Effect of Endwall Contouring on a Transonic Turbine Blade Passage: Part 1—Aerodynamic Performance

The paper presents a detailed experimental and numerical study on the effect of endwall contouring in a quasi 2D cascade, operating at transonic conditions. Aerodynamic performance of two contoured endwalls are studied and compared with a baseline (planar) endwall. The first contoured endwall was generated with the goal of reducing secondary losses (Aero-Optimized contoured endwall) and the second endwall was generated with the objective of reduced overall heat transfer to the endwall (HT-optimized contoured endwall). Midspan total pressure loss, secondary flow field and static pressure measurements on the airfoil surface were measured. The cascade exit Mach numbers range from 0.71 to 0.95 and the turning angle of the airfoil is ∼127°. The inlet span of the airfoils was reduced with respect to the outlet span with the intention of obtaining a realistic inlet/exit Mach number that is observed in a real engine. 3D viscous compressible CFD analysis was carried out to study the detailed behavior of the complex flow structures that develop as a result of endwall contouring. A 3% reduction in area averaged losses was achieved at 0.1 Cax downstream of the trailing edge and a 17% reduction in mixed out losses was achieved at 1.0 Cax downstream location with the Aero-Optimized contoured endwall.

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A Numerical Study of Aerodynamic Performance and Noise of a Bionic Airfoil Based on Owl Wing

Advances in Mechanical Engineering ◽

10.1155/2014/859308 ◽

2014 ◽

Vol 6 ◽

pp. 859308 ◽

Cited By ~ 3

Author(s):

Xiaomin Liu ◽

Xiang Liu

Keyword(s):

Reynolds Number ◽

Unsteady Flow ◽

Noise Reduction ◽

Numerical Study ◽

Aerodynamic Performance ◽

Centrifugal Fan ◽

Surface Geometry ◽

Airfoil Surface ◽

Acoustic Sources ◽

Bionic Airfoil

Noise reduction and efficiency enhancement are the two important directions in the development of the multiblade centrifugal fan. In this study, we attempt to develop a bionic airfoil based on the owl wing and investigate its aerodynamic performance and noise-reduction mechanism at the relatively low Reynolds number. Firstly, according to the geometric characteristics of the owl wing, a bionic airfoil is constructed as the object of study at Reynolds number of 12,300. Secondly, the large eddy simulation (LES) with the Smagorinsky model is adopted to numerically simulate the unsteady flow fields around the bionic airfoil and the standard NACA0006 airfoil. And then, the acoustic sources are extracted from the unsteady flow field data, and the Ffowcs Williams-Hawkings (FW-H) equation based on Lighthill's acoustic theory is solved to predict the propagation of these acoustic sources. The numerical results show that the lift-to-drag ratio of bionic airfoil is higher than that of the traditional NACA 0006 airfoil because of its deeply concave lower surface geometry. Finally, the sound field of the bionic airfoil is analyzed in detail. The distribution of the A-weighted sound pressure levels, the scaled directivity of the sound, and the distribution of dP/dt on the airfoil surface are provided so that the characteristics of the acoustic sources could be revealed.

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Blade Tip Carving Effects on the Aero-Thermal Performance of a Transonic Turbine

ASME 2013 Turbine Blade Tip Symposium ◽

10.1115/tbts2013-2028 ◽

2013 ◽

Author(s):

C. De Maesschalck ◽

S. Lavagnoli ◽

G. Paniagua

Keyword(s):

High Speed ◽

Thermal Stresses ◽

Numerical Study ◽

Three Dimensional ◽

Turbine Rotor ◽

Geometrical Parameters ◽

Complex Flow ◽

Turbine Rotor Blade ◽

Blade Tip ◽

Transonic Turbine

Tip leakage flows in unshrouded high speed turbines cause large aerodynamic penalties, induce significant thermal loads and give rise to intense thermal stresses onto the blade tip and casing endwalls. In the pursuit of superior engine reliability and efficiency, the turbine blade tip design is of paramount importance and still poses an exceptional challenge to turbine designers. The ever-increasing rotational speeds and pressure loadings tend to accelerate the tip flow velocities beyond the transonic regime. Overtip supersonic flows are characterized by complex flow patterns, which determine the heat transfer signature. Hence, the physics of the overtip flow structures and the influence of the geometrical parameters on the overtip flow require further understanding to develop innovative tip designs. Conventional blade tip shapes are not adequate for such high speed flows and hence, potential for enhanced performances lays in appropriate tip shaping. The present research aims to quantify the prospective gain offered by a fully contoured blade tip shape against conventional geometries such as a flat and squealer tip. A detailed numerical study was conducted on a modern transonic turbine rotor blade (Reynolds number is 5.5 × 105, relative exit Mach number is 0.9) by means of three-dimensional Reynolds-Averaged Navier-Stokes calculations. The novel contoured tip geometry was designed based on a 2D tip shape optimization in which only the upper 2% of the blade span was modified. This study yields a deeper insight into the application of blade tip carving in high speed turbines and provides guidelines for future tip designs with enhanced aerothermal performances.

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Study on aerodynamic optimal super/transonic turbine cascade and its geometry characteristics

Proceedings of the Institution of Mechanical Engineers Part G Journal of Aerospace Engineering ◽

10.1177/0954410016638875 ◽

2016 ◽

Vol 231 (3) ◽

pp. 435-443 ◽

Cited By ~ 2

Author(s):

Lucheng Ji ◽

Jia Yu ◽

Weiwei Li ◽

Weilin Yi

Keyword(s):

Shock Waves ◽

Adjoint Method ◽

Numerical Study ◽

Aerodynamic Performance ◽

Turbine Cascade ◽

Negative Effects ◽

Machine Dynamics ◽

Optimum Aerodynamic ◽

Transonic Turbine ◽

The Relationship

The shock waves are important phenomena in transonic turbines, which cause lots of negative effects on the aerodynamic performance. Much of attention had been paid on reducing the strength of the shock waves via modifying turbine cascade geometry, and it is highly preferred to build experiences on the relationship between the cascade aerodynamic performance and the geometric parameters. The paper presents a numerical study on the aerodynamic optimal transonic turbine cascade and its geometry characteristics. Three typical Russia transonic turbine cascades with different design conditions are selected and optimized using adjoint method at three different back pressures, respectively. Thus, the best geometry parameters for optimum aerodynamic performance can be found. Then the key geometry parameters of optimized cascades are extracted and compared with the original ones. Results show that even the best designs by hands could be less efficient than ones by computer-aided optimizations. Some experiences on how to set the key geometry parameters for a best performance are obtained. The reduced shock profiling is applied to the thermal turbomachinery and machine dynamics transonic turbine by using the adjoint method. The performance of the thermal turbomachinery and machine dynamics transonic turbine was increased significantly.

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Blade Tip Carving Effects on the Aerothermal Performance of a Transonic Turbine

Journal of Turbomachinery ◽

10.1115/1.4028326 ◽

2014 ◽

Vol 137 (2) ◽

Cited By ~ 16

Author(s):

C. De Maesschalck ◽

S. Lavagnoli ◽

G. Paniagua

Keyword(s):

High Speed ◽

Thermal Stresses ◽

Numerical Study ◽

Three Dimensional ◽

Navier Stokes ◽

Geometrical Parameters ◽

Complex Flow ◽

Leakage Flows ◽

Blade Tip ◽

Exit Mach Number

Tip leakage flows in unshrouded high speed turbines cause large aerodynamic penalties, induce significant thermal loads and give rise to intense thermal stresses onto the blade tip and casing endwalls. In the pursuit of superior engine reliability and efficiency, the turbine blade tip design is of paramount importance and still poses an exceptional challenge to turbine designers. The ever-increasing rotational speeds and pressure loadings tend to accelerate the tip flow velocities beyond the transonic regime. Overtip supersonic flows are characterized by complex flow patterns, which determine the heat transfer signature. Hence, the physics of the overtip flow structures and the influence of the geometrical parameters require further understanding to develop innovative tip designs. Conventional blade tip shapes are not adequate for such high speed flows and hence, potential for enhanced performances lays in appropriate tip shaping. The present research aims to quantify the prospective gain offered by a fully contoured blade tip shape against conventional geometries such as a flat and squealer tip. A detailed numerical study was conducted on a modern rotor blade (Reynolds number of 5.5 × 105 and a relative exit Mach number of 0.9) by means of three-dimensional (3D) Reynolds-averaged Navier–Stokes (RANS) calculations. Two novel contoured tip geometries were designed based on a two-dimensional (2D) tip shape optimization in which only the upper 2% of the blade span was modified. This study yields a deeper insight into the application of blade tip carving in high speed turbines and provides guidelines for future tip designs with enhanced aerothermal performances.

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Effect of Airfoil Shape and Turning Angle on Turbine Airfoil Aerodynamic Performance at Transonic Conditions

Volume 6: Fluids and Thermal Systems; Advances for Process Industries, Parts A and B ◽

10.1115/imece2011-62167 ◽

2011 ◽

Author(s):

Santosh Abraham ◽

Kapil Panchal ◽

Srinath V. Ekkad ◽

Wing Ng ◽

Barry J. Brown ◽

...

Keyword(s):

Mach Number ◽

Gas Turbine ◽

Turbine Blades ◽

Flow Characteristics ◽

Airfoil Surface ◽

Turning Angle ◽

Gas Turbine Blades ◽

Mach Numbers ◽

Exit Mach Number ◽

Turbine Airfoils

Performance data for high turning gas turbine blades under transonic Mach numbers is significantly lacking in literature. Performance of three gas turbine airfoils with varying turning angles at transonic flow conditions was investigated in this study. Midspan total pressure loss, secondary flow field and static pressure measurements on the airfoil surface in a linear cascade setting were measured. Airfoil curvature and true chord were varied to change the loading vs. chord for each airfoil. Airfoils A, D and E are designed to operate at different velocity triangles. Velocity triangle requirements (inlet/exit Mach number and gas angles) come from 1D and 2D models that include calibrated loss systems. One of the goals of this study was to use the experimental data to confirm/refine loss predictions for the effect of various Mach numbers and gas turning angles. The cascade exit Mach numbers were varied within a range from 0.6 to 1.1. The airfoil turning angle ranges from 120° to 138°. A realistic inlet/exit Mach number ratio, that is representative of that seen in a real engine, was obtained by reducing the inlet span with respect to the exit span of the airfoil, thereby creating a quasi 2D cascade. In order to compare the experimental results and study the detailed flow characteristics, 3D viscous compressible CFD analysis was also carried out.

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Effect of Endwall Contouring on a Transonic Turbine Blade Passage: Part 2—Heat Transfer Performance

Volume 4: Heat Transfer, Parts A and B ◽

10.1115/gt2012-68405 ◽

2012 ◽

Cited By ~ 4

Author(s):

Kapil V. Panchal ◽

Santosh Abraham ◽

Srinath V. Ekkad ◽

Wing Ng ◽

Andrew S. Lohaus ◽

...

Keyword(s):

Heat Transfer ◽

Heat Transfer Coefficient ◽

Transfer Coefficient ◽

Heat Transfer Performance ◽

Aerodynamic Performance ◽

Transfer Performance ◽

Blade Passage ◽

Transonic Turbine ◽

Endwall Contouring ◽

The Heat Transfer Coefficient

Contouring of turbine endwalls has been widely studied for aerodynamic performance improvement of turbine passages. However, it is equally important to investigate the effect of contouring on endwall heat transfer, because a substantial increase in endwall heat transfer due to contouring will render the design impractical. In this paper, the effect of contouring on endwall heat transfer performance of a high-turning HP-turbine blade passage, operating under transonic exit Mach number conditions, is reported. Three endwall geometries were experimentally investigated at three different passage exit Mach numbers, 0.71, 0.88(design) and 0.95, for their heat transfer performance. One endwall is a non-contoured baseline endwall and the other two are contoured endwall geometries. One of the contoured endwall geometry was generated with the goal of reduction in stagnation pressure losses and the other was generated with the goal of reduced overall heat transfer through the endwall. The experiments were carried out in Virginia Tech’s transient, blow down, transonic linear cascade facility. Endwall surface temperatures were measured using infrared thermography technique. Local heat transfer coefficient values were calculated using the measured temperatures. The heat transfer coefficient values were then related to the endwall geometries using a camera matrix model. The measurement technique and the methodology for the post-processing of the heat transfer coefficient data have been presented in detail. Details of the flow behavior for these endwalls were obtained using CFD simulations and have been used to assist the interpretation of the experimental results. In this study, the heat transfer performance of the contoured endwalls in comparison to the non-contoured baseline case is presented. Both the contoured endwalls demonstrated a significant reduction in the overall average heat transfer coefficient values. The surface heat transfer coefficient distributions also indicated a reduction in the level of hot spots for most of the endwall surface. However, increase in the heat transfer coefficient values was observed especially in the area near the leading edge. The results indicate that, in addition to a probable improvement in aerodynamic performance, endwall contouring may also be used to improve the heat transfer performance of turbine passages. Additionally, aerodynamic behavior of these endwalls is discussed in detail in the companion paper GT2012-68425, “Effect of endwall contouring on a transonic turbine endwall: Part 1 – Aerodynamic performance.”

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Measurement of Aerodynamic Losses for Turbine Airfoil Cascades With Varying Pitch, Operating Under Transonic Conditions

Volume 8: Turbomachinery, Parts A, B, and C ◽

10.1115/gt2012-68424 ◽

2012 ◽

Author(s):

Santosh Abraham ◽

Kapil Panchal ◽

Srinath V. Ekkad ◽

Wing Ng ◽

Andrew S. Lohaus ◽

...

Keyword(s):

Mach Number ◽

Static Pressure ◽

Leading Edge ◽

Experimental Results ◽

Turning Angle ◽

Pressure Losses ◽

Turbine Cascades ◽

Exit Mach Number ◽

End Wall ◽

Secondary Flow Field

The paper presents detailed experimental results of the midspan total pressure losses, secondary flow field, and static pressure measurements on two linear, high-turning turbine cascades at transonic conditions. The airfoils in the two cascades being studied are identical and their aerodynamic loading levels are varied by increasing the pitch of one cascade by 25% with respect to the other. The turbine cascades are referred to as B1-SP and B1-IP. Cascade B1-IP, with its increased pitch, has a Zweifel coefficient increased by 25%. The airfoils have a turning angle of ∼127°. Measurements are made at design and off-design conditions, at exit Mach numbers ranging from 0.71 to 0.95. The exit span of the airfoils are increased relative to the inlet span with the intention of obtaining a ratio of inlet Mach number to exit Mach number that is representative to that encountered in a real engine. This results in one end wall diverging from inlet to exit at a 13 degree angle, which simulates the required leading edge loading as seen in an engine. The objective of this study is to investigate the variation in airfoil loading distribution and the effect it has on aerodynamic performance in terms of pressure losses. Detailed loss measurements, both in the pitchwise as well as spanwise directions, are conducted at 0.1 Cax and 1.0 Cax locations downstream of the trailing edge. Results from 3D viscous numerical simulations have been used to assist the interpretation of experimental results.

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