WO2020199370A1 - Aube de turbine de moteur aéronautique présentant une structure de refroidissement - Google Patents

Aube de turbine de moteur aéronautique présentant une structure de refroidissement Download PDF

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Publication number
WO2020199370A1
WO2020199370A1 PCT/CN2019/091937 CN2019091937W WO2020199370A1 WO 2020199370 A1 WO2020199370 A1 WO 2020199370A1 CN 2019091937 W CN2019091937 W CN 2019091937W WO 2020199370 A1 WO2020199370 A1 WO 2020199370A1
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WIPO (PCT)
Prior art keywords
aero
turbine blade
outlet
heat exchange
engine turbine
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Application number
PCT/CN2019/091937
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English (en)
Chinese (zh)
Inventor
高晟钧
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高晟钧
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Publication date
Application filed by 高晟钧 filed Critical 高晟钧
Publication of WO2020199370A1 publication Critical patent/WO2020199370A1/fr

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Classifications

    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
    • F01D5/12—Blades
    • F01D5/14—Form or construction
    • F01D5/18—Hollow blades, i.e. blades with cooling or heating channels or cavities; Heating, heat-insulating or cooling means on blades
    • F01D5/186—Film cooling

Definitions

  • the invention relates to the technical field of aeroengine refrigeration, in particular to an aeroengine turbine blade with a cooling structure.
  • the cooling technology of turbine blades is mainly carried out from two aspects: one is to strengthen the disturbance of the cooling air inside the turbine blade, and to increase the heat exchange area inside the turbine blade; the other is to use film cooling on the blade surface to effectively block high-temperature gas from the turbine blade. Convective heat transfer.
  • the cooling effect of turbine blades is limited in either way.
  • the present invention aims to provide an aero-engine turbine blade with a cooling structure to solve the problem of insufficient cooling of existing aero-engine turbine blades.
  • an aero-engine turbine blade with a cooling structure includes a front end, a pressure surface, a suction surface and a tail end.
  • the front end, pressure surface, suction surface and tail end are enclosed in a hollow structure,
  • the hollow structure provides flow channels for cold flow;
  • At least one of the front end, the pressure surface and the suction surface is a thin-walled structure with a heat exchange cavity in the middle, and the heat exchange cavity is respectively connected with the hollow structure;
  • the heat exchange cavity and the exterior of the aero engine turbine blade are connected through an outlet reducing hole; the area of the inlet connecting the outlet reducing hole and the heat exchange cavity is smaller than the area of the outlet reducing hole and the outside of the aero engine turbine blade.
  • a heat exchange cavity is provided at the front end, and the heat exchange cavity is communicated with the outside of the aero-engine turbine blade through the air outlet reducing hole;
  • the axis of the air outlet reducing hole is perpendicular to the plane tangent to the outer surface of the front end where the air outlet reducing hole is located.
  • the diameter of the outlet reducing hole increases uniformly from the inlet to the outlet along the axial direction of the outlet reducing hole.
  • the diameter of the gas outlet reducing hole along the axial direction of the gas outlet reducing hole changes from inlet to outlet first, and then uniformly increases.
  • the pressure surface and/or the suction surface are provided with a heat exchange cavity, and the heat exchange cavity communicates with the outside of the aero-engine turbine blade through the outlet reducing hole;
  • the projection of the inlet of the outlet reducing hole on the outer surface of the aero-engine turbine blade is closer to the front end than the outlet of the outlet reducing hole.
  • the diameter of the outlet reducing hole increases uniformly from the inlet to the outlet along the axial direction of the outlet reducing hole.
  • the diameter of the gas outlet reducing hole along the axial direction of the gas outlet reducing hole changes from inlet to outlet first, and then uniformly increases.
  • a heat exchange cavity is provided at the tail end, and the heat exchange cavity at the tail end communicates with the outside of the aero-engine turbine blade through a straight air outlet;
  • the pressure surface and the suction surface are tangent at the tail end, and both are tangent to the same plane, and the axis of the straight air outlet is parallel or coincident with the plane.
  • the straight air outlet holes are uniformly arranged along the extension direction of the aero-engine turbine blades, and are evenly arranged on a plane perpendicular to the extension direction of the aero-engine turbine blades;
  • the heat exchange cavity at the tail end is provided with a supporting spoiler structure connecting the inner and outer surfaces of the aeroengine turbine blades.
  • the air outlet reducing holes are evenly arranged along the extension direction of the aero-engine turbine blades, and are evenly arranged on a plane perpendicular to the extension direction of the aero-engine turbine blades;
  • the heat exchange cavity is provided with a supporting spoiler structure connecting the inner and outer surfaces of the aeroengine turbine blades.
  • an aeroengine includes: an air inlet, a compressor, a combustion chamber, a turbine, and a tail nozzle;
  • the blades of the turbine adopt the aero-engine turbine blades with cooling structure in the above technical solution
  • the compressor is connected to the combustion chamber and shares the shell;
  • the intake port is arranged at the front end of the aero engine, and the inside of the compressor communicates with the outside through the intake port;
  • the tail nozzle is arranged at the tail of the combustion chamber, and the turbine is arranged in the tail nozzle.
  • the air outlet is set as a diameter-reducing hole with a small inside and a large outside, which can change the volume of the gas to lower the temperature inside and increase the cooling effect;
  • the air vents on the pressure surface and the suction side are biased toward the tail, and it is easier to generate an air film when the turbine blades rotate, thereby ensuring the thermal insulation effect of the air film.
  • it can avoid the air outlet with the pressure surface and the suction surface. Toward the forward direction of the turbine, thereby increasing the energy utilization rate of the turbine operation.
  • Figure 1 is an overall cross-sectional view in an embodiment of the present invention
  • Figure 2-1 is a schematic diagram of Example 1 enlarged at A of the embodiment of the present invention.
  • Example 2 2-2 is a schematic diagram of Example 2 enlarged at A of the embodiment of the present invention.
  • Figure 3-1 is a schematic diagram of Example 1 enlarged at B of the embodiment of the present invention.
  • Example 2 3-2 is a schematic diagram of Example 2 enlarged at B of the embodiment of the present invention.
  • Figure 4-1 is a schematic diagram of Example 1 enlarged at C of the embodiment of the present invention.
  • 4-2 is a schematic diagram of Example 2 enlarged at C of the embodiment of the present invention.
  • FIG. 5 is a schematic diagram of an enlarged example at D of the embodiment of the present invention.
  • an air film is formed by arranging air outlets on the surface of the turbine blade, and at the same time, the direction of the air outlet is set reasonably, which can reduce the influence of the air outlet on the fluid performance of the turbine blade as much as possible while forming the air film. Combined with the small and large cross-sectional dimensions of the vent hole, it can further ensure the high temperature outside and low temperature inside, thereby significantly improving the cooling and heat insulation effects without affecting the working efficiency of the turbine.
  • an embodiment of the present invention provides an aeroengine turbine blade with a cooling structure.
  • the aeroengine turbine blade includes: a front end 1, a pressure surface 2, a suction surface 3, and a tail end 4.
  • the front end 1 , Pressure surface 2, suction surface 3 and tail end 4 are enclosed in a hollow structure, which provides a flow channel for cold flow; at least one of front end 1, pressure surface 2 and suction surface 3 is a thin-walled structure with a heat exchange cavity in the middle , And the heat exchange cavity is respectively connected with the hollow structure; the heat exchange cavity and the outside of the aero-engine turbine blade are connected through the outlet reducing hole; the area of the inlet connecting the outlet reducing hole and the heat exchange cavity is smaller than the outlet reducing hole and the aero engine turbine blade The area of the external exit.
  • the specific shape of the turbine blade is not limited, and a common turbine blade shape may be used.
  • the flow rate and pressure of the high-temperature gas are different when the turbine blades are in operation.
  • the pressure and flow rate of the flow also need to be adjusted accordingly, so they need to be isolated from each other.
  • the cross-sectional size of the outlet diameter reducing hole is small inside and large outside.
  • the temperature difference can avoid excessive heat transfer of the high-temperature gas to the gas film, and avoid affecting the working efficiency of the turbine. At the same time, it can form a temperature-stable gas film to play a certain heat insulation effect and ensure the cooling effect.
  • the front end 1 is provided with a heat exchange cavity, and the heat exchange cavity communicates with the outside of the aero-engine turbine blade through the outlet reducing hole; the axis of the outlet reducing hole is The position of the air outlet reducing hole is perpendicular to the plane tangent to the outer surface of the front end 1. Considering that the front end 1 is on the windward side, the cold flow in the outlet hole after heat exchange needs a relatively large pressure to flow out of the outlet hole.
  • the outflow cold flow can be A gas film is directly formed at the front end 1, so the axis of the outlet reducing hole is perpendicular to the plane tangent to the outer surface of the front end 1 where the outlet reducing hole is located, that is, parallel to the flow direction of the external high temperature gas.
  • the reducing part of the air outlet reducing hole needs to be uniformly changed.
  • the larger the cone angle the more obvious the temperature reduction effect will be.
  • the temperature reduction effect will not continue to increase significantly.
  • the processing difficulty of the reduced diameter part will be greatly increased.
  • the taper angle of the side wall of the reducing part of the outlet reducing hole is 30°-120°.
  • the taper angle of the side wall is preferably 60°-90°.
  • the diameter of the outlet reducing hole increases uniformly from the inlet to the outlet along the axis of the outlet reducing hole, that is, the overall diameter is uniform; the diameter of the outlet reducing hole is from the inlet to the outlet along the axis of the outlet reducing hole.
  • the outlet is unchanged first, and then increased evenly, that is, part of the diameter is uniformly reduced and the rest is non-reducing.
  • the pressure surface 2 usually refers to the convex curved surface
  • the suction surface 3 usually refers to the concave curved surface, although one of the two is concave. It is convex, but the external high-temperature gas will not blow vertically to the outer surface of the impeller, but flow obliquely to the tail end 4 of the turbine blade. Therefore, if the vertical section of the gas outlet reducing hole is still used, a uniform gas film cannot be formed. On the contrary, it will increase the windward area, thereby reducing the efficiency of the turbine.
  • the pressure surface 2 and/or the suction surface 3 are provided with a heat exchange cavity, and the heat exchange cavity is connected to the outside of the aero-engine turbine blade through an outlet reducing hole; the inlet of the outlet reducing hole is outside the aero engine turbine blade
  • the projection of the surface is closer to the front end 1 than the outlet of the air outlet reducing hole.
  • the air outlet reducing hole inclined to the tail end 4 can make the outflowing cold flow flow toward the tail end 4, which facilitates the formation of a uniform air film, and can reduce the windward area without excessive cold flow pressure.
  • the angle between the axis of the gas outlet reducing hole and the tangent plane where it is located is 45°-90°, preferably 60°-75°.
  • the reducing parts of the outlet reducing hole on the pressure surface 2 and the suction surface 3 need to be uniformly changed.
  • the diameter of the outlet reducing hole increases uniformly from the inlet to the outlet along the axis of the outlet reducing hole, that is, the overall diameter is uniform; the diameter of the outlet reducing hole is from the inlet to the outlet along the axis of the outlet reducing hole.
  • the outlet is unchanged first, and then increased evenly, that is, part of the diameter is uniformly reduced and the rest is non-reducing.
  • the temperature of the end 4 of the turbine blade is not as high as other parts, but the end 4 can still be provided with a heat exchange cavity to further improve the cooling effect.
  • the tail end 4 is provided with a heat exchange cavity, and the heat exchange cavity of the tail end 4 communicates with the outside of the aero-engine turbine blade through the straight air outlet; the pressure surface 2 and the suction surface 3 is tangent to the tail end 4, and all tangent to the same plane, and the axis of the straight air outlet is parallel or coincident with the plane.
  • the cold flow out of the tail end 4 will directly follow the external high temperature gas to flow away in the tangential direction of the tail end 4, so there is no need to consider the problem of gas film, just ensure that the outflow direction is the same as the flow direction of the high temperature gas. Due to the difficulty of processing, the tail 4 uses straight holes directly due to its relatively low temperature.
  • the straight air outlet holes are evenly arranged along the extension direction of the aero-engine turbine blades, and are evenly arranged on a plane perpendicular to the extension direction of the aero-engine turbine blades.
  • the air outlet reducing holes are uniformly arranged along the extension direction of the aero-engine turbine blade, and are perpendicular to the extension direction of the aero-engine turbine blade. Set evenly on the plane of the
  • the outgassing The area of the variable diameter hole and the straight hole of the air outlet is 5%-20% of the area of the aero-engine turbine blade.
  • the sum of the area of the air outlet variable diameter hole and the air outlet straight hole is 5%-20% of the area of the aero engine turbine blade, preferably 7%-12%.
  • the heat exchange cavity is provided with a supporting spoiler structure that connects the inner and outer surfaces of the aeroengine turbine blades. , Can ensure the structural strength of the turbine blades, prevent structural damage during operation, and its turbulence function can also make the cold flow contact with the turbine blades more fully and improve the heat exchange efficiency.
  • an aeroengine includes: an air inlet, a compressor, a combustion chamber, a turbine, and a tail nozzle;
  • the blades of the turbine adopt the aeroengine turbine blades with cooling structure in embodiment 1;
  • the compressor is connected to the combustion chamber and shares the shell;
  • the intake port is arranged at the front end of the aero engine, and the inside of the compressor communicates with the outside through the intake port;
  • the tail nozzle is arranged at the tail of the combustion chamber, and the turbine is arranged in the tail nozzle.
  • the embodiment of the present invention provides an aero-engine turbine blade with a cooling structure.
  • the present invention sets the air outlet as a diameter-reducing hole with a small inside and a large outside, which can change the gas volume and lower the temperature inside.
  • the outer height further improves the cooling effect; in the present invention, the vent holes of the pressure surface 2 and the suction surface 3 are biased toward the tail, and it is easier to generate an air film when the turbine blades rotate, thereby ensuring the heat insulation effect of the air film.
  • it can avoid pressure
  • the air outlets of the surface 2 and the suction surface 3 face the advancing direction of the turbine, thereby improving the energy utilization rate of the turbine operation.

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  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)

Abstract

La présente invention concerne une aube de turbine de moteur aéronautique présentant une structure de refroidissement comprenant une extrémité avant (1), une surface de pression (2), une surface d'aspiration (3) et une extrémité arrière (4) ; l'extrémité avant (1), la surface de pression (2), la surface d'aspiration (3) et l'extrémité arrière (4) sont enfermées pour former une structure creuse, et la structure creuse fournit un canal d'écoulement pour un flux froid. Au moins l'une de l'extrémité avant (1), de la surface de pression (2) et de la surface d'aspiration (3) est une structure à paroi mince dotée, au niveau d'une partie centrale de celle-ci, d'une cavité d'échange de chaleur, et la cavité d'échange de chaleur est en communication avec la structure creuse. La cavité d'échange de chaleur est en communication avec l'extérieur de l'aube de turbine de moteur aéronautique par l'intermédiaire d'un trou de sortie d'air de diamètre variable, et la zone d'une entrée à travers laquelle le trou de sortie d'air de diamètre variable est raccordé à la cavité d'échange de chaleur est plus petite que la zone d'une sortie à travers laquelle le trou de sortie d'air de diamètre variable est raccordé à l'extérieur de la pale de turbine de moteur aéronautique.
PCT/CN2019/091937 2019-04-02 2019-06-19 Aube de turbine de moteur aéronautique présentant une structure de refroidissement WO2020199370A1 (fr)

Applications Claiming Priority (2)

Application Number Priority Date Filing Date Title
CN201910261827.0 2019-04-02
CN201910261827.0A CN109973154B (zh) 2019-04-02 2019-04-02 一种带有冷却结构的航空发动机涡轮叶片

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WO2020199370A1 true WO2020199370A1 (fr) 2020-10-08

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Publication number Priority date Publication date Assignee Title
CN111677557B (zh) * 2020-06-08 2021-10-26 清华大学 涡轮导向叶片及具有其的涡轮机械
CN113107611B (zh) * 2021-04-22 2022-07-12 南京航空航天大学 基于双喉道气动矢量喷管的涡轮叶片尾缘冷却结构及其尾迹控制方法

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CN86108861A (zh) * 1985-12-23 1987-08-05 联合工艺公司 薄膜冷却叶片和涡轮
US5667359A (en) * 1988-08-24 1997-09-16 United Technologies Corp. Clearance control for the turbine of a gas turbine engine
CN1717534A (zh) * 2003-11-21 2006-01-04 三菱重工业株式会社 燃气涡轮发动机的冷却叶片
US20080175714A1 (en) * 2007-01-24 2008-07-24 United Technologies Corporation Dual cut-back trailing edge for airfoils
EP2075410A2 (fr) * 2007-12-28 2009-07-01 General Electric Company Procédé de formation de trous de refroidissement et surface portante de turbine avec trous de refroidissement hybride
US7625180B1 (en) * 2006-11-16 2009-12-01 Florida Turbine Technologies, Inc. Turbine blade with near-wall multi-metering and diffusion cooling circuit

Family Cites Families (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US6287075B1 (en) * 1997-10-22 2001-09-11 General Electric Company Spanwise fan diffusion hole airfoil
CN103244196B (zh) * 2012-02-08 2015-04-22 中国科学院工程热物理研究所 一种离散气膜冷却孔型
CN205382958U (zh) * 2016-03-02 2016-07-13 中航商用航空发动机有限责任公司 涡轮叶片以及航空发动机

Patent Citations (6)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN86108861A (zh) * 1985-12-23 1987-08-05 联合工艺公司 薄膜冷却叶片和涡轮
US5667359A (en) * 1988-08-24 1997-09-16 United Technologies Corp. Clearance control for the turbine of a gas turbine engine
CN1717534A (zh) * 2003-11-21 2006-01-04 三菱重工业株式会社 燃气涡轮发动机的冷却叶片
US7625180B1 (en) * 2006-11-16 2009-12-01 Florida Turbine Technologies, Inc. Turbine blade with near-wall multi-metering and diffusion cooling circuit
US20080175714A1 (en) * 2007-01-24 2008-07-24 United Technologies Corporation Dual cut-back trailing edge for airfoils
EP2075410A2 (fr) * 2007-12-28 2009-07-01 General Electric Company Procédé de formation de trous de refroidissement et surface portante de turbine avec trous de refroidissement hybride

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CN109973154B (zh) 2019-12-06
CN109973154A (zh) 2019-07-05

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