An Axial Compressor End-Wall Boundary Layer Calculation Method

1979 ◽  
Vol 101 (2) ◽  
pp. 233-245 ◽  
Author(s):  
J. De Ruyck ◽  
C. Hirsch ◽  
P. Kool
Keyword(s):  
Experimental Data ◽  
Boundary Layer ◽  
Velocity Profile ◽  
Three Dimensional ◽  
Axial Compressor ◽  
Tip Clearance ◽  
Wall Region ◽  
Wall Boundary Layer ◽  
Wall Boundary ◽  
End Wall

An axial compressor end-wall boundary layer theory which requires the introduction of three-dimensional velocity profile models is described. The method is based on pitch-averaged boundary layer equations and contains blade force-defect terms for which a new expression in function of transverse momentum thickness is introduced. In presence of tip clearance a component of the defect force proportional to the clearance over blade height ratio is also introduced. In this way two constants enter the model. It is also shown that all three-dimensional velocity profile models present inherent limitations with regard to the range of boundary layer momentum thicknesses they are able to represent. Therefore a new heuristic velocity profile model is introduced, giving higher flexibility. The end-wall boundary layer calculation allows a correction of the efficiency due to end-wall losses as well as calculation of blockage. The two constants entering the model are calibrated and compared with experimental data allowing a good prediction of overall efficiency including clearance effects and aspect ratio. Besides, the method allows a prediction of radial distribution of velocities and flow angles including the end-wall region and examples are shown compared to experimental data.

Investigations of an Axial Compressor End-Wall Boundary Layer Prediction Method

1981 ◽  
Vol 103 (1) ◽  
pp. 20-33 ◽  
Author(s):  
J. De Ruyck ◽  
C. Hirsch
Keyword(s):  
Boundary Layer ◽  
Velocity Profile ◽  
Boundary Layers ◽  
Three Dimensional ◽  
Axial Compressor ◽  
Diffusion Factor ◽  
Profile Model ◽  
Wall Boundary Layer ◽  
Wall Boundary ◽  
End Wall

A previously developed axial compressor end-wall boundary layer calculation method which requires the introduction of three-dimensional velocity profile models is summarized. In this method the classical three-dimensional velocity profile models were shown to present inherent limitations at stall limit, with regard to the range of transverse boundary layer thicknesses they are able to represent. A corrected profile model is presented which contains no more limitations without affecting the previous found overall results. Stall limit is predicted by limiting values of shape factor and/or diffusion factor. The new profile model containing also compressibility effects allows the calculation of boundary layers in machines with shrouded blades, by simulating the jump between rotating and non rotating parts of the walls. A corrected version of a force defect correlation is presented which is shown to give better agreement at high incidences. Some results on high and low speed machines are discussed. The model is applied to obtain an end-wall blockage correlation depending on geometry, flow coefficient, AVR, aspect ratio, solidity, diffusion factor, Reynolds number, axial blade spacing, tip clearance and inlet boundary layer thickness. A quantitative estimation of the losses associated with the end-wall boundary layers can be obtained using this analysis and therefore can be a useful tool in the design of an axial compressor stage.

Modeling the 3-D Flow Effects on Deviation Angle for Axial Compressor Middle Stages

1986 ◽  
Vol 108 (1) ◽  
pp. 131-137 ◽  
Author(s):  
W. B. Roberts ◽  
G. K. Serovy ◽  
D. M. Sandercock
Keyword(s):  
Boundary Layer ◽  
Layer Thickness ◽  
Boundary Layer Thickness ◽  
Axial Compressor ◽  
Tip Clearance ◽  
Blade Element Theory ◽  
Wall Boundary Layer ◽  
Wall Boundary ◽  
Flow Effects ◽  
End Wall

A model of the spanwise variation of the 3-D flow effects on deviation is proposed for middle-stage rotors and stators. This variation is taken as the difference above or below that predicted by blade element theory at any spanwise location. It was found that the stator variation is strongly affected by the end-wall boundary-layer thickness as well as camber, solidity, and blade channel aspect ratio. Rotor variation was found to depend on end-wall boundary layer thickness and tip clearance normalized by blade span. If these parameters are known or can be calculated, the models provide a reasonable approximation to the spanwise variation of deviation for middle compressor stages operating at low to high subsonic inlet Mach numbers.

Wall Boundary Layer Development Near the Tip Region of an IGV of an Axial Flow Compressor

1984 ◽  
Vol 106 (2) ◽  
pp. 337-345
Author(s):  
B. Lakshminarayana ◽  
N. Sitaram
Keyword(s):  
Boundary Layer ◽  
Guide Vane ◽  
Three Dimensional ◽  
Axial Compressor ◽  
Axial Flow ◽  
Inlet Guide Vane ◽  
Wall Boundary Layer ◽  
Wall Boundary ◽  
Boundary Layer Development ◽  
Flow Compressor

The annulus wall boundary layer inside the blade passage of the inlet guide vane (IGV) passage of a low-speed axial compressor stage was measured with a miniature five-hole probe. The three-dimensional velocity and pressure fields were measured at various axial and tangential locations. Limiting streamline angles and static pressures were also measured on the casing of the IGV passage. Strong secondary vorticity was developed. The data were analyzed and correlated with the existing velocity profile correlations. The end wall losses were also derived from these data.

Influence of Boundary Layer Skew on the Tip Leakage Vortex of an Axial Compressor Stator

2021 ◽  
pp. 1-19
Author(s):  
Björn Koppe ◽  
Martin Lange ◽  
Ronald Mailach
Keyword(s):  
Boundary Layer ◽  
Axial Compressor ◽  
Tip Clearance ◽  
Major Influence ◽  
Wall Region ◽  
Tip Leakage Vortex ◽  
Tip Leakage ◽  
Flow Phenomena ◽  
Gap Height ◽  
End Wall

Abstract For an axial compressor stator with tip gap the boundary layer in the hub end-wall region has a significant influence on the development and progression of the tip leakage vortex. Herein the so-called boundary layer skew, which develops through relative motion of the hub, is of particular interest. Therefore, experimental and numerical investigations of a single axial compressor stator row with varying tip gap height (tip gap height/chord length = 2.0%|5.4%|6.7%) have been conducted. Comparing cases with rotating or stationary hub end-wall segments upstream of the examined vanes allowed to determine the effect of skewed and un-skewed inflow boundary layer. The steady state flow fields up- and downstream of the stator row were measured using five-hole pressure probes. For validation and to improve the understanding of the existing flow phenomena 3D-RANS CFD simulations using a commercial flow solver were carried out. Furthermore, analog cases with no tip gap were examined and considered in the comparisons to extend the knowledge on this boundary layer characteristic. The results show that the boundary layer skew has a major influence on the trajectory and size of the tip leakage vortex for the cases with tip clearance. The effect of reduction of the produced losses decreases with increasing tip gap height.

Treatment of the Annulus Wall Boundary Layer Using a Secondary Flow Hypothesis

1977 ◽  
Vol 99 (1) ◽  
pp. 29-36 ◽  
Author(s):  
J. W. Railly ◽  
P. B. Sharma
Keyword(s):  
Boundary Layer ◽  
Secondary Flow ◽  
Wire Array ◽  
Three Dimensional ◽  
Tip Clearance ◽  
Collateral Flow ◽  
Axially Symmetric ◽  
Wall Boundary Layer ◽  
Wall Boundary ◽  
Blade Row

Hitherto, theories of annulus wall boundary layer development in axial compressors have assumed an axially-symmetric flow in which the blade action has been replaced by a force field. A more rigorous treatment of the momentum equations in the annulus boundary layer by Mellor and Wood demonstrated the presence of certain terms, after the equations had been averaged in the pitch-wise direction, which arise from the truly three-dimensional character of the flow. These terms, which may be described as the gradients of apparent stresses, were not regarded by them (apart from a discussion of tip clearance) as having importance for the problem. In the present work a second equation of the annulus wall boundary layer is obtained by consideration of the work of these apparent stresses. By integration of the system of equations over a single blade row, two equations are obtained relating various integral quantities at inlet to and exit from the row. Each equation contains terms which depend upon apparent stresses connected with the relative velocity field at the exit plane. An experiment is described in which the six turbulent stresses in the stationary frame downstream of a single rotor, determined by means of a multiple hot wire array, are used to evaluate each term of the aforementioned equations. The integral quantities thus determined are shown to be reasonably consistent with the predictions from the two equations, in particular, for the case of the hub boundary layer. Theoretical solutions of the two integral equations require a secondary flow hypothesis so that the departure from collateral flow at blade row exit is determined by the solution.

An Axial Compressor End-Wall Boundary Layer Theory

1971 ◽  
Vol 93 (2) ◽  
pp. 300-314 ◽  
Author(s):  
G. L. Mellor ◽  
G. M. Wood
Keyword(s):  
Boundary Layer ◽  
Boundary Layers ◽  
Axial Compressor ◽  
Jump Condition ◽  
Present Theory ◽  
Boundary Layer Theory ◽  
Tip Clearance ◽  
Main Stream ◽  
Wall Boundary ◽  
End Wall

The essential ingredient missing in existing prediction methods for the performance of multistage axial compressors is that which would account for the effect of end-wall boundary layers. It is, in fact, believed that end-wall boundary layers play a major role in compressor performance and the absence of an adequate theory represents a handicap to turbomachinery designers that might be likened to the handicap that designers of wings, for example, would face if Prandtl had not introduced the idea of a boundary layer. In this paper a new theory is developed which retains all elements of classical boundary layer theory; for example, we discuss variables such as momentum thickness and wall shear stress. However, the present theory introduces new concepts such as axial and tangential defect force thickness, a rotor exit-stator inlet “jump condition” and the importance of these concepts is demonstrated. Inherent in the derivation is an identification of the role of secondary flow and tip clearance flow. A proper means of matching the boundary layer calculations to conventional main stream calculations is suggested. Independent of empirical parametization it appears that the theory is capable of correctly modeling boundary layer blockage, losses, and end-wall stall. Near stall, the main stream-boundary layer interaction is very strong.

Modelling of spanwise mixing in compressor through-flow computations

1997 ◽  
Vol 211 (3) ◽  
pp. 243-251 ◽  
Author(s):  
J Dunham
Keyword(s):  
Boundary Layer ◽  
Three Dimensional ◽  
Axial Compressor ◽  
Navier Stokes ◽  
Streamline Curvature ◽  
Wall Boundary Layer ◽  
Through Flow ◽  
Compressor Design ◽  
Multistage Compressors ◽  
End Wall

Although three-dimensional Navier-Stokes computations are coming into use more and more, streamline curvature through-flow computations are still needed, especially for multistage compressors, and where codes which run in minutes rather than hours are preferred. These methods have been made more realistic by taking account of end-wall effects and spanwise mixing by four aerodynamic mechanisms: turbulent diffusion, turbulent convection by secondary flow, spanwise migration of aerofoil boundary layer fluid and spanwise convection of fluid in blade wakes. This paper describes the models adopted in the DRA streamline curvature method for axial compressor design and analysis. Previous papers are summarized briefly before describing the new part of the model—that accounting for aerofoil boundary layers and wakes. Other changes to the previously published annulus wall boundary layer model have been made to enable it to cater for separations and end bends. The resulting code is evaluated against a range of experimental and computational results.

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