Turbine guide vane with composite cooling structure and guide vane cooling method

Abstract

The invention discloses a turbine guide vane with a composite cooling structure. The interior of a guide vane cavity of the turbine guide vane is divided into a front edge impact area and a tail edge impact area by a partition plate; truss structures are arranged in the first double-layer wall structure of the front edge impact area and the second double-layer wall structure of the tail edge impact area; the air film holes are formed in the outer walls of the guide vanes at the intersection points, and the impact holes are formed in the first inner wall and the second inner wall in a matrix mode. The invention further provides a cooling method of the turbine guide vane. According to the structure, heat exchange in the guide vane is enhanced, the air film cooling effect of the outer wall of the guide vane is improved, and meanwhile, the truss matrix structure enables a double-layer wall structure in a cold air cavity of the guide vane to have good mechanical performance.

Description

Technical Field

[0001] The present invention relates to the field of gas turbine engines, and in particular to a thermal protection cooling structure and a cooling method for turbine guide vanes. Background Art

[0002] Increasing the turbine inlet temperature is an important way to improve engine efficiency and thrust-to-weight ratio. However, the increase in turbine inlet temperature will increase the thermal load of the blades and reduce their service life. At present, the combustion chamber outlet temperature of advanced aircraft engines has exceeded the heat resistance limit of high-temperature resistant materials, so it is necessary to take efficient thermal protection measures for the downstream turbine guide vanes.

[0003] Film cooling with a number of discrete film holes on the blade surface and rib-shaped structures in the cooling channels of the blades are widely used as thermal protection means for turbine blades. However, with the development of advanced high-performance gas turbine engines, the gas temperature faced by the turbine continues to increase, and the consumption of cold air for turbine cooling has increased significantly, which has an adverse effect on the overall performance of the engine. It is necessary to apply a cooling structure with high-efficiency cooling capacity to the turbine guide vanes, that is, to achieve better cooling effect with as little cold air as possible.

[0004] The double-wall composite cooling structure has the advantages of small size and large heat exchange area, and is the development direction of cooling technology in the future. Nealy's research shows that compared with the traditional cooling structure, the comprehensive cooling efficiency of the layer cooling structure can reach 0.7 and reduce the cooling air consumption by 67%. Kim used infrared temperature measurement method to study the comprehensive cooling efficiency of layer cooling and single-layer air film cooling under the same blowing ratio. The results showed that under different blowing ratios, the comprehensive cooling efficiency of layer cooling was 47-141% higher than that of single-layer air film cooling. These studies have demonstrated the great potential of layer cooling in improving cooling efficiency and reducing cooling air consumption.

[0005] As an efficient cooling structure, the double-layer wall can combine multiple single cooling forms. The typical layer plate cooling structure is: a cylindrical or prismatic vertical spoiler structure is arranged in the impact plate and the air film plate, and the impact hole and the air film hole are circular holes. However, there are few studies on the use of double-layer cooling wall structures on turbine blades, especially the failure to fully consider the synergistic effect of spoiler structures, impingement heat transfer and air film cooling. In other words, the double-layer wall composite cooling structure has not and cannot be directly used on turbine guide vanes. It must be redesigned according to the structural characteristics of the turbine guide vanes to adapt to the cooling of the turbine guide vanes. Summary of the invention

[0006] The object of the present invention is to avoid the shortcomings of the prior art and provide a turbine guide vane and a guide vane cooling method with a composite cooling structure having a high-efficiency cooling capability and being suitable for turbine guide vanes.

[0007] To achieve the above object, the technical solution adopted by the present invention is: a turbine guide vane with a composite cooling structure, wherein the guide vane cavity of the turbine guide vane is divided into a leading edge impact zone and a trailing edge impact zone by a partition;

[0008] The leading edge impact zone includes a first cold air cavity provided in a cavity from the leading edge of the guide vane to the middle of the guide vane, and a first inner wall of the first cold air cavity and an outer wall of the guide vane form a first double-layer wall structure of the leading edge impact zone;

[0009] The trailing edge impact zone includes a second cold air cavity and a trailing edge exhaust cavity sequentially arranged in the cavity between the partition and the trailing edge of the guide vane, and the second inner wall of the second cold air cavity and the outer wall of the guide vane form a second double-layer wall structure of the trailing edge impact zone;

[0010] Truss structures are arranged in matrix in the double-wall cavities of the first double-wall structure and the second double-wall structure respectively; the truss structure is composed of four truss rods, one ends of the four truss rods intersect with each other to form an intersection, the intersection and the other ends of the four truss rods are respectively fixedly connected to the outer wall of the guide vane and the first inner wall or the second inner wall, and four connection points are formed on the inner wall;

[0011] The air film holes are arranged on the outer wall of the guide vane at the intersection, and the impact holes are arranged in a matrix on the first inner wall and the second inner wall.

[0012] Furthermore, the truss structure matrix provided in the second double-wall structure is arranged in the double-wall cavity on one side of the blade disk, and the intersection of the truss structure and the other ends of the four truss rods are respectively fixedly connected to the outer wall of the guide vane on the blade basin side and the inner wall of the second cold air cavity, and four connection points are formed on the inner wall, and the distances from the four connection points to the intersection points are equal;

[0013] The impact holes are arranged outside the four connection points and on the line connecting the intersection point and the center points of the four connection points.

[0014] Furthermore, the distances from four connection points of the truss structure provided in the double-wall cavity of the first double-wall structure on the first inner wall to the intersection point are equal;

[0015] The impact holes are arranged outside the four connection points and on the line connecting the intersection point and the center points of the four connection points.

[0016] Furthermore, the air film hole is a hole arranged around the intersection, including a discontinuous annular hole composed of 2 to 4 fan ring holes and fan ring partitions for connecting two fan ring holes, and the center point of the discontinuous annular hole coincides with the center point of the intersection.

[0017] Furthermore, the inner diameter of the ring formed by the air film hole is d, the outer diameter is D, and the distance between the outer ring and the inner ring is the width B of the fan ring hole, then:

[0018] On the leading edge of the guide vane, there is an air film hole with a width B < 0.5D, which is used to increase the outlet momentum of the air film hole and overcome the air flow stagnation and large pressure load;

[0019] On the blade body of the turbine guide vane, the width B of the air film hole is taken as: 0.1D < B < 0.2D, which is used to obtain the cold air coverage effect.

[0020] Furthermore, the fan-shaped ring hole is a normal through hole or a normal backward expanding hole on the outer wall of the guide vane;

[0021] When it is a normal through hole on the outer wall of the guide vane, a pair of inner inclination angles α1 = α2 = 90° corresponding to the cross section of the air film hole are set;

[0022] When it is a normal backward expanding hole, the sum of a pair of inner inclination angles corresponding to the cross section of the air film hole is 90 degrees, and:

[0023] 30° ≤ α1 ≤ 90°, 30° ≤ α2 ≤ 90°;

[0024] The fan-shaped ring partition has a pair of arc connecting edges and a pair of parallel or inclined edges;

[0025] When it is the parallel edge, the distance L between the two edges is 0.05 - 0.25 mm;

[0026] When it is the inclined edge, the central angle corresponding to the inclined edge is θ, and the range of θ is 15 - 45 degrees.

[0027] Furthermore, the impact hole is a normal through hole on the first inner wall or the second inner wall, and the diameter of the impact hole is 0.5 - 0.75 mm;

[0028] The opening area of the air film hole is 1.4 - 2.4 times that of the impact hole.

[0029] Furthermore, the matrix arrangement of the truss structure is specifically: taking Sx as the arrangement spacing of the truss structure in the guide vane flow direction, Sz as the arrangement spacing of the truss structure in the guide vane height direction, H as the distance between the outer wall of the guide vane and the first inner wall or the second inner wall, D' as the diameter of the truss rod, and d' as the diameter of the round hole with the same area as the fan-shaped ring hole, then there are:

[0030] H ≥ Sx ≥ 2D', H ≥ Sz ≥ 2D';

[0031] π·(d' / 2) 2 / Sx × Sz = 2.5 - 5%;

[0032] D' / d' = 0.5 - 3.5;

[0033] H / d' = 1 - 2.5;

[0034] The truss structures are sequentially extended and arranged in the double-wall cavities of the first double-wall structure and the second double-wall structure according to a set distance.

[0035] Furthermore, the wall normal angle between the truss structure and the first inner wall or the second inner wall is β, β is 30-60°; the length of the truss rod constituting the truss structure is l, H is the distance between the air film orifice plate and the impact orifice plate, then l=H / cosβ.

[0036] The present invention also provides a guide vane cooling method for the turbine guide vane, specifically:

[0037] The high-pressure cold air flow C1 drawn out from the compressor enters the first double-wall structure or the second double-wall structure through the impact hole, and forms an impact cooling air flow on the wall surface of the first inner wall or the second inner wall directly opposite to it. The cooling air flow is divided into two parts, one part of the cooling air flow collides with the adjacent impact jet to form a reverse air flow C2; the other part of the cooling air flow forms a cold air flow C4 on the truss surface of the truss structure, and the cold air flow C4 is sucked into the air film hole after heat exchange, and the cold air flow C3 flowing directly against the wall also enters the air film hole at the same time; the air flow C5 flowing out of the air film hole forms a cold air film C6 covering the surface of the turbine guide vane under the action of the mainstream of the gas.

[0038] The beneficial effects of the present invention are as follows: the turbine guide vane with a cooling structure provided by the present invention can also be used as a variable geometry turbine adjustable guide vane; the truss structure of the present invention is arranged in a cooling structure within a double-wall structure within the cooling air cavity of the turbine guide vane, and at the same time, the air film holes used by the guide vane are changed from a single circular hole to four 1 / 4 annular holes, but the total area of ​​the air film holes remains unchanged. Compared with the traditional double-wall heat shield with cylindrical spoilers, on the one hand, the spoiler of the structure of the present invention is an inclined spoiler column, which increases the heat exchange area and avoids the low heat exchange area caused by the collision of the air flow at the tail of the straight column spoiler; on the other hand, the form of external air film cooling is changed, the air film coverage area of ​​the four annular air film holes is increased, and the air film has better wall adhesion, which significantly improves the air film cooling effect. At the same time, the tetrahedral truss structure can effectively fix the double-wall structure, enhance its structural performance, and improve the service life and reliability of the turbine guide vane.

[0039] Therefore, the structure provided by the present invention not only strengthens the internal heat exchange of the guide vane, but also improves the air film cooling effect of the outer wall of the guide vane. In addition, the tetrahedral truss matrix structure enables the double-layer wall structure in the guide vane cooling cavity to have better mechanical properties. BRIEF DESCRIPTION OF THE DRAWINGS

[0040] Figure 1 It is a schematic diagram of the top view structure of the present invention;

[0041] Figure 2 It is a schematic diagram of the three-dimensional structure of the present invention;

[0042] Figure 3 It is a schematic diagram of the truss structure matrix structure of the present invention;

[0043] Figure 4 The truss structure of the present invention Figure 3 The enlarged schematic diagram of the monomer structure in the middle A part;

[0044] Figure 5 It is a schematic diagram of the relative positions of the air film hole and the impact hole of the present invention;

[0045] Figure 6 It is a schematic diagram of the relative position of the impact hole and the truss structure of the present invention;

[0046] Figure 7 Schematic diagram of the cross-sectional structure of the air film hole of the present invention;

[0047] Figure 8 is a top view of the air film hole of the present invention;

[0048] Fig. 9 It is a heat exchange cloud diagram of the straight column spoiler and the truss structure of the present invention impacting the target surface;

[0049] Fig.10 It is a streamline diagram of the straight column spoiler and the double-layer wall of the truss structure monomer of the present invention;

[0050] Fig.11 It is the streamline diagram of the straight column spoiler and the double-layer wall of the truss structure matrix of the present invention;

[0051] Fig.12 This is a comparison diagram of the straight column spoiler during mainstream impact and the span-wise average air film cooling efficiency of the truss structure matrix of the present invention;

[0052] Fig.13 This is a comparison diagram of the cooling efficiency of the straight column spoiler and the span-wise average air film of the truss structure matrix of the present invention when the mainstream passes by;

[0053] Fig.14 It is a comparison diagram of the cooling efficiency of the straight column spoiler and the span-wise average air film of the truss structure matrix of the present invention when the mainstream passes by. DETAILED DESCRIPTION

[0054] The principles and features of the present invention are described below in conjunction with the accompanying drawings. The examples given are only used to explain the present invention and are not used to limit the scope of the present invention.

[0055] In order to achieve the above object, the present invention provides the following specific implementation methods:

[0056] Example 1: Figure 1-8 As shown, a turbine guide vane with a composite cooling structure, wherein the guide vane cavity of the turbine guide vane is divided into a leading edge impact zone 1 and a trailing edge impact zone 2 by a partition plate 5;

[0057] like Figure 1 , Figure 3 As shown, the leading edge impact zone 1 includes a first cold air cavity 11 provided in the cavity from the leading edge of the guide vane to the middle of the guide vane, and the first inner wall 12 of the first cold air cavity 11 and the outer wall 3 of the guide vane form a first double-layer wall structure of the leading edge impact zone 1;

[0058] The trailing edge impact zone 2 includes a second cold air cavity 21 and a trailing edge exhaust cavity 22 which are sequentially arranged in the cavity between the partition plate 5 and the trailing edge of the guide vane. The second inner wall 23 of the second cold air cavity 21 and the outer wall 3 of the guide vane form a second double-layer wall structure of the trailing edge impact zone 2.

[0059] like Figure 3 , Figure 4 As shown, a truss structure 4 is provided in a matrix in the double-wall cavity of the first double-wall structure and the second double-wall structure respectively; the truss structure 4 is composed of four truss rods, one ends of the four truss rods intersect with each other to form an intersection 41, and the intersection 41 and the other ends of the four truss rods are respectively fixedly connected to the guide vane outer wall 3 and the first inner wall 12 or the second inner wall 23, and four connection points 42 are formed on the inner wall;

[0060] The air film holes 6 are arranged on the outer wall 3 of the guide vane at the intersection 41 , and the impact holes 7 are arranged in a matrix on the first inner wall 12 and the second inner wall 23 .

[0061] The truss structure 4 provided in the double-wall cavity of the first double-wall structure has four connection points 42 on the first inner wall 12 at equal distances from the intersection point 41; the impact hole 7 is provided on the first inner wall 12 outside the four connection points 42 and on the line connecting the intersection point 41 and the center points of the four connection points 42.

[0062] like Figure 2 As shown, the truss structure 4 provided in the second double-wall structure is arranged in a matrix in the double-wall cavity on one side of the blade disk, and the intersection 41 of the truss structure 4 and the other ends of the four truss rods are respectively fixedly connected to the outer wall 3 of the guide vane on the blade basin side and the inner wall 23 of the second cold air cavity 21, and four connection points 42 are formed on the inner wall 23, and the distances from the four connection points 42 to the intersection 41 are equal; the impact holes 7 are arranged on the second inner wall 23 outside the four connection points 42 and on the line connecting the center points of the intersection 41 and the four connection points 42.

[0063] like Figure 6 As shown, the matrix arrangement of the truss structure 4 is specifically as follows: Sx is the arrangement spacing of the truss structure 4 in the flow direction of the guide vane, Sz is the arrangement spacing of the truss structure 4 in the guide vane blade height direction, H is the distance between the guide vane outer wall 3 and the first inner wall 12 or the second inner wall 23, D' is the diameter of the truss rod, and d' is the diameter of the circular hole with the same area as the fan ring hole 61, then:

[0064] H ≥ Sx ≥ 2D’, H ≥ Sz ≥ 2D’;

[0065] π·(d’ / 2) 2 / Sx × Sz = 2.5 - 5%;

[0066] D’ / d’ = 0.5 - 3.5;

[0067] H / d’ = 1 - 2.5;

[0068] The truss structure 4 extends and arranges in sequence at a set distance within the double-wall cavities of the first double-wall structure and the second double-wall structure.

[0069] The normal angle of the truss structure 4 with the wall surface of the first inner wall 12 or the second inner wall 23 is β, where β is 30 - 60°; the length of the truss rod forming the truss structure 4 is l, and H is the distance between the film hole plate 1 and the impact hole plate 2, then l = H / cosβ.

[0070] Such as Figure 5 、 Figure 7 、 Figure 8 As shown, the film hole 6 is a hole arranged around the intersection point 41, including 2 - 4 sector ring holes 61 and sector ring partitions 62 for connecting two adjacent sector ring holes 61 to form a discontinuous annular hole, and the center point of the discontinuous annular hole coincides with the center point of the intersection point 41.

[0071] Such as Figure 8 As shown, the inner diameter of the annular inner ring surrounded by the film hole 6 is d, the outer diameter is D, and the distance between the outer ring and the inner ring is the width B of the sector ring hole 51, then there is:

[0072] On the leading edge 11 of the guide vane, the width B of the film hole 2 < 0.5D is provided to increase the outlet momentum of the film hole and overcome the air flow stagnation and large pressure load;

[0073] On the blade body of the turbine guide vane 1, the value of the width B of the film hole 2 is: 0.1D < B < 0.2D, which is used to obtain the cold air coverage effect.

[0074] Such as Figure 7 As shown, the sector ring hole 61 is a normal through hole or a normal backward expanding hole on the outer wall 3 of the guide vane;

[0075] And when it is a normal through hole on the outer wall 3 of the guide vane, a pair of inner inclination angles α1 = α2 = 90° corresponding to the cross-section of the film hole 6 are set;

[0076] When it is a normal backward expanding hole, the sum of a pair of inner inclination angles corresponding to the cross-section of the film hole 6 is 90 degrees, and:

[0077] 30° ≤ α1 ≤ 90°, 30° ≤ α2 ≤ 90°;

[0078] The fan ring partition 62 has a pair of arc-shaped connecting edges and a pair of parallel or inclined edges;

[0079] When the sides are parallel, the distance L between the two sides is 0.05 to 0.25 mm;

[0080] When it is an inclined edge, the central angle corresponding to the inclined edge is θ, and the range of θ is 15 to 45 degrees.

[0081] The impact hole 7 is a normal through hole of the first inner wall 12 or the second inner wall 23, and the diameter of the impact hole 7 is 0.5-0.75 mm;

[0082] The opening area of ​​the air film hole 6 is 1.4 to 2.4 times the opening area of ​​the impact hole 7.

[0083] In order to achieve a comprehensive cooling efficiency of not less than 0.5 on the entire surface, this embodiment uses the leading edge truss structure matrix row number n1≥2, the guide vane suction surface truss structure matrix row number n2≥3, and the pressure surface truss structure matrix row number n3≥5 to perform simulation, and the simulation results are as follows:

[0084] Fig. 9 The heat transfer coefficient cloud diagram of the double-wall impact target surface of the internal cooling structure matrix is ​​shown. Compared with the double-wall structure of the circular straight column flow, there are low heat transfer areas on both sides of the tail of the straight column spoiler, and the truss structure 4 and the air film hole 6 provided by the present invention can effectively avoid the appearance of the low heat transfer area.

[0085] Fig.11 The streamline distribution diagram of two composite cooling structure monomers in the leading edge area of ​​the blade is shown. Fig.11 As shown in (a), the cooling medium B1 of the double-wall structure of the circular straight column flow is injected vertically to the impact target surface along the axis of the impact hole, forming several rotating vortices B2 near the impact point. After the cooling medium B1 collides with the circular straight column, it diffuses upward along the surface of the column to form the inlet air flow B3 of the air film hole. After the cooling medium flows out of the air film hole, it forms a cold air film B4 attached to the blade surface under the action of the mainstream, forming air film cooling.

[0086] like Fig.11 As shown in (b), the cooling medium C1 of the composite cooling structure provided by the present invention is divided into two parts. One part enters the air film hole close to the upper wall surface to form a cold air film C6 with higher coverage, air flow rate, and more stable and uniformity. The other part is fully disturbed by the truss spoiler column to effectively cool the double-wall structure.

[0087] It can be seen that compared with the cylindrical spoiler layer, the composite cooling structure provided by the present invention has lower flow resistance and is more conducive to taking into account both internal cooling and external air film cooling.

[0088] Fig.12The spanwise average film cooling efficiency of the two is demonstrated in the figure. The double-layer truss wall provided by the present invention effectively reduces the flow resistance of the turbulent structure and improves the film cooling efficiency of the blade surface. The film cooling efficiency of the double-layer truss wall is increased by 21.84%.

[0089] In the blade area (suction and pressure surface areas), the main flow of gas passes over the blade surface, and the angle between the secondary flow jet and the main flow of gas is less than 90°, and there is no hedging factor between the flow directions of the two. Based on the same air film plate opening rate, impact plate opening rate and spoiler column diameter, and the same flow parameters, the tetrahedron truss sandwich double-wall structure is compared with the traditional circular straight column spoiler double-wall structure. The air film cooling effect, comprehensive cooling effect and heat transfer distribution of the impact target surface are shown in the figure. Figure 13-14 As shown. It can be seen that compared with the double-wall structure of the circular straight column flow, the air film cooling efficiency and comprehensive cooling efficiency of the truss sandwich double-wall structure provided by the present invention are significantly improved. For target surface heat exchange, there are low heat exchange areas on both sides of the tail of the straight column spoiler, and the truss structure 4 provided by the present invention can effectively avoid the appearance of this low heat exchange area.

[0090] In summary, although the cylindrical spoiler-circular air film hole composite cooling structure improves the internal cooling effect, the air film cooling effect is reduced due to the excessive flow resistance. The truss structure 4 and air film hole 6 structure provided by the present invention improve the effects of internal cooling and external cooling at the same time, and have better mechanical properties.

[0091] Example 2: Fig.10 As shown, the present invention also provides a guide vane cooling method for the above turbine guide vane, specifically:

[0092] The high-pressure cold air flow C1 drawn out from the compressor enters the first double-wall structure or the second double-wall structure through the impact hole 7, and forms an impact cooling air flow on the wall surface of the first inner wall 12 or the second inner wall 23 directly opposite it. The cooling air flow is divided into two parts, one part of the cooling air flow collides with the adjacent impact jet to form a reverse air flow C2; the other part of the cooling air flow forms a cold air flow C4 on the truss surface of the truss structure 4, and the cold air flow C4 is sucked into the air film hole 6 after heat exchange, and the cold air flow C3 flowing directly against the wall also enters the air film hole 6 at the same time; the air flow C5 flowing out of the air film hole 6 forms a cold air film C6 covering the surface of the turbine guide vane under the action of the mainstream of the fuel gas.

[0093] like Fig.10As shown in (a), the cold air flow C1 of the cylindrical spoiler-annular film hole composite cooling structure monomer enters the double-wall structure through the impact hole, and forms impact cooling on the inner wall surface of the film hole wall opposite to it, wherein a part of the gas collides with the adjacent impact jet to form a reverse air flow C2; most of the cold air flow C4 flows around the truss surface for heat exchange and is sucked into the film hole, while the cold air flow C3 that flows directly along the wall and enters the film hole is less; the air flow C5 flowing out of the film hole is 180° with the mainstream of the gas, and under the action of the mainstream of the gas, a cold air film C6 is formed to cover the surface of the turbine guide vane.

[0094] like Fig.10 As shown in (b), the cold air flow C1 in the method of the present invention enters the double-wall structure through the impact hole, and forms impact cooling on the inner wall surface of the film hole wall opposite to it, wherein a part of the gas collides with the adjacent impact jet to form a reverse air flow C2; the air flow C3 flows directly along the wall into the film hole, and the other part of the air flow C4 flows around the truss surface for heat exchange, and then is sucked into the film hole, and the two are in a similar proportion. The air flow C5 flowing out of the film hole is 180° with the main flow of the gas, and under the action of the main flow of the gas, a cold air film C6 is formed to cover the surface of the turbine guide vane.

[0095] Fig.10 (a) and Fig.10 (b) Comparison shows that: due to the large internal flow resistance of the double-wall, gas intrusion is prone to occur at the air film hole. Compared with the double-wall structure of the circular straight column flow, the composite cooling structure provided by the present invention has a smaller rotating vortex scale, and part of the cold air does not participate in the formation of the rotating vortex and flows out through the air film hole, so it has a lower flow resistance, so that the cold air at the outlet of the air film hole has a larger normal momentum. This is conducive to reducing the flow resistance of the double-wall structure, thereby taking into account both internal cooling and external air film cooling.

[0096] The above description is only a preferred embodiment of the present invention and is not intended to limit the present invention. Any modifications, equivalent substitutions, improvements, etc. made within the spirit and principle of the present invention should be included in the protection scope of the present invention.

Claims

1. A turbine guide vane with a composite cooling structure, characterized in that: The guide vane cavity of the turbine guide vane is divided into a leading edge impact zone (1) and a trailing edge impact zone (2) by a partition (5); The leading edge impact zone (1) comprises a first cold air cavity (11) provided in a cavity from the leading edge of the guide vane to the middle of the guide vane, and a first inner wall (12) of the first cold air cavity (11) and an outer wall (3) of the guide vane form a first double-layer wall structure of the leading edge impact zone (1); The trailing edge impact zone (2) comprises a second cold air cavity (21) and a trailing edge exhaust cavity (22) which are sequentially arranged in the cavity between the partition plate (5) and the trailing edge of the guide vane, and a second inner wall (23) of the second cold air cavity (21) and the outer wall (3) of the guide vane form a second double-layer wall structure of the trailing edge impact zone (2); Truss structures (4) are arranged in a matrix in the double-wall cavities of the first double-wall structure and the second double-wall structure, respectively; the truss structure (4) is composed of four truss rods, one ends of the four truss rods intersect with each other to form an intersection (41), the intersection (41) and the other ends of the four truss rods are respectively fixedly connected to the guide vane outer wall (3) and the first inner wall (12) or the second inner wall (23), and four connection points (42) are formed on the inner wall; The air film holes (6) are arranged on the outer wall (3) of the guide vane at the intersection (41), and the impact holes (7) are arranged in a matrix on the first inner wall (12) and the second inner wall (23).

2. The turbine guide vane with a composite cooling structure according to claim 1, characterized in that: The truss structure (4) provided in the second double-wall structure is arranged in a matrix in the double-wall cavity on one side of the blade disk, the intersection point (41) of the truss structure (4) and the other ends of the four truss rods are respectively fixedly connected to the outer wall (3) of the guide vane on the blade basin side and the inner wall (23) of the second cold air cavity (21), and four connection points (42) are formed on the inner wall (23), and the distances from the four connection points (42) to the intersection point (41) are equal; The impact hole (7) is arranged outside the four connection points (42) and on a line connecting the intersection point (41) and the center points of the four connection points (42).

3. The guide vane cooling method of a turbine guide vane according to claim 1, characterized in that: The distances from four connection points (42) of the truss structure (4) provided in the double-wall cavity of the first double-wall structure on the first inner wall (12) to the intersection point (41) are equal; The impact hole (7) is arranged outside the four connection points (42) and on a line connecting the intersection point (41) and the center points of the four connection points (42).

4. The turbine guide vane with a composite cooling structure according to claim 1, characterized in that: The air film hole (6) is a hole arranged around the intersection (41), including a discontinuous annular hole composed of 2 to 4 fan ring holes (61) and a fan ring partition (62) for connecting two fan ring holes (61), and the center point of the discontinuous annular hole coincides with the center point of the intersection (41).

5. The turbine guide vane with a composite cooling structure according to claim 4, characterized in that: The inner diameter of the annular ring formed by the air film hole (6) is d, the outer diameter is D, and the distance between the outer ring and the inner ring is the width B of the fan ring hole (51), then: An air film hole (2) with a width B<0.5D is provided on the guide vane leading edge (11) to increase the outlet momentum of the air film hole and overcome air flow stagnation and large pressure load; The width B of the film holes (2) provided on the blade body of the turbine guide vane (1) takes a value of: 0.1D < B < 0.2D, which is used to obtain the cold air coverage effect.

6. The turbine guide vane with a composite cooling structure according to claim 4, characterized in that: The fan-shaped ring hole (61) is a normal through hole or a normal backward-expanded hole on the outer wall (3) of the guide vane; When it is a normal through hole on the outer wall (3) of the guide vane, a pair of inner inclination angles α1 = α2 = 90° corresponding to the cross-section of the film hole (6) are set; When it is a normal backward-expanded hole, the sum of a pair of inner inclination angles corresponding to the cross-section of the film hole (6) is 90 degrees, and: 30°≤α1≤90°，30°≤α2≤90°； The fan-shaped ring partition (62) has a pair of arc-shaped connecting sides and a pair of parallel or inclined sides; When they are parallel sides, the distance L between the two sides is 0.05 - 0.25 mm; When they are inclined sides, the central angle corresponding to the inclined sides is θ, and the range of θ is 15 - 45 degrees.

7. The turbine guide vane with a composite cooling structure according to claim 1, characterized in that: The impingement hole (7) is a normal through hole on the first inner wall (12) or the second inner wall (23), and the diameter of the impingement hole (7) is 0.5 - 0.75 mm; The opening area of the film hole (6) is 1.4 - 2.4 times that of the impingement hole (7).

8. The turbine guide vane with a composite cooling structure according to claim 1, characterized in that: The matrix arrangement of the truss structure (4) is specifically as follows: Let Sx be the arrangement pitch of the truss structure (4) in the flow direction of the guide vane, Sz be the arrangement pitch of the truss structure (4) in the blade height direction of the guide vane, H be the distance between the outer wall (3) of the guide vane and the first inner wall (12) or the second inner wall (23), D' be the diameter of the truss rod, and d' be the diameter of a circular hole with the same area as the fan-shaped ring hole (61), then: H ≥ Sx ≥ 2D', H ≥ Sz ≥ 2D'; π·(d' / 2) 2 / Sx×Sz=2.5~5%; D' / d' = 0.5 - 3.5; H / d' = 1 - 2.5; The truss structure (4) is sequentially extended and arranged at a set distance in the double-wall cavity of the first double-wall structure and the second double-wall structure.

9. The heat shield for strengthening heat exchange between double-layer walls according to any one of claims 1 to 8, characterized in that: The normal angle of the truss structure (4) with the wall surface of the first inner wall (12) or the second inner wall (23) is β, and β is 30 - 60°; the length of the truss rod forming the truss structure (4) is l, and H is the distance between the film hole plate (1) and the impingement hole plate (2), then l = H / cosβ.

10. The guide vane cooling method of the turbine guide vane according to claims 1 - 9, characterized in that The high-pressure cold air flow C1 led out from the compressor enters the first double-wall structure or the second double-wall structure through the impingement hole (7), forms an impinging cooling air flow on the wall surface of the first inner wall (12) or the second inner wall (23) directly opposite to it. The cooling air flow is divided into two parts. One part of the cooling air flow collides with the adjacent impinging jets to form a reverse air flow C2; the other part of the cooling air flow forms a cold air bypass flow C4 on the surface of the truss of the truss structure (4). After heat exchange, the cold air bypass flow C4 is sucked into the film hole (6), and the cold air flow C3 flowing directly along the wall also enters the film hole (6) at the same time; the air flow C5 flowing out from the film hole (6) forms a cold air film C6 covering the surface of the turbine guide vane under the action of the mainstream of the combustion gas.

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