Ceramic matrix composite turbine vane with cooling channels and method of making

Abstract

Provided is a method for manufacturing a ceramic matrix composite turbine blade with a cooling channel, which uses a fiber cloth to wrap an inner cavity core mold, forms an inner cavity fiber preform, forms a recessed area in a local area on the outer side of the inner cavity fiber preform; places a filling fiber preform in the recessed area, which includes an internal embedded body, uses a fiber cloth to wrap the inner cavity fiber preform and the filling fiber preform, forms a blade fiber preform with a blade profile shape, and performs post-processing on the blade fiber preform to remove the inner cavity core mold and the internal embedded body, thereby forming a cold gas inner cavity and a cooling channel of the turbine blade, respectively. The above method can be used to manufacture a turbine blade with good cooling performance. A turbine blade assembly is also provided.

Description

Technical Field

[0001] This invention relates to the field of aero-engines, and more specifically to the field of turbine blades. Background Technology

[0002] Ceramic matrix composites are commonly used in the fabrication of turbine blades. Generally, turbine blades are affected by internal and external heat transfer, resulting in varying heat loads in different regions. Temperature differences between the inside and outside of the blade can easily lead to temperature gradients. Due to manufacturing constraints, typical ceramic matrix composite turbine blades usually employ a uniform wall thickness, making it difficult to form the intricate cooling channels that enhance internal heat transfer, as is possible with conventional high-temperature metals. Introducing cooling channels into cured ceramic matrix composite turbine blades through machining is difficult and can easily damage the inherent structure of the ceramic matrix composite. Summary of the Invention

[0003] One object of the present invention is to provide a method for preparing a ceramic matrix composite turbine blade, so as to prepare a turbine blade with better cooling performance.

[0004] The turbine blade manufacturing method for achieving the above objectives includes the following steps: wrapping an inner cavity mandrel with fiber cloth to form an inner cavity fiber preform; forming a recessed area in a local region on the outer side of the inner cavity fiber preform; placing a filling fiber preform in the recessed area, the filling fiber preform including an internal embedded body; wrapping the inner cavity fiber preform and the filling fiber preform with fiber cloth to form a blade fiber preform with a blade shape; and performing post-processing on the blade fiber preform to remove the inner cavity mandrel and the internal embedded body to form the cold air cavity and cooling channel of the turbine blade, respectively.

[0005] In one or more embodiments, a plurality of inner cavity fiber preforms are obtained, and the recessed area is prepared on at least one of the inner cavity fiber preforms; the plurality of inner cavity fiber preforms are spliced ​​together; and the filling fiber preform and the plurality of inner cavity fiber preforms are wrapped simultaneously with fiber cloth to form a blade fiber preform.

[0006] In one or more embodiments, a local high-temperature zone is identified in the blade, and a recessed area is prepared in the region where the inner cavity fiber preform is located within the local high-temperature zone.

[0007] In one or more embodiments, the local high-temperature zone includes the leading edge region of the blade, the trailing edge region of the blade, the pressure side region, and the suction side region.

[0008] In one or more embodiments, the recessed area includes an irregular groove surface.

[0009] In one or more embodiments, a connecting hole is machined on the turbine blade, the connecting hole connecting the cooling gas cavity and the outside of the blade, or connecting the cooling channel and the outside of the blade.

[0010] In one or more embodiments, the inner cavity core mold is provided with a recessed portion, and the recessed portion is wrapped with a fiber cloth to form the recessed area.

[0011] In one or more embodiments, the recessed area is filled with whole or shredded fiber cloth, such that the internal pre-embedded body is located within the fiber cloth, forming the filled fiber preform.

[0012] In one or more embodiments, the internal embedded body extends from the blade rim to the blade tip in the radial direction of the blade to form a radial cooling channel.

[0013] In one or more embodiments, the internal embedded body includes a turbulence cavity perpendicular to the blade wall and a matrix slurry, the matrix slurry filling the turbulence cavity.

[0014] Another object of the present invention is to provide a turbine blade assembly, including a turbine blade and a metal sleeve, wherein the turbine blade includes a cooling gas cavity and a cooling channel, the metal sleeve is disposed in the cooling gas cavity, and the distance between the center of the cooling channel and the outer wall surface of the turbine blade is 1 / 2 to 1 / 3 of the turbine blade wall thickness.

[0015] In one or more embodiments, the turbine blades are prepared by the above method.

[0016] In one or more embodiments, the cooling channel is located in a localized high-temperature zone of the blade.

[0017] In one or more embodiments, the local high-temperature zone includes one or a portion of the blade leading edge region, blade trailing edge region, pressure side region, and suction side region.

[0018] In one or more embodiments, the turbine blade includes a plurality of the cooling gas cavities, and the cooling channel is disposed in the connection area of ​​adjacent cooling gas cavities.

[0019] In one or more embodiments, the turbine blades further include film cooling holes communicating with the cold air cavity and / or the cooling channel.

[0020] The aforementioned turbine blade manufacturing method is based on the manufacturing process of ceramic matrix composite materials. It designs the internal cooling gas cavity layup scheme to introduce a cooling channel close to the outer wall, thereby making the cooling gas in the cooling channel closer to the outer wall of the blade, providing a reasonable temperature distribution layout for the blade while ensuring the blade strength. Attached Figure Description

[0021] The above and other features, properties and advantages of the present invention will become more apparent from the following description taken in conjunction with the accompanying drawings and embodiments, wherein:

[0022] Figure 1 This is a cross-sectional view of the first embodiment of the blade fiber preform;

[0023] Figure 2 This is a cross-sectional view of a second embodiment of the blade fiber preform;

[0024] Figure 3 This is a cross-sectional view of the third embodiment of the blade fiber preform;

[0025] Figure 4 This is a cross-sectional view of the first embodiment of the turbine blade assembly;

[0026] Figure 5 This is a longitudinal cross-sectional view of a first embodiment of the turbine blade assembly;

[0027] Figure 6 This is a cross-sectional view of the fourth embodiment of the blade fiber preform;

[0028] Figure 7 This is a cross-sectional view of a second embodiment of the turbine blade assembly;

[0029] Figure 8 This is a cross-sectional view of the fifth embodiment of the blade fiber preform;

[0030] Figure 9 This is a cross-sectional view of a third embodiment of the turbine blade assembly;

[0031] Figure 10 This is a schematic diagram of the turbulence cavity and the matrix slurry;

[0032] Figure 11 This is an enlarged cross-sectional view of the cooling channel. Detailed Implementation

[0033] The present invention will be further described below with reference to specific embodiments and accompanying drawings. More details are set forth in the following description in order to provide a full understanding of the present invention. However, the present invention can obviously be implemented in many other ways different from those described herein. Those skilled in the art can make similar extensions and derivations based on actual application situations without departing from the spirit of the present invention. Therefore, the scope of protection of the present invention should not be limited by the content of this specific embodiment.

[0034] It should be noted that these and other accompanying drawings are merely examples and are not drawn to scale, and should not be construed as limiting the scope of protection of the present invention.

[0035] Figure 4 and 5 The basic structure of a turbine blade assembly 1 is shown, including a turbine blade 2 and a sleeve 3. The turbine blade 2 is made of ceramic matrix composite material, and the sleeve 3 is made of metal material. The sleeve 3 is placed inside the cold zone cavity 21 of the turbine blade 2. The sleeve 3 has an impact hole 7 that penetrates the outer wall surface 32 and the inner wall surface 33 of the sleeve, connecting the air supply cavity 31 of the sleeve and the cold air cavity 21 of the blade. Cold air enters the air supply cavity 31 of the sleeve from the cold air inlet 8, and then flows into the cold air cavity 21 of the blade through the impact hole 7, providing impact cooling to the inner wall surface 29 of the blade. The turbine blade 2 has film cooling holes 6. The film cooling holes 6 can be arranged on the leading edge 22, pressure surface 23, suction surface 24 and trailing edge 25 of the blade, for providing film cooling to the outer wall surface 28 of the blade. The film cooling holes 6 penetrate the outer wall surface 28 and the inner wall surface 29 of the blade, allowing the cold air in the cold air cavity 21 of the blade to flow out through the film cooling holes 6.

[0036] Typical ceramic matrix composite turbine blades are prone to uneven temperature distribution and excessive temperature gradients, resulting in localized overheating areas that adversely affect blade structure and lifespan. Therefore, this disclosure proposes a method for manufacturing ceramic matrix composite turbine blades with radial cooling channels. This method constructs cooling channels targeting localized high-temperature areas of the blade, thereby optimizing the temperature gradient and reducing blade temperature.

[0037] Reference Figures 1 to 3 The preparation method includes the following steps: wrapping the inner cavity mandrel 40 with fiber cloth 41 to form an inner cavity fiber preform 42; forming a recessed area 46 in a local area on the outside of the inner cavity fiber preform 42, such as setting a recessed part on the surface of the inner cavity mandrel 40 and wrapping the recessed part with fiber cloth 41 to form the recessed area 46; placing a filling fiber preform 44, which includes an internal embedded body 45, in the recessed area 46; wrapping the inner cavity fiber preform 42 and the filling fiber preform 44 with fiber cloth 41 to form a blade fiber preform 43 with a blade blade shape; performing post-processing on the blade fiber preform 43, such as curing, carbonization, silicon infiltration, etc., and removing the inner cavity mandrel 40 and the internal embedded body 45 to form the cooling air cavity 21 and cooling channel 26 of the turbine blade, respectively. Figure 4 As shown.

[0038] The extension direction of the cooling channel 26 is determined by adjusting the extension direction of the internal embedded body 45 of the fiber-filled preform 44. In some embodiments, the recessed area 46 is filled with whole fiber cloth or shredded fiber cloth, so that the internal embedded body 45 is inserted into the fiber cloth 41 to form the aforementioned fiber-filled preform 44.

[0039] Preferably, the internal embedded body 45 extends radially from the blade edge 20 to the blade tip, penetrating the blade to form a radial cooling channel. Alternatively, the cooling channel 26 may penetrate the blade root but not the blade tip, or it may penetrate the blade tip but not the blade root. The cooling channel 26 forms a radial cooling channel air supply port 5 at its penetration point at the blade root and / or tip.

[0040] Furthermore, the fiber preform 44 may include multiple internal embedded bodies 45, forming multiple cooling channels. The cross-section of a single cooling channel 26 can be circular, rectangular, trapezoidal, or other regular or irregular shapes, such as... Figures 1 to 3 As shown.

[0041] In some embodiments, when preparing the fiber-filled preform 44, a cooling channel-shaped internal pre-embedded body 45 can be made of polymer materials such as nylon and polyimide. During the high-temperature process of subsequent curing and densification of the blade, the polymer material internal pre-embedded body 45 is heated into gas and loose carbon, and removed from the preform to form a cooling channel 26.

[0042] The fiber cloth 41 can be a fiber-reinforced material, including unidirectional tape, two-dimensional cloth, three-dimensional fabric, etc. Multiple layers of fiber cloth 41 are wound around the inner cavity mandrel 40 to form an inner cavity fiber preform 42. A recessed region 46 is prepared in the region of the inner cavity fiber preform 42 located in the local high-temperature zone A. The local high-temperature zone A can be obtained through previous simulation methods or based on experience. The local high-temperature zone includes, but is not limited to, the leading edge region of the blade, the trailing edge region of the blade, the pressure side region, and the suction side region.

[0043] After splicing the inner fiber preform 42 with the filling fiber preform 44, multiple layers of fiber cloth 41 are wrapped and laid on the outside to obtain the integral fiber preform 43 of the blade with the designed blade shape. The gap left after removing the original internal pre-embedded body 45 forms a cooling channel 26. The multiple layers of fiber cloth spliced ​​and contacting between adjacent cooling air cavities inside the blade form a partition 27.

[0044] In some embodiments, the internal embedded body 45 includes a turbulence cavity 9 perpendicular to the blade wall and a matrix slurry 91, the matrix slurry 91 filling the turbulence cavity 9. For example... Figure 10As shown, an internal pre-embedded body 45 is made of polymer materials such as nylon and polyimide, and a turbulence cavity 9 perpendicular to the corresponding blade wall is processed radially in the internal pre-embedded body 45, and filled into it with a toughened matrix slurry 91 made of particles, chopped fibers, etc. During the high-temperature process of subsequent curing and densification of the blade, the polymer material is heated into gas and loose carbon, which is removed from the fiber-filled pre-embedded body 44 to form a cooling channel 26; the toughened matrix slurry 91 made of particles, chopped fibers, etc., is densified to form a turbulence structure (not shown in the figure) perpendicular to the corresponding blade wall and connecting the inner wall of the cooling channel 26, which is used to enhance the heat transfer process in the radial cooling microchannel 26.

[0045] The shape of the recessed area 46 is determined based on a combination of factors, including the blade profile and the area of ​​the localized high-temperature zone. The recessed area may include irregular grooves; the groove 460 of the recessed area refers to the surface portion that constitutes the interior of the recessed area, such as... Figure 6 or Figure 8 As shown, the blade fiber preform 43 formed by splicing the filling fiber preform 44 located in the recessed area 46 with the inner cavity fiber preform 42 has the designed blade shape.

[0046] In some embodiments, the turbine blade includes multiple cooling gas cavities 21. During fabrication, different cavity mandrels 40 are wrapped with fiber cloth 41 to obtain multiple cavity fiber preforms 42 (42'), such as... Figure 8 As shown. After determining the local high-temperature zone of the blade, a recessed area 46 is prepared on at least one inner cavity fiber preform where the local high-temperature zone is located. Multiple inner cavity fiber preforms are spliced ​​together, and fiber cloth is used to simultaneously wrap and fill the fiber preform 44 and multiple inner cavity fiber preforms 42 (42') to form a blade fiber preform 43, and a ceramic matrix composite turbine blade 2 is obtained after a post-processing process.

[0047] After obtaining the ceramic matrix composite turbine blade 2, connecting holes are machined on the turbine blade. These connecting holes either connect the cooling gas cavity 21 and the outside of the blade, forming a film cooling hole 6 on the turbine blade; or connect the cooling channel 26 and the outside of the blade, forming a film cooling hole 6'. Figure 4 As shown. The impact hole 7 on the metal sleeve 3 and the air supply chamber 31 on the sleeve form a cooling flow path. In some embodiments, the impact hole 7 on the metal sleeve 3 is arranged in the corresponding area of ​​the cooling channel 26, which can cool the cooling channel 26.

[0048] exist Figure 4 and Figure 5In the embodiment shown, during cooling, all the cold air C enters the sleeve air supply cavity 31 from the cold air inlet 8, flows into the blade cold air cavity 21 through the impact hole 7, and then a portion of the cold air flows out from the air film hole 6 that penetrates the inner wall surface 29 of the blade; the other portion of the cold air flows to the blade tip, then flows into the cooling channel 26 from the radial cooling channel air supply port 5, and finally flows out from the air film hole 6 that penetrates the cooling channel 26.

[0049] Preferably, the number of blade cooling air cavities 21 and the location of cooling channels 26 are determined according to the blade's heat load distribution and cooling requirements. Based on the fit between the recessed area 46 and the fiber preform 44, the blade wall thickness between the inner wall of the cooling channel 26 and the outer wall of the blade 28 is less than the wall thickness between the inner wall of the blade 29 and the outer wall of the blade 28, and the blade wall thickness between the inner wall of the cooling channel 26 and the inner wall of the blade 29 is less than the wall thickness between the inner wall of the blade 29 and the outer wall of the blade 28. For example, the distance L1 from the center O of the cooling channel to the outer wall of the turbine blade is 1 / 2 to 1 / 3 of the turbine blade wall thickness. The turbine blade wall thickness is as follows: Figure 11 As shown in L0, this represents the wall thickness between the inner wall surface 29 and the outer wall surface 28 of the blade formed by only wrapping the fiber cloth 41, without the filling fiber preform 44. The turbine blade wall thickness is approximately equal in different regions of the blade. Therefore, the cooling channel 26 can effectively reduce the temperature of the surrounding blade body and improve the temperature gradient distribution.

[0050] Therefore, by setting the filling fiber preform in the recessed area, the above method can construct a cooling channel closer to the blade surface, which can improve the cooling performance of the blade while ensuring the blade strength. In addition, the recessed area is set based on the high heat load area of ​​the blade, so that the local high temperature area can be cooled in a targeted manner, so that the blade has a more reasonable temperature gradient distribution.

[0051] The turbine blade manufacturing method described in this disclosure will be described in detail below through three embodiments.

[0052] Figures 1 to 5A first embodiment of a turbine blade assembly is shown. The turbine blade 2 has a cooling gas cavity 21, and a multilayer fiber cloth 41 is wound around an inner cavity mandrel 40 to form an inner cavity fiber preform 42. The inner cavity fiber preform 42 has an inwardly recessed area 46, which in this embodiment is located at the blade leading edge 22, and a filler fiber preform 44 is placed within the recessed area 46. The filler fiber preform 44 has an internal embedded part 45. After splicing the inner cavity fiber preform 42 and the filler fiber preform 44, a multilayer fiber cloth 41 is wound and laid on the outside to obtain an integral fiber preform 43 with the designed blade shape. The blade fiber preform 43 is processed accordingly, and the internal embedded part 45 is removed to obtain a ceramic matrix composite blade 2. The gap left after removing the original internal embedded part 45 forms a radial cooling channel 26 located at the blade leading edge 22. Optionally, the number of radial cooling channels 26 on the leading edge 22 of the blade can be one or more, and the cross-section of the radial cooling channel 26 includes, but is not limited to, a circle, a rectangle or a trapezoid.

[0053] A connecting hole is machined into the turbine blade 2 as a film cooling hole 6 (6'). The film cooling hole 6 (6') can be arranged on the leading edge 22, pressure surface 23, suction surface 24, and trailing edge 25 of the blade to provide film cooling to the outer wall surface 28 of the blade. The film cooling hole 6 penetrates the outer wall surface 28 and the inner wall surface 29 of the blade, connecting the cooling gas cavity 2 and the outside of the blade; the film cooling hole 6' penetrates the outer wall surface 28 of the blade and the cooling channel 26, connecting the cooling channel 26 and the outside of the blade. A sleeve 3 is assembled inside the cooling gas cavity 21 of the ceramic matrix composite blade 2 to provide cooling gas to the blade. The sleeve 3 has a gas supply cavity 31. The sleeve 3 has an impact hole 7, which penetrates the outer wall surface 32 and the inner wall surface 33 of the sleeve, connecting the sleeve gas supply cavity 31 and the blade cooling gas cavity 21, to provide impact cooling to the inner wall surface 29 of the blade.

[0054] In the first embodiment of the turbine blade assembly, cool air enters the air supply cavity 31 from the cool air inlet 8, flows into the cool air cavity 21 of the blade through the impact hole 7, and then a portion of the cool air flows out from the film cooling hole 6 that penetrates the inner wall surface 29 of the blade, while the other portion of the cool air flows to the blade tip and then flows into the radial cooling microchannel 26 from the radial cooling channel air supply port 6, and finally flows out from the film cooling hole 6 that penetrates the radial cooling channel 26.

[0055] Alternatively, some of the cold air enters the air supply cavity 31 of the sleeve from the cold air inlet 8, flows into the cold air cavity 21 of the blade through the impact hole 7, and then flows out from the film air hole 5 that penetrates the inner wall surface 29 of the blade; another part of the cold air flows directly into the radial cooling channel 26 from the radial cooling channel air supply port 5 at the tip of the blade, and finally flows out from the film air hole 6 that penetrates the radial cooling channel 26.

[0056] Figures 6 to 7A second embodiment of the turbine blade assembly is shown. The turbine blade 2 has a cooling gas cavity 21. The inner cavity fiber preform 42 of the second embodiment has two inwardly recessed areas 46, located on the pressure side and suction side, respectively. Thus, the prepared blade has a cooling channel 26 on the pressure side 23 and suction side 24, respectively. A sleeve 3 is assembled inside the cooling gas cavity 21 of the ceramic matrix composite blade 2 to provide cooling gas to the blade. The sleeve 3 has an air supply cavity 31. The sleeve 3 has an impact hole 7 that penetrates the outer wall surface 32 and the inner wall surface 33 of the sleeve, connecting the air supply cavity 31 of the sleeve and the cooling gas cavity 21 of the blade, providing impact cooling to the inner wall surface 29 of the blade.

[0057] Figures 8 to 9 A third embodiment of the turbine blade assembly is shown. The turbine blade 2 has two cooling gas cavities 21: a front cooling gas cavity and a rear cooling gas cavity, where front and rear refer to the front and rear along the chord length direction. The front cooling gas cavity is close to the leading edge of the blade, and the rear cooling gas cavity is close to the trailing edge of the blade. A partition 27 exists between the front and rear cooling gas cavities. The rear cooling gas cavity is surrounded by radial cooling channels 26 of the pressure surface 23 and the suction surface 24. During fabrication, two cavity mandrels are prepared to form a front cavity fiber preform 42 and a rear cavity fiber preform 42', respectively. Two recessed areas 46 are formed on the rear cavity fiber preform 42', located on the pressure side and suction side of the front cavity fiber preform 42, respectively. Filler fiber preforms are placed in the two recessed areas, and the two filler fiber preforms, the front cavity fiber preform, and the rear cavity fiber preform are spliced ​​together and wrapped with fiber cloth to form the blade fiber preform 43. Obtained after post-processing Figure 9 The blade structure shown.

[0058] The front cooling air chamber has a sleeve 3. Cooling air enters the supply chamber 31 from the cooling air inlet 8, flows into the cooling air chamber 21 of the blade through the impact hole 7, and then flows out from the film cooling hole 6 that penetrates the inner wall surface 29 of the blade. The rear cooling air chamber does not have a sleeve 3, and the outlet of the film cooling hole 6 of the rear cooling air chamber 21 is located at the trailing edge 25 of the blade, used to cool the trailing edge 25 of the blade. Cooling air enters the rear cooling air chamber 21 directly through the cooling air inlet 8, and then flows out directly from the corresponding film cooling hole 6 at the trailing edge 25 of the blade. The radial cooling microchannels 26 on both sides of the rear cooling air chamber 21 are used to cool the corresponding outer wall surface 28 of the blade and provide cooling protection for the rear cooling air chamber 21, reducing the temperature rise of the cooling air in the rear cooling air chamber 21.

[0059] In the third embodiment, since the radial cooling channel 26 is designed to be located in the middle and rear part of the blade, the cooling effect in the middle and rear part of the blade is enhanced. In conjunction with the rear cooling air cavity, the quality of the tail cooling air is guaranteed, and the tail cooling effect is further enhanced.

[0060] Furthermore, the blade wall thickness between the wall of the radial cooling channel 26 and the outer wall of the blade 28 is relatively thin. For example, the distance between the center O of the cooling channel 26 and the outer wall of the turbine blade 28 is 1 / 2 to 1 / 3 of the turbine blade wall thickness L0, so that the cool air inside the blade can get closer to the outer wall of the blade 28, thereby effectively reducing the temperature of the blade.

[0061] The turbine blades obtained by the above method, under certain cooling gas volume constraints, achieve radial cooling microchannels in the high-heat-load region of the blade by utilizing pre-embedded bodies during the curing process of the ceramic matrix composite blade. These pre-embedded bodies are then removed in subsequent manufacturing processes, forming cooling channel cavities. This reduces the wall thickness between the channels and the inner and outer walls of the blade, and enhances heat transfer on the blade wall through convective heat transfer and internal impact cooling, thereby effectively reducing the blade temperature and temperature gradient. The design of splicing local grooves in the fiber cloth layup with the filling fiber preform avoids damage to the fiber cloth layup structure, ensuring blade strength.

[0062] Based on the above description of the method, it can also be understood that a turbine blade assembly 1 includes a turbine blade 2 prepared by the above method and a metal sleeve 3. The metal sleeve 3 is disposed in the cold air cavity 21 of the turbine blade 2, and the turbine blade 2 is provided with a film gas hole 6 communicating with the cold air cavity and / or cooling channel 26.

[0063] The distance L1 from the center O of the cooling channel 26 to the outer wall surface 28 of the turbine blade is 1 / 2 to 1 / 3 of the turbine blade wall thickness L0. Figure 2 , Figure 6 , Figure 11 As shown.

[0064] For a symmetrical cooling channel 26, the center O is located at the midpoint of the symmetry line of the hole. For an asymmetrical, irregularly shaped cooling channel 26, the center O is located at an equidistant distance from the outer inner wall surface and the inner inner wall surface of the cooling channel 26. The outer inner wall surface 261 refers to the inner wall surface near the outer side of the blade, and the inner inner wall surface 262 refers to the inner wall surface near the inner side of the blade. Figure 11 As shown, the vertical distance L1 from the center O to the outer wall surface 28 of the turbine blade is 1 / 2 to 1 / 3 of the turbine blade wall thickness L0. In locations where the blade does not have cooling channels 26, the turbine blade wall thickness L0 is generally uniform, formed by winding multiple layers of fiber cloth around a mandrel. This design allows the cool air inside the blade to get closer to the outer wall surface 28, thereby effectively reducing the blade temperature.

[0065] While ensuring blade strength, the cooling channel is located in a localized high-temperature zone of the blade. This localized high-temperature zone includes one or a portion of the leading edge region, trailing edge region, pressure side region, and suction side region. Those skilled in the art will understand that the localized high-temperature zone includes, but is not limited to, the above embodiments. In other blade structures, the localized high-temperature zone refers to a region with a significant temperature difference from the surrounding area, which can be obtained, for example, through experimental data or numerical simulation.

[0066] For turbine blades that include multiple cooling gas chambers, cooling channels are located in the connecting area between adjacent cooling gas chambers, such as... Figure 9 As shown, this enhances the cooling effect in the middle and rear of the blades, and in conjunction with the rear air cooling cavity, ensures the quality of the tail air cooling, further strengthening the tail cooling effect.

[0067] The aforementioned turbine blade assembly structure has superior cooling performance.

[0068] This application uses specific terms to describe embodiments of the application. Terms such as "an embodiment," "one embodiment," and / or "some embodiments" refer to a particular feature, structure, or characteristic associated with at least one embodiment of the application. Therefore, it should be emphasized and noted that references to "an embodiment," "one embodiment," or "an alternative embodiment" in different locations throughout this specification do not necessarily refer to the same embodiment. Furthermore, certain features, structures, or characteristics in one or more embodiments of the application can be appropriately combined.

[0069] While the present invention has been disclosed above with reference to preferred embodiments, it is not intended to limit the invention. Any variations and modifications can be made by those skilled in the art without departing from the spirit and scope of the invention. Therefore, any modifications, equivalent changes, and alterations made to the above embodiments based on the technical essence of the present invention, without departing from the scope of the invention, fall within the protection scope defined by the claims of the present invention.

Claims

1. A method for preparing ceramic matrix composite turbine blades with cooling channels, characterized in that, Includes the following steps: The inner cavity core mold is wrapped with fiber cloth to form an inner cavity fiber preform; This creates a recessed area on the outer side of the inner fiber preform. The fiber-filled preform is placed in the recessed area, and the fiber-filled preform includes an internal embedded body. The inner cavity fiber preform and the filling fiber preform are wrapped with fiber cloth to form a blade fiber preform with a leaf shape. The blade fiber preform is post-processed to remove the inner cavity core mold and the internal pre-embedded body, forming the turbine blade's cold air cavity and cooling channel, respectively.

2. The method as described in claim 1, characterized in that, A plurality of inner cavity fiber preforms are obtained, and the recessed region is prepared on at least one of the inner cavity fiber preforms; Assemble the multiple inner cavity fiber preforms; The blade fiber preform is formed by simultaneously wrapping the filling fiber preform and the multiple inner cavity fiber preforms with fiber cloth.

3. The method as described in claim 1, characterized in that, Identify the local high-temperature zone of the blade, and prepare a recessed area in the region where the inner cavity fiber preform is located within the local high-temperature zone.

4. The method as described in claim 3, characterized in that, The localized high-temperature zone includes the leading edge region of the blade, the trailing edge region of the blade, the pressure side region, and the suction side region.

5. The method as described in claim 1, characterized in that, The recessed area includes an irregular groove surface.

6. The method as described in claim 1, characterized in that, A connecting hole is machined on the turbine blade, the connecting hole connecting the inner cavity of the cooling air and the outside of the blade, or connecting the cooling channel and the outside of the blade.

7. The method as described in claim 1, characterized in that, The inner cavity core mold is provided with a recessed portion, and the recessed portion is wrapped with fiber cloth to form the recessed area.

8. The method as described in claim 1, characterized in that, The recessed area is filled with whole or shredded fiber cloth, so that the internal pre-embedded body is located inside the fiber cloth, forming the filled fiber preform.

9. The method as described in claim 1 or 8, characterized in that, The internal pre-embedded body extends from the blade edge plate to the blade tip along the radial direction of the blade to form a radial cooling channel.

10. The method as described in claim 1, characterized in that, The internal pre-embedded body includes a turbulence cavity perpendicular to the blade wall and a matrix slurry, with the matrix slurry filling the turbulence cavity.

11. A turbine blade assembly, characterized in that, The device includes turbine blades and a metal sleeve. The turbine blades include a cold air cavity and a cooling channel. The metal sleeve is disposed in the cold air cavity. The distance between the center of the cooling channel and the outer wall of the turbine blade is 1 / 2 to 1 / 3 of the turbine blade wall thickness.

12. The turbine blade assembly as claimed in claim 11, characterized in that, The turbine blades are prepared by the method as described in any one of claims 1-10.

13. The turbine blade assembly as claimed in claim 11, characterized in that, The cooling channel is located in the localized high-temperature zone of the blade.

14. The turbine blade assembly as claimed in claim 13, characterized in that, The local high-temperature zone includes one or a portion of the blade leading edge region, blade trailing edge region, pressure side region, and suction side region.

15. The turbine blade assembly as claimed in claim 11, characterized in that, The turbine blades include multiple cooling gas cavities, and the cooling channels are located in the connecting area between adjacent cooling gas cavities.

16. The turbine blade assembly as claimed in claim 11, characterized in that, The turbine blades also include film vents that communicate with the cold air cavity and / or the cooling channel.

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