Improved Blade Profile Loss and Deviation Angle Models for Advanced Transonic Compressor Bladings: Part II—A Model for Supersonic Flow

1996 ◽  
Vol 118 (1) ◽  
pp. 81-87 ◽  
Author(s):  
W. M. Ko¨nig ◽  
D. K. Hennecke ◽  
L. Fottner
Keyword(s):  
Present Report ◽  
Axial Flow ◽  
Deviation Angle ◽  
New Model ◽  
Transonic Compressor ◽  
Supersonic Inlet ◽  
Blade Row ◽  
Flow Situation ◽  
Inlet Conditions ◽  
Profile Loss

New blading concepts as used in modern transonic axial-flow compressors require improved loss and deviation angle correlations. The new model presented in this paper incorporates several elements and treats blade-row flows having subsonic and supersonic inlet conditions separately. The second part of the present report focuses on the extension of a well-known correlation for cascade losses at supersonic inlet flows. It was originally established for DCA bladings and is now modified to reflect the flow situation in blade rows having low-cambered, arbitrarily designed blades including precompression blades. Finally, the steady loss increase from subsonic to supersonic inlet-flow velocities demonstrates the matched performance of the different correlations of the new model.

Improved Blade Profile Loss and Deviation Angle Models for Advanced Transonic Compressor Bladings: Part II — A Model for Supersonic Flow

1994 ◽  
Author(s):  
W. M. König ◽  
D. K. Hennecke ◽  
L. Fottner
Keyword(s):  
Present Report ◽  
Axial Flow ◽  
Deviation Angle ◽  
New Model ◽  
Transonic Compressor ◽  
Supersonic Inlet ◽  
Blade Row ◽  
Flow Situation ◽  
Inlet Conditions ◽  
Profile Loss

New blading concepts as used in modern transonic axial-flow compressors require improved loss and deviation angle correlations. The new model presented in this paper incorporates several elements and treats separately blade-row flows having subsonic and supersonic inlet conditions. The second part of the present report focuses on the extension of a well-known correlation for cascade losses at supersonic inlet-flows. It was originally established for DCA-bladings and is now modified to reflect the flow situation in blade-rows having low-cambered, arbitrarily designed blades including precompression blades. Finally, the steady loss increase from subsonic to supersonic inlet-flow velocities demonstrates the matched performance of the different correlations of the new model.

Improved Blade Profile Loss and Deviation Angle Models for Advanced Transonic Compressor Bladings: Part I—A Model for Subsonic Flow

1996 ◽  
Vol 118 (1) ◽  
pp. 73-80 ◽  
Author(s):  
W. M. Ko¨nig ◽  
D. K. Hennecke ◽  
L. Fottner
Keyword(s):  
Axial Flow ◽  
Deviation Angle ◽  
Simple Method ◽  
Subsonic Flows ◽  
New Model ◽  
Transonic Compressor ◽  
Blade Row ◽  
Inlet Conditions ◽  
Improved Accuracy ◽  
Profile Loss

New blading concepts as used in modern transonic axial-flow compressors require improved loss and deviation angle correlations. The new model presented in this paper incorporates several elements and treats blade-row flows having subsonic and supersonic inlet conditions separately. In the first part of this paper two proved and well-established profile loss correlations for subsonic flows are extended to quasi-two-dimensional conditions and to custom-tailored blade designs. Instead of a deviation angle correlation, a simple method based on singularities is utilized. The comparison between the new model and a recently published model demonstrates the improved accuracy in prediction of cascade performance achieved by the new model.

Improved Blade Profile Loss and Deviation Angle Models for Advanced Transonic Compressor Bladings: Part I — A Model for Subsonic Flow

1994 ◽  
Author(s):  
W. M. König ◽  
D. K. Hennecke ◽  
L. Fottner
Keyword(s):  
Axial Flow ◽  
Deviation Angle ◽  
Simple Method ◽  
Subsonic Flows ◽  
New Model ◽  
Transonic Compressor ◽  
Blade Row ◽  
Inlet Conditions ◽  
Improved Accuracy ◽  
Profile Loss

New blading concepts as used in modern transonic axial-flow compressors require improved loss and deviation angle correlations. The new model presented in this paper incorporates several elements and treats separately blade-row flows having subsonic and supersonic inlet conditions. In the first part of this paper two proved and well-established profile loss correlations for subsonic flows are extended to quasi twodimensional conditions and to custom-tailored blade designs. Instead of a deviation angle correlation a simple method based on singularities is utilized. The comparison between the new model and a recently published model demonstrates the improved accuracy in prediction of cascade performance achieved by the new model.

Stator-Rotor Interactions in a Transonic Compressor— Part 1: Effect of Blade-Row Spacing on Performance

2003 ◽  
Vol 125 (2) ◽  
pp. 328-335 ◽  
Author(s):  
Steven E. Gorrell ◽  
Theodore H. Okiishi ◽  
William W. Copenhaver
Keyword(s):  
Mass Flow Rate ◽  
Mass Flow ◽  
Flow Rate ◽  
Pressure Ratio ◽  
Axial Flow ◽  
Transonic Compressor ◽  
Blade Row ◽  
Spacing Variation ◽  
Stator Blade ◽  
Flow Compressor

Usually less axial spacing between the blade rows of an axial flow compressor is associated with improved efficiency. However, mass flow rate, pressure ratio, and efficiency all decreased as the axial spacing between the stator and rotor was reduced in a transonic compressor rig. Reductions as great as 3.3% in pressure ratio, and 1.3 points of efficiency were observed as axial spacing between the blade rows was decreased from far apart to close together. The number of blades in the stator blade-row also affected stage performance. Higher stator blade-row solidity led to larger changes in pressure ratio efficiency, and mass flow rate with axial spacing variation. Analysis of the experimental data suggests that the drop in performance is a result of increased loss production due to blade-row interactions. Losses in addition to mixing loss are present when the blade-rows are spaced closer together. The extra losses are associated with the upstream stator wakes and are most significant in the midspan region of the flow.

Experimental Cascade Analysis of a Transonic Compressor Rotor Blade Section

1984 ◽  
Vol 106 (2) ◽  
pp. 288-294 ◽  
Author(s):  
H. A. Schreiber ◽  
H. Starken
Keyword(s):  
Rotor Blade ◽  
Flow Behavior ◽  
Density Ratio ◽  
Axial Flow ◽  
Flow Angle ◽  
Transonic Compressor ◽  
Compressor Rotor ◽  
Blade Row ◽  
Blade Section ◽  
Transonic Cascade

A transonic compressor rotor blade cascade was tested in order to elucidate the flow behavior in the transonic regime and to determine the performance characteristic in the whole operating range of a rotor blade section. The experiments have been conducted in a transonic cascade wind tunnel, which enables tests even at sonic inlet velocities. The influence of the upstream Mach number between 0.8 and 1.1 and the inlet flow angle between choking and stalling of the blade row was investigated. The effect of the axial velocity density ratio (AVDR) could be studied by applying an endwall suction device. Furthermore, the level of the shock losses was determined from a wake analysis. A final comparison of cascade losses and those of the corresponding rotor blade element shows good agreement which underlines the applicability of the cascade model in the design of axial flow turbomachines.

Stator-Rotor Interactions in a Transonic Compressor: Part 1 — Effect of Blade-Row Spacing on Performance

2002 ◽  
Author(s):  
Steven E. Gorrell ◽  
Theodore H. Okiishi ◽  
William W. Copenhaver
Keyword(s):  
Mass Flow Rate ◽  
Mass Flow ◽  
Flow Rate ◽  
Pressure Ratio ◽  
Axial Flow ◽  
Transonic Compressor ◽  
Blade Row ◽  
Spacing Variation ◽  
Stator Blade ◽  
Flow Compressor

Usually less axial spacing between the blade rows of an axial flow compressor is associated with improved efficiency. However, mass flow rate, pressure ratio, and efficiency all decreased as the axial spacing between the stator and rotor was reduced in a transonic compressor rig. Reductions as great as 3.3% in pressure ratio and 1.3 points of efficiency were observed as axial spacing between the blade-rows was decreased from far apart to close together. The number of blades in the stator blade-row also affected stage performance. Higher stator blade-row solidity led to larger changes in pressure ratio, efficiency, and mass flow rate with axial spacing variation. Analysis of the experimental data suggests that the drop in performance is a result of increased loss production due to blade-row interactions. Losses in addition to mixing loss are present when the blade-rows are spaced closer together. The extra losses are associated with the upstream stator wakes and are most significant in the mid-span region of the flow.

Experimental Evaluation of an Axial-Flow Transonic Compressor

1976 ◽  
Author(s):  
T. Tamaki ◽  
S. Nagano
Keyword(s):  
Rotational Speed ◽  
Velocity Ratio ◽  
Axial Flow ◽  
Rotating Stall ◽  
Test Results ◽  
Transonic Compressor ◽  
Multi Stage ◽  
Blade Row ◽  
Compressor Surge ◽  
Overall Performance

A five-stage transonic compressor was designed and tested to obtain overall performance and surge limits over a wide rotational speed range. Compressor surge limits are evaluated using individual stage performance characteristics with the aid of a single-stage test. These show that surging at low rotational speed occurs when rotating stall occurs in the first stage and that the wider operating range multi-stage compressor can be realized with the lower aspect ratio blades. On the basis of these test results, a simple model is constructed in order to evaluate the effect of the axial velocity ratio on the flow in a rotor blade row.

Experimental Cascade Analysis of a Transonic Compressor Rotor Blade Section

1983 ◽  
Author(s):  
H. A. Schreiber ◽  
H. Starken
Keyword(s):  
Rotor Blade ◽  
Density Ratio ◽  
Axial Flow ◽  
Flow Angle ◽  
Transonic Compressor ◽  
Compressor Rotor ◽  
Blade Row ◽  
Blade Section ◽  
Transonic Cascade ◽  
Good Agreement

A transonic compressor rotor blade cascade was tested in order to elucidate the flow behaviour in the transonic regime and to determine the performance characteristic in the whole operating range of a rotor blade section. The experiments have been conducted in a transonic cascade wind tunnel, which enables tests even at sonic inlet velocities. The influence of the upstream Mach number between 0.8 and 1.1 and the inlet flow angle between choking and stalling of the blade row was investigated. The effect of the axial velocity density ratio (AVDR) could be studied by applying an endwall suction device. Furthermore the level of the shock losses was determined from a wake analysis. A final comparison of cascade losses and those of the corresponding rotor blade element shows good agreement which underlines the applicability of the cascade model in the design of axial flow turbomachines.

scholarly journals Effects of Tip Leakage Flow on the Aerodynamic Performance and Stability of an Axial-Flow Transonic Compressor Stage

Energies ◽  
2021 ◽  
Vol 14 (14) ◽  
pp. 4168
Author(s):  
Botao Zhang ◽  
Xiaochen Mao ◽  
Xiaoxiong Wu ◽  
Bo Liu
Keyword(s):  
Boundary Layer ◽  
Shock Wave ◽  
Axial Flow ◽  
Leading Edge ◽  
Rotating Stall ◽  
Tip Clearance ◽  
Leakage Flow ◽  
Tip Leakage Flow ◽  
Tip Leakage ◽  
Transonic Compressor

To explain the effect of tip leakage flow on the performance of an axial-flow transonic compressor, the compressors with different rotor tip clearances were studied numerically. The results show that as the rotor tip clearance increases, the leakage flow intensity is increased, the shock wave position is moved backward, and the interaction between the tip leakage vortex and shock wave is intensified, while that between the boundary layer and shock wave is weakened. Most of all, the stall mechanisms of the compressors with varying rotor tip clearances are different. The clearance leakage flow is the main cause of the rotating stall under large rotor tip clearance. However, the stall form for the compressor with half of the designed tip clearance is caused by the joint action of the rotor tip stall caused by the leakage flow spillage at the blade leading edge and the whole blade span stall caused by the separation of the boundary layer of the rotor and the stator passage. Within the investigated varied range, when the rotor tip clearance size is half of the design, the compressor performance is improved best, and the peak efficiency and stall margin are increased by 0.2% and 3.5%, respectively.

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