Casing Wall Boundary-Layer Development Through an Isolated Compressor Rotor

1982 ◽  
Vol 104 (4) ◽  
pp. 805-817 ◽  
Author(s):  
I. H. Hunter ◽  
N. A. Cumpsty
Keyword(s):  
Boundary Layer ◽  
Large Scale ◽  
Outer Wall ◽  
Tip Clearance ◽  
Pressure Probe ◽  
Wall Boundary Layer ◽  
Compressor Rotor ◽  
Wall Boundary ◽  
Boundary Layer Development ◽  
Layer Development

Detailed measurements were made of the casing wall boundary layer development across a large-scale, low-speed axial compressor rotor blade row. An important feature of the work was the use of blading which allowed the tip clearance to be varied. A conventional pressure probe was used to obtain time-averaged measurements of the outer-wall boundary layer downstream of the rotor whilst a hot-wire anenometry technique yielded the three-dimensional, blade to blade structure of the flow. The downstream boundary layer was found to thicken as the rotor loading and blade-end clearance were increased, with fluid tending to accumulate towards the pressure side of the passage. By its pronounced effects upon wall boundary layer development, tip clearance had a deleterious effect upon the performance of the compressor.

Laser-Doppler Velocimeter Measurement of Annulus Wall Boundary Layer Development in a Compressor Rotor

1988 ◽  
Vol 110 (3) ◽  
pp. 377-385 ◽  
Author(s):  
B. Lakshminarayana ◽  
K. N. S. Murthy
Keyword(s):  
Boundary Layer ◽  
Laser Doppler ◽  
Laser Doppler Velocimeter ◽  
Leakage Flow ◽  
Detailed Measurement ◽  
Wall Boundary Layer ◽  
Compressor Rotor ◽  
Wall Boundary ◽  
Boundary Layer Development ◽  
Layer Development

Detailed measurement of the flow field in the tip region of a compressor rotor was carried out using a laser-Doppler velocimeter (LDV). The axial and tangential components of relative velocities were measured upstream, inside the passage, and at the exit of the rotor, up to about 20 percent of the blade span from the blade tip. In addition, the relative stagnation pressures were measured from a Kiel probe; static pressures were derived from this and from the LDV measurement. The data are interpreted to understand annulus-wall boundary layer development inside the rotor, leakage flow, and losses in the tip region. The annulus wall boundary layer is well behaved at the leading edge and far downstream of the rotor. But inside the passage, complex interactions between the leakage flow and the annulus-wall boundary layer result in unconventional profiles with wide deviations from models employed for analyses.

Investigation of the Tip Clearance Flow Inside and at the Exit of a Compressor Rotor Passage—Part I: Mean Velocity Field

1983 ◽  
Vol 105 (1) ◽  
pp. 1-12 ◽  
Author(s):  
A. Pandya ◽  
B. Lakshminarayana
Keyword(s):  
Boundary Layer ◽  
Tip Clearance ◽  
Mean Flow ◽  
Ensemble Averaging ◽  
Mean Velocity ◽  
Tip Clearance Flow ◽  
Wall Boundary Layer ◽  
Compressor Rotor ◽  
Wall Boundary ◽  
Clearance Flow

This paper reports on an experimental study of the nature of the tip clearance flow in a moderately loaded compressor rotor. The measurements reported were obtained using a stationary two-sensor, hot-wire probe in combination with an ensemble averaging technique. The flow field was surveyed at various radial locations and at ten axial locations, four of which were inside the blade passage in the clearance region and the remaining six outside the passage. Variations of the mean flow properties in the tangential and the radial directions at various axial locations were derived from the data. Variation of leakage velocity at different axial stations and the annulus-wall boundary layer profiles from passage-averaged mean velocities were also estimated. The results indicate that there exists a region of strong interaction of the leakage flow with the annulus-wall boundary layer at half-chord. The profiles are well-behaved beyond this point. The rotor exit flow is found to be uniform beyond 3/4 blade chord downstream of the rotor trailing edge.

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.

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.

The Influence of Tip Clearance, Stage Loading, and Wall Roughness on Compressor Casing Boundary Layer Development

1982 ◽  
Author(s):  
J. L. Bettner ◽  
C. Elrod
Keyword(s):  
Boundary Layer ◽  
Layer Growth ◽  
Tip Clearance ◽  
Experimental Program ◽  
Wall Roughness ◽  
Stall Margin ◽  
Boundary Layer Development ◽  
Blade Tip ◽  
Boundary Layer Growth ◽  
Layer Development

An experimental program was conducted in a highly loaded, single-stage, low-speed research compressor that featured variations in blade tip clearances, shroud wall roughness, and stage loading levels. The results showed that tip clearances and stage loading levels exerted a very strong influence on casing boundary layer growth. Shroud wall roughness not only showed some effect on casing boundary layer development, but also influenced stage stall margin. The results point out the need to understand the basic flow processes that occur if realistic mathematical models are to be developed so that casing boundary layer development and its influence on compressor stage performance can be accurately predicted.

Investigations of the Boundary Layer on a Highly Loaded Compressor Cascade With Wake-Induced Transition

2006 ◽  
Author(s):  
Jens Iseler ◽  
Lothar Hilgenfeld ◽  
Michael Pfitzner
Keyword(s):  
Boundary Layer ◽  
Flow Field ◽  
Large Scale ◽  
Quality Data ◽  
Experimental Investigations ◽  
Generation Process ◽  
Unsteady Boundary Layer ◽  
Compressor Cascade ◽  
Boundary Layer Development ◽  
Layer Development

The flow field through a turbomachinery compressor cascades is significantly affected by the unsteady flow originating from the upstream blade rows. The interaction is caused by the wakes from the upstream blades, which affect the properties of the boundary layer of the downstream blades. In addition, pressure fluctuations exist between upstream and downstream blades. These phenomenona play a significant role in the loss generation process on turbomachinery blades because it influences the onset of transition in the boundary layer and has the potential to suppress a boundary layer separation in some cases. Extensive experimental investigations have been performed at the Institute of Jet Propulsion in Neubiberg, where these effects where studied in detail. The measurements were performed on a large scale compressor cascade called V103-220. The chord length of l = 220 mm chosen allowed the unsteady boundary layer development to be studied in great detail and provided high quality data for this complex flow, which can be used for the validation of CFD codes. Unsteady CFD calculations were performed using the RANS-code TRACE developed at DLR Cologne. A modern variant of the Wilcox k-ω turbulence model in combination with a newly implemented transition model was used, allowing a better determination of multimode transition. A multiblock grid with an O-type grid around the blade and a boundary layer resolution of y+<1 was used. Experimental and numerical results confirm that wake passing has a large influence on the unsteady boundary layer development also in this compressor flow case. The premature forced transition is followed by a stable calmed region, which partially suppresses laminar separation due to its higher shear stress level and delays the onset of transition in the path between wakes. In addition, it was found that the leakage from two slots, which are opened in the rig when the wake generator device is installed, changes the flow field considerably. This effect is not fully reproduced by the CFD calculations. To study this effect in more detail, three-dimensional steady and unsteady CFD calculations were undertaken and are being continued.

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.

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