Turbine guide vane with suction surface provided with step slot cooling structure

Abstract

The invention discloses a novel slot gas film cooling structure used for aircraft engines and turbine guide vanes of gas turbines. The cooling structure comprises a vane basal body, a gas film slot, a connecting rib and step surfaces. The cooling structure is characterized in that by means of a step-shaped slot gas film outflow structure composed of an inner piece, an outer piece and the connecting rib, cooling gas can flow out in the tangential direction of the surface of the vane, wherein the inner piece and the outer piece are located on the basal body of the suction surface of the vane. By means of the cooling outflow structure of the novel vane, a uniform and consistent cooling gas film can be formed on the surface of the vane, the stability of the outflow gas film is higher, and the eddy current effect of the outflow cooling gas is weakened; the cover area of the cooling gas film is large, and the cooling efficiency is high; meanwhile, due to the fact that the flow in the tangential direction is mixed with high-temperature main flow, the aerodynamic loss of the vane can be reduced, and the vane is made to have the efficient cooling effect and the effect aerodynamic performance; furthermore, by means of the structure, the vane can be manufactured in a continuous-piece mode, and machining and the maintenance technique of the vane can be simplified.

Description

technical field [0001] The invention relates to a novel cooling structure for high-temperature components of an aero-engine, in particular to a cooling structure suitable for blades of a turbine deflector. This structure not only has better cooling performance for the turbine guide blade, but also has higher aerodynamic efficiency than the traditional smooth blade shape in terms of aerodynamic performance. The gas turbine engine turbine guide adopting this cooling aerodynamic structure The blade can have a large performance advantage over conventional blades in terms of cooling performance and aerodynamic performance. Background technique [0002] Turbine blades are important hot-end parts of aero-engines, especially the guide blades are directly washed by the high-temperature gas at the outlet of the combustion chamber. At present, the temperature of the front inlet of advanced engines has reached about 2000K, which is 400K higher than the melting point of the metal materi...

AI Technical Summary

Problems solved by technology

Since the existing air film holes need to flow out from the cold air cavity through the cooling holes on the wall surface to the outside of the high temperature wall surface, they must pass through holes of a certain diameter on the high temperature wall surface, which means that the use of this structure will inevitably The outflow of the gas film has a flow perpendicular to the normal direction of the high-temperature wall surface, which will definitely generate a horseshoe vortex that causes cooling and gas mixing, which reduces the cooling effect of the cold air

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