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

2017 ◽  
Author(s):  
Subbaramu Shivaramaiah ◽  
Quamber H. Nagpurwala ◽  
Mahesh K. Varpe ◽  
H. K. Narahari
Keyword(s):  
Pressure Loss ◽  
Total Pressure ◽  
Axial Compressor ◽  
Suction Side ◽  
Total Pressure Loss ◽  
Pressure Side ◽  
Compressor Stage ◽  
Tip Leakage ◽  
Compressor Rotor ◽  
Overall Performance

Winglets are plane surfaces with certain thickness and different shapes. Winglets are used in aircraft to reduce wing tip vortex which is created due to differential pressure in between pressure surface and suction surface. In transonic axial compressor, rotor tip leakage vortex interaction with shock layer and shroud boundary layer leads to total pressure loss and initiation of stall phenomenon. Effect of tip winglets are investigated in compressor rotor cascade. Cascade investigation shows that rotor tip winglet are able to reduce total pressure loss due to tip leakage flow and blade passage secondary flow. Cascade studies are performed with winglet on blade suction side, pressure side and combination of both. From cascade studies it is revealed that suction side winglet are aerodynamically better than pressure side and combined winglets. Owing to favorable results of tip winglet on compressor cascade performance, it was assumed that tip winglets would enhance overall performance of transonic compressor stage with rotating rotor. Results of present CFD simulations have predicted both positive and negative effects of winglets. Effect of different winglet configurations on pressure side and suction side of rotor blade tip are investigated to analyze the compressor stage overall performance. Rotor tip winglets are able to increase stage total pressure ratio compare to the baseline stage without winglet. Stage with winglets have shown better performance in choke region. Winglets are able to vary rotor blade loading from hub to tip region. Presence of winglet has shown ability to reduce to total pressure loss in trailing edge wake region. Stall margin is decreased in compressor stage with winglets due to more blockage towards trailing edge in tip region.

Effect of Tip Gap Variation on the Performance of the Transonic Fan Stage With Tandem Stator

2019 ◽  
Author(s):  
K. Ananthakrishnan ◽  
Shyama Prasad Das ◽  
B. V. S. S. S. Prasad
Keyword(s):  
Pressure Loss ◽  
Total Pressure ◽  
Axial Compressor ◽  
Compressor Stage ◽  
Tip Leakage ◽  
Pressure Loss Coefficient ◽  
Overall Performance ◽  
Transonic Fan ◽  
Total Pressure Loss Coefficient ◽  
Transonic Rotor

Abstract The main goal of modern axial compressor development is to increase the power to weight ratio with higher efficiency. In the present investigation, highly loaded single stage axial compressor with tandem stator vanes is used. Tandem vanes help in attaining the compact compressor stage along with high pressure loading. It is designed for a stage pressure ratio of 2, mass flow rate of 9.02 kg/s operating at 30800 rpm resulting in transonic flow field. The aerodynamic performance of this compressor detoriates due to the tip leakage and secondary flows. Steady-state numerical investigation is carried out to study the flow structures near the tip region of transonic rotor and how different tip gaps influence the overall performance of the compressor. Further the effects of tip leakage flow variation on the performance of tandem vanes are also highlighted. Transonic fan stage with baseline tip gap of 0.5mm is analyzed along with different tip clearance values ranging from 0 % to 3 % of axial chord. Three-dimensional viscous Reynolds Averaged Navier Stokes (RANS) equations are solved using SST k-ω turbulence model. Computational domain discretized with high quality hexahedral elements (Y+ < 2) in AUTOGRID, Numeca. The numerical procedure is verified against the experimental results of Rotor37 transonic rotor test case. Tip leakage losses contribute a substantial amount to the total loss of stage. Overall performance and the stall characteristics for the compressor stage has been evaluated for different tip gap variations.. Further, the topological properties are exploited to visualize the critical points and separation lines on rotor and tandem vanes. Increase in rotor total pressure loss coefficient is observed with increasing tip gap. In contrary, overall total pressure loss coefficient improves for smaller tip gap values and then detoriates. It is observed optimum tip gap height lies close to the 1.125mm, 2% of baseline design value.

Aerodynamic optimisation of a winglet-cavity tip in a high-pressure axial turbine cascade

2016 ◽  
Vol 232 (4) ◽  
pp. 649-663 ◽  
Author(s):  
Zhihua Zhou ◽  
Shaowen Chen ◽  
Songtao Wang
Keyword(s):  
High Pressure ◽  
Mass Flow ◽  
Pressure Loss ◽  
Total Pressure ◽  
Suction Side ◽  
Tip Clearance ◽  
Total Pressure Loss ◽  
Pressure Side ◽  
Tip Clearance Flow ◽  
Clearance Flow

Tip clearance flow between rotating blades and the stationary casing in high-pressure turbines is very complex and is one of the most important factors influencing turbine performance. The rotor with a winglet-cavity tip is often used as an effective method to improve the loss resulting from the tip clearance flow. In this study, an aerodynamic geometric optimisation of a winglet-cavity tip was carried out in a linear unshrouded high-pressure axial turbine cascade. For the purpose of shaping the efficient winglet geometry of the rotor tip, a novel parameterisation method has been introduced in the optimisation procedure based on the computational fluid dynamics simulation and analysis. The reliability of a commercial computational fluid dynamics code with different turbulence models was first validated by contrasting with the experimental results, and the numerical total pressure loss and flow angle using the Baseline k-omega Model (BSL κ-ω model) shows a better agreement with the test data. Geometric parameterisation of blade tips along the pressure side and suction side was adopted to optimise the tip clearance flow, and an optimal winglet-cavity tip was proven to achieve lower tip leakage mass flow rate and total pressure loss than the flat tip and cavity tip. Compared to the numerical results of flat tip and cavity tip, the optimised winglet-cavity design, with the winglet along the pressure side and suction side, had lower tip leakage mass flow rate and total pressure loss. It offered a 35.7% reduction in the change ratio [Formula: see text]. In addition, the optimised winglet along pressure side and suction side, respectively, by using the parameterisation method was studied for investigating the individual effect of the pressure-side winglet and suction-side winglet on the tip clearance flow. It was found that the suction-side extension of the optimal winglet resulted in a greater reduction of aerodynamic loss and leakage mass flow than the pressure-side extension of the optimal winglet. Moreover, with the analysis based on the tip flow pattern, the numerical results show that the pressure-side winglet reduced the contraction coefficient, and the suction-side winglet reduced the aerodynamic loss effectively by decreasing the driving pressure difference near the blade tips, the leakage flow velocity, and the interaction between the leakage flow and the main flow. Overall, a better aerodynamic performance can be obtained by adopting the pressure-side and suction-side winglet-cavity simultaneously.

An Experimental and Numerical Investigation Into the Effects of Squealer Blade Tip Modifications on Aerodynamic Performance

2013 ◽  
Author(s):  
G. A. Ledezma ◽  
J. Allen ◽  
R. S. Bunker
Keyword(s):  
Pressure Loss ◽  
Total Pressure ◽  
Turbine Blades ◽  
Suction Side ◽  
Aerodynamic Performance ◽  
Tip Clearance ◽  
Total Pressure Loss ◽  
Pressure Side ◽  
Numerical Predictions ◽  
Blade Tip

Gas turbine blades using the so-called squealer tip configuration represent a majority of the high-pressure first stage blades in service. The squealer tip in its most basic format is simply a two-tooth labyrinth seal projecting from the blade tip towards the stationary shroud or casing. As with all blade tip configurations, the geometry is a compromise between aerodynamics, cooling, mechanical stress, durability, and repair. While many proposed blade tip innovations involve more complex geometries, this study seeks to determine if a simpler geometry, other than a flat tip, can provide equivalent aerodynamic performance with a reasonable chance of satisfying all other design factors. Using an annular sector blade cascade, total pressure loss surveys are measured with three blade tip geometries, the standard squealer tip, a single-sided suction side seal strip, and the single-sided strip with a pressure side winglet added. The same cascade is modeled numerically as a periodic passage for each of the geometries tested. Experiment and simulation both utilize all blade tip cooling flow injection locations and nominal magnitudes, as well as a constant tip clearance above the suction side seal strip. Experimental data show that the removal of the pressure side seal strip reduces the area-averaged total pressure loss slightly, while the addition of a winglet returns the performance to the baseline result. Numerical predictions indicate essentially equal performance for all geometries. The numerical results provide insight into the loss mechanisms of both the tip leakage flows and the coolant injection flows. This study, when combined with literature data on heat transfer and cooling, concludes that the simpler single-sided suction seal strip is better overall than the commonly employed squealer tip.

Tip Leakage Flows Near Partial Squealer Rims in an Axial Flow Turbine Stage

2003 ◽  
Author(s):  
Cengiz Camci ◽  
Debashis Dey ◽  
Levent Kavurmacioglu
Keyword(s):  
Total Pressure ◽  
Axial Flow ◽  
Suction Side ◽  
Leakage Flow ◽  
Pressure Side ◽  
Turbine Stage ◽  
Tip Leakage Flow ◽  
Edge Zone ◽  
Tip Leakage ◽  
Partial Length

This paper deals with an experimental investigation of aerodynamic characteristics of full and partial-length squealer rims in a turbine stage. Full and partial-length squealer rims are investigated separately on the pressure side and on the suction side in the “Axial Flow Turbine Research Facility” (AFTRF) of the Pennsylvania State University. The streamwise length of these “partial squealer tips” and their chordwise position are varied to find an optimal aerodynamic tip configuration. The optimal configuration in this cold turbine study is defined as the one that is minimizing the stage exit total pressure defect in the tip vortex dominated zone. A new “channel arrangement” diverting some of the leakage flow into the trailing edge zone is also studied. Current results indicate that the use of “partial squealer rims” in axial flow turbines can positively affect the local aerodynamic field by weakening the tip leakage vortex. Results also show that the suction side partial squealers are aerodynamically superior to the pressure side squealers and the channel arrangement. The suction side partial squealers are capable of reducing the stage exit total pressure defect associated with the tip leakage flow to a significant degree.

Effects of Tip Clearance Size on Flowfield in a Low-Speed Axial Compressor Rotor

2014 ◽  
Vol 543-547 ◽  
pp. 158-163 ◽  
Author(s):  
Zhen Zhen Duan ◽  
Yang Wei Liu ◽  
Li Peng Lu
Keyword(s):  
Flow Condition ◽  
Axial Compressor ◽  
Tip Clearance ◽  
Primary Role ◽  
Tip Leakage Vortex ◽  
Low Speed ◽  
Tip Leakage ◽  
Compressor Rotor ◽  
Overall Performance ◽  
Corner Vortex

The simulations of a low-speed axial compressor rotor with two tip clearance sizes, 0.5% chord and 1.5% chord, were performed in the study. Overall performance and detailed flow fields at near stall condition are analyzed. The results show that the rotor stall occurs at higher mass flow condition with large tip clearance. For the small tip clearance the tip leakage vortex and the corner vortex both contribute significantly to the rotor stall, and the interaction between the vortices promotes the stall generation. While for the large tip clearance the tip leakage vortex plays a primary role, and the vortices interaction is ignorable.

Unsteady Simulation of an Axial Compressor Stage With Casing and Blade Passive Treatments

2009 ◽  
Vol 131 (2) ◽  
Author(s):  
Nicolas Gourdain ◽  
Francis Leboeuf
Keyword(s):  
Boundary Layer ◽  
Passive Control ◽  
Axial Compressor ◽  
Suction Side ◽  
Boundary Layer Separation ◽  
Leakage Flow ◽  
Compressor Stage ◽  
Tip Leakage Flow ◽  
Tip Leakage ◽  
Good Ability

This paper deals with the numerical simulation of technologies to increase the compressor performances. The objective is to extend the stable operating range of an axial compressor stage using passive control devices located in the tip region. First, the behavior of the tip leakage flow is investigated in the compressor without control. The simulation shows an increase in the interaction between the tip leakage flow and the main flow when the mass flow is reduced, a phenomenon responsible for the development of a large flow blockage region at the rotor leading edge. A separation of the rotor suction side boundary layer is also observed at near stall conditions. Then, two approaches are tested in order to control these flows in the tip region. The first one is a casing treatment with nonaxisymmetric slots. The method showed a good ability to control the tip leakage flow but failed to reduce the boundary layer separation on the suction side. However, an increase in the operability was observed but with a penalty for the efficiency. The second approach is a blade treatment that consists of a longitudinal groove built in the tip of each rotor blade. The simulation pointed out that the device is able to control partially all the critical flows with no penalty for the efficiency. Finally, some recommendations for the design of passive treatments are presented.

Unsteady Flow Behavior due to Breakdown of Tip Leakage Vortex in an Axial Compressor Rotor at Near-Stall Condition

2000 ◽  
Author(s):  
Masato Furukawa ◽  
Kazuhisa Saiki ◽  
Kazutoyo Yamada ◽  
Masahiro Inoue
Keyword(s):  
Unsteady Flow ◽  
Three Dimensional ◽  
Axial Compressor ◽  
Surface Boundary ◽  
Pressure Side ◽  
Suction Surface ◽  
Tip Leakage Vortex ◽  
Tip Leakage ◽  
Pressure Surface ◽  
Compressor Rotor

The unsteady flow nature caused by the breakdown of the tip leakage vortex in an axial compressor rotor at near-stall conditions has been investigated by unsteady three-dimensional Navier-Stokes flow simulations. The simulations show that the spiral-type breakdown of the tip leakage vortex occurs inside the rotor passage at the near-stall conditions. Downstream of the breakdown onset, the tip leakage vortex twists and turns violently with time, thus interacting with the pressure surface of the adjacent blade. The motion of the vortex and its interaction with the pressure surface are cyclic. The vortex breakdown causes significant changes in the nature of the tip leakage vortex, which result in the anomalous phenomena in the time-averaged flow fields near the tip at the near-stall conditions: no rolling-up of the leakage vortex downstream of the rotor, disappearance of the casing wall pressure trough corresponding to the leakage vortex, large spread of the low-energy fluid accumulating on the pressure side, and large pressure fluctuation on the pressure side. As the flow rate is decreased, the movement of the tip leakage vortex due to its breakdown becomes so large that the leakage vortex interacts with the suction surface as well as the pressure one. The interaction with the suction surface gives rise to the three-dimensional separation of the suction surface boundary layer.

Aerodynamic Performance of Suction-Side Gill Region Film Cooling

2010 ◽  
Vol 132 (3) ◽  
Author(s):  
Justin Chappell ◽  
Phil Ligrani ◽  
Sri Sreekanth ◽  
Terry Lucas ◽  
Edward Vlasic
Keyword(s):  
Mach Number ◽  
Film Cooling ◽  
Pressure Loss ◽  
Total Pressure ◽  
Suction Side ◽  
Total Pressure Loss ◽  
Pressure Loss Coefficient ◽  
Blowing Ratio ◽  
Exit Mach Number ◽  
Total Pressure Loss Coefficient

The performance of suction-side gill region film cooling is investigated using the University of Utah transonic wind tunnel and a simulated turbine vane in a two-dimensional cascade. The effects of film cooling hole orientation, shape, and number of rows, and their resulting effects on the aerodynamic losses, are considered for four different hole configurations: round axial (RA), shaped axial (SA), round radial (RR), and round compound (RC). The mainstream Reynolds number based on axial chord is 500,000, exit Mach number is 0.35, and the tests are conducted using the first row of holes, or both rows of holes at blowing ratios of 0.6 and 1.2. Carbon dioxide is used as the injectant to achieve density ratios of 1.77–1.99 similar to values present in operating gas turbine engines. Presented are the local distributions of total pressure loss coefficient, local normalized exit Mach number, and local normalized exit kinetic energy. Integrated aerodynamic losses (IAL) increase anywhere from 4% to 45% compared with a smooth blade with no film injection. The performance of each hole type depends on the airfoil configuration, film cooling configuration, mainstream flow Mach number, number of rows of holes, density ratio, and blowing ratio, but the general trend is an increase in IAL as either the blowing ratio or the number of rows of holes increase. In general, the largest total pressure loss coefficient Cp magnitudes and the largest IAL are generally present at any particular wake location for the RR or SA configurations, regardless of the film cooling blowing ratio and number of holes. The SA holes also generally produce the highest local peak Cp magnitudes. IAL magnitudes are generally lowest with the RA hole configuration. A one-dimensional mixing loss correlation for normalized IAL values is also presented, which matches most of the both rows data for RA, SA, RR, and RC hole configurations. The equation also provides good representation of the RA, RC, and RR first row data sets.

Experimental Investigation of the Clocking Effect in a 1.5-Stage Axial Turbine—Part I: Time-Averaged Results

2009 ◽  
Vol 131 (2) ◽  
Author(s):  
Sven König ◽  
Bernd Stoffel ◽  
M. Taher Schobeiri
Keyword(s):  
Pressure Loss ◽  
Total Pressure ◽  
Separation Bubble ◽  
Measurement Techniques ◽  
Suction Side ◽  
Total Pressure Loss ◽  
Experimental Investigations ◽  
Flow Angle ◽  
Factors Affecting ◽  
Lp Turbine

Comprehensive experimental investigations were conducted to get deeper insight into the physics of stator clocking in turbomachines. Different measurement techniques were used to investigate the influence of varying clocking positions on the highly unsteady flow field in a 1.5-stage axial low-pressure (LP) turbine. A Reynolds number typical for LP turbines as well as a two-dimensional blade design were chosen. Stator 2 was developed as a high-lift profile with a separation bubble on the suction side. This paper presents the results that were obtained by means of static pressure tappings and five-hole probes as well as the time-averaged results of unsteady x-wire measurements. The probes were traversed in different measuring planes for ten clocking positions. Depending on the clocking position, a variation in total pressure loss for Stator 2, a change of the rotor exit flow angle, and a dependency of the Stator 2 exit flow angle were found. The influence of these parameters on turbine efficiency was studied. Three main factors affecting the total pressure loss could be separated: the size of the separation bubble, the production of turbulent kinetic energy, and the strength of the periodic fluctuations downstream of Stator 2.

Design and Numerical Investigation of Advanced Radial Inlet for a Centrifugal Compressor Stage

2004 ◽  
Author(s):  
Yunbae Kim ◽  
Jay Koch
Keyword(s):  
Centrifugal Compressor ◽  
Pressure Loss ◽  
Total Pressure ◽  
Aerodynamic Performance ◽  
Total Pressure Loss ◽  
Compressor Stage ◽  
Flow Distortion ◽  
Centrifugal Compressor Stage ◽  
Initial Design ◽  
Wide Range

The performance of a centrifugal compressor stage can be seriously affected by inlet flow distortions due to an unsatisfactory inlet configuration and the resulting flow structure. In this study, two radial inlets were designed for a centrifugal compressor stage and investigated numerically using a commercially available 3D viscous Navier-Stokes code. The intent of the design was to minimize the total pressure loss across the inlet while distributing the flow as equally and uniformly as possible to the impeller inlet. For each inlet model, the aerodynamic performance was calculated from the simulation results and then the results from both models were evaluated and compared. The second radial inlet design outperformed the initial design in terms of total pressure loss, flow distortion and uniformity at the impeller inlet. Furthermore, the aerodynamic performance of the second radial inlet was insensitive to a wide range of mass flow rates compared to the initial design due to the distinctive geometric features implemented for the second inlet design.

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