Composite cooling structure for gas turbine blades
A turbine blade and composite cooling technology, applied in the direction of blade support components, mechanical equipment, engine components, etc., can solve the problem of low comprehensive cooling efficiency, achieve the effects of improving comprehensive cooling efficiency, reducing wall surface temperature, and enhancing heat transfer effect
Active Publication Date: 2022-05-17
BEIHANG UNIV
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Problems solved by technology
[0003] The object of the present invention is to provide a composite cooling structure for gas turbine blades to solve the problem of low comprehensive cooling efficiency of high-temperature components such as existing gas turbine blades
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Embodiment Construction
[0019] The structure, principle and features of the present invention will be described below in conjunction with the accompanying drawings, and the examples given are only used to explain the present invention, and are not intended to limit the scope of the present invention.
[0020] Such as Figure 1 to Figure 8 As shown, a composite cooling structure of a gas turbine blade includes an upper wall surface 1, a lower wall surface 2, an air film hole I3, an air film hole II4, a cross elliptical channel 5 and a side wall 6; the air film hole I3 and the air film hole II4 are all It is a cylindrical oblique hole, and the diameter of the air film hole I3 is d 1 , the diameter of gas film hole II4 is d 2 The center of air film hole I3 has air film hole I central axis 7, and the center of air film hole II4 has air film hole II central axis 8; Cross ellipse channel 5 is composed of cross ellipse channel inlet section 9, cross ellipse channel buffer section 10, The middle section 11...
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A composite cooling structure for the turbine blade of a gas turbine, comprising an upper wall surface, a lower wall surface, a gas film hole I, a gas film hole II, a cross elliptical channel and a side wall; , the middle section of the intersecting elliptical channel and the outlet section of the intersecting elliptical channel are sequentially connected; the middle section of the intersecting elliptical channel is arranged alternately in pairs with 5 to 10 groups of elliptical sections I and elliptical sections II whose major axes are perpendicular to each other; during operation, the cooling gas Enter the gas film hole I, the gas film hole II and the cross elliptical channel respectively. The gas at the exit of the gas film hole forms a gas film covering and cooling on the upper wall. The longitudinal vortex is generated, and by improving the synergistic effect of the velocity vector field and the temperature gradient field, the turbulent heat exchange effect between the cooling gas and the upper wall surface 1 is enhanced, which is complementary to the area not covered by the gas film on the upper wall surface, and the comprehensive cooling effect is improved.
Description
technical field [0001] The invention relates to the field of gas turbine engines, in particular to a composite cooling structure for gas turbine blades. Background technique [0002] As the temperature before the turbine of the gas turbine continues to increase, the temperature of the turbine operation has far exceeded the limit that the blade material can withstand, and it must be effectively cooled. In order to control the cooling gas flow rate to ensure the efficiency of the whole machine under the premise of ensuring a certain temperature drop, taking safe, effective and reasonable cooling measures for the turbine blades is the key to prolonging the service life of the turbine blades and improving the working stability of the turbine. Air film cooling is currently an external cooling method widely used in gas turbine engines. Traditional air film cooling technology faces the defect of insufficient lateral heat transfer coverage. It is urgent to adopt other cooling method...
Claims
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IPC IPC(8): F01D5/18
CPCF01D5/186F01D5/187Y02T50/60
Inventor 柳阳威高文唐雨萌张晏通
Owner BEIHANG UNIV