Through type continuous folded plate structure suitable for tail edge part of turbine blade

Abstract

The invention discloses a through type continuous folded plate structure suitable for a turbine blade tail edge part. The through type continuous folded plate structure comprises a U-shaped cooling channel and a blade tail edge cooling cavity which are arranged in a blade, wherein the U-shaped cooling channel is positioned at the front edge and the middle part of the blade, the blade tail edge cooling cavity is positioned at the tail edge of the blade, the U-shaped cooling channel and the blade tail edge cooling cavity are separated through a side wall face, an opening is formed in the bottom of the side wall face and serves as an outlet of the U-shaped cooling channel, a plurality of air holes are formed in the side wall face, a plurality of rows of through type continuous folded plates are arranged in the blade tail edge cooling cavity, and a tail edge crack is formed in the blade tail edge. The heat exchange can be effectively improved on the premise that the flow resistance is not large, and meanwhile the machinability is good. Due to the fact that the blocking and separating effects on cold air are obviously reduced through the folded plate type structure, the flow guiding performance is good, the turbulence degree of the cold air at an outlet of the blade tail edge cooling channel is low, the mixing loss of the cold air and mainstream fuel gas is reduced, and the aerodynamic efficiency of the blade is improved.

Description

technical field [0001] The invention belongs to the technical field of turbine blade cooling of a gas turbine, and in particular relates to a through-type continuous folded plate cooling structure suitable for the trailing edge channel of a turbine blade. Background technique [0002] In aviation gas turbines, thrust and efficiency are the two main performance indicators, and increasing the gas temperature before the turbine is one of the most direct and effective ways to improve engine performance. The gas temperature before the turbine of today's advanced aviation gas turbines is as high as 1700°C, which is much higher than the melting point temperature of 1000°C for high-performance blade materials. This requires an effective cooling structure to ensure the safe and stable operation of the blades. In turbine blades, internal cooling can reduce the mixing of cold air and gas, thereby reducing the impact on the aerodynamic performance of the turbine. It is a cooling method ...

AI Technical Summary

Problems solved by technology

However, this structure has a significant impact on the cold air, the flow loss is still relatively large, and there is still a large room for improvement in heat transfer performance; at the same time, there is no connection between the cooling channel at the trailing edge and the cooling channel at the middle and rear of the incoming blade, and the outlet. Effective combination of mainstreams makes it difficult to achieve uniform and stable inlet airflow and low outlet turbulence

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