Development of Hub Corner Stall and Its Influence on the Performance of Axial Compressor Blade Rows

1999 ◽  
Vol 121 (1) ◽  
pp. 67-77 ◽  
Author(s):  
C. Hah ◽  
J. Loellbach
Keyword(s):  
Numerical Solutions ◽  
Stokes Equations ◽  
Flow Direction ◽  
Three Dimensional ◽  
Suction Side ◽  
Compressor Blade ◽  
Suction Surface ◽  
Simple Diffusion ◽  
Transonic Compressor ◽  
Compressor Rotor

A detailed investigation has been performed to study hub corner stall phenomena in compressor blade rows. Three-dimensional flows in a subsonic annular compressor stator and in a transonic compressor rotor have been analyzed numerically by solving the Reynolds-averaged Navier–Stokes equations. The numerical results and the existing experimental data are interrogated to understand the mechanism of compressor hub corner stall. Both the measurements and the numerical solutions for the stator indicate that a strong twisterlike vortex is formed near the rear part of the blade suction surface. Low-momentum fluid inside the hub boundary layer is transported toward the suction side of the blade by this vortex. On the blade suction surface near the hub, this vortex forces fluid to move against the main flow direction and a limiting stream surface is formed near the hub. The formation of this vortex is the main mechanism of hub corner stall. When the aerodynamic loading is increased, the vortex initiates further upstream, which results in a larger corner stall region. For the transonic compressor rotor studied in this paper, the numerical solution indicates that a mild hub corner stall exists at 100 percent rotor speed. The hub corner stall, however, disappears at the reduced blade loading, which occurs at 60 percent rotor design speed. The present study demonstrates that hub corner stall is caused by a three-dimensional vortex system and that it does not seem to be correlated with a simple diffusion factor for the blade row.

Development of Hub Corner Stall and its Influence on the Performance of Axial Compressor Blade Rows

1997 ◽  
Author(s):  
Chunill Hah ◽  
James Loellbach
Keyword(s):  
Numerical Solutions ◽  
Stokes Equations ◽  
Flow Direction ◽  
Three Dimensional ◽  
Suction Side ◽  
Compressor Blade ◽  
Suction Surface ◽  
Simple Diffusion ◽  
Transonic Compressor ◽  
Compressor Rotor

A detailed investigation has been performed to study hub corner stall phenomena in compressor blade rows. Three-dimensional flows in a subsonic annular compressor stator and in a transonic compressor rotor have been analyzed numerically by solving the Reynolds-averaged Navier-Stokes equations. The numerical results and the existing experimental data are interrogated to understand the mechanism of compressor hub corner stall. Both the measurements and the numerical solutions indicate that a strong twister-like vortex is formed near the rear part of the blade suction surface. Low momentum fluid inside the hub boundary layer is transported toward the suction side of the blade by this vortex. On the blade suction surface near the hub, this vortex forces fluid to move against the main flow direction and a limiting stream surface is formed near the hub. The formation of this vortex is the main mechanism of hub corner stall. When the aerodynamic loading is increased, the vortex initiates further upstream, which results in a larger corner stall region. For the transonic compressor rotor studied in this paper, the numerical solution and the measured data indicate that a mild hub corner stall exists at 100 percent rotor speed. The hub corner stall, however, disappears at the reduced blade loading which occurs at 60 percent rotor design speed. The present study demonstrates that hub corner stall is caused by a three-dimensional vortex system and that it does not seem to be correlated with a simple diffusion factor for the blade row.

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.

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.

scholarly journals Navier-Stokes Solutions of Unsteady Flow in a Compressor Rotor

1986 ◽  
Author(s):  
James N. Scott ◽  
Wilbur L. Hankey
Keyword(s):  
Unsteady Flow ◽  
Turbulence Model ◽  
Stokes Equations ◽  
Numerical Procedure ◽  
Navier Stokes ◽  
Suction Surface ◽  
Blade Passage ◽  
Radial Location ◽  
Transonic Compressor ◽  
Compressor Rotor

In order to achieve more accurate predictions of unsteady flow in a transonic compressor rotor an existing numerical approach has been modified by incorporating a turbulence model. The computations are performed by solving the complete time-dependent compressible Navier-Stokes equations using MacCormack’s explicit finite difference algorithm. These equations are solved for the flow through two adjacent rotor blades at a stream surface near the blade tip subjected to the wakes emitted from upstream stators. At this radial location the flow enters the blade passage at an absolute Mach number of 0.66. The high blade curvature at this radial location produces a large region of separated flow on the suction surface with laminar flow. To more accurately resolve the features of this flow separation the Baldwin-Lomax algebraic eddy-viscosity turbulence model is incorporated into the numerical procedure in regions, near the blade surface. The unsteady flow features are represented at the inflow boundary through the use of characteristic variables involving the upstream and downstream running Riemann invariants and the entropy variation expressed in terms of the total pressure profile. At the outflow boundary the concept of a “second throat” or choke point is implemented in conjunction with supersonic outflow conditions. The results are compared with numerical results obtained without the use of a turbulence model (laminar) for a single blade passage. Improved agreement with limited experimental data is also noted.

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.

Three-Dimensional Inverse Method for Turbomachinery Blading Design

1998 ◽  
Vol 120 (2) ◽  
pp. 247-255 ◽  
Author(s):  
A. Demeulenaere ◽  
R. Van den Braembussche
Keyword(s):  
Pressure Distribution ◽  
Three Dimensional ◽  
Suction Side ◽  
Flow Angle ◽  
Transonic Compressor ◽  
Compressor Rotor ◽  
Lean Angle ◽  
Outlet Flow ◽  
Target Pressure ◽  
Euler Solver

An iterative procedure for three-dimensional blade design is presented, in which the three-dimensional blade shape is modified using a physical algorithm, based on the transpiration model. The transpiration flux is computed by means of a modified Euler solver, in which the target pressure distribution is imposed along the blade surfaces. Only a small number of modifications is needed to obtain the final geometry. The method is based on a high-resolution three-dimensional Euler solver. An upwind biased evaluation of the advective fluxes allows for a very low numerical entropy generation, and sharp shock capturing. Non-reflecting boundary conditions are applied along the inlet/outlet boundaries. The capabilities of the method are illustrated by redesigning a transonic compressor rotor blade, to achieve, for the same mass flow and outlet flow angle, a shock-free deceleration along the suction side. The second example concerns the design of a low aspect ratio turbine blade, with a positive compound lean to reduce the intensity of the passage vortices. The final blade is designed for an optimized pressure distribution, taking into account the forces resulting from the blade lean angle.

Navier–Stokes Solutions of Unsteady Flow in a Compressor Rotor

1986 ◽  
Vol 108 (2) ◽  
pp. 206-215 ◽  
Author(s):  
J. N. Scott ◽  
W. L. Hankey
Keyword(s):  
Unsteady Flow ◽  
Turbulence Model ◽  
Stokes Equations ◽  
Numerical Procedure ◽  
Navier Stokes ◽  
Suction Surface ◽  
Blade Passage ◽  
Radial Location ◽  
Transonic Compressor ◽  
Compressor Rotor

In order to achieve more accurate predictions of unsteady flow in a transonic compressor rotor an existing numerical approach has been modified by incorporating a turbulence model. The computations are performed by solving the complete time-dependent compressible Navier–Stokes equations using MacCormack’s explicit finite difference algorithm. These equations are solved for the flow through two adjacent rotor blades at a streamsurface near the blade tip subjected to the wakes emitted from upstream stators. At this radial location the flow enters the blade passage at an absolute Mach number of 0.66. The high blade curvature at this radial location produces a large region of separated flow on the suction surface with laminar flow. To more accurately resolve the features of this flow separation the Baldwin–Lomax algebraic eddy-viscosity turbulence model is incorporated into the numerical procedure in regions near the blade surface. The unsteady flow features are represented at the inflow boundary through the use of characteristic variables involving the upstream and downstream running Riemann invariants and the entropy variation expressed in terms of the total pressure profile. At the outflow boundary the concept of a “second throat” or choke point is implemented in conjunction with supersonic outflow conditions. The results are compared with numerical results obtained without the use of a turbulence model (laminar) for a single blade passage. Improved agreement with limited experimental data is also noted.

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.

Pre-Deformation Method for Manufactured Compressor Blade Based on Load Incremental Approach

2020 ◽  
Vol 37 (3) ◽  
pp. 259-265
Author(s):  
Kang Da ◽  
Wang Yongliang ◽  
Zhong Jingjun ◽  
Liu Zihao
Keyword(s):  
Compressor Blade ◽  
Aerodynamic Load ◽  
Deformation Method ◽  
Transonic Compressor ◽  
Incremental Approach ◽  
Compressor Rotor ◽  
Centrifugal Load ◽  
Blade Shape ◽  
Stationary Shape ◽  
Deformation Procedure

AbstractThe blade deformation caused by aerodynamic and centrifugal loads during operating makes blade configurations different from their stationary shape. Based on the load incremental approach, a novel pre-deformation method for cold blade shape is provided in order to compensate blade deformation under running. Effect of nonlinear blade stiffness is considered by updating stiffness matrix in response to the variation of blade configuration when calculating deformations. The pre-deformation procedure is iterated till a converged cold blade shape is obtained. The proposed pre-deformation method is applied to a transonic compressor rotor. Effect of load conditions on blade pre-deformation is also analyzed. The results show that the pre-deformation method is easy to implement with fast convergence speed. Neither the aerodynamic load nor centrifugal load can be neglected in blade pre-deformation.

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.

Sign in / Sign up

Export Citation Format

Share Document