Arrangement Effects of 30 deg Inclined Teardrop-Shaped Dimples on Film Cooling Flow Over Dimpled Cutback Surface at Airfoil Trailing Edge Investigated by 2D3C-PTV

2020 ◽  
Vol 142 (4) ◽  
Author(s):  
Shohei Yamamoto ◽  
Akira Murata ◽  
Hiroshi Saito ◽  
Kaoru Iwamoto
Keyword(s):  
Nusselt Number ◽  
Cross Sections ◽  
Film Cooling ◽  
High Speed ◽  
Reynolds Shear Stress ◽  
Line Arrangement ◽  
Recirculation Flow ◽  
Cooling Flow ◽  
Staggered Arrangement ◽  
Component Particle

Abstract The objective of this study is to determine the differences in flow fields between the 30 deg in-line and staggered arrangements of teardrop-shaped dimples, and to explain why the surface-averaged Nusselt number with the 30 deg in-line arrangement was 28.7% higher than that with the 30 deg staggered arrangement in our previous study. Measurements of the instantaneous velocity fields over the dimpled cutback surfaces in the two arrangements were performed at five spanwise cross sections using two-dimensional three-component particle tracking velocimetry (2D3C-PTV). Recirculation flows were observed only inside the dimples in the in-line arrangement, and the region above the recirculation flows exhibited a higher Reynolds shear stress. In this region, turbulent mixing between the high-speed cooling-flow and the low-speed recirculation-flow can be promoted. Streamlines of the time-averaged velocities showed that approximately half the fluid flowing out of a teardrop-shaped dimple in the in-line arrangement hardly flowed into the ones downstream. The remainder of the fluid mostly flowed into the dimple immediately downstream, and the inflow of the fluid into further downstream dimples decreased gradually. From the PTV results, we can deduce that the fluid motions in the in-line arrangement leads to a larger temperature-difference between the dimple wall and the fluid because the inflow of fluid heated inside upstream dimples into the downstream ones is less than in the staggered arrangement. Consequently, the Nusselt number in the in-line arrangement was higher.

Experimental Investigation Into the Impact of Crossflow on the Coherent Unsteadiness Within Film Cooling Flows

2011 ◽  
Author(s):  
Richard J. Fawcett ◽  
Andrew P. S. Wheeler ◽  
Li He ◽  
Rupert Taylor
Keyword(s):  
Film Cooling ◽  
High Speed ◽  
High Speed Photography ◽  
Cooling Flows ◽  
Pressure Surface ◽  
Cooling Flow ◽  
Cooling Hole ◽  
Crossflow Velocity ◽  
The Impact ◽  
First Time

It is known that the mixing of a film cooling flow with the main turbine passage flow is an unsteady process, with coherent unsteady features occurring across a range of blowing ratios. Upon an aero engine the cooling holes on a turbine blade commonly have a crossflow at the hole inlet. Previous work has shown that crossflow at the hole inlet modifies the time-mean flowfield downstream of a cooling hole compared to the case without crossflow. The current paper investigates the impact of spanwise orientated crossflow on the coherent unsteadiness within film cooling flows. Both cylindrical and fan-shaped holes, located on a blade pressure surface, are studied. The range of blowing ratios considered is 0.7 to 1.8 and the crossflow velocity is up to 0.8 times the bulk jet velocity. High Speed Photography and Hot Wire Anemometry are used to observe the presence of coherent unsteadiness, both immediately downstream of the hole exit and within the cooling hole tube. The results show that the coherent unsteadiness downstream of the hole exit is persistent and its occurrence is not significantly affected by the magnitude of spanwise crossflow. Within the cooling hole tube the existence of coherent unsteadiness is presented for the first time, inside both cylindrical and fan-shaped holes, with a Strouhal number of 0.6 to 0.8. The pattern of this in-hole coherent unsteadiness is seen to change with increasing the crossflow velocity.

Aerothermal Performance of a Cooled Winglet at Engine Representative Mach and Reynolds Numbers

2012 ◽  
Vol 135 (1) ◽  
Author(s):  
D. O. O’Dowd ◽  
Q. Zhang ◽  
L. He ◽  
B. C. Y. Cheong ◽  
I. Tibbott
Keyword(s):  
Nusselt Number ◽  
Film Cooling ◽  
High Speed ◽  
Tip Clearance ◽  
Test Facility ◽  
Pressure Probe ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Aerodynamic Loss ◽  
Loss Coefficients

This paper presents an experimental investigation of the aerothermal performance of a cooled winglet tip under transonic conditions (exit Mach number of 1.0, and an exit Reynolds number of 1.27 × 106, based on axial chord). Spatially resolved heat transfer data and film cooling effectiveness data are obtained using the transient infrared thermography technique in the Oxford High-Speed Linear Cascade test facility. Aerodynamic loss data are obtained by traversing a specially made and calibrated three-hole pressure probe and a single-hole probe one axial chord downstream of the blade. Detailed contours of Nusselt number show that for an increase in tip clearance, with and without film cooling, and for coolant injection, for both tip clearances, the Nusselt number increases. Also the smaller tip clearance observes higher film cooling effectiveness overall. Detailed distributions of kinetic energy losses as well as pitch-wise averaged loss coefficients and loss coefficients at a mixed-out plane indicate that the size of the loss core associated with the over-tip leakage vortex decreases with cooling injection.

Aero-Thermal Performance of a Cooled Winglet at Engine Representative Mach and Reynolds Numbers

2011 ◽  
Author(s):  
D. O. O’Dowd ◽  
Q. Zhang ◽  
L. He ◽  
B. C. Y. Cheong ◽  
I. Tibbott
Keyword(s):  
Nusselt Number ◽  
Thermal Performance ◽  
Film Cooling ◽  
High Speed ◽  
Tip Clearance ◽  
Test Facility ◽  
Pressure Probe ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Loss Coefficients

This paper presents an experimental investigation of the aero-thermal performance of a cooled winglet tip, under transonic conditions (exit Mach number of 1.0, and an exit Reynolds number of 1.27×106, based on axial chord). Spatially-resolved heat transfer data and film cooling effectiveness data are obtained using the transient infrared thermography technique in the Oxford High-Speed Linear Cascade test facility. Aerodynamic loss data are obtained by traversing a specially-made and calibrated three-hole pressure probe and a single-hole probe one axial chord downstream of the blade. Detailed contours of Nusselt number show that for an increase in tip clearance, with and without film cooling, and for coolant injection, for both tip clearances, the Nusselt number increases. Also the smaller tip clearance observes higher film cooling effectiveness overall. Detailed distributions of kinetic energy losses as well as pitch-wise averaged loss coefficients and loss coefficients at a mixed-out plane indicate that the size of the loss core associated with the over-tip leakage vortex decreases with cooling injection.

scholarly journals Investigations on the effect of different influencing factors on film cooling effectiveness under the injection of synthetic coolant

2021 ◽  
Vol 11 (1) ◽  
Author(s):  
Jianlong Chang ◽  
Xinlei Duan ◽  
Yang Du ◽  
Baoquan Guo ◽  
Yutian Pan
Keyword(s):  
Film Cooling ◽  
Elliptical Hole ◽  
Synthetic Jet ◽  
Cooling Effect ◽  
Flow Structures ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Cooling Flow ◽  
Cooling Method ◽  
Strong Controllability

AbstractBy combining the synthetic jet and film cooling, the incident cooling flow is specially treated to find a better film cooling method. Numerical simulations of the synthetic coolant ejected are carried out for analyzing the cooling performance in detail, under different blowing ratios, hole patterns, Strouhal numbers, and various orders of incidence for the two rows of holes. By comparing the flow structures and the cooling effect corresponding to the synthetic coolant and the steady coolant fields, it is found that within the scope of the investigations, the best cooling effect can be obtained under the incident conditions of an elliptical hole with the aspect ratio of 0.618, the blow molding ratio of 2.5, and the Strouhal number St = 0.22. Due to the strong controllability of the synthetic coolant, the synthetic coolant can be controlled through adjusting the frequency of blowing and suction, so as to change the interaction between vortex structures for improving film cooling effect in turn. As a result, the synthetic coolant ejection is more advisable in certain conditions to achieve better outcomes.

Numerical simulation of swirling flow effect on the first stage vane film cooling distribution

2016 ◽  
Vol 07 (03) ◽  
pp. 1650031
Author(s):  
Hong Yin
Keyword(s):  
Flow Field ◽  
Film Cooling ◽  
Swirling Flow ◽  
Cooling System ◽  
Leading Edge ◽  
Suction Side ◽  
Local Area ◽  
Swirl Flow ◽  
Flow Effect ◽  
Recirculation Flow

In advanced gas turbine technology, lean premixed combustion is an effective strategy to reduce peak temperature and thus, NO[Formula: see text] emissions. The swirler is adopted to establish recirculation flow zone, enhancing mixing and stabilizing the flame. Therefore, the swirling flow is dominant in the combustor flow field and has impact on the vane. This paper mainly investigates the swirling flow effect on the turbine first stage vane cooling system by conducting a group of numerical simulations. Firstly, the numerical methods of turbulence modeling using RANS and LES are compared. The computational model of one single swirl flow field is considered. Both the RANS and LES results give reasonable recirculation zone shape. When comparing the velocity distribution, the RANS results generally match the experimental data but fail to at some local area. The LES modeling gives better results and more detailed unsteady flow field. In the second step, the RANS modeling is incorporated to investigate the vane film cooling performance under the swirling inflow boundary condition. According to the numerical results, the leading edge film cooling is largely altered by the swirling flow, especially for the swirl core-leading edge aligned case. Compared to the pressure side, the suction side film cooling is more sensitive to the swirling flow. Locally, the film cooling jet is lifted and turned by the strong swirling flow.

scholarly journals Aerodynamic Losses in Turbines with and without Film Cooling, as Influenced by Mainstream Turbulence, Surface Roughness, Airfoil Shape, and Mach Number

2012 ◽  
Vol 2012 ◽  
pp. 1-28 ◽  
Author(s):  
Phil Ligrani
Keyword(s):  
Surface Roughness ◽  
Mach Number ◽  
Film Cooling ◽  
Total Pressure ◽  
High Speed ◽  
Density Ratio ◽  
Aerodynamic Loss ◽  
Pressure Losses ◽  
Loss Coefficients ◽  
Exit Mach Number

The influences of a variety of different physical phenomena are described as they affect the aerodynamic performance of turbine airfoils in compressible, high-speed flows with either subsonic or transonic Mach number distributions. The presented experimental and numerically predicted results are from a series of investigations which have taken place over the past 32 years. Considered are (i) symmetric airfoils with no film cooling, (ii) symmetric airfoils with film cooling, (iii) cambered vanes with no film cooling, and (iv) cambered vanes with film cooling. When no film cooling is employed on the symmetric airfoils and cambered vanes, experimentally measured and numerically predicted variations of freestream turbulence intensity, surface roughness, exit Mach number, and airfoil camber are considered as they influence local and integrated total pressure losses, deficits of local kinetic energy, Mach number deficits, area-averaged loss coefficients, mass-averaged total pressure loss coefficients, omega loss coefficients, second law loss parameters, and distributions of integrated aerodynamic loss. Similar quantities are measured, and similar parameters are considered when film-cooling is employed on airfoil suction surfaces, along with film cooling density ratio, blowing ratio, Mach number ratio, hole orientation, hole shape, and number of rows of holes.

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