Endwall Losses and Flow Unsteadiness in a Turbine Blade Cascade

1992 ◽  
Vol 114 (1) ◽  
pp. 191-197 ◽  
Author(s):  
L. Adjlout ◽  
S. L. Dixon
Keyword(s):  
Boundary Layer ◽  
Flow Visualization ◽  
Turbine Blade ◽  
Turbulence Intensity ◽  
Pressure Loss ◽  
Total Pressure ◽  
Free Stream ◽  
Total Pressure Loss ◽  
Pressure Distributions ◽  
Blade Cascade

The purpose of this paper is to describe an investigation of the flow within and downstream of a turbine blade cascade of high aspect ratio. A detailed experimental investigation into the changes in the endwall boundary layer in the cascade (100 deg camber angle) and total pressure loss downstream of the cascade was carried out. Flow visualization was used in order to obtain detailed photographs of the flow patterns on the endwall and for exhibiting the trailing edge vortices. Pressure measurements were carried out using a miniature cranked Kiel probe for three planes downstream of the cascade, with two levels of turbulence intensity of the free stream. Pressure distributions on the blade were measured at three spanwise locations, namely 4, 12, and 50 percent of the full span from the wall. Hot-wire anenometry combined with a spectrum analyzer program was used to determine the frequencies of the flow oscillations. The change in turbulence level of the free stream has a significant influence on all three pressure distributions. The striking difference between two of the pressure distributions is in the aft half of the suction side where the distribution with the lower turbulence intensity has the larger lift. The oil flow visualization reveals what appear to be two separation lines within the passage and are believed to originate from the horseshoe vortex. The pitchwise-averaged total pressure loss coefficient increases with the distance of the measurement plane downstream of the cascade blades. A substantial part of this loss increase close to the wall is caused by the high rate of shear of the new boundary layer on the endwall.

Endwall Losses and Flow Unsteadiness in a Turbine Blade Cascade

1990 ◽  
Author(s):  
L. Adjlout ◽  
S. L. Dixon
Keyword(s):  
Boundary Layer ◽  
Turbine Blade ◽  
Turbulence Intensity ◽  
Pressure Loss ◽  
Total Pressure ◽  
Free Stream ◽  
Total Pressure Loss ◽  
Flow Visualisation ◽  
Pressure Distributions ◽  
Blade Cascade

The purpose of this paper is to describe an investigation of the flow within and downstream of a turbine blade cascade of high aspect ratio. A detailed experimental investigation into the changes in the endwall boundary layer in the cascade (100deg camber angle) and total pressure loss downstream of the cascade was carried out. Flow visualisation was used in order to obtain detailed photographs of the flow patterns on the endwall and for exhibiting the trailing edge vortices. Pressure measurements were carried out using a miniature cranked Kiel probe for three planes downstream of the cascade, with two levels of turbulence intensity of the free-stream. Pressure distribution on the blade were measured at three spanwise locations, namely 4%, 12%, and 50% of the full-span from the wall. Hot wire anenometry combined with a spectrum analyser program was used to determine the frequencies of the flow oscillations. The change in turbulence level of the free stream has a significant influence on all three pressure distributions. The striking difference between two of the pressure distributions is in the aft half of the suction side where the distribution with the lower turbulence intensity has the larger lift. The oil flow visualisation reveals what appears to be two separation lines within the passage and are believed to originate from the horseshoe vortex. The pitchwise-averaged total pressure loss coefficient increases with the distance of the measurement plane downstream of the cascade blades. A substantial part of this loss increase close to the wall is caused by the high rate of shear of the new boundary layer on the endwall.

Location Effect of Boundary Layer Suction on Compressor Hub-Corner Separation

2014 ◽  
Author(s):  
Ping-Ping Chen ◽  
Wei-Yang Qiao ◽  
Karsten Liesner ◽  
Robert Meyer
Keyword(s):  
Boundary Layer ◽  
Pressure Loss ◽  
Total Pressure ◽  
Three Dimensional ◽  
Base Flow ◽  
Total Pressure Loss ◽  
Compressor Cascade ◽  
Suction Surface ◽  
Corner Separation ◽  
Boundary Layer Suction

The large secondary flow area in the compressor hub-corner region usually leads to three-dimensional separation in the passage with large amounts of total pressure loss. In this paper numerical simulations of a linear high-speed compressor cascade, consisting of five NACA 65-K48 stator profiles, were performed to analyze the flow mechanism of hub-corner separation for the base flow. Experimental validation is used to verify the numerical results. Active control of the hub-corner separation was investigated by using boundary layer suction. The influence of the selected locations of the endwall suction slot was investigated in an effort to quantify the gains of the compressor cascade performance. The results show that the optimal chordwise location should contain the development section of the three-dimensional corner separation downstream of the 3D corner separation onset. The best pitchwise location should be close enough to the vanes’ suction surface. Therefore the optimal endwall suction location is the MTE slot, the one from 50% to 75% chord at the hub, close to the blade suction surface. By use of the MTE slot with 1% suction flow ratio, the total-pressure loss is substantially decreased by about 15.2% in the CFD calculations and 9.7% in the measurement at the design operating condition.

Influence of a Novel 3D Leading Edge Geometry on the Aerodynamic Performance of Low Pressure Turbine Blade Cascade Vanes

2014 ◽  
Author(s):  
A. Asghar ◽  
W. D. E. Allan ◽  
M. LaViolette ◽  
R. Woodason
Keyword(s):  
Turbine Blade ◽  
Pressure Loss ◽  
Total Pressure ◽  
Separated Flow ◽  
Leading Edge ◽  
Low Pressure ◽  
Aerodynamic Performance ◽  
Total Pressure Loss ◽  
Low Pressure Turbine ◽  
Edge Geometry

This paper addresses the issue of aerodynamic performance of a novel 3D leading edge modification to a reference low pressure turbine blade. An analysis of tubercles found in nature and used in some engineering applications was employed to synthesize new leading edge geometry. A sinusoidal wave-like geometry characterized by wavelength and amplitude was used to modify the leading edge along the span of a 2D profile, rendering a 3D blade shape. The rationale behind using the sinusoidal leading edge was that they induce streamwise vortices at the leading edge which influence the separation behaviour downstream. Surface pressure and total pressure measurements were made in experiments on a cascade rig. These were complemented with computational fluid dynamics studies where flow visualization was also made from numerical results. The tests were carried out at low Reynolds number of 5.5 × 104 on a well-researched profile representative of conventional low pressure turbine profiles. The performance of the new 3D leading edge geometries was compared against the reference blade revealing a downstream shift in separated flow for the LE tubercle blades; however, total pressure loss reduction was not conclusively substantiated for the blade with leading edge tubercles when compared with the performance of the baseline blade. Factors contributing to the total pressure loss are discussed.

Effect of Turbine Blade tip shape on the Total Pressure Loss of a Turbine Cascade

2009 ◽  
Vol 12 (2) ◽  
pp. 39-45 ◽  
Author(s):  
Ki-Seon Lee ◽  
Seoung-Duck Park ◽  
Young-Chul Noh ◽  
Hak-Bong Kim ◽  
Jae-Su Kwak ◽  
...  
Keyword(s):  
Turbine Blade ◽  
Pressure Loss ◽  
Total Pressure ◽  
Total Pressure Loss ◽  
Turbine Cascade ◽  
Blade Tip

Accurate Estimation of Profile Losses and Analysis of Loss Generation Mechanisms in a Turbine Cascade

2017 ◽  
Vol 139 (12) ◽  
Author(s):  
D. Lengani ◽  
D. Simoni ◽  
M. Ubaldi ◽  
P. Zunino ◽  
F. Bertini ◽  
...  
Keyword(s):  
Boundary Layer ◽  
Pressure Loss ◽  
Total Pressure ◽  
Loss Coefficient ◽  
Total Pressure Loss ◽  
Accurate Estimation ◽  
Loss Mechanism ◽  
Pressure Loss Coefficient ◽  
Generation Mechanisms ◽  
Total Pressure Loss Coefficient

The paper analyzes losses and the loss generation mechanisms in a low-pressure turbine (LPT) cascade by proper orthogonal decomposition (POD) applied to measurements. Total pressure probes and time-resolved particle image velocimetry (TR-PIV) are used to determine the flow field and performance of the blade with steady and unsteady inflow conditions varying the flow incidence. The total pressure loss coefficient is computed by traversing two Kiel probes upstream and downstream of the cascade simultaneously. This procedure allows a very accurate estimation of the total pressure loss coefficient also in the potential flow region affected by incoming wake migration. The TR-PIV investigation concentrates on the aft portion of the suction side boundary layer downstream of peak suction. In this adverse pressure gradient region, the interaction between the wake and the boundary layer is the strongest, and it leads to the largest deviation from a steady loss mechanism. POD applied to this portion of the domain provides a statistical representation of the flow oscillations by splitting the effects induced by the different dynamics. The paper also describes how POD can dissect the loss generation mechanisms by separating the contributions to the Reynolds stress tensor from the different modes. The steady condition loss generation, driven by boundary layer streaks and separation, is augmented in the presence of incoming wakes by the wake–boundary layer interaction and by the wake dilation mechanism. Wake migration losses have been found to be almost insensitive to incidence variation between nominal and negative (up to −9 deg) while at positive incidence, the losses have a steep increase due to the alteration of the wake path induced by the different loading distribution.

Impact of a Combination of Micro-Vortex Generator and Boundary Layer Suction on Performance in a High-Load Compressor Cascade

2018 ◽  
Author(s):  
Shan Ma ◽  
Wuli Chu ◽  
Haoguang Zhang ◽  
Chuanle Liu
Keyword(s):  
Boundary Layer ◽  
Pressure Loss ◽  
Total Pressure ◽  
Vortex Generator ◽  
Aerodynamic Performance ◽  
Total Pressure Loss ◽  
Control Methods ◽  
Compressor Cascade ◽  
Boundary Layer Suction ◽  
Micro Vortex Generator

The performance of a compressor cascade is considerably influenced by flow control methods. In this paper, the synergistic effects of combination between micro-vortex generators (MVG) and boundary layer suction (BLS) are discussed in a high-load compressor cascade. Seven cases, which are grouped by a kind of micro-vortex generator and boundary layer suction with three locations, are investigated to control secondary flow effects and enhance the aerodynamic performance of the compressor cascade. The MVG is mounted on the end-wall in front of the passage. The rectangle suction slot with three radial positions is installed on the blade suction surface near the trailing edge. The numerical results show that: at the design condition, the total pressure loss is effectively decreased as well as the static pressure coefficient increase when the combined MVG and SBL method (COM) is used, which is superior to MVG in an aerodynamic performance. At the stall condition, the induced vortex coming from MVG could mix the low-energy fluid and mainstream, which result in the reduced separation, and the total pressure loss decreased by 11.54% when the suction flow ratio is 1.5%. The total pressure loss decreases by 14.59% when the COM control methods are applied.

A Comprehensive Investigation of Blade Row Interaction Effects on Stator Loss Utilizing Vane Clocking

2017 ◽  
Author(s):  
Natalie R. Smith ◽  
Nicole L. Key
Keyword(s):  
Boundary Layer ◽  
Flow Field ◽  
Pressure Loss ◽  
Total Pressure ◽  
Boundary Layer Transition ◽  
Total Pressure Loss ◽  
Tip Leakage Flow ◽  
Inlet Flow ◽  
Blade Row ◽  
Vane Clocking

Blade row interactions drive the unsteady performance of high pressure compressors. Vane clocking is the relative circumferential positioning of consecutive stationary vane rows with the same vane count. By altering the upstream vane wake’s path with respect to the downstream vane, vane clocking changes the blade row interactions and results in a change in steady total pressure loss on the downstream vane. The open literature lacks a conclusive discussion of the flow physics governing these interactions in compressors. This paper presents the details of a comprehensive vane clocking study on the embedded stage of the Purdue 3-stage axial compressor. The steady loss results, including radial total pressure profiles and surface flow visualization, suggest a shift in the Stator 2 corner separations occurs between clocking configurations associated with the maximum and minimum total pressure loss. To better understand the flow mechanisms driving the vane clocking effects on the steady Stator 2 performance, time-resolved interrogations of the Stator 2 inlet flow field, surface pressure unsteadiness, and boundary layer response were conducted. The Stator 2 surface flows, both pressure unsteadiness and boundary layer transition, are influenced by vane clocking and interactions between Rotor 1 and Rotor 2, but neither of these results indicate a cause for the change in steady total pressure loss. Moreover, they are a result of upstream changes in the flow field: the interaction between the Stator 1 wake and Rotor 2 results in a circumferentially varying pattern which alters the inlet flow field for the downstream row, including the unsteadiness and frequency content in the tip and hub regions. Therefore, under different clocking configurations, Stator 2 experiences significantly different inlet blockage and unsteadiness from the Rotor 2 tip leakage flow and hub corner separation, which, in turn, shifts the radial blade loading distribution and subsequent loss development of Stator 2.

scholarly journals The Effect of Inlet Boundary Layer Thickness on the Flow Within an Annular S-Shaped Duct

1998 ◽  
Author(s):  
Toyotaka Sonoda ◽  
Toshiyuki Arima ◽  
Mineyasu Oana
Keyword(s):  
Boundary Layer ◽  
Flow Pattern ◽  
Pressure Loss ◽  
Total Pressure ◽  
Three Dimensional ◽  
Flow Characteristics ◽  
Vortex Pair ◽  
Separated Flow ◽  
Total Pressure Loss ◽  
Centrifugal Impeller

Experimental and numerical investigations were carried out to gain a better understanding of the flow characteristics within an annular S-shaped duct, including the effect of the inlet boundary layer (IBL) on the flow. A duct with six struts and the same geometry as that used to connect compressor spools on our experimental small two-spool turbofan engine was investigated. A curved downstream annular passage with a similar meridional flow path geometry to that of the centrifugal compressor has been fitted at the exit of S-shaped duct. Two types of the IBL (i.e. thin and thick IBL) were used. Results showed that large differences of flow pattern were observed at the S-Shaped duct exit between two types of the IBL, though the value of “net” total pressure loss has not been remarkably changed. According to “overall” total pressure loss, which includes the IBL loss, the total pressure loss was greatly increased near the hub as compared to that for a thin one. For the thick IBL, a vortex pair related to the hub-side horseshoe vortex and the separated flow found at the strut trailing edge has been clearly captured in the form of the total pressure loss contours and secondary flow vectors, experimentally and numerically. The high-pressure loss regions on either side of the strut wake near the hub may act on a downstream compressor as a large inlet distortion, and strongly affect the downstream compressor performance. There is a much-distorted three-dimensional flow pattern at the exit of S-Shaped duct. This means that the aerodynamic sensitivity of S-Shaped duct to the IBL thickness is very high. Therefore, sufficient carefulness is needed to design not only downstream aerodynamic component (for example centrifugal impeller) but also upstream aerodynamic component (LPC OGV).

Control of Mach Reflection by a Single Laser Pulse

2012 ◽  
Vol 271-272 ◽  
pp. 1516-1520
Author(s):  
Dian Kai Wang ◽  
Yan Ji Hong
Keyword(s):  
Pressure Loss ◽  
Total Pressure ◽  
Free Stream ◽  
Laser Energy ◽  
Mach Reflection ◽  
Pulsed Laser ◽  
Total Pressure Loss ◽  
Mach Stem ◽  
Stem Height ◽  
Free Stream Mach Number

In the supersonic engine inlet, Mach Reflection probably appears when a supersonic flow goes through the symmetric wedges, causing a great total pressure loss. A single pulsed laser energy deposition leads a decrease of the Mach stem height and reduces the total pressure loss. By solving the two-dimensional RANS equations, with the condition of symmetric wedges at 22 degrees, and the free stream Mach number 3.45, influences of the deposition location and the magnitude of pulsed laser energy in Mach Reflection are investigated. The results indicate that when laser energy rises from 70mJ to 270mJ, the height of Mach stem changes and a vale value is obtained. The deposition position is also optimized.

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