Turbine blade double wall cooling structure

Abstract

The invention discloses a double-wall cooling structure for turbine blades, including a blade body and a plurality of outer walls, wherein the blade body has a front end and a rear end along its extension direction, and a front cavity and a rear cavity are sequentially arranged on the blade body from the front end to the rear end, and the front cavity and the rear cavity are independent of each other and not connected; the outer wall is woven from CMC material, and a plurality of outer walls are arranged at intervals with the outer wall surface corresponding to the front cavity, and a first cooling channel is formed between the outer wall and the blade body, and the outer walls are arranged at intervals with each other, and a first cooling gap is formed between two adjacent outer walls, and the first cooling channel is connected with the front cavity and the first cooling gap. It is not necessary to process air film holes on the outer wall made of CMC material, and it relies on the impact cooling of the inner wall surface, the flow-enhanced heat exchange between the inner and outer wall surfaces, and the split air film cooling; the use of CMC material can effectively utilize its high temperature resistance, low density and high strength characteristics, and can effectively reduce the amount of cooling air used and the weight of the blade body.

Description

Technical Field

[0001] The invention relates to the field of turbine blade cooling structures, and in particular to a double-wall cooling structure for turbine blades. Background Art

[0002] The development of aero-engine technology and the improvement of thrust-to-weight ratio largely rely on the increase of turbine gas inlet temperature. According to estimates, for every 55K increase in turbine front inlet temperature, the engine thrust can be increased by 10% under the condition of unchanged engine size. While high temperature brings higher gas turbine efficiency, it also brings a series of problems, such as increased blade heat load, reduced service life, etc. However, the turbine inlet temperature increases at an average rate of 25°C per year, while the metal temperature resistance increases at a rate of only about 10°C per year. It can be seen that the progress of the material's temperature resistance is far behind the development of turbine engines. Therefore, it is particularly important to develop and use more advanced cooling technology, and at the same time use new materials that are resistant to high temperatures, lightweight, high specific strength, and long life to effectively cool the high-temperature components of the engine.

[0003] As a new type of strategic high-temperature structural material, tough silicon carbide ceramic matrix composites (CMC) have the characteristics of high temperature resistance, low density, high specific strength, oxidation resistance, corrosion resistance, and no catastrophic damage. Compared with the nickel-based high-temperature alloy materials used in traditional turbine blades, CMC composite materials can increase the operating temperature of turbine parts by at least 150-200°C, reduce weight by about 1 / 3-2 / 3, simplify or even eliminate the cooling system, greatly reduce the amount of cooling gas, effectively improve engine efficiency, reduce pollution emissions and reduce fuel consumption. Every 1% reduction in fuel consumption means that airlines can save more than 1 million US dollars each year. It is based on the above advantages that CMC composite materials have become an ideal material selection solution for hot end components of advanced commercial aircraft engines, and are also the development direction of future advanced aircraft engine hot end component design and manufacturing technology.

[0004] The double-wall cooling structure has the characteristics of a large number of cooling channels, small size, and large heat exchange area. The composite cooling technology that combines impact-turbine-air film can greatly enhance internal heat exchange, thereby greatly increasing the gas inlet temperature before the turbine, allowing the turbine blades to operate at the allowable temperature of the material and reducing thermal stress. Therefore, this composite cooling technology is also the future development direction of cooling technology. Due to the delicate and complex structure of high-pressure turbine blades, it is impossible to perfectly realize fine structures such as trailing edge structure and smooth air film hole structure when weaving CMC materials. If the CMC woven material is processed with a laser for air film holes, rough broken fibers will be formed on the surface of the material, blocking the outflow of air. Summary of the invention

[0005] The technical problem to be solved by the present invention is to overcome the defect that a fine structure and an air film hole structure cannot be achieved when CMC material is used for weaving in the double-wall cooling structure of turbine blades in the prior art, and to provide a double-wall cooling structure for turbine blades.

[0006] The present invention solves the above technical problems through the following technical solutions:

[0007] A double-wall cooling structure for turbine blades, characterized in that the double-wall cooling structure for turbine blades comprises:

[0008] A blade body, wherein the blade body has a front end and a rear end along its extension direction, and a front cavity and a rear cavity are sequentially arranged on the blade body from the front end to the rear end, and the front cavity and the rear cavity are independent of each other and not connected;

[0009] Multiple outer walls, the material of the outer walls is CMC material, the multiple outer walls are spaced apart from the outer wall surface corresponding to the front cavity, a first cooling channel is formed between the outer walls and the blade body, the outer walls are spaced apart from each other, the outer walls and the blade body are connected by a partition, the partition is provided with flow holes, a first cooling gap is formed between two adjacent outer walls, and the first cooling channel is connected to the front cavity and the first cooling gap.

[0010] In this solution, the outer wall made of CMC material is spaced apart from the outer wall surface corresponding to the front cavity on the blade body, and a first cooling channel is formed to cool the turbine blade. With this structure, on the one hand, it is not necessary to process air film holes on the outer wall made of CMC material, and the impact cooling of the inner wall surface, the flow-enhanced heat exchange between the inner and outer wall surfaces, and the split air film cooling are relied on; on the other hand, the use of CMC material can effectively utilize the high temperature resistance, low density and high strength characteristics of CMC material. The above-mentioned structural form and the combination with CMC material can effectively reduce the amount of cooling air used and the weight of the blade body.

[0011] Preferably, the blade body is formed by casting of single crystal material.

[0012] In this solution, the cooling performance of the double-wall cooling structure of the turbine blade is improved by using a blade body cast from a single crystal material combined with an outer wall of a CMC material. In addition, the blade body is cast from a single crystal material because the impact holes with inner walls and the flow holes on the partition cannot be processed externally. The use of a single crystal casting method using a water-soluble core can circumvent the characteristic that CMC materials cannot weave fine structures.

[0013] Preferably, a plurality of first impact holes spaced apart from each other are arranged on the side wall of the front cavity, and the first cooling channel and the cavity body are communicated with each other through the first impact holes.

[0014] In this solution, through the above structural form, the cooling air enters the front cavity from the bottom of the blade, and a part of it passes through the first impact hole on the inner wall of the leading edge to strengthen the cooling of the outer wall of the CMC material, and then flows out from the first cooling gap at the edge of the CMC and forms an air film on the surface of the blade.

[0015] Preferably, a partition plate for supporting the outer wall is provided between the outer wall and the blade body, and the partition plate is connected to the outer wall and the blade body.

[0016] In this solution, the blade body and the outer wall are connected by a partition plate instead of a spoiler column, mainly considering preventing the outer wall of the CMC material from being deformed due to heat and improving the stability of the installation.

[0017] Preferably, a plurality of partitions are provided on each of the outer walls, the partitions divide the first cooling channel into a plurality of cooling spaces, the partitions are provided with flow holes, and the plurality of cooling spaces are connected through the flow holes.

[0018] In this solution, the above structural form is adopted. On the one hand, the flow path of the cooling air can be increased, thereby improving the cooling effect; on the other hand, the connection strength between the outer wall of the CMC material and the blade body can be further improved, which is more conducive to preventing the outer wall of the CMC material from being deformed due to heat and improving the stability of the installation.

[0019] Preferably, the side wall corresponding to the rear cavity on the blade body is double-layered, with a second cooling channel between the side walls on both sides, the front end of the blade body has a third cooling channel, the second cooling channel is connected to the rear cavity and the third cooling channel, and the end of the front end has a second cooling gap, which is connected to the third cooling channel.

[0020] In this solution, cooling air enters the blade rear cavity from above the blade, flows from the rear cavity into the second cooling channel on the pressure surface and the suction surface, and finally flows into the third cooling channel and flows out of the second cooling gap, thereby achieving cooling of the blade.

[0021] Preferably, a plurality of spoiler columns spaced apart from each other are arranged in the second cooling channel and the third cooling channel.

[0022] In this solution, a plurality of mutually spaced spoiler columns are arranged in the second cooling channel and the third cooling channel. The cooling air on the pressure surface side is disturbed by the spoiler columns between the inner and outer walls to enhance heat exchange, which is beneficial to improving cooling efficiency.

[0023] Preferably, a plurality of second impact holes spaced apart from each other are arranged on the inner wall corresponding to the rear cavity, and a plurality of air film holes spaced apart from each other are arranged on the outer wall corresponding to the rear cavity.

[0024] In this scheme, cooling air enters the rear cavity of the blade from above the blade, and impact-cools the inner layer of the outer wall of the rear cavity from the second impact hole in the inner wall of the rear cavity. The cooling air on the pressure side is disturbed by the spoiler columns between the inner and outer walls to enhance heat exchange, and flows out from the air film holes to form an air film on the surface of the blade body for air film cooling. The cooling air on the suction side is disturbed by the spoiler columns between the inner and outer walls to enhance heat exchange, flows to the trailing edge of the blade, and flows out from the air film holes and the second cooling gap at the trailing edge after the disturbance of the spoiler columns at the trailing edge enhances heat exchange. The rear cavity of the blade forms a load cooling structure of impact-turbine-air film-slit cooling, which effectively improves the cooling efficiency.

[0025] Preferably, air film holes are provided on the side walls of the third cooling channel.

[0026] In this solution, through the above structural form, the cooling air forms an air film on the side wall corresponding to the third cooling channel, thereby further improving the cooling efficiency.

[0027] Preferably, at least one side surface of the first cooling gap is an inclined surface.

[0028] In this solution, through the above structural form, a split structure is formed between the first cooling gaps to improve the cooling efficiency.

[0029] On the basis of being in accordance with the common sense in the art, the above-mentioned preferred conditions can be arbitrarily combined to obtain the preferred embodiments of the present invention.

[0030] The positive and progressive effect of the present invention is that the double-wall cooling structure of the turbine blade of the present invention is to set the outer wall made of CMC material and the outer wall surface corresponding to the front cavity on the blade body at intervals, and form a first cooling channel to cool the turbine blade. With this structure, on the one hand, it is not necessary to process air film holes on the outer wall made of CMC material, and the impact cooling of the inner wall surface, the flow-enhanced heat exchange between the inner and outer wall surfaces and the split air film cooling are relied on; on the other hand, the use of CMC material can effectively utilize the characteristics of CMC material such as high temperature resistance, low density and high strength. The above-mentioned structural form and the combination with CMC material can effectively reduce the amount of cooling air used and the weight of the blade body. BRIEF DESCRIPTION OF THE DRAWINGS

[0031] Figure 1 It is a schematic structural diagram of a turbine blade with a double-wall cooling structure according to a preferred embodiment of the present invention.

[0032] Figure 2This is a schematic diagram of the three-dimensional structure of a turbine blade cut from the middle according to a preferred embodiment of the present invention.

[0033] Figure 3 This is a schematic structural diagram of the double-wall cooling structure of a turbine blade after being cut open from the middle in a preferred embodiment of the present invention.

[0034] Figure 4 Schematic diagram of the structure of the outer wall of the CMC material according to a preferred embodiment of the present invention.

[0035] Description of reference numerals:

[0036] Front cavity 1

[0037] Rear cavity 2

[0038] First impact hole 3

[0039] Leading edge CMC material outer wall 4

[0040] Suction surface CMC material outer wall 5

[0041] Pressure surface CMC material outer wall 6

[0042] Partition 7

[0043] Suction surface split 8

[0044] First slit 9

[0045] Second slit 10

[0046] Pressure surface split 11

[0047] Second impact hole 12

[0048] Spoiler Post 13

[0049] Air film hole 14

[0050] Flow hole 15

[0051] Second cooling gap 16 DETAILED DESCRIPTION

[0052] The present invention will be described more clearly and completely below by way of embodiments in conjunction with the accompanying drawings, but the present invention is not limited to the scope of the embodiments.

[0053] like Figures 1 to 4As shown, this embodiment provides a double-layer wall cooling structure for turbine blades, which includes a blade body and multiple outer walls. The blade body has a front end and a rear end along its extension direction, and a front cavity 1 and a rear cavity 2 are sequentially arranged on the blade body from the front end to the rear end, and the front cavity 1 and the rear cavity 2 are independent of each other and not connected; the outer wall is woven from CMC material, and multiple outer walls are arranged at intervals with the outer wall surface corresponding to the front cavity 1, and a first cooling channel is formed between the outer wall and the blade body, and the outer walls are arranged at intervals with each other, and a first cooling gap is formed between two adjacent outer walls, and the first cooling channel is connected with the front cavity 1 and the first cooling gap. The outer wall and the blade body are connected by a partition, and the partition divides the first cooling channel into multiple narrow channels. The partition is provided with flow holes, and the narrow channels are connected through the flow holes.

[0054] Among them, the wall thickness formed by the blade body is set to about 1mm. The outer wall made of CMC material is spaced apart from the outer wall surface corresponding to the front cavity 1 on the blade body, and a first cooling channel is formed to cool the turbine blade. With this structure, on the one hand, it is not necessary to process the air film hole 14 on the outer wall made of CMC material, and the impact cooling of the inner wall surface, the flow-enhanced heat exchange between the inner and outer wall surfaces and the split air film cooling are relied on; on the other hand, the use of CMC material can effectively utilize the high temperature resistance, low density and high strength characteristics of CMC material. The above-mentioned structural form and the combination with CMC material can effectively reduce the amount of cooling air used and the weight of the blade body, improve the efficiency of cold air use and also improve the efficiency of the engine. In this embodiment, the curvature of the outer wall made of CMC material is small, which can effectively reduce the difficulty of weaving CMC material, thereby improving the yield rate. At least one side of the first cooling slit is an inclined surface. A split structure is formed between the first cooling slits to improve the cooling efficiency.

[0055] In this embodiment, the blade body is cast by single crystal material. The cooling performance of the double-wall cooling structure of the turbine blade is improved by combining the blade body cast by single crystal material with the outer wall of CMC material. In addition, the blade body is cast by single crystal material because the impact holes of the inner wall and the flow holes on the partition 7 cannot be processed externally. The method of using single crystal casting and water-soluble core can avoid the characteristic that CMC material cannot weave fine structures.

[0056] Furthermore, a plurality of mutually spaced first impact holes 3 are provided on the side wall of the front cavity 1, and the first cooling channel and the front cavity are connected through the first impact holes 3. Through the above structural form, the cooling air enters the front cavity 1 from the bottom of the blade, and a part of it passes through the first impact holes 3 on the inner wall of the leading edge to strengthen the cooling of the outer wall 4 of the leading edge CMC material, and then flows out from the first cooling gap at the edge of the CMC and forms an air film on the surface of the blade. In this embodiment, two rows of first impact holes 3 are opened on the inner wall of the leading edge of the front cavity 1, and the number of first impact holes 3 in each row is determined according to the size of the blade; five rows of first impact holes 3 are provided on the inner wall corresponding to the suction surface of the front cavity 1; and two rows of first impact holes 3 are provided on the inner wall corresponding to the pressure surface.

[0057] A partition 7 for supporting the outer wall is provided between the outer wall and the blade body, and the partition 7 is connected to the outer wall and the blade body. The blade body and the outer wall are connected by the partition 7, rather than the spoiler column 13, mainly to prevent the outer wall of the CMC material from being deformed by heat and to improve the stability of the installation. In this embodiment, the outer wall of the CMC material is connected to the inner wall corresponding to the front cavity 1 of the blade body through the upper and lower edge plates and the partition 7. A plurality of partitions 7 are provided on each outer wall, and the partitions 7 divide the first cooling channel into a plurality of cooling spaces. The partitions 7 are provided with flow holes, and the plurality of cooling spaces are connected through the flow holes. By adopting the above structural form, on the one hand, the flow path of the cooling air can be increased, thereby improving the cooling effect; on the other hand, the connection strength between the outer wall of the CMC material and the blade body can be further improved, which is more conducive to preventing the outer wall of the CMC material from being deformed by heat and improving the stability of the installation. In this embodiment, the leading edge CMC material outer wall 4 corresponding to the leading edge of the front cavity 1 of the blade body is connected to the inner wall via upper and lower edge plates; the suction surface CMC material outer wall 5 corresponding to the suction surface of the front cavity 1 is connected to the inner wall via three partitions, the three partitions 7 divide the first cooling channel into four narrow cooling spaces, and the flow holes 15 arranged on the partitions 7 connect the cooling spaces; the pressure surface CMC material outer wall 6 corresponding to the pressure surface is connected to the inner wall via a partition 7 and upper and lower edge plates, and a partition 7 divides the first channel into two narrow cooling spaces.

[0058] In this embodiment, the first cooling gap is a split gap, and the first cooling gap between the outer wall 6 of the pressure surface CMC material and the blade body is a pressure surface split gap 11; the first cooling gap between the outer wall 4 of the leading edge CMC material and the outer wall 6 of the pressure surface CMC material is a first split gap 9, and the first cooling gap between the outer wall 4 of the leading edge CMC material and the outer wall 5 of the suction surface CMC material is a second split gap 10; the first cooling gap between the outer wall 5 of the suction surface CMC material and the blade body is a suction surface split gap 8.

[0059] Furthermore, the side wall corresponding to the rear cavity 2 on the blade body is double-layered, with a second cooling channel between the side walls on both sides, and a third cooling channel at the front end of the blade body, the second cooling channel is connected to the rear cavity 2 and the third cooling channel, and the end of the front end has a second cooling gap 16, and the second cooling gap 16 is connected to the third cooling channel. The cooling air enters the rear cavity of the blade from the top of the blade, flows from the rear cavity into the second cooling channel on the pressure surface and the suction surface, and finally flows into the third cooling channel and flows out from the second cooling gap 16, thereby realizing the cooling effect on the blade. In order to further improve the cooling efficiency, the outer surface of the outer wall of the rear cavity is covered with a thermal insulation coating.

[0060] In this embodiment, a plurality of mutually spaced spoiler columns 13 are provided in the second cooling channel and the third cooling channel. A plurality of mutually spaced spoiler columns 13 are provided in the second cooling channel and the third cooling channel, and the cooling air on the pressure surface side is disturbed by the spoiler columns 13 between the inner and outer walls to enhance heat exchange, which is beneficial to improving the cooling efficiency. Since the spoiler columns 13 and the second cooling gap are narrow and long in size, and considering that it is too difficult to weave using CMC materials, in this embodiment, only the outer wall woven with CMC materials is used on the outside of the inner wall corresponding to the front cavity 1 of the blade body. The outer wall corresponding to the rear cavity 2 and other parts are finely processed by casting. This arrangement can not only improve the blade yield, but also meet the blade use requirements through a composite cooling structure. Air film holes 14 are provided on the side walls of the third cooling channel, so that the cooling air forms an air film on the side walls corresponding to the third cooling channel, thereby further improving the cooling efficiency. There is no limit to the number of spoiler columns 13 , four rows of spoiler columns 13 are arranged between the inner and outer walls corresponding to the suction surface, and four rows of spoiler columns 13 are arranged at the trailing edge; three rows of spoiler columns 13 are arranged between the inner and outer walls corresponding to the pressure surface.

[0061] Furthermore, a plurality of mutually spaced second impact holes 12 are provided on the inner wall corresponding to the rear cavity 2, and a plurality of mutually spaced air film holes 14 are provided on the outer wall corresponding to the rear cavity 2. The cooling air enters the rear cavity of the blade from above the blade, and the inner layer of the outer wall of the rear cavity 2 is impact-cooled from the second impact holes 12 on the inner wall of the rear cavity 2. The cooling air on the pressure side is disturbed by the spoiler columns 13 between the inner and outer walls to enhance heat exchange, and flows out from the air film holes 14 to form an air film on the surface of the blade body for air film cooling. The cooling air on the suction side is disturbed by the spoiler columns 13 between the inner and outer walls to enhance heat exchange, and flows to the trailing edge of the blade. After the spoiler columns 13 at the trailing edge enhance heat exchange, the cooling air flows out from the air film holes 14 and the second cooling gap 16 at the trailing edge. The rear cavity 2 of the blade forms a composite cooling structure of impact-turbine-air film-slit cooling, which effectively improves the cooling efficiency and reduces the amount of cold air used. There is no limitation on the shapes of the air film holes 14 on the outer wall of the third cooling channel and the rear cavity 2 , and circular holes are used in this embodiment.

[0062] The cooling process of the double-wall cooling structure of the turbine blade provided in this embodiment is as follows:

[0063] Reference Figure 2 and Figure 3 It is understood that the cooling air enters the front cavity 1 of the blade body from the lower air inlet, passes through the impact holes on the inner wall of the front cavity 1, and is divided into three air flows to impact-cool the inner wall of the outer wall woven from CMC material. Among them, a stream of cooling air impact-cools the outer wall surface of the outer layer wall of the leading edge CMC material through two rows of impact holes on the leading edge, and flows out from the first slits 9 and the second slits 10 formed on both sides of the outer layer wall 4 of the leading edge CMC material, and forms an air film on the blade body; another stream of cooling air enters the first cooling channel corresponding to the suction surface through the five rows of first impact holes 3 on the suction surface to impact-cool the outer layer wall 5 of the suction surface CMC material. After the cooling air enters the first cooling channel from the first impact holes 3 for cooling, it passes through the flow holes on the partition 7 in sequence, and finally flows out from the suction surface slits 8 corresponding to the suction surface, and forms an air film cooling suction surface of the blade; the last stream of cooling air impact-cools the outer wall surface of the pressure surface CMC material through the first impact holes 3 on the pressure surface. After the cooling air enters the first cooling channel corresponding to the pressure surface from the first impact holes 3, it passes through the flow holes on the partition 7, and finally flows out from the pressure surface slits 11 corresponding to the pressure surface, forming an air film cooling pressure surface of the blade.

[0064] The cooling air enters the rear cavity 2 of the blade from the upper air inlet, and is divided into two air flows through the impact holes on the inner wall to impact and cool the inner wall of the outer wall. Among them, one cooling air impacts the inner side of the outer wall of the cooling blade through the second impact hole 12 on the suction surface, flows into the third cooling channel through the spoiler column 13 arranged between the inner and outer walls, and after passing through the spoiler column 13 at the trailing edge of the blade, a part of it flows out through the air film hole 14 on the trailing edge for air film cooling, and a part of it flows into the mainstream through the second cooling gap 16 to cool the high-temperature area of ​​the trailing edge; the other cooling air impacts the inner side of the outer wall of the cooling blade through the second impact hole 12 on the pressure surface, and disturbs the outer wall surface through the spoiler column 13 arranged between the inner and outer walls to enhance heat exchange. After that, the cooling air flows into the mainstream through the air film hole 14 on the pressure surface to form an air film, and cools the high-temperature area of ​​the pressure surface of the blade. Since the outer wall is relatively thin, a circular hole design is adopted to ensure the aspect ratio of the air film hole 14.

[0065] Although the specific embodiments of the present invention are described above, it should be understood by those skilled in the art that this is only for illustration and the protection scope of the present invention is defined by the appended claims. Those skilled in the art may make various changes or modifications to these embodiments without departing from the principles and essence of the present invention, but these changes and modifications all fall within the protection scope of the present invention.

Claims

1. A double-wall cooling structure for turbine blades, characterized in that: The turbine blade double-wall cooling structure comprises: A blade body, wherein the blade body has a front end and a rear end along its extension direction, and a front cavity and a rear cavity are sequentially arranged on the blade body from the front end to the rear end, and the front cavity and the rear cavity are independent of each other and not connected; Multiple outer walls, the material of the outer walls is CMC material, the multiple outer walls are spaced apart from the outer wall surface corresponding to the front cavity, a first cooling channel is formed between the outer walls and the blade body, the outer walls are spaced apart from each other, a first cooling gap is formed between two adjacent outer walls, and the first cooling channel is connected to the front cavity and the first cooling gap.

2. The double-wall cooling structure of a turbine blade according to claim 1, characterized in that: The blade body is formed by casting of single crystal material.

3. The double-wall cooling structure of a turbine blade according to claim 1, characterized in that: A plurality of mutually spaced first impact holes are arranged on the side wall of the front cavity, and the first cooling channel and the cavity body are connected through the first impact holes.

4. The double-wall cooling structure for turbine blades according to claim 1, characterized in that: A partition plate for supporting the outer wall is arranged between the outer wall and the blade body, and the partition plate is connected to the outer wall and the blade body.

5. The double-wall cooling structure for turbine blades according to claim 4, characterized in that: A plurality of partitions are arranged on each of the outer walls, and the partitions divide the first cooling channel into a plurality of cooling spaces. The partitions are provided with flow holes, and the plurality of cooling spaces are connected through the flow holes.

6. The double-wall cooling structure for turbine blades according to claim 1, characterized in that: The side wall corresponding to the rear cavity on the blade body is double-layered, and a second cooling channel is formed between the double-layer side walls. The front end of the blade body has a third cooling channel, and the second cooling channel is connected to the rear cavity and the third cooling channel. The end of the front end has a second cooling gap, and the second cooling gap is connected to the third cooling channel.

7. The double-wall cooling structure of a turbine blade according to claim 6, characterized in that: A plurality of spoiler columns spaced apart from each other are arranged in the second cooling channel and the third cooling channel.

8. The double-wall cooling structure for turbine blades according to claim 6, characterized in that: A plurality of second impact holes spaced apart from each other are arranged on the inner wall corresponding to the rear cavity, and a plurality of air film holes spaced apart from each other are arranged on the outer wall corresponding to the rear cavity.

9. The double-wall cooling structure for turbine blades according to claim 6, characterized in that: Air film holes are arranged on the side wall of the third cooling channel.

10. The double-wall cooling structure for turbine blades according to any one of claims 1 to 9, characterized in that: At least one side surface of the first cooling gap is an inclined surface.

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