Three-Dimensional Boundary Layer on a Compressor Rotor Blade at Peak Pressure Rise Coefficient

1987 ◽  
Vol 109 (1) ◽  
pp. 91-98 ◽  
Author(s):  
B. Lakshminarayana ◽  
P. Popovski
Keyword(s):  
Boundary Layer ◽  
Rotor Blade ◽  
Peak Pressure ◽  
Pressure Coefficient ◽  
Three Dimensional ◽  
Pressure Rise ◽  
Layer Growth ◽  
Leakage Flow ◽  
Suction Surface ◽  
Compressor Rotor

A comprehensive study of the three-dimensional turbulent boundary layer on a compressor rotor blade at peak pressure rise coefficient is reported in this paper. The measurements were carried out at various chordwise and radial locations on a compressor rotor blade using a rotating miniature V configuration hot-wire probe. The data are compared with the measurement at the design condition. Substantial changes in the blade boundary layer characteristics are observed, especially in the outer 16 percent of the blade span. The increased chordwise pressure gradient and the leakage flow at the peak pressure coefficient have a cumulative effect in increasing the boundary layer growth on the suction surface. The leakage flow has a beneficial effect on the pressure surface. The momentum and boundary layer thicknesses increase substantially from those at the design condition, especially near the outer radii of the suction surface.

scholarly journals Three-Dimensional Boundary Layer on a Compressor Rotor Blade at Peak Pressure Rise Coefficient

1986 ◽  
Author(s):  
B. Lakshminarayana ◽  
P. Popovski
Keyword(s):  
Boundary Layer ◽  
Rotor Blade ◽  
Peak Pressure ◽  
Pressure Coefficient ◽  
Three Dimensional ◽  
Pressure Rise ◽  
Layer Growth ◽  
Leakage Flow ◽  
Suction Surface ◽  
Compressor Rotor

A comprehensive study of the three-dimensional turbulent boundary layer on a compressor rotor blade at peak pressure rise coefficient is reported in this paper. The measurements were carried out at various chordwise and radial locations on a compressor rotor blade using a rotating miniature “V” configuration hot-wire probe. The data are compared with the measurement at the design condition. Substantial changes in the blade boundary layer characteristics are observed, especially in the outer sixteen percent of the blade span. The increased chordwise pressure gradient and the leakage flow at the peak pressure coefficient have a cumulative effect in increasing the boundary layer growth on the suction surface. The leakage flow has a beneficial effect on the pressure surface. The momentum and boundary layer thicknesses increase substantially from those at the design condition, especially near the outer radii of the suction surface.

Three-Dimensional Flow Field in the Tip Region of a Compressor Rotor Passage—Part I: Mean Velocity Profiles and Annulus Wall Boundary Layer

1982 ◽  
Vol 104 (4) ◽  
pp. 760-771 ◽  
Author(s):  
B. Lakshminarayana ◽  
M. Pouagare ◽  
R. Davino
Keyword(s):  
Boundary Layer ◽  
Flow Field ◽  
Three Dimensional ◽  
Leading Edge ◽  
Leakage Flow ◽  
Suction Surface ◽  
Mean Velocity ◽  
Wall Boundary Layer ◽  
Compressor Rotor ◽  
Wall Boundary

The flow field in the annulus wall and tip region of a compressor rotor was measured using a triaxial, hot-wire probe rotating with the rotor. The flow was surveyed across the entire passage at five axial locations (leading edge, 1/4 chord, 1/2 chord, 3/4 chord, and trailing edge locations) and at six radial locations inside the passage. The data derived include all three components of mean velocity. Blade-to-blade variations of the velocity components, pitch and yaw angles, as well as the passage-averaged mean properties of the annulus wall boundary layer, are derived from this data. The measurements indicate that the leakage flow starts beyond a quarter-chord and tends to roll up farther away from the suction surface than that observed in cascades. Substantial velocity deficiencies and radial inward velocities are observed in this region. The annulus wall boundary layer is well behaved up to half a chord, beyond which interactions with the leakage flow produce complex profiles.

Physical Explanations of Tip Leakage Flow Field in an Axial Compressor Rotor

1998 ◽  
Author(s):  
Masahiro Inoue ◽  
Masato Furukawa ◽  
Kazuhisa Saiki ◽  
Kazutoyo Yamada
Keyword(s):  
Boundary Layer ◽  
Vortex Core ◽  
Leakage Flow ◽  
Surface Boundary ◽  
Suction Surface ◽  
Surface Boundary Layer ◽  
Tip Leakage Vortex ◽  
Tip Leakage ◽  
Compressor Rotor ◽  
Vortex Lines

Structure of a tip leakage flow field in an axial compressor rotor has been investigated by detailed numerical simulations and appropriate post-processing. Physical explanations of the structure are made in terms of vortex-core identification, normalized helicity, vortex-lines, limiting streamlines, etc. The onset of the discrete tip leakage vortex is located on the suction surface at some distance from the leading edge. The vortex core with high vorticity is generated from a shear layer between the leakage jet flow and the main flow. The streamlines in the leakage flow are coiling around the vortex core. All the vortex-lines in the tip leakage vortex core link to ones in the suction surface boundary layer. The other vortex-lines in the suction surface boundary layer link to the vortex-lines in the pressure surface boundary layer and in the casing wall boundary layer. There are two mechanisms to reduce intensity of the tip leakage vortex: one is reduction of discharged vorticity caused by the linkage of vortex-lines between the suction surface and casing wall boundary layers, and another is diffusion of vorticity from the tip leakage vortex. Relative motion of the endwall has a substantial influence on the structure of the leakage flow field. In the case of a compressor rotor, it intensifies streamwise vorticity of the leakage vortex but reduces leakage flow loss.

Effect of Rotor Tip Gap Variation at the Rear Stages of an Axial Flow Compressor

2012 ◽  
Vol 225 ◽  
pp. 233-238
Author(s):  
A.M. Pradeep ◽  
R.N. Chiranthan ◽  
Debarshi Dutta ◽  
Bhaskar Roy
Keyword(s):  
Rotor Blade ◽  
Cross Flow ◽  
Axial Compressor ◽  
Axial Flow ◽  
Pressure Rise ◽  
Tip Clearance ◽  
Suction Surface ◽  
Design Point ◽  
Compressor Rotor ◽  
Multi Stage

In this paper, detailed analysis of the tip flow of an axial compressor rotor blade has been carried out using the commercial CFD package ANSYS CFX. The rotor blade was designed such that it is reminiscent of the rear stages of a multi-stage axial compressor. The effects of varying tip gaps are studied using CFD simulations for overall pressure rise and flow physics of the tip flow at the design point and near the peak pressure point. Rig tests of a low speed research compressor rotor with 3% tip clearance provided characteristics plots for validation of the CFD results. With increase in clearance from 1% to 4%, the rotor pressure rise at the design point was observed to decrease linearly. Increase in the clearance increases the cross flow across the tip; however, the magnitude of the average jet velocity crossing the tip decreases. The tip leakage vortex was observed to stay close to the suction surface with increase in clearance.

scholarly journals The Role of Tip Leakage Vortex Breakdown in Compressor Rotor Aerodynamics

1998 ◽  
Author(s):  
Masato Furukawa ◽  
Masahiro Inoue ◽  
Kazuhisa Saiki ◽  
Kazutoyo Yamada
Keyword(s):  
Flow Rate ◽  
Peak Pressure ◽  
Vortex Breakdown ◽  
Pressure Rise ◽  
Suction Surface ◽  
Tip Leakage Vortex ◽  
Rotor Aerodynamics ◽  
Tip Leakage ◽  
Compressor Rotor ◽  
Breakdown Region

The breakdown of tip leakage vortex has been investigated on a low-speed axial compressor rotor with moderate blade loading. Effects of the breakdown on the rotor aerodynamics are elucidated by Navier-Stokes flow simulations and visualization techniques for identifying the breakdown. The simulations show that the leakage vortex breakdown occurs inside the rotor at a lower flow rate than the peak pressure rise operating condition. The breakdown is characterized by the existence of the stagnation point followed by a bubble-like recirculation region. The onset of breakdown causes significant changes in the nature of the tip leakage vortex: large expansion of the vortex and disappearance of the streamwise vorticity concentrated in the vortex. The expansion has an extremely large blockage effect extending to the upstream of the leading edge. The disappearance of the concentrated vorticity results in no rolling-up of the vortex downstream of the rotor and the disappearance of the pressure trough on the casing. The leakage flow field downstream of the rotor is dominated by the outward radial flow resulting from the contraction of the bubble-like structure of the breakdown region. It is found that the leakage vortex breakdown plays a major role in characteristic of rotor performance at near-stall conditions. As the flow rate is decreased from the peak pressure rise operating condition, the breakdown region grows rapidly in the streamwise, spanwise and pitchwise directions. The growth of the breakdown causes the blockage and the loss to increase drastically. Then, the interaction of the breakdown region with the blade suction surface gives rise to the three-dimensional separation of the suction surface boundary layer, thus leading to a sudden drop in the total pressure rise across the rotor.

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.

An Experimental Study of the Compressor Rotor Blade Boundary Layer

1985 ◽  
Vol 107 (2) ◽  
pp. 364-372 ◽  
Author(s):  
M. Pouagare ◽  
J. M. Galmes ◽  
B. Lakshminarayana
Keyword(s):  
Boundary Layer ◽  
Rotor Blade ◽  
Radial Profile ◽  
Three Dimensional ◽  
Axial Flow ◽  
Tip Leakage Flow ◽  
Wall Boundary Layer ◽  
Compressor Rotor ◽  
Blade Tip ◽  
Flow Compressor

The three-dimensional turbulent boundary layer developing on a rotor blade of an axial flow compressor was measured using a minature “x” configuration hot-wire probe. The measurements were carried out at nine radial locations on both surfaces of the blade at various chordwise locations. The data derived includes streamwise and radial mean velocities and turbulence intensities. The validity of conventional velocity profiles such as the “power law profile” for the streamwise profile, and Mager and Eichelbrenner’s for the radial profile, is examined. A modification to Mager’s crossflow profile is proposed. Away from the blade tip, the streamwise component of the blade boundary layer seems to be mainly influenced by the streamwise pressure gradient. Near the tip of the blade, the behavior of the blade boundary layer is affected by the tip leakage flow and the annulus wall boundary layer. The “tangential blockage” due to the blade boundary layer is derived from the data. The profile losses are found to be less than that of an equivalent cascade, except in the tip region of the blade.

Prediction of Low Speed Compressor Rotor Flowfields With Large Tip Clearances

2003 ◽  
Author(s):  
Anurag Gupta ◽  
S. Arif Khalid ◽  
G. Scott McNulty ◽  
Lyle Dailey
Keyword(s):  
Three Dimensional ◽  
Axial Compressor ◽  
Pressure Rise ◽  
Tip Clearance ◽  
Navier Stokes ◽  
Leakage Flow ◽  
Low Speed ◽  
Compressor Rotor ◽  
Wide Range ◽  
Measured Pressure

Rotor tip modeling fidelity, grid resolution, and near wall modeling have been examined to determine the requirements for an accurate prediction of the effects of large tip clearance in a low-speed axial compressor rotor. The effort, using a Reynolds-Averaged Navier-Stokes (RANS) solver, aimed to obtain the most accurate predictions from a three-dimensional, steady, single blade row simulation. A recently tested, modern low speed rotor, was used as the test geometry; the measured pressure rise characteristic as well as detailed data near stall was used to evaluate the ability of different modeling strategies to capture the correct flow structure. The leakage flow was quantified to show that a wide range of tip blockage could be obtained for different simulations of the same geometry and conditions. The results show that using a square tip and gridding to fully resolve the real tip gap was better able to capture the effects of loading on the leakage flow than either of the approximate models studied. Sufficient clustering near the casing to capture the shear layers was also found to be critical. While wall integration provided the best results in simultaneously improving the prediction of pressure rise characteristics and flow range, higher fidelity wall modeling and a casing y+ of approximately 3 were found to provide similar benefits.

Stall Margin Improvement of a Single Stage Transonic Axial Flow Compressor Using Naturally Aspirated Slots

2015 ◽  
Author(s):  
Hari Krishna Borra ◽  
Dilipkumar Bhanudasji Alone
Keyword(s):  
Boundary Layer ◽  
Steady State ◽  
Rotor Blade ◽  
Three Dimensional ◽  
Axial Flow ◽  
Compressor Stage ◽  
Suction Surface ◽  
Stall Margin ◽  
Blade Tip ◽  
Flow Compressor

This paper describes the method to improve the stall margin of transonic axial flow compressor by controlling the boundary layer on the suction surface of the rotor blade tip through natural aspiration. Aspiration slots in the compressor blade are intended to energize the flow by increasing its momentum on the suction surface. This phenomenon of boundary layer control can delay the flow separation and hence results in enhancement of the stall margin of the compressor stage. Flow behavior with aspiration slots and its performance are evaluated using commercially available numerical software. Steady state RANS simulations with three dimensional implicit pressure-based coupled solver and turbulence model SST k-ω are used. The effect of natural aspiration slot on the rotor blade performance is computed numerically. The main objective of the study was to identify the optimum location of the aspiration slot along the chord of the compressor on the rotor blade. The axial location chosen for the performance evaluations were 20%,40%,50%,60% and 70% of the rotor blade axial tip chord. By comparing the numerical simulation results with the steady state behavior in the absence of the aspirated slots, the optimized location of the aspiration slot that results in maximum stall improvement is identified. At the optimized location, natural aspiration slots on the rotor blade tip improved the stall margin with the minimum reduction in efficiency and stage pressure ratio when compared to base model. After critically understanding the performance with straight aspiration slots the compressor stage performance has enhanced further by orienting the aspiration slots. The numerical three dimensional results conclude an optimal improvement in the stall margin for the slots near the trailing edge of the rotor. The prediction shows that with the inclined aspiration slots at proper location it is possible to improve the stall margin of the compressor stage and also to restore the stage efficiency.

Control of Three-Dimensional Separations in Axial Compressors by Tailored Boundary Layer Suction

2007 ◽  
Vol 130 (1) ◽  
Author(s):  
Semiu A. Gbadebo ◽  
Nicholas A. Cumpsty ◽  
Tom P. Hynes
Keyword(s):  
Boundary Layer ◽  
Static Pressure ◽  
Three Dimensional ◽  
Pressure Rise ◽  
Surface Flow ◽  
Compressor Cascade ◽  
Suction Surface ◽  
Blade Passage ◽  
Boundary Layer Suction ◽  
End Wall

One of the important ways of improving turbomachinery compressor performance is to control three-dimensional (3D) separations, which form over the suction surface and end wall corner of the blade passage. Based on the insights gained into the formation of these separations, this paper illustrates how an appropriately applied boundary layer suction of up to 0.7% of inlet mass flow can control and eliminate typical compressor stator hub corner 3D separation over a range of operating incidence. The paper describes, using computational fluid dynamics, the application of suction on the blade suction surface and end wall boundary layers and exemplifies the influence of end wall dividing streamline in initiating 3D separation in the blade passage. The removal of the separated region from the blade suction surface is confirmed by an experimental investigation in a compressor cascade involving surface flow visualization, surface static pressure, and exit loss measurements. The ensuing passage flow field is characterized by increased blade loading (static pressure difference between pressure and suction surface), enhanced average static pressure rise, significant loss removal, and a uniform exit flow. This result also enables the contribution of the 3D separation to the overall loss and passage blockage to be assessed.

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