Reynolds Number Effects in Cascades and Axial Flow Compressors

1964 ◽  
Vol 86 (3) ◽  
pp. 236-242
Author(s):  
J. H. Horlock ◽  
R. Shaw ◽  
D. Pollard ◽  
A. Lewkowicz
Keyword(s):  
Reynolds Number ◽  
Aspect Ratio ◽  
Guide Vane ◽  
Axial Flow ◽  
Compressor Cascade ◽  
Compressor Blades ◽  
Laminar Separation ◽  
Reynolds Number Effects ◽  
Number Variation ◽  
Flow Compressor

A series of tests on guide vane and compressor cascades is reported. The Reynolds number was varied in the guide vane cascade tests, and the Reynolds number and the cascade aspect ratio were varied in the compressor cascade tests. The substantial laminar separation observed in the compressor cascades at high aspect ratio (and low Reynolds number) was suppressed in the cascade tests at low aspect ratio, 2:1. Effects of Reynolds number variation on the performance of a single stage axial flow compressor are also given, and compared with predictions of performance using the cascade tests. Calculations of laminar separation points agree quite well with the experimental observations. It appears that transition due to laminar boundary layer instability is unlikely to occur on compressor blades, in the normal operating range of Reynolds number.

An Experiment With Aspect Ratio as a Means of Extending the Useful Range of a Transonic Inlet Stage of an Axial Flow Compressor

1964 ◽  
Vol 86 (3) ◽  
pp. 243-246 ◽  
Author(s):  
W. C. Swan
Keyword(s):  
Reynolds Number ◽  
Aspect Ratio ◽  
Axial Flow ◽  
Compressor Stage ◽  
Axial Flow Compressor ◽  
Flow Compressor

An experiment in unstalled range of a transonic axial flow compressor stage is discussed. Two rotors of identical design, but of differing aspect ratio, are compared. The study suggests that some criteria other than blade chord Reynolds number be used to define chord lengths of transonic stages.

Boundary Layer Optimization for the Design of High Turning Axial Flow Compressor Blades

1971 ◽  
Vol 93 (1) ◽  
pp. 147-154 ◽  
Author(s):  
K. D. Papailiou
Keyword(s):  
Boundary Layer ◽  
Axial Flow ◽  
Optimization Method ◽  
Mapping Method ◽  
Compressor Cascade ◽  
Compressor Blades ◽  
Conformal Mapping Method ◽  
Theoretical Predictions ◽  
Flow Convergence ◽  
Flow Compressor

An optimization method, based on Le Foll’s boundary layer theory and on Goldstein’s conformal mapping method is described. The parameters of optimization are circulation per blade and absolute losses. The problem is treated as an inverse problem (i.e., the best blading is found starting from the flow conditions imposed). The performance of a highly loaded compressor cascade, designed according to the method presented and tested in a low speed wind tunnel, is compared with the theoretical predictions. Some discrepancies exist which are due to the influence of flow convergence and blade curvature. A modification of the method to take into account these effects is discussed.

Development of a New Front Stage for an Industrial Axial Flow Compressor

1994 ◽  
Vol 116 (4) ◽  
pp. 597-604 ◽  
Author(s):  
B. Eisenberg
Keyword(s):  
Design Method ◽  
Guide Vane ◽  
Theoretical Calculations ◽  
Axial Flow ◽  
Operating Range ◽  
Performance Curves ◽  
Front Stage ◽  
Inverse Design Method ◽  
Flow Compressor ◽  
High Degree

Industrial axial flow compressors are specially designed to achieve a wide operating range. The analysis of an existing six-stage axial flow research compressor indicated that the front stage could be improved significantly using modern design technique. To demonstrate the advantages of such a technique a redesign of the current front stage was conducted. By controlling the diffusion inside the blade sections with an inverse design method, loading was enlarged. Higher loading normally results in a reduction of profile incidence range. For compensation a wide chord application was chosen. Compared to the original compressor version, experiments resulted in steeper characteristic curves together with larger usable operating range. Keeping the same outer and inner diameter, mass flow was increased by 6 percent. Measurements of performance curves with variable speed and for guide vane control are presented. Theoretical calculations achieve a high degree of agreement with measured performance.

Performance Estimation of Axial Flow Turbines

1970 ◽  
Vol 185 (1) ◽  
pp. 407-424 ◽  
Author(s):  
H. R. M. Craig ◽  
H. J. A. Cox
Keyword(s):  
Reynolds Number ◽  
Aspect Ratio ◽  
Gas Turbines ◽  
Three Dimensional ◽  
Axial Flow ◽  
Performance Estimation ◽  
Speed Ratio ◽  
Test Results ◽  
Wide Range ◽  
Relevant Variables

A comprehensive method of estimating the performance of axial flow steam and gas turbines is presented, based on analysis of linear cascade tests on blading, on a number of turbine test results, and on air tests of model casings. The validity of the use of such data is briefly considered. Data are presented to allow performance estimation of actual machines over a wide range of Reynolds number, Mach number, aspect ratio and other relevant variables. The use of the method in connection with three-dimensional methods of flow estimation is considered, and data presented showing encouraging agreement between estimates and available test results. Finally ‘carpets’ are presented showing the trends in efficiencies that are attainable in turbines designed over a wide range of loading, axial velocity/blade speed ratio, Reynolds number and aspect ratio.

Development of a New Front Stage for an Industrial Axial Flow Compressor

1993 ◽  
Author(s):  
B. Eisenberg
Keyword(s):  
Design Method ◽  
Guide Vane ◽  
Theoretical Calculations ◽  
Axial Flow ◽  
Operating Range ◽  
Performance Curves ◽  
Front Stage ◽  
Inverse Design Method ◽  
Flow Compressor ◽  
High Degree

Industrial axial flow compressors are specially designed to achieve a wide operating range. The analysis of an existing 6 stage axial flow research compressor indicated that the front stage could be improved significantly using modern design technique. To demonstrate the advantages of such a technique a redesign of the current front stage was conducted. By controlling the diffusion inside the blade sections with an inverse design method, loading was enlarged. Higher loading normally results in a reduction of profile incidence range. For compensation a wide chord application was chosen. Compared to the original compressor version, experiments resulted in steeper characteristic curves together with larger usable operating range. Keeping the same outer and inner diameter, mass flow was increased by 6%. Measurements of performance curves with variable speed and for guide vane control are presented. Theoretical calculations achieve a high degree of agreement with measured performance.

Wall Boundary Layer Development Near the Tip Region of an IGV of an Axial Flow Compressor

1984 ◽  
Vol 106 (2) ◽  
pp. 337-345
Author(s):  
B. Lakshminarayana ◽  
N. Sitaram
Keyword(s):  
Boundary Layer ◽  
Guide Vane ◽  
Three Dimensional ◽  
Axial Compressor ◽  
Axial Flow ◽  
Inlet Guide Vane ◽  
Wall Boundary Layer ◽  
Wall Boundary ◽  
Boundary Layer Development ◽  
Flow Compressor

The annulus wall boundary layer inside the blade passage of the inlet guide vane (IGV) passage of a low-speed axial compressor stage was measured with a miniature five-hole probe. The three-dimensional velocity and pressure fields were measured at various axial and tangential locations. Limiting streamline angles and static pressures were also measured on the casing of the IGV passage. Strong secondary vorticity was developed. The data were analyzed and correlated with the existing velocity profile correlations. The end wall losses were also derived from these data.

Experimental and Analytical Investigation of the Effects of Reynolds Number and Blade Surface Roughness on Multistage Axial Flow Compressors

1980 ◽  
Vol 102 (1) ◽  
pp. 5-12 ◽  
Author(s):  
A. Scha¨ffler
Keyword(s):  
Experimental Data ◽  
Boundary Layer ◽  
Reynolds Number ◽  
Critical Reynolds Number ◽  
Pressure Ratio ◽  
Axial Flow ◽  
Blade Surface ◽  
Laminar Separation ◽  
High Pressure Ratio ◽  
Attached Flow

The general effect of Reynolds Number on axial flow compressors operating over a sufficiently wide range is described and illustrated by experimental data for four multistage axial compressors. The wide operating range of military aircraft engines leads in the back stages of high pressure ratio compression systems to three distinctly different regimes of operation, characterized by the boundary layer conditions of the cascade flow: • laminar separation, • turbulent attached flow with hydraulically smooth blade surface, • turbulent attached flow with hydraulically rough blade surface. Two “critical” Reynolds Numbers are defined, the “lower critical Reynolds Number” below which laminar separation occurs with a definite steepening of the efficiency/Reynolds Number relation and an “upper critical Reynolds Number” above which the blade surface behaves hydraulically rough, resulting in an efficiency independant of Reynolds Number. The permissible blade surface roughness for hydraulically smooth boundary layer conditions in modern high pressure ratio compression systems is derived from experimental data achieved with blades produced by grinding, electrochemical machining and forging. A correlation between the effect of technical roughness and sand type roughness is given. The potential loss of efficiency in the back end of compression systems due to excessive blade roughness is derived from experimental results. The repeatedly experienced different sensitivity of front and back stages towards laminar separation in the low Reynolds Number regime is explained by boundary layer calculations as a Mach Number effect on blade pressure distribution, i.e. transonic versus subsonic flow.

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