Effects of Inlet Distortion on the Flow Field in a Transonic Compressor Rotor

1998 ◽  
Vol 120 (2) ◽  
pp. 233-246 ◽  
Author(s):  
C. Hah ◽  
D. C. Rabe ◽  
T. J. Sullivan ◽  
A. R. Wadia
Keyword(s):  
Boundary Layer ◽  
Flow Field ◽  
Viscous Flow ◽  
Total Pressure ◽  
Stokes Equations ◽  
Three Dimensional ◽  
Aerodynamic Loss ◽  
Transonic Compressor ◽  
Inlet Distortion ◽  
Compressor Rotor

The effects of circumferential distortions in inlet total pressure on the flow field in a low-aspect-ratio, high-speed, high-pressure-ratio, transonic compressor rotor are investigated in this paper. The flow field was studied experimentally and numerically with and without inlet total pressure distortion. Total pressure distortion was created by screens mounted upstream from the rotor inlet. Circumferential distortions of eight periods per revolution were investigated at two different rotor speeds. The unsteady blade surface pressures were measured with miniature pressure transducers mounted in the blade. The flow fields with and without inlet total pressure distortion were analyzed numerically by solving steady and unsteady forms of the Reynolds-averaged Navier–Stokes equations. Steady three-dimensional viscous flow calculations were performed for the flow without inlet distortion while unsteady three-dimensional viscous flow calculations were used for the flow with inlet distortion. For the time-accurate calculation, circumferential and radial variations of the inlet total pressure were used as a time-dependent inflow boundary condition. A second-order implicit scheme was used for the time integration. The experimental measurements and the numerical analysis are highly complementary for this study because of the extreme complexity of the flow field. The current investigation shows that inlet flow distortions travel through the rotor blade passage and are convected into the following stator. At a high rotor speed where the flow is transonic, the passage shock was found to oscillate by as much as 20 percent of the blade chord, and very strong interactions between the unsteady passage shock and the blade boundary layer were observed. This interaction increases the effective blockage of the passage, resulting in an increased aerodynamic loss and a reduced stall margin. The strong interaction between the passage shock and the blade boundary layer increases the peak aerodynamic loss by about one percent.

Effects of Inlet Distortion on the Flow Field in a Transonic Compressor Rotor

1996 ◽  
Author(s):  
Chunill Hah ◽  
Douglas C. Rabe ◽  
Thomas J. Sullivan ◽  
Aspi R. Wadia
Keyword(s):  
Boundary Layer ◽  
Flow Field ◽  
Viscous Flow ◽  
Total Pressure ◽  
Stokes Equations ◽  
Three Dimensional ◽  
Aerodynamic Loss ◽  
Transonic Compressor ◽  
Inlet Distortion ◽  
Compressor Rotor

The effects of circumferential distortions in inlet total pressure on the flow field in a low-aspect-ratio, high-speed, high-pressure-ratio, transonic compressor rotor are investigated in this paper. The flow field was studied experimentally and numerically with and without inlet total pressure distortion. Total pressure distortion was created by screens mounted upstream from the rotor inlet. Circumferential distortions of 8 periods per revolution were investigated at two different rotor speeds. The unsteady blade surface pressures were measured with miniature pressure transducers mounted in the blade. The flow fields with and without inlet total pressure distortion were analyzed numerically by solving steady and unsteady forms of the Reynolds-averaged Navier-Stokes equations. Steady three-dimensional viscous flow calculations were performed for the flow without inlet distortion while unsteady three-dimensional viscous flow calculations were used for the flow with inlet distortion. For the time-accurate calculation, circumferential and radial variations of the inlet total pressure were used as a time-dependent inflow boundary condition. A second-order implicit scheme was used for the time integration. The experimental measurements and the numerical analysis are highly complementary for this study because of the extreme complexity of the flow field. The current investigation shows that inlet flow distortions travel through the rotor blade passage and are convected into the following stator. At a high rotor speed where the flow is transonic, the passage shock was found to oscillate by as much as 20% of the blade chord, and very strong interactions between the unsteady passage shock and the blade boundary layer were observed. This interaction increases the effective blockage of the passage, resulting in an increased aerodynamic loss and a reduced stall margin. The strong interaction between the passage shock and the blade boundary layer increases the peak aerodynamic loss by about one percent.

A Numerical Analysis of the Three-Dimensional Viscous Flow in a Transonic Compressor Rotor and Comparison With Experiment

1987 ◽  
Vol 109 (1) ◽  
pp. 83-90 ◽  
Author(s):  
W. N. Dawes
Keyword(s):  
Boundary Layer ◽  
Numerical Analysis ◽  
Stokes Equations ◽  
Three Dimensional ◽  
Axial Flow ◽  
Tip Clearance ◽  
Navier Stokes ◽  
Viscous Effects ◽  
Transonic Compressor ◽  
Compressor Rotor

The numerical analysis of highly loaded transonic compressors continues to be of considerable interest. Although much progress has been made with inviscid analyses, viscous effects can be very significant, especially those associated with shock–boundary layer interactions. While inviscid analyses have been enhanced by the interactive inclusion of blade surf ace boundary layer calculations, it may be better in the long term to develop efficient algorithms to solve the full three-dimensional Navier–Stokes equations. Indeed, it seems that many phenomena of key interest, like tip clearance flows, may only be accessible to a Navier–Stokes solver. The present paper describes a computer program developed for solving the three-dimensional viscous compressible flow equations in turbomachine geometries. The code is applied to the study of the flowfield in an axial-flow transonic compressor rotor with an attempt to resolve the tip clearance flow. The predicted flow is compared with laser anemometry measurements and good agreement is found.

Flow Field Unsteadiness in the Tip Region of a Transonic Compressor Rotor

1997 ◽  
Vol 119 (1) ◽  
pp. 122-128 ◽  
Author(s):  
S. L. Puterbaugh ◽  
W. W. Copenhaver
Keyword(s):  
Flow Field ◽  
High Performance ◽  
Three Dimensional ◽  
Navier Stokes ◽  
Vortex Interaction ◽  
Tip Leakage Vortex ◽  
Tip Leakage ◽  
Transonic Compressor ◽  
Compressor Rotor ◽  
Operating Points

An experimental investigation concerning tip flow field unsteadiness was performed for a high-performance, state-of-the-art transonic compressor rotor. Casing-mounted high frequency response pressure transducers were used to indicate both the ensemble averaged and time varying flow structure present in the tip region of the rotor at four different operating points at design speed. The ensemble averaged information revealed the shock structure as it evolved from a dual shock system at open throttle to an attached shock at peak efficiency to a detached orientation at near stall. Steady three-dimensional Navier Stokes analysis reveals the dominant flow structures in the tip region in support of the ensemble averaged measurements. A tip leakage vortex is evident at all operating points as regions of low static pressure and appears in the same location as the vortex found in the numerical solution. An unsteadiness parameter was calculated to quantify the unsteadiness in the tip cascade plane. In general, regions of peak unsteadiness appear near shocks and in the area interpreted as the shock-tip leakage vortex interaction. Local peaks of unsteadiness appear in mid-passage downstream of the shock-vortex interaction. Flow field features not evident in the ensemble averaged data are examined via a Navier-Stokes solution obtained at the near stall operating point.

Numerical Investigation of Effect of Inlet Distortion on Compressor Flow Field and Stability

2017 ◽  
Author(s):  
HaoGuang Zhang ◽  
Kang An ◽  
Feng Tan ◽  
YanHui Wu ◽  
WuLi Chu
Keyword(s):  
Numerical Simulation ◽  
Flow Field ◽  
Total Pressure ◽  
Tip Clearance ◽  
Numerical Work ◽  
Stall Margin ◽  
Numerical Investigations ◽  
Inlet Distortion ◽  
Compressor Rotor ◽  
The Stability

The compressor aerodynamic design is conducted under the condition of clean inlet in general, but a compressor often operates under the condition of inlet distortion in the practical application. It has been proven by a lot of experimental and numerical investigations that inlet distortion can decrease the performance and stability of compressors. The circumferential or radial distorted inlet in mostly numerical investigations is made by changing the total pressure and total temperature in the inlet ring surface of the compressors. In most of inlet distortion experiments, distorted inlets are usually created by using wire net, flashboards, barriers or the generator of rotating distortion. The fashion of generating distorted inlet for experiment is different from that for numerical simulation. Consequently, the flow mechanism of affecting the flow field and stability of a compressor with distorted inlet for experiment is partly different than that for numerical simulation. In the numerical work reported here, the inlet distortion is generated by setting some barriers in the inlet ring surface of an axial subsonic compressor rotor. Two kinds of distorted inlet are investigated to exploring the effect of distorted range on the flow field and stability of the compressor with ten-passage unsteady numerical method. The numerical results show that the inlet distortions not only degrade the total pressure and efficiency of the compressor rotor, but also decrease the stability of the rotor. The larger the range of distorted inlet is, the stronger the adverse effect is. The comprehensive stall margin for the inlet distortion of 24 degrees and 48 degrees of ten-passages is reduced about 3.35% and 5.88% respectively. The detailed analysis of the flow field in the compressor indicates that the blockage resulted from tip clearance leakage vortex (TLV) and the flow separation near the suction surfaces of some blades tip for distorted inlet is more serious than that resulted from TLV for clean inlet. Moreover, the larger the range of distorted inlet is, the larger the range of the blockage is. The analysis of unsteady flow shows that during this process, which is that one rotor blade passes through the region affected by the distorted inlet, the range of the blockage in the rotor passage increases first, then reduces, and increases last.

Comparison of Experimental and Computational Shock Structure in a Transonic Compressor Rotor

1981 ◽  
Vol 103 (1) ◽  
pp. 78-88
Author(s):  
G. Haymann-Haber ◽  
W. T. Thompkins
Keyword(s):  
Boundary Layer ◽  
Three Dimensional ◽  
Inviscid Flow ◽  
Boundary Layer Separation ◽  
Shock Strength ◽  
Layer Separation ◽  
Transonic Compressor ◽  
Compressor Rotor ◽  
Pressure Peak ◽  
Rotor Losses

Measurement of passage shock strength in a transonic compressor rotor using a gas fluorescent technique revealed an unexpected variation in shock strength in the radial direction. An axisymmetric idealization would normally predict that the passage shock strength would gradually weaken when moving radially inward until disappearing at the sonic radius. However, the measurements indicated a sharp peak in strength at the nominal sonic radius. Blade boundary layer separation originating at this point accounts for about one half of the total rotor losses. A numerical computation of the three-dimensional inviscid flow, using time-marching techniques, has accurately predicted in general the radial and tangential variations in passage shock strength and in particular the sharp pressure peak at the nominal sonic radius. The overall shock strength was somewhat over-predicted, but this overprediction may be the result of boundary layer separation in the experiment. This paper presents comparisons between the optical density measurements and computational results and in addition a short analytical discussion which demonstrates that the sharp shock strength rise may occur in many transonic compressor rotors.

Flow Field Analysis of Inlet Distortion in a Transonic Fan With Straight and Bowed Stator Blade

2007 ◽  
Author(s):  
Peng Sun ◽  
Jingjun Zhong ◽  
Guotai Feng
Keyword(s):  
Flow Field ◽  
Total Pressure ◽  
Three Dimensional ◽  
Field Analysis ◽  
Navier Stokes ◽  
Fan Performance ◽  
Inlet Distortion ◽  
Flow Loss ◽  
Stator Blade ◽  
Transonic Fan

The performance and stability of a fan in clean and distorted inlet flow can be improved through the use of bowed stator blades. Measurements between the blade rows in transonic and supersonic flow are too complex to provide any useful insights, so 3D flow simulations are required. In this paper, a time-accurate three-dimensional Navier-Stokes solver of the unsteady flow field in a transonic fan is carried out using “Fluent-parallel” in a parallel supercomputer. Two sets of simulations are performed. The first simulation focuses on a better understanding of inlet total pressure distortion effects on a transonic fan. The second set of numerical simulation aims at studying the improvements of fan performance made by bowed stator blades. Three aspects are contained in this paper. The first is about the distortion effects on characteristics of the fan stage with straight stator. The effects of bowed stator on fan performance with inlet distortion are demonstrated secondly. One hand bowed stator increases the loss in rotor. On the other hand, it reduces the flow loss in stator. Finally, the patterns of flow loss caused by total pressure distortion with straight/bowed stator are compared. The scale of vortex in stator induced by inlet total pressure distortion is weakened by bowed blades, which decreases the stator loss.

Quantitative Density Visualization in a Transonic Compressor Rotor

1977 ◽  
Vol 99 (3) ◽  
pp. 460-475 ◽  
Author(s):  
A. H. Epstein
Keyword(s):  
Boundary Layer ◽  
Three Dimensional ◽  
Imaging Systems ◽  
Three Dimensional Flow ◽  
Transonic Compressor ◽  
Dimensional Flow ◽  
Compressor Rotor ◽  
Density Maps ◽  
Shock System ◽  
High Work

The flow in a 59-cm dia high work, transonic compressor rotor has been visualized using a fluorescent gas, 2,3, butanedione, as a tracer. The technique allows the three-dimensional flow to be imaged as a set of distinct planes. Quantitative static density maps were obtained by correcting the images for distortion and nonlinearities introduced by the illumination and imaging systems. These images and maps were used to analyze the three-dimensional nature of the blade’s boundary layer and shock system.

Tip-Clearance and Secondary Flows in a Transonic Compressor Rotor

1999 ◽  
Vol 121 (4) ◽  
pp. 751-762 ◽  
Author(s):  
G. A. Gerolymos ◽  
I. Vallet
Keyword(s):  
Stokes Equations ◽  
Three Dimensional ◽  
Main Flow ◽  
Secondary Flows ◽  
Tip Clearance ◽  
Computational Method ◽  
Tip Vortex ◽  
Computational Results ◽  
Transonic Compressor ◽  
Compressor Rotor

The purpose of this paper is to investigate tip-clearance and secondary flows numerically in a transonic compressor rotor. The computational method used is based on the numerical integration of the Favre-Reynolds-averaged three-dimensional compressible Navier–Stokes equations, using the Launder–Sharma near-wall k–ε turbulence closure. In order to describe the flowfield through the tip and its interaction with the main flow accurately, a fine O-grid is used to discretize the tip-clearance gap. A patched O-grid is used to discretize locally the mixing-layer region created between the jetlike flow through the gap and the main flow. An H–O–H grid is used for the computation of the main flow. In order to substantiate the validity of the results, comparisons with experimental measurements are presented for the NASA_37 rotor near peak efficiency using three grids (of 106, 2 X 106, and 3 X 106 points, with 21, 31, and 41 radial stations within the gap, respectively). The Launder–Sharma k–ε model underestimates the hub corner stall present in this configuration. The computational results are then used to analyze the interblade-passage secondary flows, the flow within the tip-clearance gap, and the mixing downstream of the rotor. The computational results indicate the presence of an important leakage-interaction region where the leakage-vortex after crossing the passage shock-wave mixes with the pressure-side secondary flows. A second trailing-edge tip vortex is also clearly visible.

A Comparison between Measured and Computed Flow Fields in a Transonic Compressor Rotor

1980 ◽  
Vol 102 (4) ◽  
pp. 883-889 ◽  
Author(s):  
P. W. McDonald ◽  
C. R. Bolt ◽  
R. J. Dunker ◽  
H. B. Weyer
Keyword(s):  
Mach Number ◽  
Total Pressure ◽  
Pressure Ratio ◽  
Three Dimensional ◽  
Axial Compressor ◽  
Computational Procedure ◽  
Operating Conditions ◽  
Transonic Compressor ◽  
Compressor Rotor ◽  
Pressure Profiles

The flow field within the rotor of a transonic axial compressor has been computed and compared to measurements obtained with an advanced laser velocimeter. The compressor was designed for a total pressure ratio of 1.51 at a relative tip Mach number of 1.4. The comparisons are made at 100 percent design speed (20,260 RPM) with pressure ratios corresponding to peak efficiency, near surge, and wide open discharge operating conditions. The computational procedure iterates between a blade-to-blade calculation and an intrablade through flow calculation. Calculated Mach number contours, surface pressure distributions, and exit total pressure profiles are in agreement with the experimental data demonstrating the usefulness of quasi three-dimensional calculations in compressor design.

Calculation of a Three-Dimensional Turbomachinery Rotor Flow With a Navier-Stokes Code

1987 ◽  
Author(s):  
Matthew J. Warfield ◽  
B. Lakshminarayana
Keyword(s):  
Boundary Layer ◽  
Stokes Equations ◽  
Three Dimensional ◽  
Streamwise Velocity ◽  
Reynolds Stress Model ◽  
Layer Growth ◽  
Navier Stokes ◽  
Laser Doppler Velocimeter ◽  
Eddy Viscosity Model ◽  
Compressor Rotor

This paper deals with a numerical solution of the full Navier-Stokes equations governing the flowfield in a turbomachinery rotor. The incompressible equations are solved using a pseudocompressibility time marching code. A two-equation turbulence model (k-ε) coupled with a vectorial eddy viscosity model based on an Algebraic Reynolds Stress Model is used to account for the anisotropic effects of rotation and three dimensionality. The predictions are compared with laser doppler velocimeter and hot wire data acquired in a compressor rotor passage at two different flow coefficients. The predicted blade to blade profiles of velocity at various radial locations as well as the streamwise velocity profiles in the blade boundary layer show good agreement with experimental data. The radial velocities are qualitatively predicted but good comparison with data was not achieved. Boundary layer growth is predicted reasonably well.

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