Turbine blade

Abstract

A turbine blade is used for optimizing the heat transfer performance of a ceramic matrix composite turbine blade. A leading edge wall body of the turbine blade comprises a gradient ceramic matrix composite material, the gradient ceramic matrix composite material is formed by stacking a plurality of laying layers in the first direction, and the first direction is perpendicular to the leading edge wall body and points to the outer wall face of the leading edge wall body from the inner wall face of the leading edge wall body. The second direction is the extending direction of the front edge wall body between the suction side wall of the turbine blade and the pressure side wall of the turbine blade, the second direction is perpendicular to the first direction, and at least two of the multiple laying layers have different in-plane heat conductivity of the second direction, so that in-plane heat conductivity of the multiple laying layers in the first direction is different from that of the multiple laying layers in the pressure side wall of the turbine blade. The in-plane thermal conductivity of the leading edge wall in the second direction increases.

Description

Technical Field

[0001] The invention relates to the field of heat transfer, and in particular to a turbine blade. Background Art

[0002] In order to achieve higher cycle thermal efficiency, the compressor pressure ratio and pre-turbine temperature of modern aircraft engines are constantly increasing, while the proportion of cooling air used to cool hot end components is constantly decreasing. These changes have led to increasingly harsh working environments for high-temperature components of aircraft engines, posing extremely severe challenges to heat transfer and cooling design. In order to improve the performance of materials, researchers have conducted in-depth research on high-temperature alloy manufacturing processes, such as using powder metallurgy to manufacture high-temperature alloy turbine discs and single crystal casting to manufacture high-temperature alloy turbine blades. However, the application of these special processes has reached the limit of the exploration of the performance of traditional high-temperature alloys. Therefore, in order to adapt to the ever-increasing operating temperature and meet the development needs of the next generation of aircraft engines, the use of ceramic-based composite materials with good high-temperature resistance has become a development trend in the manufacture of high-temperature components of aircraft engines in recent years.

[0003] Although the application of ceramic matrix composites has greatly increased the safe operating temperature of turbine blades, the operating temperature of the hot end components of the future new generation of aircraft engines will still far exceed the temperature limit that ceramic matrix composites can withstand. On the one hand, due to the limitations of the preparation process, and on the other hand, the need to further reduce the amount of cooling gas, ceramic matrix composite turbine blades cannot adopt the same complex and efficient cooling method as metal blades. Therefore, it is very necessary to optimize the heat transfer performance of ceramic matrix composite turbine blades with limited cooling capacity. Summary of the invention

[0004] The object of the present invention is to provide a turbine blade for optimizing the heat transfer performance of a ceramic matrix composite material turbine blade.

[0005] According to an embodiment of the present invention, the leading edge wall of the turbine blade includes a gradient ceramic matrix composite material, which is made of a plurality of plies stacked along a first direction, the first direction is perpendicular to the leading edge wall and points from the inner wall surface of the leading edge wall to the outer wall surface of the leading edge wall, the second direction is the extension direction of the leading edge wall between the suction side wall of the turbine blade and the pressure side wall of the turbine blade, the second direction is perpendicular to the first direction, and at least two of the plurality of plies have different in-plane thermal conductivities in the second direction, so that along the first direction, the in-plane thermal conductivity of the leading edge wall in the second direction increases.

[0006] In one or more embodiments, each of the plies has a different in-plane thermal conductivity in the second direction, and along the first direction, the in-plane thermal conductivity of each of the plies in the second direction increases.

[0007] In one or more embodiments, each of the plies includes a fiber reinforcement and a matrix, the thermal conductivity of the fibers is greater than the thermal conductivity of the matrix, the multiple plies include a first ply and a second ply, the in-plane thermal conductivity of the first ply in the second direction is a first thermal conductivity, the in-plane thermal conductivity of the second ply in the second direction is a second thermal conductivity, and the volume fraction of the fibers of the first ply is less than the volume fraction of the fibers of the second ply, so that the first thermal conductivity is less than the second thermal conductivity.

[0008] In one or more embodiments, the fibers extend along the second direction.

[0009] In one or more embodiments, the fibers include silicon carbide fibers, and the matrix includes a silicon carbide ceramic matrix.

[0010] In one or more embodiments, each of the plies includes a fiber reinforcement, a metal and a matrix, the thermal conductivity of the metal is greater than the thermal conductivity of the matrix, the multiple plies include a first ply and a second ply, the in-plane thermal conductivity of the first ply in the second direction is a first thermal conductivity, the in-plane thermal conductivity of the second ply in the second direction is a second thermal conductivity, and the volume fraction of the metal in the first ply is less than the volume fraction of the metal in the second ply, so that the first thermal conductivity is less than the second thermal conductivity.

[0011] In one or more embodiments, the metal is dispersed in the matrix.

[0012] In one or more embodiments, the fiber comprises silicon carbide fiber, the matrix comprises a silicon carbide ceramic matrix, and the metal comprises copper powder.

[0013] In one or more embodiments, the fibers extend along the second direction.

[0014] The embodiments of the present invention have at least the following beneficial effects:

[0015] The leading edge wall is directly facing the incoming gas flow, and is most directly impacted by the gas, with a high heat transfer coefficient. The leading edge wall is a high temperature area, and the heat transfer coefficients of the suction side wall and the pressure side wall are relatively low, and the temperature is low. Along the first direction, the in-plane thermal conductivity of the leading edge wall in the second direction increases, and the heat transfer capacity of the leading edge wall along the second direction increases from the inner wall surface of the leading edge wall to the outer wall surface of the leading edge wall 2. The portion of the leading edge wall close to the outer wall surface transfers the heat of the leading edge wall with a higher temperature to the suction side wall and the pressure side wall with a relatively lower temperature, thereby reducing the temperature of the portion of the leading edge wall close to the outer wall surface, reducing the wall temperature of the outer wall surface, reducing the amount of cooling gas used, and also reducing the heat transferred from the outer wall surface of the leading edge wall to the inner wall surface in the reverse direction of the first direction, reducing the temperature of the portion of the leading edge wall close to the inner wall surface, reducing the wall temperature of the inner wall surface, reducing the temperature gradient of the leading edge wall in the first direction, and making the temperature distribution of the leading edge wall in the first direction more uniform, reducing thermal stress, improving high temperature resistance, and further increasing the temperature before the turbine of the aircraft engine, thereby improving the engine cycle efficiency. BRIEF DESCRIPTION OF THE DRAWINGS

[0016] The above and other features, properties and advantages of the present invention will become more apparent through the following description in conjunction with the accompanying drawings and embodiments, in which:

[0017] Figure 1 is an oblique view of a turbine blade;

[0018] Figure 2 Schematic diagram of gradient ceramic matrix composite;

[0019] Figure 3 Schematic diagram of gradient ceramic matrix composite;

[0020] Figure 4 for Figure 1 A partial enlarged view of the middle A;

[0021] Reference numerals:

[0022] 1- Turbine blades;

[0023] 2-front wall;

[0024] 3-First layer;

[0025] 4- Second layer;

[0026] 5- third layer;

[0027] 6- the fourth layer;

[0028] 7-inner wall surface of the leading edge wall;

[0029] 8- the outer wall surface of the leading edge wall;

[0030] 9- Environmental barrier coating;

[0031] 10-Suction side wall;

[0032] 11-pressure side wall;

[0033] 12-fiber;

[0034] 13-matrix;

[0035] 14-Metal. DETAILED DESCRIPTION

[0036] Reference will now be made in detail to embodiments of the present invention, one or more examples of which are illustrated in the accompanying drawings. Each example is provided to explain the present invention, not to limit the present invention. In fact, it will be apparent to those skilled in the art that various modifications and variations may be made in the present invention without departing from the scope or spirit of the present invention. For example, a feature illustrated or described as part of one embodiment may be used together with another embodiment to produce yet another embodiment. Therefore, the present invention is intended to cover these modifications and variations within the scope of the appended claims and their equivalents.

[0037] It should be noted that these and other subsequent drawings are only examples and are not drawn to scale, and should not be used to limit the actual scope of protection required by the present invention.

[0038] The terms "first", "second", etc. may be used interchangeably to distinguish one feature from another, and are not intended to mean that each feature must be located as shown in the figures in each embodiment.

[0039] like Figure 1 As shown, the turbine blade 1 may be a turbine moving blade. The turbine blade 1 may also be a turbine stationary blade.

[0040] like Figure 1 As shown, the leading edge wall 2 of the turbine blade 1 is made of a gradient ceramic matrix composite material. Ceramic matrix composite materials are composite materials with ceramic as the matrix and composite reinforcement. The ceramic matrix can be high-temperature structural ceramics such as silicon nitride and silicon carbide, so that the ceramic matrix composite material has good high temperature resistance and reduces the use of cooling gas. Gradient refers to the gradient change of the composition and structure of the ceramic matrix composite material, so that the performance of the ceramic matrix composite material has a gradient change, becoming a gradient functional material. The gradient ceramic matrix composite material can provide the inner wall surface 7 and the outer wall surface 8 of the leading edge wall 2. Or, further refer to Figure 4 The outer wall surface of the turbine blade 1 is coated with an environmental barrier (EBC) coating 9, and the outer wall surface 8 of the leading edge wall 2 is provided by the environmental barrier coating 9. The environmental barrier coating 9 has good heat insulation, oxidation resistance and corrosion resistance, and resists the erosion of the turbine blade 1 wall by the high temperature water vapor environment.

[0041] Further integration Figure 2 and Figure 3 The gradient ceramic matrix composite material is made by stacking a plurality of layers 3, 4, 5, 6 along a first direction. Figure 2 and Figure 3 In the figure, the leading edge wall 2 is simplified as a straight wall and is only partially shown. The first direction is perpendicular to the leading edge wall 2 and points from the inner wall surface 7 of the leading edge wall 2 to the outer wall surface 8 of the leading edge wall 2. Figure 2 and Figure 3 It is shown with an arrow and marked with "z". It should be understood that due to the shape of the leading edge wall 2, different positions of the leading edge wall 2 have different first directions, and the first direction is often regarded as the thickness direction of the leading edge wall 2. The second direction is the extending direction of the leading edge wall 2 between the suction side wall 10 of the turbine blade 1 and the pressure side wall 11 of the turbine blade. The positive direction of the second direction is the extending direction of the leading edge wall 2 from the suction side wall 10 of the turbine blade 1 to the pressure side wall 11 of the turbine blade. The positive direction of the second direction is Figure 2 and Figure 3 It is shown with arrows and marked with "x". It should be understood that due to the shape of the leading edge wall 2, different positions of the leading edge wall 2 may have different second directions. The first direction and the second direction at the same position of the leading edge wall 2 are perpendicular to each other. The second direction is often regarded as the width direction of the leading edge wall 2. The third direction is the extension direction of the leading edge wall 2 from the blade root to the blade crown. Figure 2 and Figure 3 As shown by an arrow and marked with "y", it should be understood that due to the shape of the leading edge wall 2, different positions of the leading edge wall 2 may have different third directions. The first direction, the second direction, and the third direction at the same position of the leading edge wall 2 are perpendicular to each other, and the third direction is often regarded as the height direction of the leading edge wall 2.

[0042] like Figures 1 to 3 As shown, at least two of the multiple plies 3, 4, 5, and 6 have different in-plane thermal conductivities in the second direction, so that the in-plane thermal conductivity of the leading edge wall 2 in the second direction increases along the first direction. The composition and structure of at least two plies change in a gradient, and the in-plane thermal conductivity of the ply close to the outer wall surface 8 in the second direction is greater than the in-plane thermal conductivity of the ply close to the inner wall surface 7 in the second direction, so that the in-plane thermal conductivity of the gradient ceramic matrix composite material in the second direction increases along the first direction, and has a gradient change, thereby increasing the in-plane thermal conductivity of the leading edge wall 2 in the second direction along the first direction. The in-plane thermal conductivity in the second direction represents the heat transfer capacity of the object inside the second direction. The greater the in-plane thermal conductivity in the second direction, the stronger the heat transfer capacity of the object inside the second direction. The in-plane thermal conductivity in the second direction of the ply represents the heat transfer capacity in the ply in the second direction. The in-plane thermal conductivity in the second direction of the leading edge wall 2 represents the heat transfer capacity in the second direction of the leading edge wall 2.

[0043] The leading edge wall 2 faces the incoming gas and is most directly impacted by the gas, so the heat transfer coefficient is relatively high. The leading edge wall 2 is a high temperature area, while the suction side wall 10 and the pressure side wall 11 have relatively low heat transfer coefficients and low temperatures. Along the first direction, the in-plane thermal conductivity of the leading edge wall body 2 in the second direction increases, and the heat transfer capacity of the leading edge wall body 2 along the second direction increases from the inner wall surface 7 of the leading edge wall body 2 to the outer wall surface 8 of the leading edge wall body 2. The portion of the leading edge wall body 2 close to the outer wall surface 8 transfers the heat of the leading edge wall body 2 with a higher temperature to the suction side wall 10 and the pressure side wall 11 with relatively lower temperatures, thereby reducing the temperature of the portion of the leading edge wall body 2 close to the outer wall surface 8, reducing the wall temperature of the outer wall surface 8, reducing the amount of cooling air used, and also reducing the amount of heat transferred from the outer wall surface 8 of the leading edge wall body 2 to the inner wall surface 7 in the opposite direction of the first direction, reducing the temperature of the portion of the leading edge wall body 2 close to the inner wall surface 7, reducing the wall temperature of the inner wall surface 7, and reducing the temperature gradient of the leading edge wall body 2 in the first direction. The temperature distribution of the leading edge wall body 2 in the first direction is more uniform, reducing thermal stress, improving high temperature resistance, and further increasing the temperature before the turbine of the aircraft engine, thereby improving the engine cycle efficiency.

[0044] like Figures 1 to 3 As shown, each ply may have a different in-plane thermal conductivity in the second direction, and the in-plane thermal conductivity of each ply in the second direction increases along the first direction. The gradient ceramic matrix composite material may be made by stacking four plies 3, 4, 5, and 6 along the first direction, wherein ply 3 is referred to as the first ply 3, ply 4 is referred to as the second ply 4, ply 5 is referred to as the third ply 5, and ply 6 is referred to as the fourth ply 6. The first ply 3, the second ply 4, the third ply 5, and the fourth ply 6 are stacked in sequence along the first direction, and the in-plane thermal conductivity of the first ply 3 in the second direction is less than the in-plane thermal conductivity of the second ply 4 in the second direction, the in-plane thermal conductivity of the second ply 4 in the second direction is less than the in-plane thermal conductivity of the third ply 5 in the second direction, and the in-plane thermal conductivity of the third ply 5 in the second direction is less than the in-plane thermal conductivity of the fourth ply 6 in the second direction, so that the in-plane thermal conductivity of the gradient ceramic matrix composite material in the second direction increases along the first direction, and has a gradient change. In another embodiment, some plies may have the same in-plane thermal conductivity in the second direction. For example, the in-plane thermal conductivity of the first ply 3 in the second direction is smaller than the in-plane thermal conductivity of the second ply 4 in the second direction, the in-plane thermal conductivity of the second ply 4 in the second direction is equal to the in-plane thermal conductivity of the third ply 5 in the second direction, and the in-plane thermal conductivity of the third ply 5 in the second direction is smaller than the in-plane thermal conductivity of the fourth ply 6 in the second direction, thereby increasing the in-plane thermal conductivity of the gradient ceramic matrix composite material in the second direction along the first direction, and having a gradient change.

[0045] like Figure 2As shown, each ply may include a reinforcement of fiber 12 and a matrix 13. The thermal conductivity of fiber 12 is greater than the thermal conductivity of matrix 13. Fiber 12 may be silicon carbide fiber, and matrix 13 may be a silicon carbide ceramic matrix. Fiber 12 may extend along a second direction to increase the thermal conductivity of fiber 12 along the second direction. For two plies with different in-plane thermal conductivities in the second direction, the volume fraction of fiber 12 in one ply is smaller than the volume fraction of fiber 12 in the other ply, and the volume fraction of matrix 13 in the ply is of course greater than the volume fraction of matrix 13 in the other ply. Since the thermal conductivity of fiber 12 is greater than the thermal conductivity of matrix 13, the in-plane thermal conductivity in the second direction of the ply is smaller than the in-plane thermal conductivity in the second direction of the other ply.

[0046] Continue to refer to Figure 2 As mentioned above, the in-plane thermal conductivity of the first ply 3 in the second direction is less than the in-plane thermal conductivity of the second ply 4, the volume fraction of the fiber 12 of the first ply 3 is less than the volume fraction of the fiber 12 of the second ply 4, and the volume fraction of the matrix 13 of the first ply 3 is naturally greater than the volume fraction of the matrix 13 of the second ply 4. The in-plane thermal conductivity of the second ply 4 in the second direction is less than the in-plane thermal conductivity of the third ply 5 in the second direction, the volume fraction of the fiber 12 of the second ply 4 is less than the volume fraction of the fiber 12 of the third ply 5, and the volume fraction of the matrix 13 of the second ply 4 is naturally greater than the volume fraction of the matrix 13 of the third ply 5. The in-plane thermal conductivity of the third ply 5 in the second direction is less than the in-plane thermal conductivity of the fourth ply 6 in the second direction, the volume fraction of the fiber 12 of the third ply 5 is less than the volume fraction of the fiber 12 of the fourth ply 6, and the volume fraction of the matrix 13 of the third ply 5 is naturally greater than the volume fraction of the matrix 13 of the fourth ply 6. The first layer 3 may be composed of 15% by volume of silicon carbide fiber and 85% by volume of silicon carbide ceramic matrix, the second layer 4 may be composed of 25% by volume of silicon carbide fiber and 75% by volume of silicon carbide ceramic matrix, the third layer 5 may be composed of 35% by volume of silicon carbide fiber and 65% by volume of silicon carbide ceramic matrix, and the fourth layer 6 may be composed of 45% by volume of silicon carbide fiber and 55% by volume of silicon carbide ceramic matrix. The gradient functional silicon carbide fiber / silicon carbide ceramic matrix composite material (SiCf / SiC gradient ceramic matrix composite material) can be prepared by melt infiltration (MI) process, continuous silicon carbide fiber is made into unidirectional prepreg tape, the unidirectional prepreg tapes of silicon carbide fibers in different layers have different densities, so that each layer has a different volume fraction of silicon carbide fibers, and then a plurality of unidirectional prepreg tapes are stacked to obtain a preform, and then after the steps of curing, cracking, siliconization, etc., the gradient functional silicon carbide fiber / silicon carbide ceramic matrix composite material is obtained.

[0047] like Figure 3As shown, each ply may include a reinforcement of fiber 12, a matrix 13 and a metal 14. The thermal conductivity of the metal 14 is greater than the thermal conductivity of the matrix 13. The metal 14 may be dispersed in the matrix 13. The fiber 12 may be a silicon carbide fiber, the matrix 13 may be a silicon carbide ceramic matrix, the metal 14 may be an atom or compound such as copper, aluminum, nickel, etc. that has good compatibility with the silicon carbide ceramic matrix and high thermal conductivity, and the metal 14 may be copper powder, which is copper atoms. The fiber 12 may extend along the second direction to improve the thermal conductivity of the fiber 12 along the second direction. For two plies with different in-plane thermal conductivities in the second direction, the volume fraction of the metal 14 of one ply is less than the volume fraction of the metal 14 of the other ply, and the volume fraction of the matrix 13 of the ply is of course greater than the volume fraction of the matrix 13 of the other ply. Since the thermal conductivity of the metal 14 is greater than the thermal conductivity of the matrix 13, the in-plane thermal conductivity of the second direction of the ply is less than the in-plane thermal conductivity of the second direction of the other ply.

[0048] Continue to refer to Figure 3As mentioned above, the in-plane thermal conductivity of the first ply 3 in the second direction is smaller than the in-plane thermal conductivity of the second ply 4, the volume fraction of the metal 14 of the first ply 3 is smaller than the volume fraction of the metal 14 of the second ply 4, and the volume fraction of the matrix 13 of the first ply 3 is naturally greater than the volume fraction of the matrix 13 of the second ply 4. The in-plane thermal conductivity of the second ply 4 in the second direction is smaller than the in-plane thermal conductivity of the third ply 5 in the second direction, the volume fraction of the metal 14 of the second ply 4 is smaller than the volume fraction of the metal 14 of the third ply 5, and the volume fraction of the matrix 13 of the second ply 4 is naturally greater than the volume fraction of the matrix 13 of the third ply 5. The in-plane thermal conductivity of the third ply 5 in the second direction is smaller than the in-plane thermal conductivity of the fourth ply 6 in the second direction, the volume fraction of the metal 14 of the third ply 5 is smaller than the volume fraction of the metal 14 of the fourth ply 6, and the volume fraction of the matrix 13 of the third ply 5 is naturally greater than the volume fraction of the matrix 13 of the fourth ply 6. The first layer 3 may include 1% copper powder by volume, 30% silicon carbide fiber by volume and 69% silicon carbide ceramic matrix by volume, the second layer 4 may include 4% copper powder by volume, 30% silicon carbide fiber by volume and 66% silicon carbide ceramic matrix by volume, the third layer 5 may include 7% copper powder by volume, 30% silicon carbide fiber by volume and 63% silicon carbide ceramic matrix by volume, and the fourth layer 6 may include 10% copper powder by volume, 30% silicon carbide fiber by volume and 60% silicon carbide ceramic matrix by volume. The gradient functional silicon carbide fiber / silicon carbide ceramic matrix composite material (SiCf / SiC gradient ceramic matrix composite material) can be prepared by melt infiltration (MI) process, continuous silicon carbide fiber is made into unidirectional prepreg tape, the unidirectional prepreg tapes of silicon carbide fibers with different layers are passed through a slurry pool containing copper powder with different concentrations, so that the unidirectional prepreg tapes of each silicon carbide fiber are impregnated with slurry containing copper powder with different volume fractions, and then multiple unidirectional prepreg tapes are stacked to obtain a preform, and then after the steps of curing, cracking, siliconization, etc., the gradient functional silicon carbide fiber / silicon carbide ceramic matrix composite material is obtained, and the copper powder is dispersed in the silicon carbide ceramic matrix.

[0049] Although the present invention is disclosed as above by the embodiments, it is not intended to limit the present invention. Any person skilled in the art may make possible changes and modifications without departing from the spirit and scope of the present invention.

Claims

1. A turbine blade, characterized in that: The leading edge wall of the turbine blade includes a gradient ceramic matrix composite material, which is made of a plurality of plies stacked along a first direction, wherein the first direction is perpendicular to the leading edge wall and points from the inner wall surface of the leading edge wall to the outer wall surface of the leading edge wall, and the second direction is the extension direction of the leading edge wall between the suction side wall of the turbine blade and the pressure side wall of the turbine blade, and the second direction is perpendicular to the first direction, and at least two of the plurality of plies have different in-plane thermal conductivities in the second direction, so that along the first direction, the in-plane thermal conductivity of the leading edge wall in the second direction increases.

2. The turbine blade according to claim 1, characterized in that: Each of the plies has a different in-plane thermal conductivity in the second direction, and the in-plane thermal conductivity of each of the plies in the second direction increases along the first direction.

3. The turbine blade according to claim 1, characterized in that: Each of the plies includes a fiber reinforcement and a matrix, the thermal conductivity of the fibers is greater than the thermal conductivity of the matrix, the multiple plies include a first ply and a second ply, the in-plane thermal conductivity of the first ply in the second direction is a first thermal conductivity, the in-plane thermal conductivity of the second ply in the second direction is a second thermal conductivity, and the volume fraction of the fibers of the first ply is less than the volume fraction of the fibers of the second ply, so that the first thermal conductivity is less than the second thermal conductivity.

4. The turbine blade according to claim 3, characterized in that: The fibers extend along the second direction.

5. The turbine blade according to claim 3, characterized in that: The fibers include silicon carbide fibers, and the matrix includes a silicon carbide ceramic matrix.

6. The turbine blade according to claim 1, characterized in that: Each of the plies includes a fiber reinforcement, a metal and a matrix, the thermal conductivity of the metal is greater than the thermal conductivity of the matrix, the multiple plies include a first ply and a second ply, the in-plane thermal conductivity of the first ply in the second direction is a first thermal conductivity, the in-plane thermal conductivity of the second ply in the second direction is a second thermal conductivity, and the volume fraction of the metal in the first ply is less than the volume fraction of the metal in the second ply, so that the first thermal conductivity is less than the second thermal conductivity.

7. The turbine blade according to claim 6, characterized in that: The metal is dispersed in the matrix.

8. The turbine blade according to claim 7, characterized in that: The fiber comprises silicon carbide fiber, the matrix comprises a silicon carbide ceramic matrix, and the metal comprises copper powder.

9. The turbine blade according to claim 6, characterized in that: The fibers extend along the second direction.

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