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film cooling performance and flow structure of single-hole and double-holes with swirling jet

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Chinese Journal of Aeronautics 2022年第3期35卷 201-213页

作者： Rui ZHU Simon TERRENCE Shulei LI Gongnan XIE School of Mechanical Engineering Northwestern Polytechnical UniversityXi’an 710072China Research&Development Institute of Northwestern Polytechnical University in Shenzhen Shenzhen 518057China Department of Mechanical Engineering University of MinnesotaMinneapolisMN 55455USA School of Marine Science and Technology Northwestern Polytechnical UniversityXi’an 710072China

This paper presents the results of a numerical study of the effects of swirling flow in coolant jets on film cooling performance.Some combined-hole designs with swirling coolant flow entering the delivery hole are proposed and analyzed.Adiabatic film cooling effectiveness values for cases with various blowing ratios are compared.Detailed flow structures and underlying mechanisms are discussed.The results show that film cooling effectiveness is improved with jet swirl at high blowing ratios,and that swirl strength has significant influence on film cooling performance.Combined-hole designs can further improve film cooling performance using swirling jets due to mixing of coolant flows and interaction of vortices.The largest improvements of area-averaged film cooling effectiveness for a single-hole swirl case and a combined-hole swirl case over corresponding non-swirling case results are 157%and 173%,respectively.

关键词： Combined-hole film cooling Interactions of vortices Jet-in-cross-flow Swirling flow

Experimental Study on the film cooling Characteristics of Three Complex Tip Structures

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Journal of Thermal Science 2023年第4期32卷 1378-1392页

作者： LIN Juqiang LI Haiwang YOU Ruquan LIU Runzhou TAO Zhi LIU Song School of Energy and Power Engineering Beihang UniversityBeijing100191China Research Institute for Aero-Engines Beihang UniversityBeijing100191China National Key Laboratory of Science and Technology on Aero Engines Aero-thermodynamics Beijing100191China AECC Sichuan Gas Turbine Establishment Sichuan610500China

The turbine blades of aircrafts must be properly cooled to prevent engine failure.Thus,to investigate the influence of the tip structure on the film cooling effect,pressure-sensitive paint test technology was used to determine the adiabatic film cooling effectiveness in this study.The experiment was completed in a cascade comprising three straight blades.The effects of the blowing ratio,density ratio,tip clearance,and tip structure on film cooling efficiency were analyzed.The experimental results demonstrated that,as the blowing ratio increased,the film coverage area and film cooling efficiency increased under most experimental conditions.However,the film cooling efficiency was found to initially increase,and subsequently decrease,as the blowing ratio increased.The respective influences of the density ratio and tip clearance on the film cooling efficiency were found to be significant.The density ratio experiments revealed that a high-density ratio can result in better film coverage than the low-density-ratio air.The tip clearance experimental results indicated that a small tip clearance promotes an increase in film cooling efficiency;this is because the small tip clearance negatively affects the main stream leakage flow,which can reduce the film coverage area.Under the conditions of the Base case 2 configuration,a blowing ratio of 2.1,and a tip clearance of 0.6%h,the average film cooling efficiency of the blade tip was 0.22.Among the three blade tip structures applied in this study,Base case 2 demonstrated higher film cooling efficiency than the other two blade tip structures under the conditions of the same blowing ratio,tip clearance,and density ratio.

关键词： film cooling tip structure pressure-sensitive paint leakage flow

A numerical study of segmented cooling-stream injection in supersonic film cooling

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Chinese Journal of Aeronautics 2022年第6期35卷 156-171页

作者： Hang NI Mingjun WANG Peixue JIANG Wei PENG Yinhai ZHU Institute of Nuclear and New Energy Technology Collaborative Innovation Center of Advanced Nuclear Energy TechnologyKey Laboratory of Advanced Reactor Engineering and Safety of Ministry of EducationTsinghua UniversityBeijing 100084China School of Nuclear Science and Technology Xi’an Jiaotong UniversityXi’an 710049China Department of Energy and Power Engineering Key Laboratory for Thermal Science and Power Engineering of Ministry of EducationTsinghua UniversityBeijing 100084China Tsinghua University-Zhang Jiagang Joint Institute for Hydrogen Energy and Lithium-Ion Battery Technology Tsinghua UniversityBeijing 100084China

The present study proposes a segmented cooling-stream injection structure based on a certain coolant mass flow rate,and numerically investigates the effect of segmented cooling-stream injection on supersonic film cooling.The results indicate that without shock-wave impingement and with helium as the coolant,segmented cooling-stream injection can reduce the mixing between the mainstream and the cooling stream to produce better cooling performance than single injection,especially at larger coolant Mach numbers.However,with nitrogen as the coolant,the cooling effect of the segmented-injection system is very close to that of the single-injection system.Mixing at the impinging region is enhanced significantly when there is an incident shock wave.When the shock wave impinges between the two coolant inlets,segmented cooling-stream injection improves film cooling effectiveness in the midstream and downstream regions more than single injection because only part of the cooling stream undergoes the enhanced mixing effect of the shock wave.The advantage of segmented injection is reduced when the impinging region is behind the second coolant inlet.The further downstream the impinging region,the smaller the associated advantage.

关键词： film cooling Mach number Segmented injection Shock wave Supersonic flow

Experimental investigation of the full coverage film cooling effectiveness of a turbine blade with shaped holes

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Chinese Journal of Aeronautics 2022年第3期35卷 297-308页

作者： Cunliang LIU Fan ZHANG Shuaiqi ZHANG Qingqing SHI Hui SONG School of Power and Energy Northwestern Polytechnical UniversityXi'an 710129China Shaanxi Key Laboratory of Thermal Sciences in Aero-engine System Xi'an 710129China Beijing Institute of Radio Measurement Beijing 100854China

The film cooling effectiveness of two turbine blades at different turbulence intensities(0.62%and 16.00%)and mass flux ratios(2.91%,5.82%,8.73%and 11.63%)is studied by using the Pressure-Sensitive Paint(PSP)measurement technique.There are a baseline and an improved turbine blade in current work,and their film cooling hole position distribution is the same.But the hole shape on suction surface and pressure surface is changed from cylindrical hole(baseline)to laid-back fan-shaped hole(improved blade).Both blades have 5 rows of cylindrical holes at the leading edge and 4 rows of cooling-holes on the suction surface and the pressure surface.The experimental results show that the film cooling effectiveness of the improved blade is much better than the baseline.The increase in turbulence intensity will reduce the cooling effectiveness on the surface of turbine blade,but the effect of turbulence intensity becomes weaker with an increase in the mass flux ratio.Compared with the multiple rows of cylindrical holes,the cooling effectiveness of shaped holes is more influenced by the turbulence intensity at low mass flux ratio.

关键词： Adiabatic film cooling effectiveness film cooling Shaped holes Turbine blade Turbulence intensity

Effects of plasma actuation and hole configuration on film cooling performance

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Propulsion and Power Research 2023年第2期12卷 227-237页

作者： Jie Sun Fuxing Zhang Jin Wang Gongnan Xie Bengt Sundén School of Energy and Environmental Engineering Hebei University of TechnologyTianjin 300401China Key Laboratory of Thermo-Fluid Science and Engineering(Xi’an Jiaotong University) Ministry of EducationXi’an 710049China School of Marine Science and Technology Northwestern Polytechnical UniversityXi’an 710072China Department of Energy Sciences Division of Heat TransferLund UniversityLund SE-22100Sweden

In this paper,plasma actuators are arranged asymmetrically downstream the wall to improve film cooling performance.Effects of blowing ratio,hole configuration and applied voltage on flow characteristics and film cooling effectiveness were investigated numerically on a flat plate.Results show that highest film cooling effectiveness distribution is obtained both in the spanwise and streamwise directions under blowing ratio of 0.5.Average wall film cooling effectiveness of cylindrical hole increases by 251.9%under blowing ratio of 0.5 compared to that under blowing ratio of 1.5.The scale of the counter rotating vortex pairs(CRVP)from fan shaped hole and sister hole are significantly reduced compared to that from cylindrical hole.The console hole has an anti-counter rotating vortex pair(Anti-CRVP),which weakens the entrainment of the CRVP to the coolant air near the wall.Compared with the cylindrical hole,average wall film cooling effectivenesses for fan shaped hole,sister hole and console hole increase by 73.1%,97.5%and 119.9%.The adherent performance of the coolant air is enhanced after applying plasma actuator.The aerodynamic actuation of the plasma results in the rebound of the fluid close to the wall at 24 kV applied voltage.Average wall film cooling effectiveness of the console hole at 12 kV applied voltage is 10.6%higher than that without plasma.

关键词： Aerodynamic actuation film cooling Flow control Hole configurat

Heat Transfer and Flow Structure of a Turbine Blade's Air-cooled Leading Edge Considering Different Hole Shapes and Additional Flow Angles

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Journal of Thermal Science 2024年第4期33卷 1421-1442页

作者： QIN Runxuan ZHOU Xun WANG Songtao CAI Le School of Energy Science and Engineering Harbin Institute of Technology

A numerical study is conducted to elucidate the impact of hole shapes and additional flow angles on the flow structure of the coolant and temperature field in the leading edge area of the gas turbine rotor.Four typical hole shapes are considered for the GE-E3 blade.The impact of the additional flow angle(E) within each hole shape on the temperature field is investigated.The results indicate that for the leading edge area and suction surface,the fan-shaped hole case performs best in decreasing temperatures,with a decrease of about 43 K.This is mainly due to the fact that the fan-shaped hole has the maximum expansion in hole spanwise direction.For the pressure surface,the console hole case performs best in decreasing temperatures,with a maximum reduction of about 47.2 K.The influence of E on the surface temperature at leading edge area varied between the different hole shapes.For the cylinder hole and console hole,the E=-20°case has the lowest area-averaged temperature.Because both the fan-shaped hole and the 7-7-7 shaped hole are expansion holes,the pattern of variation of the leading edge area temperature with increasing E is similar for the fan-shaped hole case and 7-7-7 shaped hole case.The E=20°case shows the lowest spanwise-averaged temperature near the hole outlet,and the E=-20°case shows the lowest spanwise-averaged temperature further downstream.

关键词： film cooling GE-E3 turbine rotor blade conjugate heat transfer leading edge

Interpretation of gas-film cooling against aero-thermal heating for high-speed vehicles

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Applied Mathematics and Mechanics(English Edition) 2016年第12期37卷 1615-1630页

作者： Ming DONG Department of Mechanics Tianjin UniversityTianjin 300072China Tianjin Key Laboratory of Modern Engineering Mechanics Tianjin 300072China

The possible application of the film-cooling technique against aero-thermal heating for surfaces of high-speed flying vehicles is discussed. The technique has been widely used in the heat protection of gas turbine blades. It is shown in this paper that, by applying this technique to high-speed flying vehicles, the working principle is fundamentally different. Numerical simulations for two model problems axe performed to support the argument. Besides the heat protection, the appreciable drag reduction is found to be another favorable effect. For the second model problem, i.e., the gas cooling for an optical window on a sphere cone, the hydrodynamic instability of the film is studied by the linear stability analysis to observe possible occurrence of laminar-turbulent transition.

关键词： film cooling aero-thermal heating numerical simulation laminar-turbulenttransition linear stability analysis

film cooling adiabatic effectiveness measurements of pressure side trailing edge cooling configurations

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Propulsion and Power Research 2015年第4期4卷 190-201页

作者： R.Becchi B.Facchini A.Picchi L.Tarchi D.Coutandin S.Zecchi Department of Industrial Engineering University of FlorenceVia S.Marta 350139FirenzeItaly GE Avio S.r.l. Via I Maggio 9910040Rivalta di TorinoTorinoItaly

Nowadays total inlet temperature of gas turbine is far above the permissible metal temperature;as a consequence,advanced cooling techniques must be applied to protect from thermal stresses,oxidation and corrosion the components located in the high pressure stages,such as the blade trailing edge.A suitable design of the cooling system for the trailing edge has to cope with geometric constraints and aerodynamic demands;state-of-the-art of cooling concepts often use film cooling on blade pressure side:the air taken from last compressor stages is ejected through discrete holes or slots to provide a cold layer between hot mainstream and the blade surface.With the goal of ensuring a satisfactory lifetime of blades,the design of efficient trailing edge film cooling schemes and,moreover,the possibility to check carefully their behavior,are hence necessary to guarantee an appropriate metal temperature distribution.For this purpose an experimental survey was carried out to investigate the film covering performance of different pressure side trailing edge cooling systems for turbine blades.The experimental test section consists of a scaled-up trailing edge model installed in an open loop suction type test rig.Measurements of adiabatic effectiveness distributions were carried out on three trailing edge cooling system configurations.The baseline geometry is composed by inclined slots separated by elongated pedestals;the second geometry shares the same cutback configuration,with an additional row of circular film cooling holes located upstream;the third model is equipped with three rows of in-line film cooling holes.Experiments have been performed at nearly ambient conditions imposing several blowing ratio values and using carbon dioxide as coolant in order to reproduce a density ratio close to the engine conditions(DR¼1.52).To extend the validity of the survey a comparison between adiabatic effectiveness measurements and a prediction by correlative approach was performed to compare the experimental results with 1D methodologi

关键词： Turbine blades Trailing edge film cooling Pressure sensitive paint(PSP)

Influence of shock wave impinging region on supersonic film cooling

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Chinese Journal of Aeronautics 2021年第5期34卷 452-465页

作者： Xiaokai SUN Hang NI Wei PENG Peixue JIANG Yinhai ZHU Institute of Nuclear and New Energy Technology Collaborative Innovation Center of Advanced Nuclear Energy TechnologyKey Laboratory of Advanced Reactor Engineering and Safety of Ministry of EducationTsinghua UniversityBeijing 100084China China Institute of Nuclear Industry Strategy 100048China Department of Energy and Power Engineering Key Laboratory for Thermal Science and Power Engineering of Ministry of EducationTsinghua UniversityBeijing 100084China

Shock waves can significantly affect the film cooling for supersonic flow and shock waves may have different influence when impinging in different regions.The present study numerically compared the results of shock wave impinging in three different regions and analyzed the effect of impinging region.The shock wave generators were located at x/s=5,25,45 with 4°,7°and 10°shock wave incidence.The mainstream Mach number was 3.2 and the coolant Mach number was 1.2 or 1.5.The numerical results illustrated that the shock wave impinged in the further upstream region led to a larger high-pressure region and a larger vortex in the boundary layer.Moreover,placing the shock wave generator upstream resulted in the lower mass fraction of coolant in the downstream region.The velocity in the upstream part of the cooling layer was lower than the midstream and downstream part,which resulted in the less ability to resist the shock wave impingement.Therefore,the upstream impingement deteriorated the cooling performance to a greater extent.The study also manifested that the stronger shock wave had a larger effect on supersonic film cooling.Increasing the coolant inlet Mach number can increase the blowing ratio and reduce the mixing,which was of benefit to improve cooling effect.

关键词： film cooling Impinging region Mach number Shock wave Supersonic flow