Flowfield Investigation of a Compressor Cascade at High Incidence—Part 1: Pneumatic Probe Measurements

2009 ◽  
Author(s):  
Pavlos K. Zachos ◽  
Vassilios Pachidis ◽  
Bernard Charnley ◽  
Pericles Pilidis
Keyword(s):  
Pressure Loss ◽  
Total Pressure ◽  
Incidence Angle ◽  
Axial Flow ◽  
Quantitative Measure ◽  
Operating Conditions ◽  
Total Pressure Loss ◽  
Compressor Cascade ◽  
Engine Operation ◽  
Pressure Surface

The performance prediction of axial flow compressors and turbines still relies on the stationary testing of blade cascades. Most of the blade testing studies are done for operating conditions close to the design point or in off-design areas not too far from it. However, blade performance remains unexplored at very far off-design conditions, such as windmilling, characterised by operation under extremely low mass flows and rotational speeds which, in turn, imply highly negative incidence angle values. In this paper, the flow field generated by a 3-dimensional linear compressor cascade at a highly negative incidence angle and zero rotational speed is experimentally investigated using a pneumatic miniature cobra probe. The main objective of the study is to derive the total pressure loss through the blades at such a highly negative incidence angle. An overview of the blade geometry as well as of the experimental facility is given whereas the measurement strategy and the data acquisition technique are also presented. An uncertainty study taking into account the most significant factors affecting the quality of the results has been carried out. As shown by the measurements taken at specific positions downstream of the blades, the flowfield is dominated by highly separated flows on the pressure surface, which contribute to the increased values of the total pressure loss coefficient which seems to be weakly dependent on the inlet Mach number. The quantitative measure of the pressure losses at the extremely negative incidence angle examined can be considered to be a validation platform for correspondent numerical studies of similar flow conditions. Additionally, the experimental results obtained can be used to extend the applicability of the current pressure loss models, increasing the predictive capability of the through flow numerical approaches towards far off-design areas of component or whole engine operation.

Non-Axisymmetric Endwall Contouring in a Compressor Cascade With Tip Gap

2014 ◽  
Author(s):  
Mahesh K. Varpe ◽  
A. M. Pradeep
Keyword(s):  
Pressure Loss ◽  
Total Pressure ◽  
Three Dimensional ◽  
Axial Flow ◽  
Tip Clearance ◽  
Total Pressure Loss ◽  
Compressor Cascade ◽  
Flow Solver ◽  
Pressure Loss Coefficient ◽  
Flow Compressor

This paper describes the design of a non-axisymmetric hub contouring in a shroudless axial flow compressor cascade operating at near stall condition. Although, an optimum tip clearance reduces the total pressure loss, further minimization of the losses using hub contouring was achieved. The design methodology presented here combines an evolutionary principle with a three-dimensional CFD flow solver to generate different geometric profiles of the hub systematically. The total pressure loss coefficient was used as a single objective function to guide the search process for the optimum hub geometry. The resulting three dimensionally complex hub promises considerable benefits discussed in detail in this paper. A reduction of 15.2% and 16.23% in the total pressure loss and secondary kinetic energy, respectively, was achieved in the wake. The blade loading was observed to improve by about 4.53%. Other complementary benefits are also listed in the paper. The results confirm that non-axisymmetric contouring is an effective method for reducing the losses and thereby improving the performance of the cascade.

scholarly journals Wake Mixing Improvements in a Linear Compressor Cascade With Crenulated Trailing Edges

1990 ◽  
Author(s):  
J. L. Veesart ◽  
P. I. King ◽  
W. C. Elrod ◽  
A. J. Wennerstrom
Keyword(s):  
Pressure Loss ◽  
Total Pressure ◽  
Total Pressure Loss ◽  
Compressor Cascade ◽  
Suction Surface ◽  
Trailing Edge ◽  
Turbine Engine ◽  
Pressure Surface ◽  
Trailing Edges ◽  
Linear Compressor Cascade

Trailing edge crenulations offer one possible way of energizing the trailing edge wakes from a gas turbine engine compressor blade by creating small vortices as a result of the pressure differential between the suction surface and pressure surface. The effect of crenulated trailing edges on wake dissipation and mixed-out total pressure loss in a linear, subsonic, compressor cascade was investigated for three low aspect ratio blade configurations: one with no crenulations and two others with large and small crenulation patterns, respectively. The effect of crenulations was to improve the wake mixing and reduce the velocity deficit. larger crenulations dissipated the wake most rapidly, and both crenulation configurations offered an improvement in total pressure loss and some improvement in flow turning.

Benefits of Nonaxisymmetric Endwall Contouring in a Compressor Cascade With a Tip Clearance

2015 ◽  
Vol 137 (5) ◽  
Author(s):  
Mahesh K. Varpe ◽  
A. M. Pradeep
Keyword(s):  
Pressure Loss ◽  
Total Pressure ◽  
Three Dimensional ◽  
Axial Flow ◽  
Vortex Generator ◽  
Tip Clearance ◽  
Total Pressure Loss ◽  
Compressor Cascade ◽  
Pressure Loss Coefficient ◽  
Flow Compressor

This paper describes the design of a nonaxisymmetric hub contouring in a shroudless axial flow compressor cascade operating at near stall condition. Although an optimum tip clearance (TC) reduces the total pressure loss, further reduction in the loss was achieved using hub contouring. The design methodology presented here combines an evolutionary principle with a three-dimensional (3D) computational fluid dynamics (CFD) flow solver to generate different geometric profiles of the hub systematically. The resulting configurations were preprocessed by GAMBIT© and subsequently analyzed computationally using ANSYSFluent©. The total pressure loss coefficient was used as a single objective function to guide the search process for the optimum hub geometry. The resulting three dimensionally complex hub promises considerable benefits discussed in detail in this paper. A reduction of 15.2% and 16.23% in the total pressure loss and secondary kinetic energy (SKE), respectively, is achieved in the wake region. An improvement of 4.53% in the blade loading is observed. Other complimentary benefits are also listed in the paper. The majority of the benefits are obtained away from the hub region. The contoured hub not only alters the pitchwise static pressure gradient but also acts as a vortex generator in an effort to alleviate the total pressure loss. The results confirm that nonaxisymmetric contouring is an effective method for reducing the losses and thereby improving the performance of the cascade.

scholarly journals Corner Separation Control Using a New Combined Slotted Configuration in a High-Turning Compressor Cascade under Different Solidities

Energies ◽  
2021 ◽  
Vol 14 (12) ◽  
pp. 3376
Author(s):  
Hejian Wang ◽  
Yanshan Qing ◽  
Bo Liu ◽  
Xiaochen Mao
Keyword(s):  
Pressure Loss ◽  
Total Pressure ◽  
Incidence Angle ◽  
Pressure Coefficient ◽  
Weight Ratio ◽  
Total Pressure Loss ◽  
Compressor Cascade ◽  
Turning Angle ◽  
Blade Loading ◽  
Corner Separation

In order to comprehensively control the corner separation and the blade trailing edge (TE) separation in a high-turning compressor stator cascade, this research proposes a new combined slotted configuration consisting of one full-span slot and two blade-end slots. Taking into account the effect of the blade solidity, the performance of the original cascade and the combined slotted cascade was calculated and evaluated in a wide incidence angle range at two blade solidities. The results indicated that the blade loading and the corner separation range of the original cascade becomes larger as the blade solidity decreases from 1.66 to 1.36, which leads to higher total pressure loss and lower pressure diffusing capacity under positive incidence angles. The low-momentum fluid in the boundary layer can be significantly re-energized by the high-momentum blade-end and full-span slots jets, hence the combined slotted configuration can eliminate the blade TE separation and reduce the corner separation remarkably in the full incidence angle range at the two blade solidities. By adopting the combined slotted configuration, the total pressure loss, turning angle and static pressure coefficient of the original cascade can be increased by −23.2%, 2.7° and 4.7% on average, respectively, when the blade solidity is 1.66, while they can be increased by −27.7%, 3.3° and 7.6% on average, respectively, when the blade solidity is 1.36. The combined slotted configuration has a significant adaptability to the low blade solidity (or high loading) condition and it shows a certain potential in increasing the aeroengine thrust-to-weight ratio by decreasing the compressor single-stage blade number.

Probabilistic and Numerical Modelling of a Lobed Mixer at Windmilling Conditions

2013 ◽  
Author(s):  
Maxime Lecoq ◽  
Nicholas Grech ◽  
Pavlos K. Zachos ◽  
Vassilios Pachidis
Keyword(s):  
Flow Field ◽  
Response Surface ◽  
Gas Turbine ◽  
Pressure Loss ◽  
Total Pressure ◽  
Engine Performance ◽  
Operating Conditions ◽  
Total Pressure Loss ◽  
Operating Point ◽  
Mixing Losses

Aero-gas turbine engines with a mixed exhaust configuration offer significant benefits to the cycle efficiency relative to separate exhaust systems, such as increase in gross thrust and a reduction in fan pressure ratio required. A number of military and civil engines have a single mixed exhaust system designed to mix out the bypass and core streams. To reduce mixing losses, the two streams are designed to have similar total pressures. In design point whole engine performance solvers, a mixed exhaust is modelled using simple assumptions; momentum balance and a percentage total pressure loss. However at far off-design conditions such as windmilling and altitude relights, the bypass and core streams have very dissimilar total pressures and momentum, with the flow preferring to pass through the bypass duct, increasing drastically the bypass ratio. Mixing of highly dissimilar coaxial streams leads to complex turbulent flow fields for which the simple assumptions and models used in current performance solvers cease to be valid. The effect on simulation results is significant since the nozzle pressure affects critical aspects such as the fan operating point, and therefore the windmilling shaft speeds and air mass flow rates. This paper presents a numerical study on the performance of a lobed mixer under windmilling conditions. An analysis of the flow field is carried out at various total mixer pressure ratios, identifying the onset and nature of recirculation, the flow field characteristics, and the total pressure loss along the mixer as a function of the operating conditions. The data generated from the numerical simulations is used together with a probabilistic approach to generate a response surface in terms of the mass averaged percentage total pressure loss across the mixer, as a function of the engine operating point. This study offers an improved understanding on the complex flows that arise from mixing of highly dissimilar coaxial flows within an aero-gas turbine mixer environment. The total pressure response surface generated using this approach can be used as look-up data for the engine performance solver to include the effects of such turbulent mixing losses.

Endwall Contouring and Fillet Design for Reducing Losses and Homogenizing the Outflow of a Compressor Cascade

2014 ◽  
Author(s):  
Oliver Reutter ◽  
Stefan Hemmert-Pottmann ◽  
Alexander Hergt ◽  
Eberhard Nicke
Keyword(s):  
Pressure Loss ◽  
Total Pressure ◽  
Leading Edge ◽  
Total Pressure Loss ◽  
Wall Region ◽  
Optimized Design ◽  
Compressor Cascade ◽  
Outflow Region ◽  
Endwall Contouring ◽  
Operating Points

The following paper deals with the development of an optimized fillet and an endwall contour for reducing the total pressure loss and for homogenizing the outflow of a highly loaded cascade with a low aspect ratio. The NACA-65 K48 cascade profile without a fillet and without endwall contouring is used as a basis. Optimizations are performed using the DLR in-house tool AutoOpti and the RANS-solver TRACE. Three operating points at an inflow Mach number of 0.67 with different inflow angles are used to secure a wide operating range of the optimized design. At first only a fillet is optimized. The optimized fillet is small at the leading edge and rather high, wide and thick towards the trailing edge. It reduces the total pressure loss and homogenizes the outflow up to a blade height of 20 %. Following this a combined optimization of the endwall and the fillet is performed. The optimized contour leads to the development of a vortex, which changes the secondary flow in such a way, that the corner separation is reduced, which in turn significantly reduces the total pressure loss up to 16 % in the design operating point. The contour in the outflow region leads to a significant homogenization of the outflow in the near wall region.

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.

A new curvature-controlled stacking-line method for optimization design of compressor cascade considering surface smoothness

2019 ◽  
Vol 234 (5) ◽  
pp. 1061-1074
Author(s):  
Xinyi Zhang ◽  
Xiaoqing Qiang ◽  
Jinfang Teng ◽  
Wensheng Yu
Keyword(s):  
Pressure Loss ◽  
Total Pressure ◽  
Optimization Design ◽  
Parametric Method ◽  
Total Pressure Loss ◽  
Compressor Cascade ◽  
Surface Smoothness ◽  
Line Method ◽  
Radial Curvature ◽  
Bézier Spline

The paper presents an advanced parametric method of blade stacking lines in terms of sweep and lean based on controlled curvature. To the knowledge of the authors, there is no related approach reported in open literature that uses Bezier spline as the radial curvature distribution to improve the smoothness of the blade surface; most previous studies ignored the discontinuous slopes of curvature of the parametric curves. The parametric method called curvature-controlled stacking-line method (CCSLM) is performed by changing the magnitude of the sweep or lean. A fourth Bezier spline is adopted to define the curvature of spanwise stacking line directly ensuring surface smoothness. Then, the redesign cascades are created by sectional profiles stacked along the radial stacking lines which are obtained by twice integrating the Bezier spline. Then, the advanced method is conducted to optimize a high-subsonic controlled diffusion airfoil at design point, where the blade shape is generated in terms of lean. A single-objective optimization is performed using Kriging model and genetic algorithm to optimize total pressure loss, and the optimized geometry is obtained. The optimization results show that the blade design CCSLM has significant effects on the endwall flow vortex as well as radial loading distribution. The reduction of total pressure loss and secondary flow is also observed, and the aerodynamic performance is well improved compared with the original cascade.

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.

Investigation on Performance of Compressor Cascade with Tubercle Leading Edge Blade

2020 ◽  
Vol 37 (3) ◽  
pp. 295-303 ◽  
Author(s):  
Tu Baofeng ◽  
Zhang Kai ◽  
Hu Jun
Keyword(s):  
Pressure Loss ◽  
Total Pressure ◽  
Large Scale ◽  
Design Method ◽  
Leading Edge ◽  
Experimental Tests ◽  
Aerodynamic Performance ◽  
Total Pressure Loss ◽  
Compressor Cascade ◽  
Linear Compressor Cascade

AbstractIn order to improve compressor performance using a new design method, which originates from the fins on a humpback whale, experimental tests and numerical simulations were undertaken to investigate the influence of the tubercle leading edge on the aerodynamic performance of a linear compressor cascade with a NACA 65–010 airfoil. The results demonstrate that the tubercle leading edge can improve the aerodynamic performance of the cascade in the post-stall region by reducing total pressure loss, with a slight increase in total pressure loss in the pre-stall region. The tubercles on the leading edge of the blades cause the flow to migrate from the peak to the valley on the blade surface around the tubercle leading edge by the butterfly flow. The tubercle leading edge generates the vortices similar to those created by vortex generators, splitting the large-scale separation region into multiple smaller regions.

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