Effect of Common Blade Tip Squealer Designs in Terms of Tip Clearance Loss Control

2013 ◽  
Author(s):  
N. Lomakin ◽  
A. Granovskiy ◽  
V. Belkanov ◽  
J. Szwedowicz
Keyword(s):  
Numerical Data ◽  
Suction Side ◽  
Tip Clearance ◽  
Navier Stokes ◽  
Radial Clearance ◽  
Engineering Practice ◽  
Inlet Temperature ◽  
Pressure Side ◽  
Blade Tip ◽  
Loss Control

The increase of new gas turbine’s efficiency is connected with further rise of turbine inlet temperature and sometimes as well pressure. In these conditions, first cooled turbine stages of a gas turbine engine usually consist of freestanding airfoils, which do not use an integrated shroud, to avoid risk of shroud overheating. In order to better control the radial gap leakage flow between the rotating blade tip and turbine casing, special design features of the airfoil tip need to be considered in the design process to meet the best possible stage performance. In the general engineering practice, a blade tip squealer provides opportunities to control tip clearance loss. In this paper several simplified types of the tip squealer design are investigated to determine the most effective loss control. At this stage of the investigation, blade tip cooling was not taken into account, but aerodynamic effects were analysed in detail. Based on the most common designs of the blade tip in the literature, four geometry types were investigated: (i) a flat tip design as the reference baseline solution, (ii) full tip squealer, (iii) partial squealer along the pressure side (PS) wall with a cut-out at the pressure side near the trailing edge (TE) and (iv) partial squealer along the suction side (SS) wall with a cut-out at the suction side near TE. All these cases have been compared among each other for two relative radial gaps (gap to blade height) of 0.6% and 1.36%. The flow calculations were done with a full 3D Navier-Stokes CFD code. For the flat tip and for full squealer designs, numerical results were validated against well-known experimental data measured on the GE-E3 blade cascade test rig found in the open literature. By using the 3D numerical data, the special attention was considered to confirm reliability and predictive credibility of the blade tip flow obtained from the analytical model. The obtained loss values and flow details were compared for all studied cases, providing insight into turbine stage aerodynamics with respect to minimal and maximal radial clearance.

scholarly journals Effects of Tip Clearance Gap and Exit Mach Number on Turbine Blade Tip and Near-Tip Heat Transfer

2013 ◽  
Author(s):  
K. Anto ◽  
S. Xue ◽  
W. F. Ng ◽  
L. J. Zhang ◽  
H. K. Moon
Keyword(s):  
Heat Transfer ◽  
Mach Number ◽  
Turbine Blade ◽  
Leading Edge ◽  
Suction Side ◽  
Tip Clearance ◽  
Pressure Side ◽  
Engine Blade ◽  
Blade Tip ◽  
Exit Mach Number

This study focuses on local heat transfer characteristics on the tip and near-tip regions of a turbine blade with a flat tip, tested under transonic conditions in a stationary, 2-D linear cascade with high freestream turbulence. The experiments were conducted at the Virginia Tech transonic blow-down wind tunnel facility. The effects of tip clearance and exit Mach number on heat transfer distribution were investigated on the tip surface using a transient infrared thermography technique. In addition, thin film gages were used to study similar effects in heat transfer on the near-tip regions at 94% height based on engine blade span of the pressure and suction sides. Surface oil flow visualizations on the blade tip region were carried-out to shed some light on the leakage flow structure. Experiments were performed at three exit Mach numbers of 0.7, 0.85, and 1.05 for two different tip clearances of 0.9% and 1.8% based on turbine blade span. The exit Mach numbers tested correspond to exit Reynolds numbers of 7.6 × 105, 9.0 × 105, and 1.1 × 106 based on blade true chord. The tests were performed with a high freestream turbulence intensity of 12% at the cascade inlet. Results at 0.85 exit Mach showed that an increase in the tip gap clearance from 0.9% to 1.8% translates into a 3% increase in the average heat transfer coefficients on the blade tip surface. At 0.9% tip clearance, an increase in exit Mach number from 0.85 to 1.05 led to a 39% increase in average heat transfer on the tip. High heat transfer was observed on the blade tip surface near the leading edge, and an increase in the tip clearance gap and exit Mach number augmented this near-leading edge tip heat transfer. At 94% of engine blade height on the suction side near the tip, a peak in heat transfer was observed in all test cases at s/C = 0.66, due to the onset of a downstream leakage vortex, originating from the pressure side. An increase in both the tip gap and exit Mach number resulted in an increase, followed by a decrease in the near-tip suction side heat transfer. On the near-tip pressure side, a slight increase in heat transfer was observed with increased tip gap and exit Mach number. In general, the suction side heat transfer is greater than the pressure side heat transfer, as a result of the suction side leakage vortices.

Reduction of Tip Clearance Losses in an Axial Turbine by Shaped Design of the Blade Tip Region

2007 ◽  
Author(s):  
K. Kusterer ◽  
N. Moritz ◽  
D. Bohn ◽  
T. Sugimoto ◽  
R. Tanaka
Keyword(s):  
Numerical Investigation ◽  
Mass Flow ◽  
Suction Side ◽  
Tip Clearance ◽  
Leakage Flow ◽  
Pressure Side ◽  
Aerodynamic Efficiency ◽  
Blade Tip ◽  
Radial Gap ◽  
Tip Clearance Vortex

Secondary flows and leakage flows lead to complex vortex structures in the flow field inside the passages of the vanes and blades in turbo machines. These result in aerodynamic losses and, thus, reduced efficiency. One of the major vortex structures is the tip clearance vortex, which is generated on the airfoil’s suction side due to the leakage flow through the tip clearance, e.g. between rotating blades and casing. This leakage flow is induced by the pressure difference between pressure and suction side. The tip clearance vortex intensity strongly depends on the amount of tip clearance leakage. Thus, the reduction of this leakage mass flow increases the aerodynamic efficiency of a turbo-machine. In gas turbines, two ways are commonly used to influence the tip leakage flow: contouring of the radial gap either at blade tip or endwall, or changing the blade tip geometry by application of squealers or winglets on the blade tip. In this paper, a numerical investigation on the principle physics of a specific blade tip design is presented. On the pressure side the blades are extended in the tip region comparable to winglets (“hook-shaped”). With this change, the structures of the flow entering the gap between blade tip and casing are influenced to achieve a reduction of the mass flow in the radial gap. In this approach, the contour of the blade on the pressure side surface is shaped smoothly so that only a low increase of the local stresses should be expected and the blade is manufactured in one part. Furthermore, the height of the tip clearance is not affected. The new blade tip design is applied to 2nd and 3rd blade of the axial turbine in a test configuration of a KHI industrial gas turbine. Thus, a multi-stage numerical approach has been selected for the numerical investigation. The numerical model includes the flow path, vanes and blades of the 2nd and 3rd stage. The mixing plane technique is used to couple the blocks computed in stationary system of reference and rotating system of reference. The aerodynamic efficiency of the new designed blade tip in the two-stage arrangement is compared to the original design. It shows that a slight increase can be achieved in the static polytropic efficiency of the turbine configuration. The influence of the new design on the flow structures in the tip clearance region of the blades is analysed in detail to explain the mechanisms that cause the efficiency increase.

Modeling of Blade Tip Geometries in an Axial Compressor Stage

2003 ◽  
Author(s):  
Carsten Stockhaus ◽  
Werner Volgmann ◽  
Horst Stoff
Keyword(s):  
Beneficial Effect ◽  
Axial Compressor ◽  
Suction Side ◽  
Tip Clearance ◽  
Leakage Flow ◽  
Efficiency Gain ◽  
Pressure Side ◽  
Compressor Stage ◽  
Tip Leakage Flow ◽  
Blade Tip

The purpose of this paper is to investigate numerically the tip leakage flow for different blade tip geometries in an axial compressor stage under design and off-design conditions. Using flat tips, suction and pressure side squealers in combination with knife tips, a comparison of the rotor performance in terms of pressure and efficiency gain is reported. Detailed flow characteristics within the tip clearance gap, interaction of the leakage flow with the main flow and resultant turning effects at the exit of the row have been investigated. The CFD method is based on a commercially available compressible Navier-Stokes solver (STAR-CD), using a turbulent compressible high Reynolds number k-ε model. Accurate numerical comparison of different blade tip geometries is achieved by using the same grid for the various shapes. The blocking strategy with O-grid structure is presented. The numerical results show clearly the beneficial effect of cutting away material from the pressure side. The higher surface curvature of the suction side squealer affects the pressure blade loading and increases the lift in the same way. This effect is increased by increasing the squealer height and results in a lower efficiency gain near the surge line. The best modification of the blade tip shows a maximum reduction of the tip discharge coefficient of 20 %. This leads to an improved total pressure ratio of 0.29% and an improved total polytropic efficiency of 0.40% under design condition. The influences of favourable squealer geometries on stage characteristics are described along an operating line. With a simulation of IGV-setting from Δα = −15° to Δα = +20° different operating points have been investigated in a swirl performance map. The beneficial effect of the suction side squealer found for the rotor row could assign to the stator row and results in an improved static pressure gain. Furthermore, design indications are presented which help to keep the efficiency gain under surge condition as high as possible.

Stator-Averaged, Rotor Blade-to-Blade Near-Wall Flow in a Multistage Axial Compressor With Tip Clearance Variation

1992 ◽  
Vol 114 (3) ◽  
pp. 668-674 ◽  
Author(s):  
I. N. Moyle ◽  
G. J. Walker ◽  
R. P. Shreeve
Keyword(s):  
Rotor Blade ◽  
Axial Compressor ◽  
Suction Side ◽  
Tip Clearance ◽  
Pressure Side ◽  
Second Stage ◽  
Rotor Wall ◽  
Blade Tip ◽  
Wall Flow ◽  
Multistage Axial Compressor

This paper describes the effect of tip clearance changes on the pressure at the case wall of a second-stage rotor. Wall shear distributions under the rotor tip are also presented. The results show low-pressure areas extending along the rotor suction side but lying away from the blade. Pressure contours indicate the tangential loading at the tip is lower than predicted by two-dimensional calculations; however, the predicted loading is observed between the lowest pressure’s path in the passage and the blade pressure side. The results suggest that a viscous or shearing layer, due to blade-to-wall relative motion, is generated on the blade side of the tip gap, which modifies the inviscid relative flow field and produces an unloading on the blade tip.

Unsteady Analysis on the Effects of Tip Clearance Height on Hot Streak Migration Across Rotor Blade Tip Clearance

2013 ◽  
Author(s):  
Zhaofang Liu ◽  
Zhao Liu ◽  
Zhenping Feng
Keyword(s):  
Rotor Blade ◽  
High Speed ◽  
Stokes Equations ◽  
Tip Clearance ◽  
Navier Stokes ◽  
Inlet Temperature ◽  
Leakage Flow ◽  
Blade Tip ◽  
Hot Streak ◽  
Blade Tip Clearance

This paper presents an investigation on the hot streak migration across rotor blade tip clearance in a high pressure gas turbine with different tip clearance heights. The blade geometry is taken from the first stage of GE-E3 turbine engine. Three tip clearances, 1.0%, 1.5% and 2.5% of the blade span with a flat tip were investigated respectively, and the uniform and non-uniform inlet temperature profiles were taken as the inlet boundary conditions. By solving the unsteady compressible Reynolds-averaged Navier-Stokes equations, the time dependent solutions were obtained. The results indicate that the large tip clearance intensifies the leakage flow, increases the hot streak migration rate, and aggravates the heat transfer environment on blade tip. However, the reverse secondary flow dominated by the relative motion of casing is insensitive to the change of tip clearance height. Attributed to the high-speed rotation of rotor blade and the low pressure difference between both sides of blade, a reverse leakage flow zone emerges over blade tip near trailing edge. To eliminate the effects of blade profile variation due to twist along the blade span on the aerothermal performance in tip clearance, the tested rotor (straight) blade and the original rotor (twisted) blade of GE-E3 first stage with the same tip profile are compared in this paper.

scholarly journals Combustor Exit Temperature Distortion Effects on Heat Transfer and Aerodynamics Within a Rotating Turbine Blade Passage

1990 ◽  
Author(s):  
S. P. Harasgama
Keyword(s):  
Suction Side ◽  
Turbine Rotor ◽  
Secondary Flows ◽  
Navier Stokes ◽  
Inlet Temperature ◽  
Pressure Side ◽  
Radial Temperature ◽  
Pressure Surface ◽  
Hot Gas ◽  
Analytical Work

A numerical simulation of temperature distortion at inlet to a rotating turbine rotor has been performed. A 3D Navier-Stokes code the 2D boundary layer code - STAN5, have been used. The results show that the hot gas is transported to the pressure surface of the blade and that hot gas also migrates to the blade pressure side tip. At locations greater than 50–60% axial chord the hot gas enters the tip-gap and emerges over the suction side. These computations are in agreement with previous experimental results. The secondary flows within the turbine rotor are enhanced by the introduction of inlet radial temperature distortion and this is in accord with previous analytical work. It is shown that the heat flux near the tip region on the pressure side of the blade can be increased by up to 76% due to the re-distribution of the inlet temperature distortion.

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.

scholarly journals Stator Averaged, Rotor Blade-to-Blade Near Wall Flow in a Multistage Axial Compressor With Tip Clearance Variation

1991 ◽  
Author(s):  
I. N. Moyle ◽  
G. J. Walker ◽  
R. P. Shreeve
Keyword(s):  
Rotor Blade ◽  
Axial Compressor ◽  
Suction Side ◽  
Tip Clearance ◽  
Pressure Side ◽  
Second Stage ◽  
Rotor Wall ◽  
Blade Tip ◽  
Wall Flow ◽  
Multistage Axial Compressor

This paper describes the effect of tip clearance changes on the pressure at the case wall of a second stage rotor. Wall shear distributions under the rotor tip are also presented. The results show low pressure areas extending along the rotor suction side but lying away from the blade. Pressure contours indicate the tangential loading at the tip is lower than predicted by two dimensional calculations, however, the predicted loading is observed between the lowest pressure’s path in the passage and the blade pressure side. The results suggest a viscous or shearing layer, due to blade-to-wall relative motion, is generated on the blade side of the tip gap which modifies the inviscid relative flow field and produces an unloading on the blade tip.

Influence of tip clearance and cavity depth on heat transfer in a cutback squealer tip

2020 ◽  
pp. 095441002096469
Author(s):  
Weijie Wang ◽  
Shaopeng Lu ◽  
Hongmei Jiang ◽  
Qiusheng Deng ◽  
Jinfang Teng ◽  
...  
Keyword(s):  
Heat Transfer ◽  
Suction Side ◽  
High Heat ◽  
Tip Clearance ◽  
Tip Leakage Flow ◽  
Trailing Edge ◽  
Cavity Depth ◽  
High Heat Transfer ◽  
Blade Tip ◽  
High Heat Transfer Coefficient

Numerical simulations are conducted to present the aerothermal performance of a turbine blade tip with cutback squealer rim. Two different tip clearance heights (0.5%, 1.0% of the blade span) and three different cavity depths (2.0%, 3.0%, and 6.0% of the blade span) are investigated. The results show that a high heat transfer coefficient (HTC) strip on the cavity floor appears near the suction side. It extends with the increase of tip clearance height and moves towards the suction side with the increase of cavity depth. The cutback region near the trailing edge has a high HTC value due to the flush of over-tip leakage flow. High HTC region shrinks to the trailing edge with the increase of cavity depth since there is more accumulated flow in the cavity for larger cavity depth. For small tip clearance cases, high HTC distribution appears on the pressure side rim. However, high HTC distribution is observed on suction side rim for large tip clearance height. This is mainly caused by the flow separation and reattachment on the squealer rims.

Investigation of a Rotor Blade With Tip Cooling Subject to a Nonuniform Temperature Profile

2021 ◽  
Vol 143 (7) ◽  
Author(s):  
Harika S. Kahveci
Keyword(s):  
Temperature Profile ◽  
Turbine Blade ◽  
Numerical Study ◽  
Turbine Rotor ◽  
Inlet Temperature ◽  
Pressure Side ◽  
Engine Operation ◽  
Pressure Losses ◽  
Blade Tip ◽  
Temperature Levels

Abstract One of the challenges in the design of a high-pressure turbine blade is that a considerable amount of cooling is required so that the blade can survive high temperature levels during engine operation. Another challenge is that the addition of cooling should not adversely affect blade aerodynamic performance. The typical flat tips used in designs have evolved into squealer form that implements rims on the tip, which has been reported in several studies to achieve better heat transfer characteristics as well as to decrease pressure losses at the tip. This paper demonstrates a numerical study focusing on a squealer turbine blade tip that is operating in a turbine environment matching the typical design ratios of pressure, temperature, and coolant blowing. The blades rotate at a realistic rpm and are subjected to a turbine rotor inlet temperature profile that has a nonuniform shape. For comparison, a uniform profile is also considered as it is typically used in computational studies for simplicity. The effect of tip cooling is investigated by implementing seven holes on the tip near the blade pressure side. Results confirm that the temperature profile nonuniformity and the addition of cooling are the drivers for loss generation, and they further increase losses when combined. Temperature profile migration is not pronounced with a uniform profile but shows distinct features with a nonuniform profile for which hot gas migration toward the blade pressure side is observed. The blade tip also receives higher coolant coverage when subject to the nonuniform profile.

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