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.

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.

Design Point Variation of Three-Dimensional Loss and Deviation for Axial Compressor Middle Stages

1988 ◽  
Vol 110 (4) ◽  
pp. 426-433 ◽  
Author(s):  
W. B. Roberts ◽  
G. K. Serovy ◽  
D. M. Sandercock
Keyword(s):  
Boundary Layer ◽  
Layer Thickness ◽  
Boundary Layer Thickness ◽  
Three Dimensional ◽  
Axial Compressor ◽  
Tip Clearance ◽  
Flow Angle ◽  
Design Point ◽  
Blade Element Theory ◽  
The Difference

Three-dimensional spanwise pressure loss and flow angle deviation variations have been deduced from NASA, university, and industrial sources from middle-stage research compressors operating near design point. These variations are taken as the difference above or below that predicted by blade element theory at any spanwise location. It was observed that the magnitude of the three-dimensional loss and deviation in the endwall regions is affected by hub and casing boundary layer thickness, camber, solidity, and blade channel aspect ratio for stators and rotor hubs. Rotor tip variations were found to depend on casing boundary layer thickness and tip clearance. Simple design point loss models derived from these data can aid in the design of axial compressor middle stages.

Annulus Wall Boundary-Layer Measurements in a Four-Stage Compressor

1986 ◽  
Vol 108 (1) ◽  
pp. 2-6 ◽  
Author(s):  
N. A. Cumpsty
Keyword(s):  
Boundary Layer ◽  
Reynolds Number ◽  
Layer Thickness ◽  
Boundary Layers ◽  
Boundary Layer Thickness ◽  
Tip Clearance ◽  
The Other ◽  
Wall Boundary Layer ◽  
Multistage Compressors ◽  
Full Design

There are few available measurements of the boundary layers in multistage compressors when the repeating-stage condition is reached. These tests were performed in a small four-stage compressor; the flow was essentially incompressible and the Reynolds number based on blade chord was about 5 • 104. Two series of tests were performed; in one series the full design number of blades were installed, in the other series half the blades were removed to reduce the solidity and double the staggered spacing. Initially it was wished to examine the hypothesis proposed by Smith [1] that staggered spacing is a particularly important scaling parameter for boundary layer thickness; the results of these tests and those of Hunter and Cumpsty [2] tend to suggest that it is tip clearance which is most potent in determining boundary-layer integral thicknesses. The integral thicknesses agree quite well with those published by Smith.

A Review of Cascade Data on Secondary Losses in Turbines

1970 ◽  
Vol 12 (1) ◽  
pp. 48-59 ◽  
Author(s):  
J. Dunham
Keyword(s):  
Boundary Layer ◽  
Layer Thickness ◽  
Boundary Layers ◽  
Boundary Layer Thickness ◽  
Blade Loading ◽  
Axial Turbine ◽  
Wall Boundary Layer ◽  
Wall Boundary ◽  
Turbine Cascades

Theories and experiments on secondary losses in axial turbine cascades without end clearance are reviewed. A formula is given which correlates the effect of blade loading on secondary losses more successfully than hitherto. However, it is also shown that secondary losses increase with upstream wall boundary layer thickness. Only a tentative expression for that effect can be suggested. In order to predict secondary losses reliably more must be known about these wall boundary layers.

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.

Influence of the wall boundary layer thickness on a gas jet injected into a liquid crossflow

2001 ◽  
Vol 30 (4) ◽  
pp. 458-466 ◽  
Author(s):  
O. Vigneau ◽  
S. Pignoux ◽  
J. L. Carreau ◽  
F. Roger
Keyword(s):  
Boundary Layer ◽  
Layer Thickness ◽  
Boundary Layer Thickness ◽  
Gas Jet ◽  
Wall Boundary Layer ◽  
Wall Boundary
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