Propagation of Inlet Flow Distortions through an Axial Compressor Stage

1979 ◽  
Vol 101 (1) ◽  
pp. 116-124 ◽  
Author(s):  
J. Colpin
Keyword(s):  
Total Pressure ◽  
Axial Compressor ◽  
Two Dimensional ◽  
Compressor Stage ◽  
Flow Distortion ◽  
Inlet Flow ◽  
Flow Equations ◽  
Finite Difference Technique ◽  
Original Calculation ◽  
Theory And Experiments

This contribution will present an original calculation method predicting the development of an inlet flow distortion through a compressor stage. A finite difference technique is used to treat the flow equations outside the blade rows. That flow is two-dimensional, compressible and nonviscous. The blade rows are modelized using a quasi actuator disk approach, but include the unsteady transfer terms due to the rotor relative motion in a non uniform inlet flow. A set of experimental data, measured on a one stage axial compressor, submitted to a rectangular inlet total pressure distortion will be discussed and will serve as basis for a comparison between theory and experiments.

Studies on Downstream Stator With Rotor Re-Staggering and Forward Sweeping in a Subsonic Axial Compressor Stage

2011 ◽  
Author(s):  
P. V. Ramakrishna ◽  
M. Govardhan
Keyword(s):  
Flow Field ◽  
Computational Model ◽  
Flow Rate ◽  
Total Pressure ◽  
Axial Compressor ◽  
Pressure Rise ◽  
Axial Distance ◽  
Compressor Stage ◽  
Numerical Work ◽  
Stagger Angle

The present numerical work studies the flow field in subsonic axial compressor stator passages for: (a) preceding rotor sweep (b) preceding rotor re-staggering (three stagger angle changes: 0°, +3° and +5°); and (c) stator sweeping (two 20° forward sweep schemes). The following are the motives for the study: at the off-design conditions, compressor rotors are re-staggered to alleviate the stage mismatching by adjusting the rows to the operating flow incidence. Fundamental to this is the understanding of the effects of rotor re-staggering on the downstream component. Secondly, sweeping the rotor stages alters the axial distance between the successive rotor-stator stages and necessitates that the stator vanes must also be swept. To the best of the author’s knowledge, stator sweeping to suit such scenarios has not been reported. The computational model for the study utilizes well resolved hexahedral grids. A commercial CFD package ANSYS® CFX 11.0 was used with standard k-ω turbulence model for the simulations. CFD results were well validated with experiments. The following observations were made: (1) When the rotor passage is closed by re-staggering, with the same mass flow rate and the same stator passage area, stators were subjected to negative incidences. (2) Effect of stator sweeping on the upstream rotor flow field is insignificant. Comparison of total pressure rise carried by the downstream stators suggests that an appropriate redesign of stator is essential to match with the swept rotors. (3) While sweeping the stator is not recommended, axial sweeping is preferable over true sweeping when it is necessary.

scholarly journals Investigation of Flows in Rectangular Diffusers With Inlet Flow Distortion

1982 ◽  
Author(s):  
M. M. Al-Mudhafar ◽  
M. Ilyas ◽  
F. S. Bhinder
Keyword(s):  
Experimental Study ◽  
Velocity Profiles ◽  
Pressure Recovery ◽  
Two Dimensional ◽  
Flow Distortion ◽  
Inlet Flow ◽  
Inlet Velocity ◽  
Inlet Distortion ◽  
Mach Numbers ◽  
Inlet Flow Distortion

The results of an experimental study on the influence of severely distorted velocity profiles on the performance of a straight two-dimensional diffuser are reported. The data cover entry Mach numbers ranging from 0.1 to 0.6 and several inlet distortion levels. The pressure recovery progressively deteriorates as the inlet velocity is distorted.

Effect of Tip Gap Variation on the Performance of the Transonic Fan Stage With Tandem Stator

2019 ◽  
Author(s):  
K. Ananthakrishnan ◽  
Shyama Prasad Das ◽  
B. V. S. S. S. Prasad
Keyword(s):  
Pressure Loss ◽  
Total Pressure ◽  
Axial Compressor ◽  
Compressor Stage ◽  
Tip Leakage ◽  
Pressure Loss Coefficient ◽  
Overall Performance ◽  
Transonic Fan ◽  
Total Pressure Loss Coefficient ◽  
Transonic Rotor

Abstract The main goal of modern axial compressor development is to increase the power to weight ratio with higher efficiency. In the present investigation, highly loaded single stage axial compressor with tandem stator vanes is used. Tandem vanes help in attaining the compact compressor stage along with high pressure loading. It is designed for a stage pressure ratio of 2, mass flow rate of 9.02 kg/s operating at 30800 rpm resulting in transonic flow field. The aerodynamic performance of this compressor detoriates due to the tip leakage and secondary flows. Steady-state numerical investigation is carried out to study the flow structures near the tip region of transonic rotor and how different tip gaps influence the overall performance of the compressor. Further the effects of tip leakage flow variation on the performance of tandem vanes are also highlighted. Transonic fan stage with baseline tip gap of 0.5mm is analyzed along with different tip clearance values ranging from 0 % to 3 % of axial chord. Three-dimensional viscous Reynolds Averaged Navier Stokes (RANS) equations are solved using SST k-ω turbulence model. Computational domain discretized with high quality hexahedral elements (Y+ < 2) in AUTOGRID, Numeca. The numerical procedure is verified against the experimental results of Rotor37 transonic rotor test case. Tip leakage losses contribute a substantial amount to the total loss of stage. Overall performance and the stall characteristics for the compressor stage has been evaluated for different tip gap variations.. Further, the topological properties are exploited to visualize the critical points and separation lines on rotor and tandem vanes. Increase in rotor total pressure loss coefficient is observed with increasing tip gap. In contrary, overall total pressure loss coefficient improves for smaller tip gap values and then detoriates. It is observed optimum tip gap height lies close to the 1.125mm, 2% of baseline design value.

Stall Inception in Axial Compressors

1990 ◽  
Vol 112 (1) ◽  
pp. 116-123 ◽  
Author(s):  
N. M. McDougall ◽  
N. A. Cumpsty ◽  
T. P. Hynes
Keyword(s):  
Boundary Layer ◽  
Axial Compressor ◽  
Tip Clearance ◽  
Compressor Stage ◽  
Inlet Flow ◽  
Stall Inception ◽  
Axial Compressors ◽  
Multistage Compressor

Detailed measurements have been made of the transient stalling process in an axial compressor stage. The stage is of high hub-casing ratio and stall is initiated in the rotor. If the rotor tip clearance is small stall inception occurs at the hub, but at clearances typical for a multistage compressor the inception is at the tip. The crucial quantity in both cases is the blockage caused by the endwall boundary layer. Prior to stall, disturbances rotate around the inlet flow in sympathy with rotating variations in the endwall blockage; these can persist for some time prior to stall, rising and falling in amplitude before the final increase, which occurs as the compressor stalls.

Stage Performance and Radial Matching of Axial Compressor Blade Rows

1959 ◽  
Vol 81 (1) ◽  
pp. 13-23
Author(s):  
Jeffrey Watkins
Keyword(s):  
Axial Compressor ◽  
Compressor Blade ◽  
Secondary Flows ◽  
Two Dimensional ◽  
Compressor Stage ◽  
Stage Performance ◽  
Design Techniques

To obtain the ultimate in performance from an axial compressor stage it is necessary that the operation of the blade rows comprising the stage be thoroughly understood. Appreciable advances have been accomplished in recent years, but much still remains to be done. This paper aims to present practical hypotheses concerning the nature of the flow between the blades and deals with design techniques in relation to this flow. In addition, some elements of two-dimensional cascade data and a discourse on practical secondary flows are presented.

Aerodynamic Design and Analysis of a High Pressure Ratio Aspirated Compressor Stage

2000 ◽  
Author(s):  
Ali A. Merchant ◽  
Mark Drela ◽  
Jack L. Kerrebrock ◽  
John J. Adamczyk ◽  
Mark Celestina
Keyword(s):  
Boundary Layer ◽  
Total Pressure ◽  
Pressure Ratio ◽  
Three Dimensional ◽  
Axial Compressor ◽  
Adverse Pressure Gradient ◽  
Compressor Stage ◽  
High Pressure Ratio ◽  
Total Pressure Ratio ◽  
Isentropic Efficiency

The pressure ratio of axial compressor stages can be significantly increased by controlling the development of blade and endwall boundary layers in regions of adverse pressure gradient by means of boundary layer suction. This concept is validated and demonstrated through the design and analysis of a unique aspirated compressor stage which achieves a total pressure ratio of 3.5 at a tip speed of 1500 ft/s. The aspirated stage was designed using an axisymmetric through-flow code coupled with a quasi three-dimensional cascade plane code with inverse design capability. Validation of the completed design was carried out with three-dimensional Navier-Stokes calculations. Spanwise slots were used on the rotor and stator suction surfaces to bleed the boundary layer with a total suction requirement of 4% of the inlet mass flow. Additional bleed of 3% was also required on the hub and shroud near shock impingement locations. A three-dimensional viscous evaluation of the design showed good agreement with the quasi three-dimensional design intent, except in the endwall regions. The three-dimensional viscous analysis predicted a mass averaged total pressure ratio of 3.7 at an isentropic efficiency of 93% for the rotor, and a mass averaged total pressure ratio of 3.4 at an isentropic efficiency of 86% for the stage.

Compressor-Diffuser Interaction with Circumferential Flow Distortion

1976 ◽  
Vol 18 (1) ◽  
pp. 25-38 ◽  
Author(s):  
E. M. Greitzer ◽  
H. R. Griswol
Keyword(s):  
Flow Field ◽  
Total Pressure ◽  
Static Pressure ◽  
Axisymmetric Flow ◽  
Axial Compressor ◽  
Effective Area ◽  
Flow Distortion ◽  
Diffuser Flow ◽  
Theoretical Predictions ◽  
Annular Diffuser

An analytical and experimental study of axial compressor-diffuser interaction in circumferentially non-uniform flow is reported. An analysis of non-axisymmetric flow in an annular diffuser is presented, based on an inviscid rotational core flow plus the use of a diffuser effective area ratio to account for boundary layer blockage. The analysis is applied to the prediction of the diffuser flow field associated with the presence of a circumferential total pressure distortion. It is found that large static pressure non-uniformities can exist at the inlet of diffusers that are short compared with their mean circumferences, as is usually the case in turbomachinery applications. The analysis is coupled to an asymmetric compressor flow field prediction to provide a method for calculating the effect of an exit diffuser on compressor performance with distortion. It is shown that the velocity defect seen by the compressor can be substantially increased by the presence of the diffuser. The experiments were directed at assessing the method used to predict the flow in the diffuser. Measurements were carried out of the inlet static pressure distortion associated with a circumferentially non-uniform total pressure distribution. The results are found to be in good agreement with the theoretical predictions.

Design and Numerical Investigation of Advanced Radial Inlet for a Centrifugal Compressor Stage

2004 ◽  
Author(s):  
Yunbae Kim ◽  
Jay Koch
Keyword(s):  
Centrifugal Compressor ◽  
Pressure Loss ◽  
Total Pressure ◽  
Aerodynamic Performance ◽  
Total Pressure Loss ◽  
Compressor Stage ◽  
Flow Distortion ◽  
Centrifugal Compressor Stage ◽  
Initial Design ◽  
Wide Range

The performance of a centrifugal compressor stage can be seriously affected by inlet flow distortions due to an unsatisfactory inlet configuration and the resulting flow structure. In this study, two radial inlets were designed for a centrifugal compressor stage and investigated numerically using a commercially available 3D viscous Navier-Stokes code. The intent of the design was to minimize the total pressure loss across the inlet while distributing the flow as equally and uniformly as possible to the impeller inlet. For each inlet model, the aerodynamic performance was calculated from the simulation results and then the results from both models were evaluated and compared. The second radial inlet design outperformed the initial design in terms of total pressure loss, flow distortion and uniformity at the impeller inlet. Furthermore, the aerodynamic performance of the second radial inlet was insensitive to a wide range of mass flow rates compared to the initial design due to the distinctive geometric features implemented for the second inlet design.

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