scholarly journals SINGLE AND MULTIPLE CIRCUMFERENTIAL CASING GROOVE FOR STALL MARGIN IMPROVEMENT IN A TRANSONIC AXIAL COMPRESSOR

2021 ◽  
pp. 1-22
Author(s):  
Ahmad Fikri Bin Mustaffa ◽  
Vasudevan Kanjirakkad
Keyword(s):  
Axial Compressor ◽  
Stability Limit ◽  
Leading Edge ◽  
Vortex Interaction ◽  
Stall Margin ◽  
Tip Leakage ◽  
Stall Margin Improvement ◽  
Transonic Axial Compressor ◽  
The Stability ◽  
Transonic Rotor

Abstract The stability limit of a tip-stalling axial compressor is sensitive to the magnitude of the near casing blockage. In transonic compressors, the presence of the passage shock could be a major cause for the blockage. Identification and elimination of this blockage could be key to improving the stability limit of the compressor. In this paper, using numerical simulation, the near casing blockage within the transonic rotor, NASA Rotor 37, is quantified using a blockage parameter. For a smooth casing, the blockage at conditions near stall has been found to be maximum at about 20% of the tip axial chord downstream of the tip leading edge. This maximum blockage location is found to be consistent with the location of the passage shock-tip leakage vortex interaction. A datum single casing groove design that minimises the peak blockage is found through an optimisation approach. The stall margin improvement of the datum casing groove is about 0.6% with negligible efficiency penalty. Furthermore, the location of the casing groove is varied upstream and downstream of the datum location. It is shown that the stability limit of the compressor is best improved when the blockage is reduced upstream of the peak blockage location. The paper also discusses the prospects of a multi-groove casing configuration.

scholarly journals Single and Multiple Circumferential Casing Groove for Stall Margin Improvement in a Transonic Axial Compressor

2020 ◽  
Author(s):  
A. F. Mustaffa ◽  
V. Kanjirakkad
Keyword(s):  
Axial Compressor ◽  
Stability Limit ◽  
Leading Edge ◽  
Vortex Interaction ◽  
Stall Margin ◽  
Tip Leakage ◽  
Stall Margin Improvement ◽  
Transonic Axial Compressor ◽  
The Stability ◽  
Transonic Rotor

Abstract The stability limit of a tip-stalling axial compressor is sensitive to the magnitude of the near casing blockage. In transonic compressors, the presence of the passage shock could be a major cause for the blockage. Identification and elimination of this blockage could be key to improving the stability limit of the compressor. In this paper, using numerical simulation, the near casing blockage within the transonic rotor, NASA Rotor 37, is quantified using a blockage parameter. For a smooth casing, the blockage at conditions near stall has been found to be maximum at about 20% of the tip axial chord downstream of the tip leading edge. This maximum blockage location is found to be consistent with the location of the passage shock-tip leakage vortex interaction. A datum single casing groove design that minimises the peak blockage is found through an optimisation approach. The stall margin improvement of the datum casing groove is about 0.6% with negligible efficiency penalty. Furthermore, the location of the casing groove is varied upstream and downstream of the datum location. It is shown that the stability limit of the compressor is best improved when the blockage is reduced upstream of the peak blockage location. The paper also discusses the prospects of a multi-groove casing configuration.

scholarly journals Stall margin improvement in a low-speed axial compressor rotor using a blockage-optimised single circumferential casing groove

2021 ◽  
Vol 5 ◽  
pp. 79-89
Author(s):  
Ahmad Fikri Mustaffa ◽  
Vasudevan Kanjirakkad
Keyword(s):  
Axial Compressor ◽  
Pressure Rise ◽  
Leading Edge ◽  
Tip Leakage Flow ◽  
Stall Margin ◽  
Low Speed ◽  
Tip Leakage ◽  
Compressor Rotor ◽  
Stall Margin Improvement ◽  
Edge Based

The stall margin of tip-critical axial compressors can be improved by using circumferential casing grooves. From previous studies, in the literature, the stall margin improvement due to the casing grooves can be attributed to the reduction of the near casing blockage. The pressure rise across the compressor as the compressor is throttled intensifies the tip leakage flow. This results in a stronger tip leakage vortex that is thought to be the main source of the blockage. In this paper, the near casing blockage due to the tip region aerodynamics in a low-speed axial compressor rotor is numerically studied and quantified using a mass flow-based blockage parameter. The peak blockage location at the last stable operating point for a rotor with smooth casing is found to be at about 10% of the tip chord aft of the tip leading edge. Based on this information, an optimised single casing groove design that minimises the peak blockage is found using a surrogate-based optimisation approach. The implementation of the optimised groove is shown to produce a stall margin improvement of about 5%.

A Study of Influences of Inlet Total Pressure Distortions On Clearance Flow in an Axial Compressor

2021 ◽  
Author(s):  
Guoming Zhu ◽  
Xiaolan Liu ◽  
Bo Yang ◽  
Moru Song
Keyword(s):  
Total Pressure ◽  
Pressure Ratio ◽  
Axial Compressor ◽  
Leading Edge ◽  
Stall Margin ◽  
Flow Cells ◽  
Static Pressure Distribution ◽  
Speed Flow ◽  
Clearance Flow ◽  
The Stability

Abstract The rotating distortion generated by upstream wakes or low speed flow cells is a kind of phenomenon in the inlet of middle and rear stages of an axial compressor. Highly complex inflow can obviously affect the performance and the stability of these stages, and is needed to be considered during compressor design. In this paper, a series of unsteady computational fluid dynamics (CFD) simulations is conducted based on a model of an 1-1/2 stage axial compressor to investigate the effects of the distorted inflows near the casing on the compressor performance and the clearance flow. Detailed analysis of the flow field has been performed and interesting results are concluded. The distortions, such as total pressure distortion in circumferential and radial directions, can block the tip region so that the separation loss and the mixing loss in this area are increased, and the efficiency and the total pressure ratio are dropped correspondingly. Besides, the distortions can change the static pressure distribution near the leading edge of the rotor, and make the clearance flow spill out of the rotor edge more easily under near stall condition, especially in the cases with co-rotating distortions. This phenomenon can be used to explain why the stall margin is deteriorated with nonuniform inflows.

Effect of Various Trench Designs on Axial Compressor Blade Tip Aerodynamics

2019 ◽  
Author(s):  
Ashwin Ashok ◽  
Patur Ananth Vijay Sidhartha ◽  
Shine Sivadasan
Keyword(s):  
Axial Compressor ◽  
Leading Edge ◽  
Suction Side ◽  
Horseshoe Vortex ◽  
Tip Clearance ◽  
Compressor Cascade ◽  
Tip Leakage Vortex ◽  
Stall Margin ◽  
Casing Treatment ◽  
Tip Leakage

Abstract Tip clearance of axial compressor blades allows leakage of the flow, generates significant losses and reduces the compressor efficiency. The present paper aims to discuss the axial compressor tip aerodynamics for various configurations of tip gap with trench. The various configurations are obtained by varying the clearance, trench depth, step geometry and casing contouring. In this paper the axial compressor aerodynamics for various configurations of tip gap with trench have been studied. The leakage flow structure, vorticity features and entropy generations are analyzed using RANS based CFD. The linear compressor cascade comprises of NACA 651810 blade with clearance height varied from 0.5% to 2% blade span. Trail of the tip leakage vortex and the horseshoe vortex on the blade suction side are clearly seen for the geometries with and without casing treatments near the stalling point. Since the trench side walls are similar to forward/backing steps, a step vortex is observed near the leading edge as well as trailing edge of the blade and is not seen for the geometry without the casing treatment. Even though the size of the tip leakage vortex seams to be reduces by providing a trench to the casing wall over the blade, the presence of additional vortices like the step vortex leads to comparatively higher flow losses. An increase in overall total pressure loss due to the application of casing treatment is observed. However an increase in stall margin for the geometries with casing is noted.

Experimental and Numerical Investigation of Effect of Center Offset Degree on Compressor Stability With Circumferential Grooved Casing Treatment

2016 ◽  
Author(s):  
HaoGuang Zhang ◽  
Feng Tan ◽  
YanHui Wu ◽  
WuLi Chu ◽  
Wei Wang ◽  
...  
Keyword(s):  
Axial Compressor ◽  
Leading Edge ◽  
Groove Depth ◽  
Tip Clearance ◽  
Central Difference ◽  
Stall Margin ◽  
Casing Treatment ◽  
Compressor Rotor ◽  
Stall Margin Improvement ◽  
Blade Tip

For compressor blade tip stall, one effective way of extending stable operating range is with the application of circumferential grooved casing treatment and its validity was proved by a lot of experimental and numerical investigations. The emphases of most circumferential grooved investigations are focused on the influence of groove depth and groove number on compressor stability, and there is few investigations dealt with the center offset degree of circumferential grooves casing treatment. Hence, an axial compressor rotor with casing treatment (CT) was investigated with experimental and numerical methods to explore the effect of center offset degree on compressor stability and performance. In the work reported here, The center offset degree is defined as the ratio of the central difference between rotor tip axial chord and CT to the axial chord length of rotor tip. When the center of CT is located within the upstream direction of the center of rotor tip axial chord, the value of center offset degree is positive. The experimental and numerical results show that stall margin improvement gained with CT is reduced as the value of center offset degree varies from 0 to 0.33 or −0.33, and the CT with −0.33 center offset degree achieves the lowest value of stall margin improvement at 53% and 73% design rotational speed. The detailed analysis of the flow-field in compressor tip indicates that there is not positive effect made by grooves on leading edge of rotor blade tip when the value of center offset degree is −0.33. As the mass flow of compressor reduces further, tip clearance leakage flow results in the outlet blockage due to the absence of the positive action of grooves near blade tip tail when the value of center offset degree is 0.33. Blockage does not appear in rotor tip passage owing to utilizing the function of all grooves with CT of 0 center offset degree.

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.

Investigation of Passage Flow Features in a Transonic Compressor Rotor With Casing Treatments

2011 ◽  
Author(s):  
M. W. Mu¨ller ◽  
H.-P. Schiffer ◽  
Melanie Voges ◽  
Chunill Hah
Keyword(s):  
Measurement Techniques ◽  
Stability Limit ◽  
Rotating Stall ◽  
Tip Leakage Flow ◽  
Stall Margin ◽  
Tip Leakage ◽  
Operating Range ◽  
Transonic Compressor ◽  
The Stability ◽  
Casing Treatments

An experimental investigation on casing treatments in a one-stage transonic compressor is presented. The reference case consists of a radially staggered blisk and six circumferential grooves. Speedlines show that this axisymmetric treatment already provided a substantial increase in operating range with relatively small losses in efficiency. Since the onset of rotating stall in tip-critical high-speed compressors is always linked to the tip-leakage flow and the build-up of blockage within the blade passage. High-resolution measurement techniques have been employed to investigate the corresponding effects. Results with Particle Image Velocimetry show that the interaction between the tip leakage vortex and the shock front cause a blockage area. When throttled further, the blockage increases. The shock structure changes similar to the phenomena of vortex breakdown described by different researchers in the past, but a stagnation point is not present. Before reaching the stability limit, the interface line between the incoming flow and the blocked area moves towards the inlet plane of the rotor indicating spike-type stall inception. Wall pressure measurements confirmed this theory for the smooth wall, but with circumferential grooves applied, a part span stall cell develops prior to the stability limit. In order to assess the performance of circumferential grooves, two additional configurations are presented. The corresponding measurements addressed the questions whether circumferential grooves also provide an operating range extension when applied to an optimized rotor design with higher initial stall margin. Therefore, an identical casing treatment is applied to a forward swept rotor. The second question is, how circumferential grooves perform in direct comparison to a non-axisymmetric endwall structure. Axial slots have been applied to the radially staggered rotor. While the stall margin exceeds all other configurations, detrimential effects in efficiency are observed. A detailed anaylsis of probe data shows the changes of the radial profile at the rotor outlet which allows recommendations for more efficient CT designs. Parameters allowing to evaluate the CT influence are presented.

Origins and Structure of Rotating Instability: Part 1 — Experimental and Numerical Observations in a Subsonic Axial Compressor Rotor

2016 ◽  
Author(s):  
Yanhui Wu ◽  
Zhiyang Chen ◽  
Guangyao An ◽  
Jun Liu ◽  
Guowei Yang
Keyword(s):  
Axial Compressor ◽  
Stability Limit ◽  
Periodic Oscillation ◽  
Tip Leakage ◽  
Flow Unsteadiness ◽  
Compressor Rotor ◽  
Clearance Flow ◽  
The Stability ◽  
Rotating Instability ◽  
Mode Order

Rotating instability (RI) is an obvious unsteady flow phenomenon occurring in the tip region of compressors, which is potentially linked to tip clearance flow noise, blade vibration and rotating stall/surge. The existing investigations for RI indicate the origins of RI are closely related to unsteady flow behaviors in given blade passages in the rotating reference frame, which depend on the design specifics of axial machines. However, no efforts are made to set up a quantitative link between the time scale of unsteady behavior in a given passage and the characteristic parameters of RI, let alone to define the fluid dynamic processes/events which are causally linked with the RI inception. This is the motivation for the current investigations. In Part I, the experimental and numerical investigations are carried out to investigate tip flow unsteadiness in a subsonic axial compressor rotor. The measurement results show RI appears at operating points near the stability limit of the test rotor. It becomes more pronounced with mass flow rate decreased. The corresponding computational experiments show that flow unsteadiness in given passages also appears close to the stability limit with its initial origination confined to the tip region. The appearance of tip flow unsteadiness is accompanied by a phase lag pattern in different passages across the circumference similar to the detection of stall flutter. The well-developed Fourier-decomposed method is thus used to evaluate the mode characteristics of circumferential traveling waves. It turns out the circumferential traveling wave rotating against the rotor rotation direction with the mode order of 4 is prominent in the flow field with its frequency in the absolute frame equivalent to the mean frequency value of RI detected in measurements. The further analyses of the simulated flow fields indicate that tip flow unsteadiness in a given passage attributes to the periodic oscillation of “secondary clearance flow”, which induces a blockage tranfer across the passage. The mode order and propagation speed of RI depend on the blockage transfer induced by the periodic oscillation of “secondary clearance flow” between two neighbouring passages along the whole circumference. The investigation results presented in the paper implies that one of early ideas to interpret the origin of RI might be altered to such an extent that it contains any unsteady behavior associated with tip leakage flow, rather than limited to “periodical oscillation of tip leakage vortex”.

Numerical Study of Discrete Tip Injection in a Transonic Axial Compressor

2010 ◽  
Author(s):  
Hossein Khaleghi ◽  
Joao A. Teixeira
Keyword(s):  
Numerical Study ◽  
Axial Compressor ◽  
Leading Edge ◽  
Boundary Layer Separation ◽  
Rotating Stall ◽  
Low Velocity ◽  
Tip Leakage Vortex ◽  
Tip Leakage ◽  
Transonic Axial Compressor ◽  
Tip Injection

This paper reports on a time-accurate simulation of discrete tip injection upstream of a transonic axial compressor, NASA Rotor-67. Twenty two discrete injectors were placed around the casing upstream of the blade to investigate the effect of injector-rotor interaction at near-stall condition. Time-accurate simulations were performed with and without tip injection at stable and unstable operating points. Although the injected mass flow rate is very small, the range extension obtained promises the effective use of such injection in suppressing rotating stall at early stages of formation, with almost no efficiency penalty incurred. The effect of injection on the tip flow structure and unsteady response of the leakage flow are presented and discussed. Results indicate that injection periodically pushes the tip leakage vortex and passage shock rearward. The location of the leakage vortex with injection was found to be backward of that without injection, at the near-stall condition of the non-injection case. At the near-stall point of the injection case, however, the tip leakage vortex was at some situations rearward and at some situations forward, as compared with the non-injection case. In other words, the leakage vortex in the injection case oscillates around the location of the time-averaged leakage vortex without injection. This is the situation, at which, the interface between the leakage and oncoming flows tends to become parallel to the leading-edge plane. The effect of injection on the boundary layer separation from the casing wall is also investigated. The rotor operation at in-stall condition for both the injection and non-injection cases is studied and the path into instability is discussed for each case. The propagation of a low-velocity region near the pressure surface and leading-edge of the blade was found to be responsible for the detachment of the passage shock from the leading-edge and upstream movement of the leakage vortex, leading to the occurrence of the leading-edge vortex spillage, for both the injection and non-injection cases.

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