A blade force-defect model in the casing wall boundary layer of an axial flow compressor.

1988 ◽  
Vol 54 (498) ◽  
pp. 430-434
Author(s):  
Masahiro INOUE ◽  
Minoru FUKUHARA ◽  
Motoo KUROUMARU
Keyword(s):  
Boundary Layer ◽  
Axial Flow ◽  
Defect Model ◽  
Axial Flow Compressor ◽  
Wall Boundary Layer ◽  
Wall Boundary ◽  
Flow Compressor

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.

scholarly journals Computation of Secondary Flows in an Axial Flow Compressor Including Inviscid and Viscous Flow Computation

1984 ◽  
Author(s):  
Francis Leboeuf
Keyword(s):  
Boundary Layer ◽  
Viscous Flow ◽  
Axial Flow ◽  
Inviscid Flow ◽  
Secondary Flows ◽  
Computational Method ◽  
Flow Computation ◽  
Axial Flow Compressor ◽  
Flow Compressor ◽  
Highly Loaded

A computational method for secondary flows in a compressor has been extended to treat stalled flows. An integral equation is used which simulates the inviscid flow at the wall, under the viscous flow influence. We present comparisons with experimental results for a 2D stalled boundary layer, and for the secondary flow in a highly loaded stator of an axial flow compressor.

Highly-Loaded Low-Reaction Boundary Layer Suction Axial Flow Compressor

2007 ◽  
Author(s):  
Songtao Wang ◽  
Xiaoqing Qiang ◽  
Weichun Lin ◽  
Guotai Feng ◽  
Zhongqi Wang
Keyword(s):  
Boundary Layer ◽  
High Pressure ◽  
Pressure Ratio ◽  
Axial Flow ◽  
Design Concept ◽  
Axial Flow Compressor ◽  
Boundary Layer Suction ◽  
High Pressure Ratio ◽  
Flow Compressor ◽  
Highly Loaded

In order to design high pressure ratio and highly loaded axial flow compressor, a new design concept based on Highly-Loaded Low-Reaction and boundary layer suction was proposed in this paper. Then the concept’s characteristics were pointed out by comparing with the MIT’s boundary layer suction compressor. Also the application area of this design concept and its key technic were given out in this paper. Two applications were carried out in order to demonstrate the concept. The first application was to redesign a low speed duplication-stage axial flow compressor into a single stage. The second one was a feasibility analysis to decrease an 11 stage axial compressor’s stage count to 7 while not changing its aerodynamic performance. The analysis result showed that the new design concept is feasible and it can be used on high pressure stage of the aero-engine, compressor of ground gas turbine (except the transonic stage) and high total pressure ratio blower.

Effects of Stream Surface Inclination on Tip Leakage Flow Fields in Compressor Rotors

1998 ◽  
Vol 120 (4) ◽  
pp. 683-692 ◽  
Author(s):  
M. Furukawa ◽  
K. Saiki ◽  
K. Nagayoshi ◽  
M. Kuroumaru ◽  
M. Inoue
Keyword(s):  
Boundary Layer ◽  
Vortex Breakdown ◽  
Axial Flow ◽  
Flow Fields ◽  
Leakage Flow ◽  
Tip Leakage Flow ◽  
Tip Leakage Vortex ◽  
Tip Leakage ◽  
Wall Boundary Layer ◽  
Wall Boundary

Experimental and computational results of tip leakage flow fields in a diagonal flow rotor at the design flow rate are compared with those in an axial flow rotor. In the diagonal flow rotor, the casing and hub walls are inclined at 25 deg and 45 deg, respectively, to the axis of rotation, and the blade has airfoil sections with almost the same tip solidity as that of the axial flow rotor. It is found out that “breakdown” of the tip leakage vortex occurs at the aft part of the passage in the diagonal flow rotor. The “vortex breakdown” causes significant changes in the nature of the tip leakage vortex: disappearance of the vortex core, large expansion of the vortex, and appearance of low relative velocity region in the vortex. These changes result in a behavior of the tip leakage flow that is substantially different from that in the axial flow rotor: no rolling-up of the leakage vortex downstream of the rotor, disappearance of the casing pressure trough at the aft part of the rotor passage, large spread of the low-energy fluid due to the leakage flow, much larger growth of the casing wall boundary layer, and considerable increase in the absolute tangential velocity in the casing wall boundary layer. The vortex breakdown influences the overall performance, also: large reduction of efficiency with the tip clearance, and low level of noise.

Boundary-Layer Development on an Axial-Flow Compressor Stator Blade

1978 ◽  
Vol 100 (2) ◽  
pp. 287-292 ◽  
Author(s):  
R. L. Evans
Keyword(s):  
Boundary Layer ◽  
Axial Flow ◽  
Layer Growth ◽  
Ensemble Averaging ◽  
Mean Velocity ◽  
Axial Flow Compressor ◽  
Boundary Layer Development ◽  
Stator Blade ◽  
Flow Compressor ◽  
Layer Development

The boundary layer on an axial-flow compressor stator blade has been measured using an ensemble-averaging technique. Although the mean velocity profiles appear to indicate fully developed turbulent flow, ensemble-averaged instantaneous profiles show the boundary layer to be highly unsteady and transitional over much of the blade chord. At a given chordwise position, variations in boundary-layer thickness with time of up to 150 percent were recorded. When compared to boundary-layer development on a similar blade in a two-dimensional cascade the stator blade boundary-layer growth was found to be much greater. The results indicate that extreme caution should be used in attempting to predict blade boundary-layer development from cascade test results or steady calculation procedures.

An Experimental Study of the Compressor Rotor Blade Boundary Layer

1985 ◽  
Vol 107 (2) ◽  
pp. 364-372 ◽  
Author(s):  
M. Pouagare ◽  
J. M. Galmes ◽  
B. Lakshminarayana
Keyword(s):  
Boundary Layer ◽  
Rotor Blade ◽  
Radial Profile ◽  
Three Dimensional ◽  
Axial Flow ◽  
Tip Leakage Flow ◽  
Wall Boundary Layer ◽  
Compressor Rotor ◽  
Blade Tip ◽  
Flow Compressor

The three-dimensional turbulent boundary layer developing on a rotor blade of an axial flow compressor was measured using a minature “x” configuration hot-wire probe. The measurements were carried out at nine radial locations on both surfaces of the blade at various chordwise locations. The data derived includes streamwise and radial mean velocities and turbulence intensities. The validity of conventional velocity profiles such as the “power law profile” for the streamwise profile, and Mager and Eichelbrenner’s for the radial profile, is examined. A modification to Mager’s crossflow profile is proposed. Away from the blade tip, the streamwise component of the blade boundary layer seems to be mainly influenced by the streamwise pressure gradient. Near the tip of the blade, the behavior of the blade boundary layer is affected by the tip leakage flow and the annulus wall boundary layer. The “tangential blockage” due to the blade boundary layer is derived from the data. The profile losses are found to be less than that of an equivalent cascade, except in the tip region of the blade.

scholarly journals Increasing Surge Margin in an Axial Flow Compressor Using “End-Bend” Airfoils

1985 ◽  
Author(s):  
Y. J. Cai ◽  
Y. L. Zhong ◽  
L. H. Qian ◽  
S. Z. He ◽  
Q. H. Pang
Keyword(s):  
Boundary Layer ◽  
Axial Flow ◽  
Experimental Results ◽  
Axial Flow Compressor ◽  
Surge Margin ◽  
Flow Loss ◽  
Flow Compressor ◽  
Percent Efficiency

This paper concentrates solely on the effect of end bend airfoil on performance of an axial flow compressor. The purpose is to reduce flow loss in the region of endwall boundary layer and to enlarge surge margin of an axial flow compressor by employing end bend airfoil on the stator blades in the last stages of the compressor. The experimental results show that surge margin is increased by 4.9–5.2 percent, efficiency by 2.6–3.2 percent at operational speed n=1.0–1.1.

Sign in / Sign up

Export Citation Format

Share Document