CN117266938A

CN117266938A - Turbine guide vane structure

Info

Publication number: CN117266938A
Application number: CN202210672937.8A
Authority: CN
Inventors: 曹源; 张诗尧; 鲍骐力; 洪辉; 张屹尚
Original assignee: AECC Commercial Aircraft Engine Co Ltd
Current assignee: AECC Commercial Aircraft Engine Co Ltd
Priority date: 2022-06-14
Filing date: 2022-06-14
Publication date: 2023-12-22
Also published as: WO2023241450A1

Abstract

The utility model provides a turbine stator structure, this structure includes CMC part and metal cover plate, and CMC part includes blade body and rim plate, and the rim plate is used for cooperating with metal cover plate, and the axial side of at least part rim plate includes the rim plate turn-ups, and the axial side of at least part and this rim plate complex metal cover plate includes the cover plate turn-ups, and the rim plate turn-ups is used for with cover plate turn-ups face-to-face contact, and this structure still includes pin and elastic component. The pin penetrates through at least part of the flange plate flanging and the cover plate flanging and is used for realizing the constraint between the part of the flange plate flanging and the cover plate flanging; wherein at least part of the pins further comprise an extension end for hanging against the casing; the elastic piece is sleeved outside at least part of the pin and is extruded between the flange plate flanging and the cover plate flanging which are connected with the pin and used for providing axial pretightening force. The turbine guide vane structure can alleviate the problem of thermal mismatch, is simple in structure overall, and reduces the complexity of component structures.

Description

Turbine guide vane structure

Technical Field

The invention relates to the field of aeroengines, in particular to the field of CMC turbine guide vanes.

Background

Modern commercial turbofan aeroengines are continuously developed towards the directions of large bypass ratio, high thrust, low oil consumption and high safety and reliability, and hot end parts, such as guide vanes (Vane), of the modern commercial turbofan aeroengines are generally cast and molded by high-temperature alloy precision at present. With the continuous improvement of performance design indexes of advanced commercial aeroengines, the pressure ratio is increased, the temperature before the turbine is also higher and higher, and the problems of insufficient temperature resistance and the like of the traditional high-temperature alloy materials are often exposed, so that a complex air film cooling system is required to be adopted to reduce the service temperature of the traditional high-temperature alloy materials. The introduction of a large amount of cooling gas can lead to insufficient combustion, cause the emission of harmful gas, pollute the environment, reduce the thermal efficiency of the engine, and hardly meet the design requirement of an advanced aeroengine.

At present, ceramic matrix composite materials (Ceramics matrix composites, CMC) are mostly adopted on turbine stator components such as turbine guide vanes, but are limited by the preparation process of the ceramic matrix composite materials and the characteristics of the ceramic matrix composite materials, and the following problems exist in the aspects of CMC component structures and assembly with upstream and downstream components: (1) CMC component structures cannot be overly complex; (2) The CMC component has relatively low molding precision, is difficult to machine after molding, and should reduce the machining and machining amount of various types as much as possible, so as to avoid the requirement on the CMC assembly on excessively high precision; (3) Due to the large thermal expansion coefficient and metal gap, the thermal mismatch between CMC component and metal structure is severe at high temperatures.

Accordingly, there is a need for a turbine vane assembly structure that addresses the above-described issues.

Disclosure of Invention

It is an object of the present invention to provide a turbine vane structure that enables a simple assembly between CMC components and metal components and that alleviates the problem of thermal mismatch in the axial direction.

The turbine vane structure for achieving the above object comprises a CMC component and a metal cover plate, wherein the CMC component comprises a vane body and a flange plate, and the flange plate is used for being matched with the metal cover plate. At least part of the axial side face of the flange plate comprises a flange plate flanging, at least part of the axial side face of the metal cover plate matched with the flange plate comprises a cover plate flanging, the flange plate flanging is used for being in surface contact with the cover plate flanging, and the structure further comprises a pin and an elastic piece. The pin penetrates through at least part of the flange plate flanging and the cover plate flanging and is used for realizing constraint between the part of flange plate flanging and the cover plate flanging; wherein at least part of the pins further comprise an extension end for hanging against the casing; the elastic piece is sleeved outside at least part of the pins and is extruded between the flange plate flanging and the cover plate flanging which are connected with the part of the pins, and the elastic piece is used for providing axial pretightening force.

In one or more embodiments, the rim plate comprises an upper rim plate and a lower rim plate, the metal cover plate comprises an upper cover plate and a lower support plate, the upper cover plate is matched with the upper rim plate, the lower support plate is matched with the lower rim plate, and a pin with an extending end penetrates through a cover plate flanging of the upper cover plate and a rim plate flanging of the upper rim plate.

In one or more embodiments, the turbine vane structure includes a leading edge and a trailing edge, the resilient member is sleeved on a pin located between the shroud flange and the platform flange of the leading edge.

In one or more embodiments, the rim flange of the upper rim extends radially outwardly and the cover flange of the upper cover extends radially inwardly; the flange of the lower flange plate extends radially inwards, and the cover plate flange of the lower supporting plate extends radially outwards.

In one or more embodiments, the cover flange is located inboard of the rim flange.

In one or more embodiments, the cover flange and the rim flange include pin holes thereon, and the pins are in interference fit with the pin holes on the cover flange and in clearance fit with the pin holes on the rim flange.

In one or more embodiments, the pin holes are racetrack or oval shaped holes, the long axes of which coincide with the circumferential direction of the turbine vane structure.

In one or more embodiments, the pin comprises a horn pin or a stepped pin.

In one or more embodiments, the turbine vane structure further includes a connecting screw penetrating the metal cover plate and the CMC component, and a fastening locknut disposed at one end of the connecting screw for fastening the connecting screw, thereby compressing the metal cover plate and the CMC component.

In one or more embodiments, the turbine vane structure further includes a spring member radially compressed between the CMC component and the metal cover plate.

The turbine guide vane structure realizes positioning constraint by means of the surface matching of the cover plate flanging and the flange flanging, and further strengthens the connection constraint between the metal part and the CMC part by means of the connection of air flow pressure or pins; in addition, the problem of thermal mismatch between the metal component and the CMC component in the axial direction is relieved by arranging the elastic component. The structure is simple as a whole, the complexity of the CMC component structure is reduced, the excessive requirement on CMC molding precision is avoided, and the machining amount is reduced.

Drawings

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

FIG. 1 is a schematic illustration of a typical aircraft engine configuration.

FIG. 2 is a schematic view of an embodiment of a turbine vane structure.

FIG. 3 is a cross-sectional view of an embodiment of a turbine vane structure.

FIG. 4 is a cross-sectional view of one embodiment of the upper cover plate mated with the upper rim plate.

Fig. 5A is a schematic view of the position of the leading edge locating pin.

Fig. 5B is a schematic view of the location of the trailing edge locating pin.

FIG. 6 is a schematic exterior view of one embodiment of a pin hole.

FIG. 7 is a schematic view of another embodiment of a cover flange and rim flange mating.

Sign mark description

1. Nacelle (GY)

2. Fan with fan body

3. Supercharging stage

5. Air compressor

6. Combustion chamber

7. Turbine wheel

8. Caudal vertebra

700. CMC component

710. Edge plate flanging

750. Edge plate

751. Upper edge plate

752. Lower edge plate

753. Third flange plate flanging

754. First flange plate flanging

755. Second flange plate flanging

756. Blade body

801. Primary outer ring casing

811. Two-stage outer ring casing

900. Metal cover plate

901. Upper cover plate

902. Lower support plate

903. Second cover plate flanging

904. First cover plate flanging

905. Third cover plate flanging

910. Cover plate flanging

930. Pin pin

931. First pin

932. Second pin

933. Connecting screw

934. Fastening locknut

940. Pin hole

921. Elastic piece

922. Spring element

9331. Extension end

Detailed Description

The present invention will be further described with reference to specific embodiments and drawings, in which more details are set forth in the following description in order to provide a thorough understanding of the present invention, but it will be apparent that the present invention can be embodied in many other forms than described herein, and that those skilled in the art may make similar generalizations and deductions depending on the actual application without departing from the spirit of the present invention, and therefore should not be construed to limit the scope of the present invention in terms of the content of this specific embodiment.

It is noted that these and other figures are merely examples, which are not drawn to scale and should not be construed as limiting the scope of the invention as it is actually claimed.

A typical construction of a gas turbine engine is shown in FIG. 1, and includes a nacelle 1, a fan 2, a booster stage 3, a compressor 5, a combustor 6, a turbine 7, and a tailcone 8. After entering from the fan 2, the gas is pressurized by the pressurizing stage 3 and enters the combustion chamber 6, and the gas is combusted by the combustion chamber 6 to apply work to the turbine 7 to drive the fan 2 to generate thrust. In fig. 1, the X direction indicates the engine axial direction, and the Y direction indicates the radial direction.

The turbine vane structure is located at the front end of the turbine and needs to withstand higher temperatures. The ceramic matrix composite (Ceramics matrix composites, CMC) is adopted at present, so that the working temperature of parts can be increased by about 400-500 ℃, the weight can be reduced by 1/3-2/3, the part structure is simplified, the cooling gas consumption is greatly reduced, the engine efficiency is improved, and meanwhile, the oil consumption and the emission of NOx and COx can be greatly reduced. However, the CMC turbine guide vane also needs to be connected with metal components such as a casing. Due to the large difference in thermal expansion coefficients of CMC materials and metallic materials, the thermal mismatch between CMC components and metallic structures is severe at high temperatures. In addition, the CMC component structure cannot be too complex, and mechanical design, machining amount and mechanical cooperation of various types should be reduced as much as possible, so that the CMC assembly is prevented from being subjected to excessive precision requirements.

The turbine guide vane structure can alleviate the problem of thermal mismatch between CMC components and metal structures, and can effectively reduce the complexity and the forming accuracy of CMC component structures and reduce the machining amount under the premise of guaranteeing the connection effect.

As will be understood from fig. 2 to 4, the X direction in the drawings represents the axial direction of the engine, and the air flow flows from left to right, i.e., from-X to +x; the Y direction represents the radial direction of the turbine, and the Z direction perpendicular to the X direction and the Y direction represents the circumferential direction of the turbine, that is, the circumferential direction. In fig. 2, the direction of the incoming airflow, i.e., the leading edge direction, is indicated in the-X direction, and the direction of the outgoing airflow, i.e., the trailing edge direction, is indicated in the +x direction.

The turbine vane structure includes a CMC component 700 and a metallic cover plate 900, the CMC component 700 including a blade airfoil 756 and a platform 750, the platform 750 being configured to mate with the metallic cover plate 900. Specifically, rim plate 750 includes upper rim plate 751 and lower rim plate 752, metal cover plate 900 includes upper cover plate 901 and lower tray 902, upper cover plate 901 mates with upper rim plate 751, lower tray 902 mates with lower rim plate 752, and blade 756 connects upper rim plate 751 and lower rim plate 752. The CMC material upper and lower rims 751, 752, 756 may be integrally manufactured, with the fibers being continuous; the fibers can also be assembled together through a mortise and tenon structure, and the fibers are discontinuous.

The turbine vane structure further comprises a connecting screw 933 and a fastening locknut 934, wherein the connecting screw 933 penetrates through the metal cover plate 900 and the CMC component 700, and the fastening locknut 934 is arranged at one end of the connecting screw 933 and used for fastening the connecting screw 933 so as to tightly press the metal cover plate 900 and the CMC component 700. The turbine vane structure also includes a spring member 922 radially compressed between the CMC component 700 and the metal cover plate 900 for solving the thermal mismatch problem between the metal and CMC component in the radial direction. The connecting screw 933 is integrally connected with the metal upper cover plate 901 by welding or the like, and the fastening locknut 934 connects the connecting screw 933 and the lower plate 902 together. The bolt pretightening force is initially applied, and the metal upper cover plate 901 and the lower supporting plate 902 are pressed between the upper edge plate 751 and the lower edge plate 752 through the precompression of the spring member 922, so that the elastic deformation of the spring member 922 can provide a movable allowance, and the problem of thermal mismatch in the radial direction is effectively relieved.

For the fit between CMC rim 750 and metal cover plate 900, at least a portion of the axial sides of rim 750 include rim flange 710 and at least a portion of the axial sides of the metal cover plate that mates with the rim include cover plate flange 910, rim flange 710 for surface-to-surface contact with cover plate flange 910. As shown in fig. 3, the upper and lower flanges 751, 752, which are approximately parallel to the plane ZOX, are folded over in the Y-direction on the axial sides to form flange flanges 710.

Specifically, in the embodiment shown in fig. 2 and 3, the flange turn 710 includes a first flange turn 754, a second flange turn 755, and a third flange turn 753. Corresponding to the flange 710, the upper cover 901 and the lower pallet 902, which are approximately parallel to the ZOX plane, are folded in the Y direction on the axial side to form the cover flange 910. Specifically, in the embodiment shown in fig. 2 and 3, the cover flange 910 includes a first cover flange 904 corresponding to the first flange 754, a second cover flange 903 corresponding to the second flange 755, and a third cover flange 905 corresponding to the third flange 753. Each rim flange 710 is in surface-to-surface contact with each cover flange 910.

It will be appreciated by those skilled in the art that the number of turns is not limited to the embodiments described above, and in other embodiments, the metallic cover plate 900 or CMC cap 750 may have 1-4 number of turns, with the particular number being determined by the particular operating environment of the turbine vane.

Continuing with FIG. 2, the turbine vane structure further includes a pin 930 and an elastic member 921. Pins 930 extend through at least a portion of the rim flange 710 and the cover flange 910 for effecting restraint between the portion of the rim flange and the cover flange. Wherein at least part of the pin 930 further comprises an extension 9331 for hooking against the casing, thus achieving both axial and radial constraint. For example, the first outer ring casing 801 and the second outer ring casing 811 positioned on two sides of the turbine guide vane structure in fig. 4, the extending ends 9331 of the pins 930 penetrate through the first outer ring casing 801 and the second outer ring casing 811, so as to fix the radial position of the turbine guide vane structure on the outer ring casing. The pins 930 themselves axially pass through the flange rims 710 and the cover rims 910 to axially connect the metal and CMC components.

As in one embodiment shown in fig. 2, pins 930 are present at upper cover plate 901 and upper rim plate 751, and contact of lower rim plate 752 with the flange of lower tray 902 does not use pins 930. This is because, during engine operation, aerodynamic forces act directly on CMC vane blade 756 from left to right in the direction of airflow G flow, thereby pushing second flange turn-ups 755 located on lower flange 752 against second cover flange turn-ups 903 located on lower plate 902, the surface contact ensuring the constraint between lower flange 752 and lower plate 902, and being able to transfer the forces of CMC components to metal components.

With continued reference to the embodiment shown in fig. 2. A pin 930 having an extended end 9331 extends through the deck flange 910 of the upper deck 901 and the flange 710 of the upper flange 751. Since the radially outer side of the turbine vane structure, i.e. the upper flange plate 751 and the upper cover plate 901 portions need to be fixed, while the opposite radially inner side, a certain play between the lower flange plate 752 and the lower support plate 902 may be allowed, pins 930 with extended ends 9331 may provide fixation points at the upper cover plate 901 and the upper flange plate 751. It will be appreciated that in other embodiments, such as turbine vane structures requiring only one side to be secured to the casing, such as only the leading edge side to be secured to the casing, the second flange 755 and the second cover flange 903 may employ pins 930 without extensions 9331, where the pins 930 serve only a connection constraint and not an external connection.

As will be further appreciated in connection with fig. 2, the turbine vane structure further comprises an elastic member 921, the elastic member 921 being arranged around at least a portion of the pin 930 and being compressed between the flange and the cover flange to which the portion of the pin is connected, for providing an axial preload, such as the elastic member 921 being located between the third cover flange 905 and the third flange 753. The elastic member 921 has an initial pre-compression and remains in a certain pre-compressed state during the high temperature phase. The elastic member 921 is suitably precompressed to generate a pressing force during installation, and tension of the elastic member 921 is directly applied to the third flange 753 of the leading edge, so that the first flange 754 on the trailing edge of the upper flange 751 is pressed against the first cover flange 904 of the upper cover 901, and the two flanges are in surface contact. The second outer ring casing 811 further axially constrains and positions the first flange 754 and the first cover flange 904 via the extending ends 9331 of the pins 930. The contact surface between the three is used as an assembly reference surface. The interface of the primary outer ring casing 801 to the third cover flange 905 at the leading edge may serve as an auxiliary axial restraint.

In the high temperature stage, because the metal cover plate 900 and upstream and downstream metal components, such as the primary outer ring casing 801 and the secondary outer ring casing 811, have thermal expansion in the engine axial direction much greater than the CMC component 700, the thermal mismatch problem in the axial direction is alleviated by the length variation of the axial elastic member 921. Along with the spring member 922 in the radial direction, the problem of thermal mismatch of the turbine guide vane structure in the radial direction and the axial direction is relieved, and the structure is simple, so that the complexity of the whole structure is reduced, and the machining amount is effectively reduced.

The resilient member 921 and/or the spring member 922 includes, but is not limited to, a coil spring using a high temperature nickel-based alloy, such as an X750 spring, a ceramic spring, a high temperature nickel-based combination Bellville washers, or a high temperature nickel-based machined spring, among others. The metallic structural material can be made of various high-temperature alloys, including but not limited to GH4169/GH3536/K417/DD6 and other metallic materials.

Since the turbine vane structure includes a leading edge and a trailing edge, in one embodiment, the resilient member 921 is sleeved over the pin 930 between the cover flange 910 and the cover flange 710 of the leading edge, as shown in FIGS. 2 and 5A. Because the elastic pre-tightening force of the elastic piece 921 can be directly transmitted to the first flange turn-up 754 of the trailing edge through the upper flange 751, the adhesion degree of the first flange turn-up 754 and the first cover plate turn-up 904 can be enhanced, and the fixation of the upper cover plate 901 and the upper flange 751 at the trailing edge can be ensured.

It will be appreciated that the resilient members 921 may also be added between the first rim flange 754 and the first cover flange 904, and the second rim flange 755 and the second cover flange 903 to further alleviate thermal mismatch problems in the axial direction at multiple locations.

In one or more embodiments, the rim flange 710 of the upper rim 751 extends radially outward and the cover flange 910 of the upper cover 901 extends radially inward, here inward in the-Y direction, and outward in the +y direction. The rim flange 710 of the lower rim plate 752 extends radially inward and the cover flange 910 of the lower tray 902 extends radially outward. The flange 750 is embedded into a U-shaped groove formed by flanging of the metal cover plate 900 to realize matching.

In one or more embodiments, as shown in FIG. 7, the cover flange 910 is located inboard of the rim flange 710, i.e., axially inboard. The flange 710 of CMC material is disposed outside the metal material and can protect the metal cover flange 910 from external high temperatures. In the high temperature stage, the expansion rate of the metal cover flange 910 is larger than that of the CMC component, so that the flange 710 made of external CMC material can be extruded, so that the assembly of the metal cover flange and the flange is more reliable, the precompression amount of the elastic piece 921 can be reduced, and the stress of the CMC flange 710 is further reduced.

Based on the above embodiments, the cover flange 910 and the flange 710 include pin holes 940 thereon, and the pins 930 are in interference fit with the pin holes 940 on the cover flange 910 and in clearance fit with the pin holes 940 on the flange 710.

Specifically, as will be appreciated with reference to fig. 5A to 6, the lower edge of the third cover flange 905 is hooked on the first outer ring casing 801 by the first pin 931; the first pin 931 connects the third cap flange 905 at the leading edge of the upper cap and the third cap flange 753 at the leading edge of the upper cap of the CMC vane. Wherein the first pin 931 is in interference fit with the pin hole 940 on the third cap flange 905 and in clearance fit with the pin hole 940 of the third cap flange 753. The interference fit can avoid the pin from falling off. The second pin 932 connects the first rim flange 754, the first cover flange 904, and the secondary outer ring casing 811 together. Wherein the second pin 932 is an interference fit with the pin hole in the first cover flange 904 and a clearance fit with the pin hole in the first flange 754.

In one embodiment, the pin holes 940 are racetrack or oval shaped holes, with the long axes of the pin holes 940 aligned with the circumferential direction of the turbine vane structure to provide suitable clearance for the pins, mitigating thermal mismatch issues.

Further, in one or more embodiments, the pins 930 include a horn pin or a stepped pin. Referring to FIG. 5A, the portion of the first pin 931 located within the third platform turn-up 753 is a horn-like structure to enhance axial constraint of the vane. Referring to fig. 5B, a portion of the second pin 932 located in the second outer ring casing 811 and the first cover flange 904 has a larger diameter, and a portion located in the first flange 754 has a smaller diameter, so that a stepped pin structure is formed, and when the first flange 754 moves axially to the right, the second pin 932 can be blocked by a step of the pin, thereby reinforcing axial constraint.

The turbine guide vane structure realizes positioning constraint by means of the surface matching of the cover plate flanging and the flange flanging, and further strengthens the connection constraint between the metal part and the CMC part by means of the connection of air flow pressure or pins; in addition, the problem of thermal mismatch between the metal component and the CMC component in the radial direction and the axial direction is relieved by arranging the elastic component. The structure is simple as a whole, the complexity of the CMC component structure is reduced, the excessive requirement on CMC molding precision is avoided, and the machining amount is reduced.

It should be noted that, in the foregoing description, the terms "first", "second", etc. are used to define the components, and are merely for convenience in distinguishing the corresponding components, and the terms have no special meaning unless otherwise stated, so they should not be construed as limiting the scope of protection of the present application.

Meanwhile, the present application uses specific words to describe embodiments of the present application. Reference to "one embodiment," "an embodiment," and/or "some embodiments" means that a particular feature, structure, or characteristic is associated with at least one embodiment of the present application. Thus, it should be emphasized and should be appreciated that two or more references to "an embodiment" or "one embodiment" or "an alternative embodiment" in various positions in this specification are not necessarily referring to the same embodiment. Furthermore, certain features, structures, or characteristics of one or more embodiments of the present application may be combined as suitable.

While the invention has been described in terms of preferred embodiments, it is not intended to be limiting, but rather to the invention, as will occur to those skilled in the art, without departing from the spirit and scope of the invention. Therefore, any modification, equivalent variation and modification of the above embodiments according to the technical substance of the present invention fall within the protection scope defined by the claims of the present invention.

Claims

1. A turbine vane structure comprising a CMC component (700) and a metallic cover plate (900), the CMC component (700) comprising a blade body (756) and a platform (750), the platform (750) being adapted to mate with the metallic cover plate (900), characterized in that,

at least part of the axial side surface of the flange plate (750) comprises a flange plate flanging (710), at least part of the axial side surface of the metal cover plate matched with the flange plate comprises a cover plate flanging (910), the flange plate flanging (710) is used for being in surface-to-surface contact with the cover plate flanging (910),

the structure further comprises:

a pin (930) extending through at least a portion of the rim flange (710) and the cover flange (910) for effecting a constraint between the portion of the rim flange and the cover flange; wherein at least part of the pin (930) further comprises an extension (9331) for hanging against the casing;

and the elastic piece (921) is sleeved outside at least part of the pin (930) and is extruded between the flange plate flanging and the cover plate flanging which are connected with the pin, and is used for providing axial pretightening force.

2. The turbine vane structure of claim 1, said platform (750) comprising an upper platform (751) and a lower platform (752), said metal cover plate (900) comprising an upper cover plate (901) and a lower platform (902), said upper cover plate (901) mated with said upper platform (751), said lower platform (902) mated with said lower platform (752),

a pin (930) having the extension end (9331) penetrates a cover flange (910) located on the upper cover (901) and a flange (710) located on the upper flange (751).

3. Turbine vane structure according to claim 1 or 2, characterized in that it comprises a leading edge and a trailing edge, said elastic element (921) being fitted over a pin (930) between said cover flange (910) and said flange (710) on the leading edge side.

4. The turbine vane structure of claim 2, characterized in that a rim flange (710) of the upper rim plate (751) extends radially outwardly and a cover flange (910) of the upper cover plate (901) extends radially inwardly; the rim flange (710) of the lower rim (752) extends radially inward and the cover flange (910) of the lower tray (902) extends radially outward.

5. The turbine vane structure of claim 1 or 4, characterized in that the shroud flange (910) is located inside the shroud flange (710).

6. The turbine vane structure of claim 5, characterized in that the cover flange (910) and the rim flange (710) include pin holes (940) thereon, the pins (930) being in interference fit with the pin holes (940) on the cover flange (910) and in clearance fit with the pin holes (940) on the rim flange (710).

7. The turbine vane structure of claim 6, characterized in that the pin holes (940) are racetrack or oval holes, the long axes of the pin holes (940) being coincident with the circumferential direction of the turbine vane structure.

8. The turbine vane structure of claim 1, characterized in that the pin (930) is a horn pin or a stepped pin.

9. The turbine vane structure of claim 1, further comprising a connecting screw (933) and a fastening locknut (934), the connecting screw (933) penetrating through the metal cover plate (900) and the CMC component (700), the fastening locknut (934) being disposed at one end of the connecting screw (933) for fastening the connecting screw (933) and thereby compressing the metal cover plate (900) and the CMC component (700).

10. The turbine vane structure of claim 9, further comprising a spring member (922) radially compressed between the CMC component (700) and the metallic cover plate (900).