Numerical Investigation on Propagation Characteristics of Inlet Total-Pressure Distortion in a Centrifugal Compressor

2021 ◽  
Author(s):  
Mingyi Wang ◽  
Zhiheng Wang ◽  
Guang Xi ◽  
Yurun Li
Keyword(s):  
Centrifugal Compressor ◽  
Total Pressure ◽  
Three Dimensional ◽  
Leading Edge ◽  
Low Pressure ◽  
Propagation Characteristics ◽  
Suction Surface ◽  
Axial Section ◽  
Flow Angle ◽  
Rotation Direction

Abstract The propagation characteristics of inlet total-pressure distortion in a centrifugal compressor are investigated by full-annulus unsteady three-dimensional numerical simulation. The inlet distortions considered in the paper are the total-pressure distortions covering a 60-deg sector (60deg distortion) and three 20-deg sectors (3*20deg distortion), respectively. One is the classical distortion form, and the other is to simulate the downstream flow of the axial section of a centrifugal-axial combined compressor. By analyzing the distributions of flow parameters, the propagation of the total-pressure distortion in the centrifugal compressor is interpreted. The results show that, with the distortion propagating to the downstream, the low-pressure region produces a phase deviation along the streamwise direction relative to the opposite direction of impeller rotation direction, and the range of distortion region is reduced. Additionally, the propagation of the inlet distortion makes the three-dimensional characteristics of airflow more complex. The flow angle increases with different amplitudes along the direction of blade height corresponding to the distorted sector. The distortion region affects the location of blades which are in a low-pressure area, and the intensity of the distortion affects the increase of the flow angle. The distortion region causes more local relative flow losses, especially near the leading edge of blade suction surface.

Three-Dimensional Relief in Turbomachinery Blading

1990 ◽  
Vol 112 (4) ◽  
pp. 587-596 ◽  
Author(s):  
A. R. Wadia ◽  
B. F. Beacher
Keyword(s):  
High Speed ◽  
Three Dimensional ◽  
Axial Flow ◽  
Leading Edge ◽  
Two Dimensional ◽  
Suction Surface ◽  
Airfoil Surface ◽  
Flow Angle ◽  
Edge Loading ◽  
Cascade Analysis

The leading edge region of turbomachinery blading in the vicinity of the endwalls is typically characterized by an abrupt increase in the inlet flow angle and a reduction in total pressure associated with endwall boundary layer flow. Conventional two-dimensional cascade analysis of the airfoil sections at the endwalls indicates large leading edge loadings, which are apparently detrimental to the performance. However, experimental data exist that suggest that cascade leading edge loading conditions are not nearly as severe as those indicated by a two-dimensional cascade analysis. This discrepancy between two-dimensional cascade analyses and experimental measurements has generally been attributed to inviscid three-dimensional effects. This article reports on two and three-dimensional calculations of the flow within two axial-flow compressor stators operating near their design points. The computational results of the three-dimensional analysis reveal a significant three-dimensional relief near the casing endwall that is absent in the two-dimensional calculations. The calculated inviscid three-dimensional relief at the endwall is substantiated by airfoil surface static pressure measurements on low-speed research compressor blading designed to model the flow in the high-speed compressor. A strong spanwise flow toward the endwall along the leading edge on the suction surface of the airfoil is responsible for the relief in the leading edge loading at the endwall. This radial migration of flow results in a more uniform spanwise loading compared to that predicted by two-dimensional calculations.

The Influence of Different Endwall Contouring Locations on the Secondary Flow Losses in a Highly Loaded Low Pressure Turbine

2018 ◽  
Author(s):  
Wenhua Duan ◽  
Weiyang Qiao ◽  
Zuojun Wei ◽  
Jian Liu ◽  
Haoyi Cheng
Keyword(s):  
Secondary Flow ◽  
Separation Bubble ◽  
Three Dimensional ◽  
Leading Edge ◽  
Low Pressure ◽  
Suction Surface ◽  
Surface Separation ◽  
Flow Losses ◽  
Endwall Contouring ◽  
Highly Loaded

A detailed experimental and numerical investigation of the effect of endwall contouring on the secondary flow was performed in a highly loaded low pressure linear cascade. A comparison between a planar and three contoured configurations has been performed, and the three-dimensional endwall secondary flow structures and the secondary flow losses are here analyzed and discussed. For the contoured endwall, three locations of the same contouring were investigated, first one with the contouring starting upstream the airfoil and ending just before the leading edge of the airfoil, secondary one with the contouring starting upstream the airfoil and ending in the middle of the airfoil passage, the last one with the contouring starting just from the leading edge of the airfoil and continuing through the airfoil passage just ending at the trailing edge. The results suggest that the numerical results offer a reliable prediction for the endwall secondary flow structure together with suction surface separation bubble. It was found that all the three locations of contouring could reduce the secondary flow losses effectively. The location through the airfoil passage got the most benefit in the reduction of secondary flow losses whereas the all contouring upstream the airfoil location got the least. It was also found that the profile losses was affected by the contoured endwall.

Three-Dimensional Relief in Turbomachinery Blading

1989 ◽  
Author(s):  
A. R. Wadia ◽  
B. F. Beacher
Keyword(s):  
High Speed ◽  
Three Dimensional ◽  
Axial Flow ◽  
Leading Edge ◽  
Two Dimensional ◽  
Suction Surface ◽  
Airfoil Surface ◽  
Flow Angle ◽  
Edge Loading ◽  
Cascade Analysis

The leading edge region of turbomachinery blading in the vicinity of the endwalls is typically characterized by an abrupt increase in the inlet flow angle and a reduction in total pressure associated with endwall boundary layer flow. Conventional two-dimensional cascade analysis of the airfoil sections at the endwalls indicates large leading edge loadings, apparently detrimental to the performance. However, experimental data exist that suggest that cascade leading edge loading conditions are not nearly as severe as those indicated by a two-dimensional cascade analysis. This discrepancy between two-dimensional cascade analyses and experimental measurements has generally been attributed to inviscid three-dimensional effects. This article reports on two- and three-dimensional calculations of the flow within two axial flow compressor Stators operating near their design points. The computational results of the three-dimensional analysis reveal a significant three-dimensional relief near the casing endwall absent in the two-dimensional calculations. The calculated inviscid three-dimensional relief at the endwall is substantiated by airfoil surface static pressure measurements on low speed research compressor blading designed to model the flow in the high speed compressor. A strong spanwise flow towards the endwall along the leading edge on the suction surface of the airfoil is responsible for the relief in the leading edge loading at the endwall. This radial migration of flow results in a more uniform spanwise loading compared to that predicted by two-dimensional calculations.

Flowfield in a Film-Cooled Three-Dimensional Contoured Endwall Passage: Aerodynamic Measurements

2007 ◽  
Author(s):  
Ross Gustafson ◽  
Gazi I. Mahmood ◽  
Sumanta Acharya
Keyword(s):  
Film Cooling ◽  
Total Pressure ◽  
Three Dimensional ◽  
Leading Edge ◽  
Suction Side ◽  
Axial Direction ◽  
Turbine Rotor ◽  
Suction Surface ◽  
Pressure Loss Coefficient ◽  
Profile Height

The influence of endwall film cooling on the aerodynamic performance of a linear blade cascade employing an asymmetric contoured endwall is measured. Cylindrical coolant holes are strategically located on the contoured endwall to provide full coverage film cooling. The endwall contour is varied along the pitch direction, and is elevated near the pressure side and depressed near the suction surface. The profile height also varies in the axial direction from the inlet to exit. Measurements of total pressure, vorticity, and velocity components are obtained at different axial locations inside the passage for six inlet blowing ratios ranging from 1.0 to 2.4. The results are compared with the measured data on the contoured endwall without any film cooling flow (uncooled case). All the tests are performed in a low speed cascade facility employing a scaled up two-dimensional blade profile of the GE-E3 turbine rotor section. The Reynolds number based on the cascade inlet velocity and blade actual chord is 2.30×105. The results near the leading edge show that the suction side-leg vortex on the uncooled endwall weakens when the film cooling jets are employed. The axial vorticity and velocity vectors near the exit plane indicate that the passage vortex is located nearer to the mid-pitch location and higher above the endwall for the film-cooled endwall than for the uncooled contoured endwall. While the overall total pressure loss coefficient across the passage decreases at the high blowing ratios compared to the uncooled case, the overall cascade loss increases with the blowing ratio.

Secondary Flow and Loss Distribution in a Radial Compressor With Untwisted Backswept Vanes

1990 ◽  
Author(s):  
G. Sipos
Keyword(s):  
Three Dimensional ◽  
Leading Edge ◽  
Suction Side ◽  
Meridional Flow ◽  
Specific Work ◽  
Suction Surface ◽  
Inlet Flow ◽  
Loss Distribution ◽  
Flow Angle ◽  
Radial Compressor

The unshrouded impeller and the vaneless diffuser of a single-stage radial compressor have been investigated at three flow rates. Three-dimensional velocities and pressures were measured at a tip speed of 84 m/s by an L2F-velocimeter, a slanted single hot-wire probe and piezoresistive pressure transducers. The measurements show that upstream the blading the averaged meridional inlet flow angle is about 54 degree and a periodical variation of the meridional flow angle of about 25 degree occurs near the casing wall. Further, an inlet vortex of clockwise direction appears and an initial whirl is induced. The specific work of the initial whirl corresponds to approximately 12% of the enthalpy losses between inlet pipe and diffuser outlet. In the beginning of the passage, the inlet vortex is suppressed and a solid body vortex of counterclockwise direction can be observed. At the outlet, a heavy flow deceleration at the blade suction side with subsequent separation can be seen. Increasing the flow rate decreases the wake and causes a more uniform loss distribution in this area. The measured secondary vortex flow and rotary stagnation pressure gradients are compared with test results from impellers with inducer. The incidence of the investigated impeller is greater than that of the impellers with inducer, but the wake-jet outlet flows are very similar. Inlet losses could be reduced by improving incidence angles by matching the blade angles to the inlet flow angles. Smaller blade angles at the shroud would reduce or eliminate separation at the leading edge, and the resulting reduction in low momentum fluid along the suction surface would help to avoid separation on that surface near the outlet.

Passive Boundary Layer Control on a Highly Loaded Low Pressure Turbine Cascade

2010 ◽  
Author(s):  
Markus Martinstetter ◽  
Reinhard Niehuis ◽  
Matthias Franke
Keyword(s):  
Total Pressure ◽  
Three Dimensional ◽  
Reynolds Numbers ◽  
Low Pressure ◽  
Boundary Layer Control ◽  
Loss Reduction ◽  
Suction Surface ◽  
Low Pressure Turbine ◽  
Layer Control ◽  
Highly Loaded

The objective of this work is to investigate the performance of a highly loaded Low Pressure Turbine (LPT)-Cascade over a wide range of turbomachinery realistic Reynolds numbers, while maintaining a constant real engine Mach number. Fundamental investigations show a strong increase in total pressure loss at low Reynolds number due to an enlarged separation region on the suction surface. The aim of this paper is to reduce total pressure loss through the application of passive turbulators on the suction surface. Both, steady inflow conditions, as well as effects of rotor-stator-interaction, are considered. To identify the importance of turbulator geometry on loss reduction, three dimensional turbulator elements have been compared to a two dimensional trip. Both passive devices were located on the suction surface and the three dimensional elements performed better at high Reynolds numbers. Based on the experimental investigations, the use of passive boundary layer control can be recommended as a promising approach for loss reduction on highly loaded low pressure turbine profiles.

Investigation of an Inversely Designed Centrifugal Compressor Stage—Part I: Design and Numerical Verification

2004 ◽  
Vol 126 (1) ◽  
pp. 73-81 ◽  
Author(s):  
M. Zangeneh ◽  
M. Schleer ◽  
F. Pløger ◽  
S. S. Hong ◽  
C. Roduner ◽  
...  
Keyword(s):  
Centrifugal Compressor ◽  
Design Method ◽  
Three Dimensional ◽  
Leading Edge ◽  
Inverse Design ◽  
Numerical Verification ◽  
Appreciable Effect ◽  
Tip Leakage Flow ◽  
Flow Angle ◽  
Blade Geometry

In this paper the three-dimensional inverse design code TURBOdesign-1 is applied to the design of the blade geometry of a centrifugal compressor impeller with splitter blades. In the design of conventional impellers the splitter blades normally have the same geometry as the full blades and are placed at mid-pitch location between the two full blades, which can usually result in a mismatch between the flow angle and blade angles at the splitter leading edge. In the inverse design method the splitter and full blade geometry is computed independently for a specified distribution of blade loading on the splitter and full blades. In this paper the basic design methodology is outlined and then the flow in the conventional and inverse designed impeller is compared in detail by using computational fluid dynamics (CFD) code TASCflow. The CFD results confirm that the inverse design impeller has a more uniform exit flow, better control of tip leakage flow and higher efficiency than the conventional impeller. The results also show that the shape of the trailing edge geometry has a very appreciable effect on the impeller Euler head and this must be accurately modeled in all CFD computations to ensure closer match between CFD and experimental results. Detailed measurements are presented in part II of the paper.

Secondary Flow and Loss Distribution in a Radial Compressor With Untwisted Backswept Vanes

1991 ◽  
Vol 113 (4) ◽  
pp. 686-695 ◽  
Author(s):  
G. Sipos
Keyword(s):  
Three Dimensional ◽  
Leading Edge ◽  
Suction Side ◽  
Periodic Variation ◽  
Specific Work ◽  
Suction Surface ◽  
Inlet Flow ◽  
Loss Distribution ◽  
Flow Angle ◽  
Radial Compressor

The unshrouded impeller and the vaneless diffuser of a single-stage radial compressor have been investigated at three flow rates. Three-dimensional velocities and pressures were measured at a tip speed of 84 m/s by an L2F-velocimeter, a slanted single hotwire probe, and piezoresistive pressure transducers. The measurements show that upstream of the blading the averaged meridional inlet flow angle is about 54 deg and a periodic variation of the meridional flow angle of about 25 deg occurs near the casing wall. Further, an inlet vortex in the clockwise direction appears and an initial whirl is induced. The specific work of the initial whirl corresponds to approximately 12 percent of the enthalpy losses between inlet pipe and diffuser outlet. In the beginning of the passage, the inlet vortex is suppressed and a solid body vortex in the counterclockwise direction can be observed. At the outlet, a heavy flow deceleration at the blade suction side with subsequent separation can be seen. Increasing the flow rate decreases the wake and causes a more uniform loss distribution in this area. The measured secondary vortex flow and rotary stagnation pressure gradients are compared with test results from impellers with inducer. The incidence of the investigated impeller is greater than that of the impellers with inducer, but the wake-jet outlet flows are very similar. Inlet losses could be reduced by improving incidence angles by matching the blade angles to the inlet flow angles. Smaller blade angles at the shroud would reduce or eliminate separation at the leading edge, and the resulting reduction in low-momentum fluid along the suction surface would help to avoid separation on that surface near the outlet.

Effect of Vaned Diffuser on a Modified Turbocharger Centrifugal Compressor Performance

2014 ◽  
Vol 663 ◽  
pp. 347-353
Author(s):  
Layth H. Jawad ◽  
Shahrir Abdullah ◽  
Zulkifli R. ◽  
Wan Mohd Faizal Wan Mahmood
Keyword(s):  
Centrifugal Compressor ◽  
Numerical Study ◽  
Pressure Ratio ◽  
Three Dimensional ◽  
Aerodynamic Characteristics ◽  
Width Ratio ◽  
Outlet Section ◽  
Suction Surface ◽  
Vaned Diffuser ◽  
Minimum Width

A numerical study that was made in a three-dimensional flow, carried out in a modified centrifugal compressor, having vaned diffuser stage, used as an automotive turbo charger. In order to study the influence of vaned diffuser meridional outlet section with a different width ratio of the modified centrifugal compressor. Moreover, the performance of the centrifugal compressor was dependent on the proper matching between the compressor impeller along the vaned diffuser. The aerodynamic characteristics were compared under different meridional width ratio. In addition, the velocity vectors in diffuser flow passages, and the secondary flow in cross-section near the outlet of diffuser were analysed in detail under different meridional width ratio. Another aim of this research was to study and simulate the effect of vaned diffuser on the performance of a centrifugal compressor. The simulation was undertaken using commercial software so-called ANSYS CFX, to predict numerically the performance charachteristics. The results were generated from CFD and were analysed for better understanding of the fluid flow through centrifugal compressor stage and as a result of the minimum width ratio the flow in diffuser passage tends to be uniformity. Moreover, the backflow and vortex near the pressure surface disappear, and the vortex and detachment near the suction surface decrease. Conclusively, it was observed that the efficiency was increased and both the total pressure ratio and static pressure for minimum width ratio are increased.

Experimental Studies of Leading Edge Contouring Influence on Secondary Losses in Transonic Turbines

2012 ◽  
Author(s):  
Ranjan Saha ◽  
Jens Fridh ◽  
Torsten Fransson ◽  
Boris I. Mamaev ◽  
Mats Annerfeldt
Keyword(s):  
Gas Turbine ◽  
Guide Vane ◽  
Experimental Studies ◽  
Three Dimensional ◽  
Leading Edge ◽  
Secondary Flows ◽  
Noticeable Effect ◽  
Flow Angle ◽  
Wide Range ◽  
Contour Plots

An experimental study of the hub leading edge contouring using fillets is performed in an annular sector cascade to observe the influence of secondary flows and aerodynamic losses. The investigated vane is a three dimensional gas turbine guide vane (geometrically similar) with a mid-span aspect ratio of 0.46. The measurements are carried out on the leading edge fillet and baseline cases using pneumatic probes. Significant precautions have been taken to increase the accuracy of the measurements. The investigations are performed for a wide range of operating exit Mach numbers from 0.5 to 0.9 at a design inlet flow angle of 90°. Data presented include the loading, fields of total pressures, exit flow angles, radial flow angles, as well as profile and secondary losses. The vane has a small profile loss of approximately 2.5% and secondary loss of about 1.1%. Contour plots of vorticity distributions and velocity vectors indicate there is a small influence of the vortex-structure in endwall regions when the leading edge fillet is used. Compared to the baseline case the loss for the filleted case is lower up to 13% of span and higher from 13% to 20% of the span for a reference condition with Mach no. of 0.9. For the filleted case, there is a small increase of turning up to 15% of the span and then a small decrease up to 35% of the span. Hence, there are no significant influences on the losses and turning for the filleted case. Results lead to the conclusion that one cannot expect a noticeable effect of leading edge contouring on the aerodynamic efficiency for the investigated 1st stage vane of a modern gas turbine.

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