Numerical Investigation of Endwall Boundary Layer Removal on Highly-Loaded Axial Compressor Blade Rows

2005 ◽  
Author(s):  
V. Gu¨mmer ◽  
M. Goller ◽  
M. Swoboda
Keyword(s):  
Boundary Layer ◽  
Flow Field ◽  
Reverse Flow ◽  
Axial Compressor ◽  
Flow Path ◽  
Compressor Blade ◽  
Main Flow ◽  
Local Flow ◽  
Layer Removal ◽  
Highly Loaded

This paper presents results of numerical investigations carried out to explore the benefit of endwall boundary layer removal from critical regions of highly-loaded axial compressor blade rows. At the loading level of modern aero-engine compressors the performance is primarily determined by three-dimensional flow phenomena occuring in the endwall regions. 3DNS simulations were conducted on both a rotor and a stator test case in order to evaluate basic effects and the practical value of bleeding air from specific locations at the casing endwall. The results of the numerical survey demonstrated substantial benefits of relatively small bleed rates to the local flow field and to the performance of the two blade rows. On the rotor, boundary layer fluid was removed from the main flow path through an axisymmetric slot in the casing over the rotor tip. This proved to give some control over the tip leakage vortex flow and the associated loss generation. On the stator, boundary layer fluid was taken from the flow path through a single bleed hole within the passage. Two alternative off-take configurations were evaluated, revealing a large impact of the bleed hole shape and location on the cross-passage flow and the suction side corner separation. On both blade rows investigated, rotor and stator, boundary layer removal resulted in a reduction of local reverse flow, blockage and losses in the respective near-casing region. This paper gives insight into changes occuring in the 3D passage flow field near the casing and summarises the effects on the radial matching and pitchwise-averaged performance parameters, namely loss and deviation of the rotor and stator when suction is active. Primary focus is put on the aerodynamics in the blade rows in the main flow path; details of the internal flow structure within the bleed off-take cavities/ports are not discussed here.

Numerical Investigation of End Wall Boundary Layer Removal on Highly Loaded Axial Compressor Blade Rows

2008 ◽  
Vol 130 (1) ◽  
Author(s):  
V. Gümmer ◽  
M. Goller ◽  
M. Swoboda
Keyword(s):  
Boundary Layer ◽  
Three Dimensional ◽  
Axial Compressor ◽  
Flow Path ◽  
Compressor Blade ◽  
Main Flow ◽  
Layer Removal ◽  
Wall Boundary Layer ◽  
End Wall ◽  
Highly Loaded

This paper presents results of numerical investigations carried out to explore the benefit of end wall boundary layer removal from critical regions of highly loaded axial compressor blade rows. At the loading level of modern aero engine compressors, the performance is primarily determined by three-dimensional (3D) flow phenomena occurring in the end wall regions. Three-dimensional Navier–Stokes simulations were conducted on both a rotor and a stator test case in order to evaluate the basic effects and the practical value of bleeding air from specific locations at the casing end wall. The results of the numerical survey demonstrated substantial benefits of relatively small bleed rates to the local flow field and to the performance of the two blade rows. On the rotor, the boundary layer fluid was removed from the main flow path through an axisymmetric slot in the casing over the rotor tip. This proved to give some control over the tip leakage vortex flow and the associated loss generation. On the stator, the boundary layer fluid was taken from the flow path through a single bleed hole within the passage. Two alternative off-take configurations were evaluated, revealing a large impact of the bleed hole shape and the location on the cross-passage flow and the suction side corner separation. On both blade rows investigated, rotor and stator, the boundary layer removal resulted in a reduction of the local reverse flow, blockage, and losses in the respective near-casing region. This paper gives insight into changes occurring in the 3D passage flow field near the casing and summarizes the effects on the radial matching and pitchwise-averaged performance parameters, namely loss and deviation of the rotor and stator when suction is active. Primary focus is put on the aerodynamics in the blade rows in the main flow path; details of the internal flow structure within the bleed off-take cavities/ports are not discussed here.

Effect of Boundary Layer Suction on the Corner Separation in a Highly Loaded Axial Compressor Cascade

2021 ◽  
Vol 143 (6) ◽  
Author(s):  
Tian Liang ◽  
Bo Liu ◽  
Stephen Spence
Keyword(s):  
Boundary Layer ◽  
Flow Field ◽  
Axial Compressor ◽  
Compressor Cascade ◽  
Flow Uniformity ◽  
Suction Surface ◽  
Corner Separation ◽  
Boundary Layer Suction ◽  
End Wall ◽  
Highly Loaded

Abstract Control of corner separation in axial compressor blade rows has attracted much interest due to its potential to improve compressor efficiency and the energy utilization in turbomachinery. This paper investigates the effectiveness and mechanisms of boundary layer suction in controlling the corner separation of a highly loaded axial compressor cascade. Numerical simulations have been carried out to investigate the effect of different suction schemes on the loss downstream of the cascade and the change in incidence characteristics with the variation of the suction flowrate. The results show that the effectiveness of flow suction in controlling the flow separation depends heavily on the proportion of the blade for which it is applied. It was found that suction along part of the blade span on the suction surface could effectively remove the separation at the region of the span influenced by the suction slot. However, this resulted in a deterioration of the flow field at other parts of the span. The full-span suction scheme on the suction surface not only eliminated the separation of the boundary layer in the middle of the blade but also significantly improved the flow uniformity near the end-wall. Despite the improvement in flow uniformity using the full-span suction scheme, a three-dimensional (3D) corner separation still existed due to the strong cross-passage pressure gradient. To improve the flow field uniformity further, two combined suction schemes with one spanwise slot on the suction surface and another slot on the end-wall were designed in order to fully remove both the separated flow on the blade suction surface and the 3D corner separation. It was found that the total pressure loss coefficient was reduced significantly by 63.8% with suction flowrates of 1.88% and 0.82% for the slots on the suction surface and the end-wall, respectively. Further work showed that the behavior of the loss coefficient is different as the combination of suction flowrates is changed for different incidence. The cascade loss at high incidence operation can be more effectively reduced with suction control on the end-wall. When implementing combined suction, it is necessary to determine the best combination of suction flowrate according to the incidence level.

Numerical Investigation of a Cantilevered Compressor Stator at Varying Clearance Sizes

2015 ◽  
Author(s):  
Chengwu Yang ◽  
Xingen Lu ◽  
Yanfeng Zhang ◽  
Shengfeng Zhao ◽  
Junqiang Zhu
Keyword(s):  
Flow Field ◽  
Axial Compressor ◽  
High Flow ◽  
Leakage Flow ◽  
Streamwise Direction ◽  
Stall Margin ◽  
Axial Compressors ◽  
Pressure Surface ◽  
The Impact ◽  
Highly Loaded

The clearance size of cantilevered stators affects the performance and stability of axial compressors significantly. Numerical calculations were carried out using the commercial software FINE/Turbo for a 2.5-stage highly loaded transonic axial compressor, which is of cantilevered stator for the first stage, at varying hub clearance sizes. The aim of this work is to improve understanding of the impact mechanism of hub clearance on the performance and the flow field in high flow turning conditions. The performance of the front stage and the compressor with different hub clearance sizes of the first stator has been analyzed firstly. Results show that the efficiency decreases as clearance size varies from 0 to 3% of hub chordlength, but the operating range has been extended. For the first stage, the efficiency decreases about 0.5% and the stall margin is extended. The following analysis of detailed flow field in the first stator shows that the clearance leakage flow and elimination of hub corner separation is responsible for the increasing loss and stall margin extending respectively. The effects of hub clearance on the downstream rotor have been discussed lastly. It indicates that the loss of the rotor increases and the flow deteriorates due to increasing of clearance size and hence the leakage mass flow rate, which mainly results from the interaction of upstream leakage flow with the passage flow near pressure surface. The affected region of rotor passage flow field expands in spanwise and streamwise direction as clearance size grows. The hub clearance leakage flow moves upward in span as it flows toward downstream.

scholarly journals Optimization Design of a 2.5 Stage Highly Loaded Axial Compressor with a Bezier Surface Modeling Method

2020 ◽  
Vol 10 (11) ◽  
pp. 3860
Author(s):  
Song Huang ◽  
Jinxin Cheng ◽  
Chengwu Yang ◽  
Chuangxin Zhou ◽  
Shengfeng Zhao ◽  
...  
Keyword(s):  
Flow Field ◽  
Dimensional Space ◽  
Three Dimensional ◽  
Axial Compressor ◽  
High Dimensions ◽  
Peak Efficiency ◽  
Stall Margin ◽  
Second Stage ◽  
Highly Loaded ◽  
Three Dimensional Space

Due to the complexity of the internal flow field of compressors, the aerodynamic design and optimization of a highly loaded axial compressor with high performance still have three problems, which are rich engineering design experience, high dimensions, and time-consuming calculations. To overcome these three problems, this paper takes an engineering-designed 2.5-stage highly loaded axial flow compressor as an example to introduce the design process and the adopted design philosophies. Then, this paper verifies the numerical method of computational fluid dynamics. A new Bezier surface modeling method for the entire suction surface and pressure surface of blades is developed, and the multi-island genetic algorithm is directly used for further optimization. Only 32 optimization variables are used to optimize the rotors and stators of the compressor, which greatly overcome the problem of high dimensions, time-consuming calculations, and smooth blade surfaces. After optimization, compared with the original compressor, the peak efficiency is still improved by 0.12%, and the stall margin is increased by 2.69%. The increase in peak efficiency is mainly due to the rotors. Compared with the original compressor, for the second-stage rotor, the adiabatic efficiency is improved by about 0.4%, which is mainly due to the decreases of total pressure losses in the range of above 30% of the span height and 10%–30% of the chord length. Besides, for the original compressor, due to deterioration of the flow field near the tip region of the second-stage stator, the large low-speed region eventually evolves from corner separation into corner stall with three-dimensional space spiral backflow. For the optimized compressor, the main reason for the increased stall margin is that the flow field of the second-stage stator with a span height above 50% is improved, and the separation area and three-dimensional space spiral backflow are reduced.

A Study on the Flowfield of an Innovative Z-Flowpath Gas Turbine Combustor

2010 ◽  
Author(s):  
Zongming Yu ◽  
Yong Huang ◽  
Fang Wang
Keyword(s):  
Gas Turbine ◽  
Reverse Flow ◽  
Flow Path ◽  
Main Flow ◽  
Wall Area ◽  
Pressure Losses ◽  
Micro Gas Turbine ◽  
Primary Zone ◽  
Commercial Code ◽  
The Stability

Reverse flow combustors were widely used in small and micro gas turbine engines. The wall area of this type of combustors was quite large. And there were two flow turning points in their flow-path. Thus the wall cooling and main flow dilution were two intrinsic problems for them. Apart from that, their high pressure losses and heavy weight were also two problems which seriously deteriorate the performance of the engines. Moreover, their primary hole jets on opposite walls were non-symmetrical, which would affect the stability and intensity of the recirculation flows. In order to improve the combustion performance, a new conceptual Z-flowpath combustor was proposed. The new combustor consisted of two 45 degree yawing instead of returning in the main flow-path. The flowfield of the new combustor was predicted by the commercial code FLUENT, after a validation for the flowfield in a model reverse flow combustor with previous experimental results. The prediction showed that the flowfield of the primary zone in the Z-flowpath combustor was highly symmetrical, the size and the intensity of the recirculation zone were about 10 and 2 times greater than the normal reverse flow combustor, respectively, while the pressure loss and the total area of the flame tube wall of the Z-flowpath combustor were decreased dramatically to be 69.4% and 51% of that in the reverse flow combustor, respectively.

Simulation and Analysis of Flow Field Control in Bionic Jet Surface for Drag Reduction

2013 ◽  
Vol 461 ◽  
pp. 746-750
Author(s):  
Zhao Gang ◽  
Fang Li ◽  
Jun Wei Du ◽  
Muhammad Farid ◽  
Dong Yang Zang
Keyword(s):  
Boundary Layer ◽  
Flow Field ◽  
Drag Reduction ◽  
Velocity Gradient ◽  
Reverse Flow ◽  
Reduction Rate ◽  
Frictional Resistance ◽  
Longitudinal Direction ◽  
Bottom Layer ◽  
Control Behavior

Numerical simulation was used with SST turbulence model on the drag reduction characteristics of bionic jet surface, which clarified the reason that the bionic jet surface could reduce the frictional resistance and the control behavior to the flow field near the wall. Results show that when the area of the jet hole is constant, the higher the ratio of the length along the longitudinal direction of jet hole and that of jet surface is, the better the drag reduction effect is. With the jet speed and jet flux increasing, the drag reduction rate will increase gradually until the maximum of 35.97%. The frictional resistance of bionic jet surface will decrease by increasing the area of reverse flow and decreasing the velocity gradient of the wall; the control behavior of jet surface to boundary layer embodies the shear stress in the bottom of boundary layer caused by the reverse flow in the back flow surface is opposite to the main flow field direction when the shear flow near the wall converges the jet impedance, which causes the low speed reverse rotating vortex pair in the downstream of jet hole, the secondary vortex near the wall caused by the extent of reverse vortex towards the downstream can increase the boundary bottom layer thickness and decrease the velocity gradient and frictional resistance.

Investigation of the Surge Behavior of a Multi-Stage Axial Compressor With a Multi-Sensor Probe

2011 ◽  
Author(s):  
P. Waniczek ◽  
P. Jeschke ◽  
H. Schoenenborn ◽  
T. Metzler
Keyword(s):  
Flow Field ◽  
Velocity Vector ◽  
Reverse Flow ◽  
Axial Compressor ◽  
Leading Edge ◽  
Field Structure ◽  
Pressure Probe ◽  
High Time Resolution ◽  
Trailing Edge ◽  
Flow Angle

The surge behavior of the first rotor of an eight-stage aero engine high pressure compressor has been investigated experimentally. For that purpose, a new multi-hole pressure probe was developed and adapted to the axial compressor test rig. Due to the high time resolution measurements (more than 45000 measuring points per surge cycle) it is possible to investigate the dynamic flow field of a surge cycle in a time-accurate manner. The results especially show the complex flow field structure at the surge inception. At the rotor leading edge the flow shows perturbations with high amplitudes and initiates the surge event, whereas the flow at the rotor trailing edge is less influenced. The inflow vector turns around the leading edge of the blade relatively slowly. During that turn around three different characteristic flow conditions have been identified. These are ‘zero rotor turning’, ‘turbine-like flow’ and ‘no flow’. ‘No flow’ means, that the absolute velocity vector reaches a flow angle where it consists of a pure tangential velocity component. That is the point where the reverse flow phase is initiated. A 180° shift of the flow direction at the rotor trailing edge is the consequence. After a quasi-steady reverse flow the acceleration of the flow starts. In total, this paper gives new and fundamental insights into the unsteady flow field phenomena during various surge cycles. Especially the transient velocity vector imparts a good idea of the flow field structure of a surging compressor.

scholarly journals Impact of Fluid Substitution on the Performance of an Axial Compressor Blade Cascade Working with Supercritical Carbon Dioxide

2019 ◽  
Vol 142 (1) ◽  
Author(s):  
Carlos Tello ◽  
Alejandro Muñoz ◽  
David Sánchez ◽  
Timoleon Kipouros ◽  
Mark Savill
Keyword(s):  
Carbon Dioxide ◽  
Supercritical Carbon Dioxide ◽  
Axial Compressor ◽  
Computational Design ◽  
Compressor Blade ◽  
Working Fluid ◽  
Local Flow ◽  
Variable Boundary ◽  
Turbomachinery Design ◽  
Supercritical Carbon

Abstract Recent research on turbomachinery design and analysis for supercritical carbon dioxide (sCO2) power cycles has relied on computational fluid dynamics. This has produced a large number of works whose approach is mostly case-specific, rather than of general application to sCO2 turbomachinery design. As opposed to such approach, this work explores the aerodynamic performance of compressor blade cascades operating on air and supercritical CO2 with the main objective to evaluate the usual aerodynamic parameters of the cascade for variable boundary conditions and geometries, enabling “full” or “partial” similarity. The results present both the global performance of the cascades and certain features of the local flow (trailing edge and wake). The discussion also highlights the mechanical limitations of the analysis (forces exerted on the blades), which is the main restriction for applying similarity laws to extrapolate the experience gained through decades of work on air turbomachinery to the new working fluid. This approach is a step toward the understanding and appropriate formulation of a multi-objective optimization problem for the design of such turbomachinery components where sCO2 is used as the operating fluid. With this objective, the paper aims to identify and analyze what would be expected if a common description of such computational design problems similar to those where air is the working fluid were used.

scholarly journals Impact of Fluid Substitution on the Performance of an Axial Compressor Blade Cascade Working With Supercritical Carbon Dioxide

2019 ◽  
Author(s):  
Alejandro Muñoz ◽  
David Sánchez ◽  
Mark Savill ◽  
Timoleon Kipouros ◽  
Carlos Tello-Castillo
Keyword(s):  
Carbon Dioxide ◽  
Supercritical Carbon Dioxide ◽  
Axial Compressor ◽  
Computational Design ◽  
Compressor Blade ◽  
Working Fluid ◽  
Local Flow ◽  
Variable Boundary ◽  
Turbomachinery Design ◽  
Supercritical Carbon

Abstract Recent research on turbomachinery design and analysis for supercritical Carbon Dioxide (sCO2) power cycles has relied on Computational Fluid Dynamics. This has produced a large number of works whose approach is mostly case-specific, rather than of general application to sCO2 turbomachinery design. As opposed to such approach, this work explores the aerodynamic performance of compressor blade cascades operating on air and supercritical CO2 with the main objective to evaluate the usual aerodynamic parameters of the cascade for variable boundary conditions and geometries, enabling ‘full’ or ‘partial’ similarity. The results present both the global performance of the cascades and certain features of the local flow (trailing edge and wake). The discussion also highlights the mechanical limitations of the analysis (forces exerted on the blades), which is the main restriction to applying similarity laws to extrapolate the experience gained through decades of work on air turbomachinery to the new working fluid. This approach is a step towards the understanding and appropriate formulation of a multi-objective optimisation problem for the design of such turbomachinery components where sCO2 is used as the operating fluid. With this objective, the paper aims to identify and analyse what would be expected if a common description of such computational design problems similar to those where air is the working fluid were used.

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