The Effect of Two Different Endwall-Penny Concepts for Variable Stator Vanes in a Compressor Cascade

2012 ◽  
Author(s):  
Marcel Gottschall ◽  
Konrad Vogeler ◽  
Ronald Mailach
Keyword(s):  
Flow Field ◽  
Reynolds Numbers ◽  
Compressor Cascade ◽  
Reference Case ◽  
Flow Angle ◽  
Field Development ◽  
Pressure Losses ◽  
Angle Deviation ◽  
Stagger Angle ◽  
Variable Stator

The article describes investigations on the influence of two possible endwall-penny concepts for variable stator vanes to secondary flow field development and the performance of high pressure compressors. Concept I uses a penny covering the whole vane chord with no radial clearances, whilst the concept II applies a piecewise flattened polygonal hub or casing to achieve constant radial gaps. Both approaches were converted to a linear cascade of modern stator profiles. Measurements were conducted with a 5-hole-probe in planes up- and downstream the cascade as well as with pressure tappings on the airfoil and the endwall. Additional 3D numerical calculations were carried out to gain information about the flow field inside the cascade. These analyses were performed at three stagger angles and two characteristic Reynolds numbers with a constant cascade turning to model the adjusting range of aircraft engines. Compared to a reference case without endwall contour and no clearance the results indicate slightly increased efficiency due to smaller total pressure losses for both concepts. The penny edges as well as the polygonal endwall at the cascade inlet are responsible for a higher turbulence in the inlet boundary layer, which results in a smaller endwall separation that is detected with outlet loss distributions. Considering the cascade loading in terms of flow turning, a small overall benefit was achieved with both configurations caused by much higher overturning near the wall. These tendencies increase with the stagger angle at part-load conditions of the compressor. Furthermore, the results were compared with a part gap configuration of a variable vane with benefits concerning pressure losses and flow turning. For lower cascade loadings the concept I still reaches the highest overall flow turning. This benefit compared to the part gap configuration nearly disappears at the high stagger angle. Additionally, both concepts feature a much higher flow angle deviation along the blade height and significant increased losses.

The Effect of Four Part Gap Geometry Configurations for Variable Stator Vanes in a Compressor Cascade

2012 ◽  
Author(s):  
Marcel Gottschall ◽  
Konrad Vogeler ◽  
Ronald Mailach
Keyword(s):  
Flow Field ◽  
Separated Flow ◽  
Reynolds Numbers ◽  
Compressor Cascade ◽  
Field Development ◽  
Operating Range ◽  
Rear Part ◽  
Axis Position ◽  
Stagger Angle ◽  
Variable Stator

The article describes numerical investigations on the influence of four different endwall clearance topologies for variable stator vanes to secondary flow field development and the performance of high pressure compressors. The aim of this work is to quantify the characteristics of different clearance configurations depending on the penny-axis position and the penny diameter for a typical operating range. All clearance configurations were implemented to a linear cascade of modern stator profiles. The analysis was introduced using a relative clearance size of 1.3% chord at three stagger angles and two characteristic Reynolds numbers to model the operating range on aircraft engines. 3D numerical calculations were carried out to gain information about the flow field inside the cascade. They were compared with measurements of a 5-hole-probe as well as pressure tappings on the airfoil and the endwall. The CFD shows the clearance characteristics in good agreement with the measurements for the lower and the nominal stagger angle. Small gaps in the rear part of the vane have a beneficial effect on the flow field. In contrast, a clearance in the higher loaded front part of the vane always resulted in increased losses. Otherwise, the significant enhanced performance of a rear part gap, which was measured at the higher stagger angle, was not reflected by the CFD. The reduced mixing losses and the higher averaged flow turning even compared to a configuration without a clearance are not verified with the calculations. Large flow separations at the high stagger angle result in a two to four times higher underturning of the CFD in comparison to the experiments. The clearance effects to the characteristic radial loss distribution up to 40 % bladeheight also deviate from the measurements due to heavy mixing of clearance and reversed separated flow.

Investigation of Tip Clearance Phenomena in an Axial Compressor Cascade Using Euler and Navier–Stokes Procedures

1993 ◽  
Vol 115 (3) ◽  
pp. 453-467 ◽  
Author(s):  
R. F. Kunz ◽  
B. Lakshminarayana ◽  
A. H. Basson
Keyword(s):  
Flow Field ◽  
Axial Compressor ◽  
Numerical Procedure ◽  
Tip Clearance ◽  
Tip Vortex ◽  
Navier Stokes ◽  
Experimental Investigations ◽  
Compressor Cascade ◽  
Flow Angle ◽  
Physical Phenomena

Three-dimensional Euler and full Navier–Stokes computational procedures have been utilized to simulate the flow field in an axial compressor cascade with tip clearance. An embedded H-grid topology was utilized to resolve the flow physics in the tip gap region. The numerical procedure employed is a finite difference Runge-Kutta scheme. Available measurements of blade static pressure distributions along the blade span, dynamic pressure and flow angle in the cascade outlet region, and spanwise distributions of blade normal force coefficient and circumferentially averaged flow angle are used for comparison. Several parameters that were varied in the experimental investigations were also varied in the computational studies. Specifically, measurements were taken and computations were performed on the configuration with and without: tip clearance, the presence of an endwall, inlet endwall total pressure profiles and simulated relative casing rotation. Additionally, both Euler and Navier–Stokes computations were performed to investigate the relative performance of these approaches in reconciling the physical phenomena considered. Results indicate that the Navier–Stokes procedure, which utilizes a low Reynolds number k–ε model, captures a variety of important physical phenomena associated with tip clearance flows with good accuracy. These include tip vortex strength and trajectory, blade loading near the tip, the interaction of the tip clearance flow with passage secondary flow, and the effects of relative endwall motion. The Euler computation provides good but somewhat diminished accuracy in resolution of some of these clearance phenomena. It is concluded that the level of modeling embodied in the present approach is sufficient to extract much of the tip region flow field information useful to designers of turbomachinery.

Meridional Flow Calculation Using Advanced CFD Techniques

1992 ◽  
Author(s):  
P. Kiousis ◽  
P. Chaviaropoulos ◽  
K. D. Papailiuou
Keyword(s):  
High Speed ◽  
Industrial Applications ◽  
Meridional Flow ◽  
Finite Difference Schemes ◽  
Governing Equations ◽  
Flow Angle ◽  
Flow Equations ◽  
Pressure Losses ◽  
Angle Deviation ◽  
Calculation Results

Working experience on traditional Meridional Flow Solvers has revealed difficulties concerning both convergence and accuracy of the solution. These difficulties have been observed for instance in certain industrial applications where steep gradients of flow and/or geometrical quantities are present. Transonic flow conditions can cause extra difficulties. All these difficulties may be circumvented when advanced CFD techniques are utilized. A computational tool, suitable for the solution of the Meridional Through Flow equations for Turbomachinery applications is presented. Assuming that the flow is compressible and inviscid the governing equations are obtained using a streamfunction formulation for the pitch-averaged flow equations. Viscous corrections have been incorporated in the inviscid model in terms of flow angle deviation and total pressure losses. Governing equations, are discretized using body fitted finite difference schemes. An artificial density upwinding scheme assures convergence in the transonic region. Particular attention has been paid to the numerical integration procedure which is based on a preconditioned gradient method (GMRES). Calculation results for low and high-speed turbomachines are presented and discussed.

Influence of Surface Roughness on the Profile and End-Wall Losses in Low Pressure Turbines

2011 ◽  
Author(s):  
Rau´l Va´zquez ◽  
Diego Torre ◽  
Fernando Partida ◽  
Leyre Arman˜anzas ◽  
Antonio Antoranz
Keyword(s):  
Surface Roughness ◽  
Total Pressure ◽  
Reynolds Numbers ◽  
Low Pressure ◽  
Pressure Flow ◽  
Flow Angle ◽  
Pressure Losses ◽  
Wide Range ◽  
Wall Losses ◽  
End Wall

The influence of surface roughness on the profile and end-wall total pressure losses in Low Pressure Turbines was investigated experimentally in a turbine high-speed rig. The rig consisted of a rotor-stator configuration. Both rows of airfoils are high lift, high aspect ratio and high turning blades that are characteristic of state of the art Low Pressure Turbines. The stator airfoils (both vanes and platforms) were casted and afterwards they were barreled to improve their surface finish up to 1.73 μm Ra. Then they were assembled in the rig and tested. The stator was traversed upstream and downstream with miniature pneumatic probes to obtain total pressure, flow angle and static pressure flow fields. Once this test was completed the rig was disassembled and the stator airfoils were polished to achieve a roughness size of 0.72 μm Ra, characteristic of Low Pressure Turbine polished airfoils. Once again, the stators were assembled in the rig and tested to carry out a back-to-back comparison between the two different surface roughnesses. The total pressure profile and end-wall losses were measured for a wide range of Reynolds numbers, extending from 8×104 to 2.4×105, based on suction surface length (Res∼1.5 ReCx) and exit Mach number of 0.61. Experimental results are presented and compared in terms of area average, radial pitchwise average distributions and exit plane contours of total pressure losses, flow angles and helicity. The results agree with previous studies of roughness in Turbines, a beneficial effect of surface roughness was found at very low Reynolds numbers, in stagnation pressure losses.

Influence of an Adaptive Blade System on Unsteady Flow Conditions in a 2D Compressor Cascade

2017 ◽  
Author(s):  
Julija Peter ◽  
David Konstantin Tilcher ◽  
Robert Meyer ◽  
Paul Uwe Thamsen
Keyword(s):  
Phase Shift ◽  
Flow Field ◽  
Unsteady Flow ◽  
High Speed ◽  
Optical Measurement ◽  
Water Channel ◽  
Tip Clearance ◽  
Flow Conditions ◽  
Compressor Cascade ◽  
Flow Angle

The flow field inside a compressor is characterized by highly unsteady flow effects. Consequently, the performance of a compressor is significantly influenced by the complex flow field. Especially at off-design conditions, flow separation and tip clearance flow cause vortex structures and thus increased losses. The objective of this paper is to give an insight into the effect mechanism of the movable stator vanes as an adaptive system to affect unsteady flow conditions. The experiments were conducted in a stator cascade in a water channel at a Reynolds number of Re = 500 000. Inlet guide vanes with movable flaps were used to simulate the periodic variation of the inlet flow angle. As parameters, the mean stagger angle of the stator cascade as well as the phase shift between the sinusoidal movement of the stator and the inlet guide cascade were varied. By using the optical measurement technique High-Speed Particle Image Velocimetry (HS-PIV), the flow fields upstream and downstream of the stator cascade were captured. Overall, the results revealed that the loss coefficient is strongly dependent on the phase shift between the inlet guide cascade and the stator cascade. Using certain phase shifts, a reduction in losses of up to 20% was achieved by the movable stator cascade.

Impact of Surface Roughness on Compressor Cascade Performance

2010 ◽  
Vol 132 (6) ◽  
Author(s):  
Seung Chul Back ◽  
June Hyuk Sohn ◽  
Seung Jin Song
Keyword(s):  
Gas Turbines ◽  
Velocity Ratio ◽  
Compressor Blade ◽  
Compressor Cascade ◽  
Roughness Effects ◽  
Flow Angle ◽  
Angle Deviation ◽  
Blade Row ◽  
Rough Region ◽  
Linear Compressor Cascade

This paper presents an experimental investigation of roughness effects on aerodynamic performance in a low-speed linear compressor cascade. Equivalent sandgrain roughnesses of 12 μm, 180 μm, 300 μm, 425 μm, and 850 μm have been tested. In nondimensional terms, these roughnesses represent compressor blade roughnesses found in actual gas turbines. Downstream pressure, velocity, and angle have been measured with a five-hole probe at 0.3 chord downstream of the blade trailing edge. For the tested roughnesses of 180 μm, 300 μm, 425 μm, and 850 μm, the axial velocity ratio across the blade row decreases by 0.1%, 2.1%, 2.5%, and 5.4%, respectively. For the same cases, the exit flow angle deviation increases by 24%, 38%, 51%, and 70%, respectively. Finally, the mass-averaged total pressure loss increases by 12%, 44%, 132%, and 217%, respectively. Also, the loss increases more rapidly in the transitionally rough region. Thus, among the three parameters, the loss responds most sensitively to changes in compressor blade roughness.

Study of Relative Endwall Motion Effects in a Compressor Cascade Through Direct Numerical Simulations

2020 ◽  
Vol 143 (1) ◽  
Author(s):  
Jordi Ventosa-Molina ◽  
Martin Lange ◽  
Ronald Mailach ◽  
Jochen Fröhlich
Keyword(s):  
Numerical Simulations ◽  
Secondary Flow ◽  
Relative Motion ◽  
Direct Numerical Simulations ◽  
Compressor Cascade ◽  
Flow Angle ◽  
Pressure Losses ◽  
Wake Effects ◽  
Secondary Flow Structure ◽  
Linear Compressor Cascade

Abstract Linear cascades are commonly used as surrogate geometries when performing fundamental studies of turbomachinery blading. Several effects are not accounted for in linear cascades, such as the relative motion between blade and endwall. In this study, three different relative endwall velocities are analyzed. The effect of the relative motion between endwall and blade in a linear compressor cascade is studied through direct numerical simulations. Results show a significant change in the secondary flow structure within the passage. Most notably, the tip leakage vortex is displaced away from the blade. Still, the blade spanwise range affected by the secondary flow field is similar to the case without relative endwall motion. At the outlet plane, a stratification of the total pressure losses and the exit flow angle is found, which overshadows any blade wake effects near the endwall.

Investigations of Three-Dimensional Flow Field Development in an Axial Compressor Cascade

2019 ◽  
Author(s):  
Saeed A. El-Shahat ◽  
Hesham M. El-Batsh ◽  
Ali M. A. Attia ◽  
Guojun Li ◽  
Lei Fu
Keyword(s):  
Flow Field ◽  
Three Dimensional ◽  
Computational Results ◽  
Compressor Cascade ◽  
Linear Compressor ◽  
Blade Passage ◽  
Field Development ◽  
Hole Pressure ◽  
Sensor Module ◽  
Linear Compressor Cascade

Abstract This paper presents a complete study about three-dimensional (3-D) flow field development in a linear compressor cascade where flow field in the blade passage has been studied experimentally as well as numerically. In the experimental work, a linear compressor cascade test section was installed in an open loop wind tunnel. The experimental data was acquired for a Reynolds number of 2.98 × 105 based on the blade chord and the inlet flow conditions. The flow field characteristics in blade passage including 3-D flow velocity and velocity magnitude have been measured by using calibrated five and seven-hole pressure probes connected to ATX sensor module data acquisition system (DAQ). To investigate flow development in the blade passage, velocity coefficient through streamwise planes has been calculated from the measured data. The computational fluid dynamics (CFD) study of the flow field was performed to gain a better understanding of the flow features. Present computational study was first validated with previous experimental and numerical work to check mesh accuracy and give confidence for computational results. Then, two turbulence models, Spalart-Allmaras (S-A) and shear stress transport SST (k-ω) were used for the present work. From both parts of study, the flow field development through the cascade have been investigated and compared. Moreover, the received data demonstrated a good agreement between the experimental and computational results. The predicted flow streamlines by numerical calculations showed regions characterized by flow separation and recirculation zones such as corner separation that could be used to enhance the understanding of the loss mechanism in compressor cascades. All measurements taken by the two probes, 5 and 7-hole pressure probes, have been analyzed and compared. The 5-hole pressure probe measurements have showed more agreements with computational results than 7-hole probe. Furthermore S-A turbulence model calculations showed more consistencies with experimental results than SST (k-ω) model.

Study of Relative Endwall Motion Effects in a Compressor Cascade Through Direct Numerical Simulations

2020 ◽  
Author(s):  
Jordi Ventosa-Molina ◽  
Martin Lange ◽  
Ronald Mailach ◽  
Jochen Fröhlich
Keyword(s):  
Numerical Simulations ◽  
Secondary Flow ◽  
Relative Motion ◽  
Direct Numerical Simulations ◽  
Compressor Cascade ◽  
Flow Angle ◽  
Pressure Losses ◽  
Wake Effects ◽  
Secondary Flow Structure ◽  
Linear Compressor Cascade

Abstract Linear cascades are commonly used as surrogate geometries when performing fundamental studies of turbomachinery blading. Several effects are not accounted for in linear cascades, among them the relative motion between blade and endwall. In this study three different relative endwall velocities are analysed. The effect of the relative motion between endwall and blade in a linear compressor cascade is studied through Direct Numerical Simulations. Results show a significant change in the secondary flow structure within the passage. Most notably, the tip leakage vortex is displaced away from the blade. Still, the blade spanwise range affected by the secondary flow field is similar to the case without relative endwall motion. At the outlet plane, a stratification of the total pressure losses and the exit flow angle is found, which overshadows any blade wake effects near the endwall.

scholarly journals The Effect of the Reynolds Number on the Three-Dimensional Flow in a Straight Compressor Cascade

1988 ◽  
Author(s):  
Václav Cyrus
Keyword(s):  
Reynolds Number ◽  
Aspect Ratio ◽  
Three Dimensional ◽  
Reynolds Numbers ◽  
Compressor Cascade ◽  
Simple Correlation ◽  
Three Dimensional Flow ◽  
Flow Angle ◽  
Low Reynolds Numbers ◽  
Dimensional Flow

A straight compressor cascade of aspect ratio 2 was tested in a low speed tunnel within Reynolds number Re1 = 45 000 – 150 000 and inlet flow angle α1 = 35° – 48°. The profile of the blade was NACA 65-12-10. The purpose of the paper was to obtain data on three–dimensional flow in a straight cascade at low Reynolds numbers. Some experimental results on secondary flow have been made into simple correlation relations.

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