Effect of Airfoil Vortex Generator on the Performance and Stability of a Transonic Axial Compressor

2021 ◽  
Author(s):  
Subbaramu Shivaramaiah ◽  
Mahesh K. Varpe
Keyword(s):  
Pressure Ratio ◽  
Research Work ◽  
Vortex Generator ◽  
Leading Edge ◽  
Leakage Flow ◽  
Compressor Stage ◽  
Tip Leakage Flow ◽  
Stall Margin ◽  
Tip Leakage ◽  
Transonic Compressor

Abstract In the present research work, effect of airfoil vortex generator on performance and stability of transonic compressor stage is investigated through CFD simulations. In turbomachines vortex generators are used to energize boundary and generated vortex is made to interact with tip leakage flow and secondary flow vortices formed in rotor and stator blade passage. In the present numerical investigation symmetrical airfoil vortex generator is placed on rotor casing surface close to leading edge, anticipating that vortex generated will be able to disturb tip leakage flow and its interaction with rotor passage core flow. Six different vortex generator configuration are investigated by varying distance between vortex generator trailing edge and rotor leading edge. Particular vortex generator configuration shows maximum improvement of stall margin and operating range by 5.5% and 76.75% respectively. Presence of vortex generator alters flow blockage by modifying flow field in rotor tip region and hence contributes to enhancement of stall margin. As a negative effect, interaction of vortex generator vortices and casing causes surface friction and high entropy generation. As a result compressor stage pressure ratio and efficiency decreases.

Effect of Rotor Tip Casing Treatment on the Performance and Stability of a Transonic Axial Compressor

2021 ◽  
Author(s):  
Subbaramu Shivaramaiah ◽  
Mahesh K. Varpe ◽  
Mohammed Afzal
Keyword(s):  
Research Work ◽  
Stability Margin ◽  
Leakage Flow ◽  
Compressor Stage ◽  
Tip Leakage Flow ◽  
Stall Margin ◽  
Casing Treatment ◽  
Tip Leakage ◽  
Transonic Compressor ◽  
Stage Efficiency

Abstract In a transonic compressor rotor, tip leakage flow interacts with passage shock, casing boundary layer and secondary flow. This leads to increase in total pressure loss and reduction of compressor stability margin. Casing treatment is one of the passive endwall geometry modification technique to control tip leakage flow interaction. In the present investigation effect of rotor tip casing treatment is investigated on performance and stability of a NASA 37 transonic compressor stage. Existing literature reveals, that endwall casing treatment slots i.e., porous casing treatment, axial slots axially skewed slots, circumferential grooves, recirculating casing treatment etc. are able to improve compressor stability margin with penalty on stage efficiency. Turbomachinery engineers and scientists are still focusing their research work to identify an endwall casing treatment configuration with improves both compressor stall margin as well as stage efficiency. Hence in the current work, as an innovative idea, effect of casing treatment slot along rotor tip mean camber line is investigated on NASA 37 compressor stage. Casing treatment slot with rectangular cross-section was created along the rotor tip mean camber line. Four different casing treatment configurations were created by changing number of slots on rotor casing surface. In all four configurations casing treatment slot width and height remains same. Flow simulation of NASA 37 compressor stage was performed with all these four casing treatment configurations. A maximum stall margin improvement of 3% was achieved with a particular slot configuration, but without any increase in compressor stage efficiency.

scholarly journals Effects of Tip Leakage Flow on the Aerodynamic Performance and Stability of an Axial-Flow Transonic Compressor Stage

Energies ◽  
2021 ◽  
Vol 14 (14) ◽  
pp. 4168
Author(s):  
Botao Zhang ◽  
Xiaochen Mao ◽  
Xiaoxiong Wu ◽  
Bo Liu
Keyword(s):  
Boundary Layer ◽  
Shock Wave ◽  
Axial Flow ◽  
Leading Edge ◽  
Rotating Stall ◽  
Tip Clearance ◽  
Leakage Flow ◽  
Tip Leakage Flow ◽  
Tip Leakage ◽  
Transonic Compressor

To explain the effect of tip leakage flow on the performance of an axial-flow transonic compressor, the compressors with different rotor tip clearances were studied numerically. The results show that as the rotor tip clearance increases, the leakage flow intensity is increased, the shock wave position is moved backward, and the interaction between the tip leakage vortex and shock wave is intensified, while that between the boundary layer and shock wave is weakened. Most of all, the stall mechanisms of the compressors with varying rotor tip clearances are different. The clearance leakage flow is the main cause of the rotating stall under large rotor tip clearance. However, the stall form for the compressor with half of the designed tip clearance is caused by the joint action of the rotor tip stall caused by the leakage flow spillage at the blade leading edge and the whole blade span stall caused by the separation of the boundary layer of the rotor and the stator passage. Within the investigated varied range, when the rotor tip clearance size is half of the design, the compressor performance is improved best, and the peak efficiency and stall margin are increased by 0.2% and 3.5%, respectively.

Numerical Analysis of Transonic Compressor With Various Tip Clearance Gap

2016 ◽  
Author(s):  
Yasunori Sakuma ◽  
Toshinori Watanabe ◽  
Takehiro Himeno
Keyword(s):  
Tip Clearance ◽  
Leakage Flow ◽  
Numerical Instability ◽  
Tip Leakage Flow ◽  
Trailing Edge ◽  
Stall Margin ◽  
Tip Leakage ◽  
Transonic Compressor ◽  
Blade Tip ◽  
Edge Separation

Computational analysis has been conducted on the NASA Rotor 37 transonic compressor with various tip clearance gap heights. Using steady rotor-only analysis, the change in overall performance, basic flow characteristics, and near-casing phenomena have been carefully observed. The results have clarified that the peak efficiency of the compressor decreases almost linearly with the increase in gap height. Meanwhile, the stall margin was prone to deterioration in cases of significantly small or significantly large clearance gaps. The peak stall margin was attained when the gap was set to 75% of the original height. Focusing on the flow structures, the tip leakage flow and tip leakage vortex seemed to be dominant loss sources in the case of a large tip clearance gap. On the other hand, trailing edge separation at the blade tip was the major loss source in case of a small tip clearance gap. The difference in the near-casing flow structure also determined the onset process of numerical instability. In case of a large tip clearance gap, the advance of the interface between the main flow and tip leakage flow seemed to cause an accumulation of blockage in the region near the casing, possibly triggering the tip-initiated stall. In the case of a small tip clearance gap, interaction among the wall separation, blade tip trailing edge separation, and shockwave /boundary layer interaction was significant. These phenomena appeared to play a major role in the onset of numerical instability in the blade tip region.

Design of Casing Treatment on a Mixed-Flow Compressor

2017 ◽  
Author(s):  
Juan Du ◽  
Joerg R. Seume
Keyword(s):  
Pressure Ratio ◽  
Leading Edge ◽  
Area Ratio ◽  
Open Area ◽  
Tip Leakage Flow ◽  
Mixed Flow ◽  
Stall Margin ◽  
Casing Treatment ◽  
Tip Leakage ◽  
Flow Compressor

Casing treatments (CTs) have been proved to beneficially affect the behavior of tip clearance flow and compressor stability. This paper presents the design of casing treatment for a mixed-flow compressor with a very small tip gap of 0.1mm. In the first part, the potential of applying two traditional types of casing treatments, i.e. circumferential grooves and axial slots, to enhance the stability of a mixed-flow compressor is investigated. The flow details in the reference compressor with smooth casing are examined first. It is found that a separating vortex is formed due to the reversed flow on the blade suction side near the rotor trailing edge at the near-stall point. It is supposed to be responsible for the decrease in total pressure ratio when the compressor approaches to stall. The numerical stall, i.e. the breakdown of the simulation, is initiated from the spillage of tip leakage flow over the rotor blade leading edge. The effect of circumferential grooves on the compressor performances is not remarkable. The implement of axial slots ameliorates the total pressure ratio and extend the flow range substantially, but with higher efficiency penalty than the circumferential grooves. The recirculation formed in the axial skewed slots eliminates the separation vortex near the trailing edge and suppresses the spillage of the tip leakage flow forward the rotor leading edge simultaneously. The axial skewed slots are further designed and optimized numerically by DoE (Design of Experiments). As DoE factors the axial length, the height, the open area ratio, and the number per blade passage of the slots are varied. Their effects on the two target values stall margin and polytropic efficiency are investigated. The plot of stall margin improvement (SMI) with a function of the peak efficiency improvement (PEI) indicates that the SMI changes reversely with the PE. There are two trends in the correlation curves of SMI with PE. For the configurations with the open area ratio of 20%, the SMI is changed from 9% to 23% with 1% decrease in PE by varying other three factors. For the CTs with the open area ratio of 60% the augment in SMI from 17.8% to 26.3% produces extra efficiency loss of 4.2%.

Prediction of Unsteady Tip Leakage Flow of a Transonic Compressor Rotor by Reynolds-Stress-Constrained Large Eddy Simulation

2018 ◽  
Author(s):  
Yueqing Zhuang ◽  
Hui Liu
Keyword(s):  
Large Eddy Simulation ◽  
Reynolds Stress ◽  
Rotor Blade ◽  
Leading Edge ◽  
Leakage Flow ◽  
Tip Leakage Flow ◽  
Tip Leakage ◽  
Transonic Compressor ◽  
Eddy Simulation ◽  
Large Eddy

Since the unsteadiness of tip leakage flow has profound effects on both aerodynamic performance and stall margin of axial compressors, it is important to accurately predict the transient tip flow at affordable computational cost. Limited by the high requirement of grid resolution of wall turbulence flow, large eddy simulation (LES) method is greatly restricted in engineering application. In the present work, a Reynolds-stress-constrained large eddy simulation (CLES) method has been introduced, in which the whole domain is simulated using LES while Reynolds stress constraint is enforced on the subgrid-scale (SGS) stress model for near-wall regions aiming at reducing the near-wall grid resolution. The CLES simulations have been performed to investigate the flow behaviors of the unsteady tip leakage flow in a transonic compressor NASA Rotor 67 at near-stall conditions. Reliability assessments have been conducted through comparisons of experimental measurements and numerical results obtained by RANS, DES, CLES as well as LES, respectively. Both the total pressure ratio and isentropic efficiency calculated by CLES agree well with experiment. The turbulence statistical results show three distinct high flow fluctuation regions near the blade tip. The first one is a long and narrow strip ahead of the leading edge of the rotor caused by the movement of the passage shock wave. The second one is formed on the suction side from the leading-edge of the rotor blade due to the oscillation of the tip leakage vortex. And the third one, which occupies most of the blade passage from the middle part of the rotor blade, is generated under multiple factors. The frequency characteristic of the unsteady tip leakage flow has been analyzed. The energy spectrums of the local transient pressure signals are highly related with the local unsteady flow features. The originating mechanisms of the flow unsteadiness in the rotor tip leakage flow have also been discussed, and the results show that the flow unsteadiness is mainly caused by a combined interaction effect of the double leakage flow, the tip leakage vortex flow spilled from the adjacent blade passage, as well as the involved main flow.

Numerical Investigation on Slot Casing Treatment in a Transonic Axial Compressor Stage: Part 1 — Casing Treatment Design

2016 ◽  
Author(s):  
Mingmin Zhu ◽  
Xiaoqing Qiang ◽  
Jinfang Teng
Keyword(s):  
Axial Compressor ◽  
Leakage Flow ◽  
Compressor Stage ◽  
Tip Leakage Flow ◽  
Suction Surface ◽  
Stall Margin ◽  
Rotating Speed ◽  
Casing Treatment ◽  
Tip Leakage ◽  
Stall Margin Improvement

Slot-type casing treatment generally has a great potential of enhancing the operating range for tip-critical compressor rotors, however, with remarkable efficiency drop. Part I of this two-part paper was committed to develop a slot configuration with desired stall margin improvement and minimized efficiency loss. Steady simulation was carried out in a 1.5 transonic axial compressor stage at part design rotating speed. At this rotating speed this compressor stage operated at a subsonic condition and showed a rather narrow operating range, which needed to be improved badly. Flow fields analysis at peak efficiency and near stall point showed that the development of tip leakage vortex and resulting blockage near casing resulted in numerical stall. Three kinds of skewed slots with same rotor exposure and casing porosity were designed according to the tip flow field and some empirical strategies. Among three configurations, arc-curved skewed slot showed minimum peak efficiency drop with considerable stall margin improvement. Then rotor exposure and casing porosity were varied based on the original arc-curved skewed slot, with a special interest in detecting their impact on the compressor stability and overall efficiency. Result showed that smaller rotor exposure and casing porosity leaded to less efficiency drop. But meanwhile, effectiveness of improving compressor stability was weakened. The relation between efficiency drop and stall margin improvement fell on a smooth continuous curve throughout all slots configurations, indicating that the detrimental effect of casing treatment on compressor was inevitable. Flow analysis was carried out for cases of smooth casing and three arc-curved configurations at smooth casing near stall condition. The strength of suction/injection, tip leakage flow behavior and removal of blockage near casing were detailed examined. Larger rotor tip exposure and slots number contributed to stronger injection flow. The loss generated within the mixing process of injection flow with main flow and leakage flow is the largest source of entropy increase. Further loss mechanisms were interpreted at eight axial cuts, which were taken through the blade row and slots to show the increase in entropy near tip region. Entropy distributions manifested that loss generations with smooth casing were primarily ascribed to low-momentum tip leakage flow/vortex and suction surface separation at leading edge. CU0 slot, the arc-curved slots with 50% rotor tip exposure, was capable of suppressing the suction surface separation loss. Meanwhile, accelerated tip leakage flow brought about additional loss near casing and pressure surface. Upstream high entropy flow would be absorbed into the rear portion of slots repeatedly, resulting in further loss.

Effect of inlet distortion on the performance of axial transonic contra-rotating compressor

2016 ◽  
Vol 232 (1) ◽  
pp. 42-54 ◽  
Author(s):  
Hanru Liu ◽  
Yangang Wang ◽  
Songchuan Xian ◽  
Wenbin Hu
Keyword(s):  
Total Pressure ◽  
Pressure Ratio ◽  
Leading Edge ◽  
Tip Leakage Flow ◽  
Stall Margin ◽  
Tip Leakage ◽  
Inlet Distortion ◽  
Flow Mixing ◽  
Inflow Conditions ◽  
Total Pressure Ratio

The present paper numerically conducted full-annulus investigation on the effects of circumferential total pressure inlet distortion on the performance and flow field of the axial transonic counter-rotating compressor. Results reveal that the inlet distortion both deteriorates the performance of the upstream and downstream rotors resulting in reduction of total pressure ratio, efficiency and stall margin of the transonic contra-rotating compressor. Regarding the development of distortion inside compressor, the downstream rotor reinforces the air-flow mixing effects and, thus, attenuates the distortion intensity significantly. Under the distorted inflow conditions, the detached shockwave at the leading edge of downstream rotor interacts with the tip leakage flow and causes the blockage of the blades passage, which is one important reason for the transonic contra-rotating compressor stall.

Application of Casing Circumferential Grooves to Counteract the Influence of Tip Clearance

2008 ◽  
Author(s):  
Victor Mileshin ◽  
Igor Brailko ◽  
Andrew Startsev
Keyword(s):  
Mass Flow Rate ◽  
Mass Flow ◽  
Flow Rate ◽  
Pressure Ratio ◽  
Leading Edge ◽  
Tip Clearance ◽  
Leakage Flow ◽  
Tip Leakage Flow ◽  
Blade Height ◽  
Transonic Compressor

Widening of surge margin of a transonic compressor stage is the main objective of the paper. This stage is a typical middle stage of a modern high pressure compressor (HPC) with decreased number of stages. Hot tip clearance of the stage being integrated into a six-stage HPC providing total pressure ratio π* HPC ≥ 12 and mass flow-rate < 16 kg/sec is estimated at 2.5 – 3% of blade height and is classified as a large tip clearance. In this paper experimental and 3D viscous numerical performances of the stage are obtained for two values of rotor tip clearance — equal to 0.76% (small size) and 2.66% (large size) of blade height. In doing so, tip clearance enlargement from 0.76% to 2.66% has been made by increase of casing (shroud) radius. This increase is manufactured as a circumferential trench (recess) with axial width 30% larger than rotor axial chord. Below this tip clearance is called “recessed” tip clearance. A distinguishing feature of leakage flow in case of large tip clearance is a formation of reversed flow near rotor casing. This backflow being intensified by throttling causes increase of incidence at the rotor leading edge and development of rotor stall. Casing treatments are intended to inhibit and delay the process. Among them circumferential grooves is the simplest casing treatment. Investigated in this paper casing circumferential grooves cover 82% of rotor axial chord. Numerical visualization of the near-casing streamlines demonstrates that tip leakage flow drains into the casing grooves giving rise to extended domains of positive axial velocity. Calculated mass flow-rate through groove’s cross-section demonstrates maximum over the rotor blade tip (flow into the groove) and minimum at mid-pitch (flow out of the groove). Amplitude of this variation depends on the groove location and stage throttling.

A Proposal for Passive Wall Treatment Applied in High Performance Axial Flow Compressors

2020 ◽  
Author(s):  
Rubén Bruno Díaz ◽  
Jesuino Takachi Tomita ◽  
Cleverson Bringhenti ◽  
Francisco Carlos Elizio de Paula ◽  
Luiz Henrique Lindquist Whitacker
Keyword(s):  
Stokes Equations ◽  
Axial Compressor ◽  
Axial Flow ◽  
Tip Clearance ◽  
Leakage Flow ◽  
Smooth Wall ◽  
Tip Leakage Flow ◽  
Viscous Effects ◽  
Stall Margin ◽  
Tip Leakage

Abstract Numerical simulations were carried out with the purpose of investigating the effect of applying circumferential grooves at axial compressor casing passive wall treatment to enhance the stall margin and change the tip leakage flow. The tip leakage flow is pointed out as one of the main contributors to stall inception in axial compressors. Hence, it is of major importance to treat appropriately the flow in this region. Circumferential grooves have shown a good performance in enhancing the stall margin in previous researches by changing the flow path in the tip clearance region. In this work, a passive wall treatment with four circumferential grooves was applied in the transonic axial compressor NASA Rotor 37. Its effect on the axial compressor performance and the flow in the tip clearance region was analyzed and set against the results attained for the smooth wall case. A 2.63% increase in the operational range of the axial compressor running at 100%N, was achieved, when compared with the original smooth wall casing configuration. The grooves installed at compressor casing, causes an increase in the flow entropy generation due to the high viscous effects in this gap region, between the rotor tip surface and casing with grooves. These viscous effects cause a drop in the turbomachine efficiency. For the grooves configurations used in this work, an efficiency drop of 0.7% was observed, compared with the original smooth wall. All the simulations were performed based on 3D turbulent flow calculations using Reynolds Averaged Navier-Stokes equations, and the flow eddy viscosity was determined using the two-equation SST turbulence model. The details of the grooves geometrical dimensions and its implementation are described in the paper.

Tip Leakage Flow Instability in a Centrifugal Compressor

2021 ◽  
Vol 143 (4) ◽  
Author(s):  
Teng Cao ◽  
Tadashi Kanzaka ◽  
Liping Xu ◽  
Tobias Brandvik
Keyword(s):  
Shear Layer ◽  
Centrifugal Compressor ◽  
Large Scale ◽  
Flow Instability ◽  
Leading Edge ◽  
Main Stream ◽  
Leakage Flow ◽  
Tip Leakage Flow ◽  
Unstable Flow ◽  
Tip Leakage

Abstract In this paper, an unsteady tip leakage flow phenomenon is identified and investigated in a centrifugal compressor with a vaneless diffuser at near-stall conditions. This phenomenon is associated with the inception of a rotating instability in the compressor. The study is based on numerical simulations that are supported by experimental measurements. The study confirms that the unstable flow is governed by a Kelvin–Helmholtz type instability of the shear layer formed between the main-stream flow and the tip leakage flow. The shear layer instability induces large-scale vortex roll-up and forms vortex tubes, which propagate circumferentially, resulting in measured pressure fluctuations with short wavelength and high amplitude which rotate at about half of the blade speed. The 3D vortex tube is also found to interact with the main blade leading edge, causing the reduction of the blade loading identified in the experiment. The paper also reveals that the downstream volute imposes a once-per-rev circumferential nonuniform back pressure at the impeller exit, inducing circumferential loading variation at the impeller inducer, and causing circumferential variation in the unsteady tip leakage flow.

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