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.

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.

Three-Dimensional Turbulent Flow in the Tip Region of an Axial Compressor Rotor Passage at a Near Stall Condition

2001 ◽  
Author(s):  
Hongwei Ma ◽  
Haokang Jiang
Keyword(s):  
Turbulent Flow ◽  
Large Scale ◽  
Three Dimensional ◽  
Axial Flow ◽  
Leading Edge ◽  
Rotating Stall ◽  
Suction Surface ◽  
Tip Leakage Vortex ◽  
Tip Leakage ◽  
Compressor Rotor

This paper presents an experimental study of the three-dimensional turbulent flow field in the tip region of an axial flow compressor rotor passage at a near stall condition. The investigation was conducted in a low-speed large-scale compressor using a 3-component Laser Doppler Velocimetry and a high frequency pressure transducer. The measurement results indicate that a tip leakage vortex is produced very close to the leading edge, and becomes the strongest at about 10% axial chord from the leading edge. Breakdown of the vortex periodically occurs at about 1/3 chord, causing very strong turbulence in the radial direction. Flow separation happens on the tip suction surface at about half chord, prompting the corner vortex migrating toward the pressure side. Tangential migration of the low-energy fluids results in substantial flow blockage and turbulence in the rear of a rotor passage. Unsteady interactions among the tip leakage vortex, the separated vortex and the corner flow should contribute to the inception of the rotating stall in a compressor.

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.

Mass Flow Limitation of Supersonic Blade Rows due to Leading Edge Blockage

1984 ◽  
Author(s):  
Hans Starken ◽  
Zhong Yongxing ◽  
Heinz-Adolf Schreiber
Keyword(s):  
Mach Number ◽  
Axial Flow ◽  
Leading Edge ◽  
Incidence Relation ◽  
Suction Surface ◽  
Flow Angle ◽  
Blade Row ◽  
Flow Compressor ◽  
Theoretical Results ◽  
Surface Angle

The supersonic flow passing through the blade row of an axial flow compressor depends on the magnitude of the axial inlet Mach number. If the upstream stream tube is convergent, the axial Mach number remains subsonic and the unique incidence relation holds between inlet Mach number and inlet flow angle. This paper presents the influence of leading edge thickness and suction surface angle on the unique incidence and choking of supersonic blade rows. The theoretical results were experimentally verified in supersonic cascade wind tunnel tests.

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.

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.

Evaluation of RANS and ZDES Methods for the Prediction of Three-Dimensional Separation in Axial Flow Compressors

2015 ◽  
Author(s):  
Ashley D. Scillitoe ◽  
Paul G. Tucker ◽  
Paolo Adami
Keyword(s):  
Pressure Distribution ◽  
Separation Bubble ◽  
Computational Cost ◽  
Three Dimensional ◽  
Axial Flow ◽  
Small Time ◽  
Suction Surface ◽  
Flow Angle ◽  
Corner Separation ◽  
Displacement Thickness

Regions of three-dimensional separations are an inherent flow feature of the corner formed by the suction surface and endwall of axial compressors. RANS turbulence models, common in industrial CFD codes, often struggle in these regions. This paper investigates the use of two hybrid RANS/LES methods as alternatives to pure RANS methods. SA and SST based Zonal DES (ZDES) are applied to a linear blade cascade case, studied experimentally by Gbadebo [1]. The time-averaged results are compared to steady SA, SST and RSM RANS results. SA model corrections for streamline curvature, anisotropy and non-equilibrium effects are also examined. For the ZDES computations the solver is modified to reduce dissipation at low Mach numbers. Significant uncertainty is observed in the RANS results, with the origin of the suction surface corner separation occurring too far upstream, and the extent of the corner separation significantly over-predicted. The laminar separation bubble and the turbulent reattachment are also missed. Consequently the surface pressure distribution, exit flow angle and total pressure loss predictions are poor. Conversely, the ZDES results were encouraging; with much better predictions of the pressure distribution, exit flow angle and trailing edge boundary layer displacement thickness. Some RANS corrections proved effective, such as the SA model with Rotation/Curvature correction (SA-RC), however all had deficiencies in some areas. Although the ZDES results are encouraging it is noted that these computations were two orders of magnitude more computaionally expensive due to the high mesh densities and small time-steps required. For the ZDES results quality indexes are examined in order to determine whether the computational mesh used is sufficient in different flow regions. Mesh generation strategies based on using a pre-cursor RANS solution to obtain a modelled energy spectrum and various turbulent length scales to guide mesh refinement are considered. These can provide a quick estimate of the potential computational cost of LES or hybrid RANS/LES computations from a RANS solution.

Three-Dimensional Flows and Loss Reduction in Axial Compressors

1987 ◽  
Vol 109 (3) ◽  
pp. 354-361 ◽  
Author(s):  
Y. Dong ◽  
S. J. Gallimore ◽  
H. P. Hodson
Keyword(s):  
Three Dimensional ◽  
Axial Compressor ◽  
Axial Flow ◽  
Tracer Gas ◽  
Suction Surface ◽  
Axial Compressors ◽  
Small Clearance ◽  
Oil Flow ◽  
Stator Blade ◽  
Flow Compressor

Measurements have been performed in a low-speed high-reaction single-stage axial compressor. Data obtained within and downstream of the rotor, when correlated with the results of other investigations, provide a link between the existence of suction surface–hub corner separations, their associated loss mechanisms, and blade loading. Within the stator, it has been shown that introducing a small clearance between the stator blade and the stationary hub increases the efficiency of the stator compared to the case with no clearance. Oil flow visualizaton indicated that the leakage reduced the extensive suction surface–hub corner separation that would otherwise exist. A tracer gas experiment showed that the large radial shifts of the surface streamlines indicated by the oil flow technique were only present close to the blade. The investigation demonstrates the possible advantages of including hub clearance in axial flow compressor stator blade rows.

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.

The Impact of Geometric Variation on Compressor Two-Dimensional Incidence Range

2014 ◽  
Vol 137 (2) ◽  
Author(s):  
Martin N. Goodhand ◽  
Robert J. Miller ◽  
Hang W. Lung
Keyword(s):  
Design Process ◽  
Design Space ◽  
Leading Edge ◽  
Critical Value ◽  
Two Dimensional ◽  
Suction Surface ◽  
Design Intent ◽  
Manufacture Process ◽  
The Impact ◽  
Geometric Variation

An important question for a designer is how, in the design process, to deal with the small geometric variations which result from either the manufacture process or in-service deterioration. For some blade designs geometric variations will have little or no effect on the performance of a row of blades, while in others their effects can be significant. This paper shows that blade designs which are most sensitive are those which are susceptible to a distinct switch in the fluid mechanisms responsible for limiting blade performance. To demonstrate this principle, the sensitivity of compressor 2D incidence range to manufacture variations is considered. Only one switch in mechanisms was observed, the onset of flow separation at the leading edge. This switch is only sensitive to geometric variations around the leading edge, 0–3% of the suction surface. The consequence for these manufacture variations was a 10% reduction in the blade's positive incidence range. For this switch, the boundary in the design space is best defined in terms of the blade pressure distribution. Blade designs where the acceleration exceeds a critical value just downstream of the leading edge are shown to be robust to geometric variation. Two historic designs, supercritical blades and blades with sharp leading edges, though superior in design intent, are shown to sit outside this robust region and thus, in practice, perform worse. The improved understanding of the robust, region of the design space is then used to design a blade capable of a robust, 5% increase in operating incidence range.

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