A Corrugated Rib Flow Guide Structure Applied to Half Slits on the Trailing Edge of Turbine Blades

Abstract

The invention relates to a corrugated rib diversion structure and a diversion method applied to the half-slit of the trailing edge of a turbine blade. Seam downstream wall, partition ribs, continuous corrugated ribs. The pressure surface of the turbine blade close to the trailing edge cuts off a part of the wall and forms multiple trailing edge half-slit structures with the partition ribs. After the cooling gas flows out from the outlet, a cooling gas film is formed on the half-slit wall and the downstream wall to isolate the high-temperature mainstream gas and reduce wall temperature. By arranging the corrugated ribs on the downstream wall of the half-slit to create a turbulence effect on the cooling air film, the unevenness of the distribution of the air film downstream of the half-slit is significantly reduced, and the span-wise coverage and cooling efficiency of the cooling air film are improved to achieve the highest trailing edge. Reduction of temperature and temperature gradient. The present invention arranges a continuous corrugated rib structure on the downstream wall of the half-slit, has the characteristics of simple structure, convenient processing and good cooling effect, and can be applied to various half-slit structures of the trailing edge of turbine blades.

Description

technical field [0001] The invention belongs to the technical field of cooling the turbine blades of gas turbines, and in particular relates to a corrugated rib guide structure applied to the semi-slits of the trailing edge of the turbine blades. Background technique [0002] Gas turbine engine is a thermal power device based on Brayton cycle, which has been widely used in modern military and industry relying on its powerful output power and high thermal efficiency. Experience shows that under the premise of the same engine size, the thrust of the gas turbine can be increased by 8-13% and the cycle efficiency can be increased by 2-4% for every 56K increase in the turbine inlet temperature. At present, the temperature before the turbine of an advanced aero-engine has exceeded 2000K, but the temperature resistance limit of the turbine blade material is much lower than the turbine inlet temperature, so efficient cooling technology must be used to ensure its normal operation. B...

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Problems solved by technology

The mechanism of the above two existing technologies is to enhance the convective heat transfer coefficient and heat transfer area at the wall through the spoiler structure, thereby increasing the heat transfer between the upstream air film and the blade wall to reduce the temperature of this part of the wall. , however, for the downstream blade wall of the turbine blade trailing edge half-slit, this part of the blade wall is not covered by air film due to the blocking effect of the upstream half-slit rib

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