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Turbine cooling vane of gas turbine engine

By designing cooling channels, film cooling holes, impact cooling components and spherical components in the cooling blades of small gas turbine engines, the problems of low cooling efficiency and backflow are solved, and efficient cooling effects and structural simplicity are achieved.

Inactive Publication Date: 2006-01-04
MITSUBICHI HEAVY IND AERO ENGINES LTD
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  • Summary
  • Abstract
  • Description
  • Claims
  • Application Information

AI Technical Summary

Problems solved by technology

Therefore, the working fluid flowing into the inside of the cooling turbine blade B flows to the downstream side through the space between the insert 92 and the inner wall 901 of the cooling passage 91, thereby greatly reducing the cooling effect of impingement cooling.

Method used

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Examples

Experimental program
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Effect test

Embodiment Construction

[0032] Hereinafter, an embodiment of the present invention will be described with reference to the drawings. Figure 1A It is a perspective view of an example of a cooling blade of a small gas turbine engine according to an embodiment of the present invention. Figure 1B yes Figure 1A X-X sectional view of a cooling blade of a small gas turbine engine shown. Figure 1C yes Figure 1B An enlarged cross-sectional view of the circled portion shown. Figure 1D is when the blade wall surface is intercepted as a rectangle, Figure 1A A front view of the inner wall surface of the cooling turbine blade is shown.

[0033] gas turbine engine with Image 6The prior art gas turbine engine shown is of substantially the same construction. That is, the gas turbine engine GT includes: an inlet fan Kf installed at the air inlet port; a compressor Cp that compresses the gas introduced; a combustion chamber Bs that burns fuel by using the gas compressed by the compressor Cp; The turbine T...

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PUM

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Abstract

A cooled blade of a small gas turbine engine can enhance cooling performance without increasing the amount of cooling air in a simple configuration. A cooled turbine blade is provided with a cooling passageway 12 formed inside thereof, making the cooling air flow therein. Film-cooling holes 13 penetrate from an inner wall surface 111 to an external wall surface 112 and form a cooling film. An impingement cooling member 2 has a multiple number of small holes 21 ejecting the cooling air. A gap t made by the inner wall surface 111 and the impingement cooling member 2 has a sealing portion 14 mounted thereon which divides the gap in a blade chord direction.

Description

technical field [0001] The present invention relates to cooling blades used for small gas turbine engines of aircraft and the like, and particularly relates to cooling blades as turbine blades. Background technique [0002] Today, gas turbines are used as power sources for various machines and equipment. For example, it is used in a power generation device application by connecting its main shaft to a generator, or as an engine using a gas turbine as a transportation power source such as an aircraft. [0003] Image 6 is a schematic diagram of a gas turbine engine. Image 6 The gas turbine engine GT shown includes: an inlet fan Kf installed at an air inlet port, a compressor Cp that compresses incoming air, a combustion chamber Bs that burns fuel by using the air compressed by the compressor, The turbine Tb is driven by the combustion gas ejected from the combustion chamber Bs, and the nozzle Nz is generated by ejecting the combustion gas to generate thrust. The inlet fan ...

Claims

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Application Information

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Patent Type & Authority Applications(China)
IPC IPC(8): F01D5/18F01D9/02
CPCY02T50/676F05D2260/201F01D5/186F01D5/189F05B2260/201Y02T50/60
Owner MITSUBICHI HEAVY IND AERO ENGINES LTD