance were that the seed material be dispersed as single particles and that the liquid carrier used be completely evaporated before entering the compressor.
Wasserbauer, Charles A.↗
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.↗
Effect of Various Blade Modifications on Performance of A16-Stage Axial-Flow Compressor. V - Effect on Over-All Performance Characteristics of a 20-Percent Reduction in Solidity of the Fourteenth Through Sixteenth Stage Rotors
The performance of a 16-stage axial-flow compressor, in which the mean-radius solidity was reduced from 1.28 to 1.02 in the fourteenth through sixteenth stage rotors was determined. The performance of this modification was compared with that of the compressor with original rotors. The reduced solidity resulted in slightly improved performance.
Medeiros, Arthur A.↗
Effect of Various Blade Modifications on Performance of a 16-Stage Axial-Flow Compressor. III - Effect on Over-All Performance Characteristics on Increasing Stator-Blade Angles in Inlet Stages
The stator-blade angles in the first four stages of a 16-stage axial-flow compressor were increased in order to decrease the angles of attack of these stages, and thereby to improve part-speed performance. The performance of this modified compressor was compared with that of the same compressor with original blade angles.
Medeiros, Arthur A.↗
Experimental and theoretical distribution of flow produced by inlet guide vanes of an axial-flow compressor
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COMPRESSORS - BLADES - CASCADES↗