Turbine blade tail edge turbulence half-crack cooling structure with spherical surface convex block

Abstract

The invention discloses a turbine blade tail edge turbulence half-crack cooling structure with a spherical surface convex block; the spherical surface convex block structure is applied to a half-crack wall surface; and under the precondition of not increasing the outlet flow of an air film, the convection heat exchange coefficient and the heat exchange area of the air film are increased through a turbulence structure, and the convection heat exchange strength in self-crack air film cooling is reinforced, so that the comprehensive cooling effect of blade tail edges is improved. The turbine blade tail edge turbulence half-crack cooling structure with the spherical surface convex block is to cut off one part of the wall surface on a blade tail edge pressure surface; the wall surface for retaining one side of a blade tail edge suction surface and spaced separation ribs form multiple half-crack structures; cooling air currents are sprayed out from a cold current outlet to cover the self-crack wall surface to form a cooling air film, so that the structure is simple, the highest temperature and the average temperature of the suction surface are effectively lowered, and the high-temperature ablation of the suction furnace of turbine blades is prevented; and the spherical surface convex block structure is arranged on the half-crack wall surface to achieve excellent heat transfer characteristic and machining exploitativeness.

Description

technical field [0001] The invention belongs to the technical field of gas turbine blade cooling, and in particular relates to a turbine blade trailing edge turbulence half-slit cooling structure with spherical bumps. Background technique [0002] Increasing the turbine inlet temperature is an effective way to improve the thrust and efficiency of gas turbines, but the increase in turbine inlet temperature will cause the turbine blades to bear a greater heat load, and excessive temperature and thermal stress may cause the turbine blades to fail to work properly. The inlet temperature of modern gas turbine design has far exceeded the temperature limit of the materials used, so complex cooling technology must be used to ensure the normal operation of the turbine under high temperature conditions. The trailing edge of the turbine blade is often a high-temperature part, and is also the most vulnerable to heat corrosion and damage. The main reason is that the gas side flow at the ...

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Problems solved by technology

[0004] In the literature "Experimental Research on the Air Film Cooling Characteristics of Turbine Blade Trailing Slits" ("Journal of Engineering Thermophysics", 2016, No. 37, pp. 1988-1993), the thermochromic liquid crystal transient heat transfer measurement was used to study The effects of two different split rib structures, straight rib and chamfered rib, on the film cooling characteristics of the split ribs in the trailing edge laminate cooling structure are studied, although the conclusion shows that the air film cooling effect of the half-slit trailing edge of the blade is affected by the split ribs. However, with the continuous increase of the gas temperature of the aero-engine, the cooling capacity of the traditional half-slit structure gradually tends to the limit, and the ablation phenomenon of the trailing edge of the high-pressure turbine blade often occurs on the suction side of the trailing edge of the blade, so the development of And the innovative efficient cooling structure of the trailing edge of the turbine blades can further improve the comprehensive cooling effect without increasing the amount of cooling air, which is very necessary and meaningful for the development of advanced high-performance aero-engines

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