Turbine guide vane assembly structure

Abstract

According to the turbine guide vane assembly structure, on the premise that small stress is kept, a gap between a guide vane and a high-temperature state can be reduced. The two supports are connected at the hollow position of the turbine guide vane, the turbine guide vane is clamped between the two supports, so that axial and circumferential loads borne by the turbine guide vane are transmitted to upstream and downstream components of the turbine guide vane through the two supports, and a gap is formed in the connecting position of the two supports in the radial direction. The gap is enough to absorb the thermal mismatch amount generated by the two brackets due to thermal expansion relative to the turbine guide vane; one end of the pulling plate is connected with one of the two supports, the other end of the pulling plate is connected with the other one of the two supports, and tensioning force is exerted on the two supports so that the two supports can be pressed on the two edge plates of the turbine guide vane.

Description

Technical Field

[0001] The invention relates to a turbine guide vane assembly structure, in particular to a turbine guide vane assembly structure of an aircraft engine or a gas turbine, which comprises a turbine guide vane with a smaller thermal expansion coefficient and a bracket with a larger thermal expansion coefficient. Background Art

[0002] The development trend of civil aviation engines is that the temperature before the turbine is getting higher and higher, and the service life requirements are getting longer and longer. For example, the design life of the hot end components of the mainstream commercial aviation engines currently under development is not less than 10,000 flight cycles (20,000 hours), and the high-temperature takeoff turbine gas temperature has reached 1978K.

[0003] For high-temperature turbine stators, the use temperature and service performance of traditional high-temperature alloy materials have reached their limits under the conditions of using traditional cooling technology and thermal barrier coating technology, making it difficult to meet the design requirements of the next generation of advanced aircraft engines. Using ceramic-based composites to replace traditional high-temperature alloy materials is the best way to improve the temperature resistance and efficiency of aircraft engine hot end components.

[0004] Compared with traditional high-temperature alloys, ceramic matrix composites (CMC) have the following advantages:

[0005] (1) High temperature resistance, reducing cooling volume - High temperature resistant materials can reduce the cooling volume of turbine components, thereby improving the efficiency of turbine components;

[0006] (2) Corrosion resistance: Ceramics such as SiC and Si3N4 can form a silicon oxide protective layer on the surface under high temperature conditions, which can meet the high temperature anti-oxidation requirements below 1600°C;

[0007] (3) Low density - The density of CMC is only 1 / 3 to 1 / 4 of that of high-temperature alloys, which can reduce the weight of the engine and further improve the efficiency of the engine;

[0008] (4) Increase combustion temperature and thermal efficiency, thereby reducing pollutant emissions by 75%.

[0009] Based on these advantages, the civil aviation engine's need to reduce weight and the subsequent need to increase the turbine front inlet temperature makes the use of CMC materials in turbine stators such as turbine guide vanes an inevitable choice for future aviation engines.

[0010] In the traditional turbine guide vane assembly structure, the aerodynamic load is transmitted to the front and rear casings or other metal parts through the upper and lower edge plates. At this time, the guide vane stress is concentrated on the edge plate and the blade body chamfer. However, due to process conditions, there is a possibility of reduced strength near the chamfer of the CMC component relative to the matrix. A known countermeasure is that the ceramic matrix composite turbine guide vane assembly structure includes a turbine guide vane, an upper bracket, and a lower bracket. The turbine guide vane is mainly made of CMC material. The upper bracket and the lower bracket are mainly made of metal material. The upper bracket and the lower bracket are connected at the hollow part of the turbine guide vane, and the turbine guide vane is clamped between the upper bracket and the lower bracket. The load on the turbine guide vane is transmitted to the upstream and downstream components through the upper bracket and the lower bracket. However, the thermal expansion coefficient of ceramic matrix composites is small, generally (4~5)*10 -6 / K, while nickel-based metal materials are (10~20)*10 -6 / K. The two are prone to thermal mismatch under high-temperature working conditions, resulting in a large gap between CMC components and metal components at high temperatures. Summary of the invention

[0011] An object of the present invention is to provide a turbine guide vane assembly structure, which can reduce the gap between the guide vanes and the turbine under high temperature conditions while maintaining low stress.

[0012] According to one embodiment of the present invention, a turbine guide vane assembly structure includes a turbine guide vane and two brackets, the thermal expansion coefficient of the material of the two brackets is greater than the thermal expansion coefficient of the material of the turbine guide vane, the two brackets are respectively matched with two edge plates of the turbine guide vane, the two brackets are connected at the hollow part of the turbine guide vane, and the turbine guide vane is clamped between the two brackets, so that the axial and circumferential loads on the turbine guide vane are transmitted to the upstream and downstream components of the turbine guide vane through the two brackets, wherein

[0013] There is a gap in the radial direction at the connection between the two brackets, and the gap is sufficient to absorb the thermal mismatch between the two brackets and the turbine guide vane caused by thermal expansion;

[0014] The turbine guide vane assembly structure also includes a pull plate, one end of which is connected to one of the two brackets, and the other end of which is connected to the other of the two brackets. A tensioning force is applied to the two brackets to press the two brackets against the two edge plates of the turbine guide vane, and the thermal expansion coefficient of the material of the pull plate is sufficiently close to the thermal expansion coefficient of the turbine guide vane, so that the two brackets can be kept pressed against the two edge plates of the turbine guide vane even when heated.

[0015] In one embodiment, a connection structure between one end of the pull plate and one of the two brackets provides a pre-tightening force applied by the pull plate.

[0016] In one embodiment, the connection structure includes a connecting member with two ends connected to the bracket and the middle connected to the pull plate, and elastic gaskets for providing the preload force are arranged between the two ends of the connecting member and the bracket.

[0017] In one embodiment, the connecting member is a metal spring.

[0018] In one embodiment, at least one of the two brackets can be elastically deformed under the force of the pulling plate when it expands due to heat.

[0019] In one embodiment, the turbine guide vane and the pull plate are made of CMC material.

[0020] In one embodiment, the fiber preform of the pull plate is formed into a long runway shape by winding CMC continuous fibers, and the bends are used for inserting fasteners.

[0021] In one embodiment, at the connection point of the two brackets, one of them provides a pin and the other provides a pin hole. The pin and the pin hole are slidably matched. The depth of the pin hole is greater than the matching length of the pin to provide the gap. Only radial relative movement is allowed between the pin and the pin hole, and only axial and circumferential force transmission is allowed.

[0022] In one embodiment, one of the two brackets includes a plate and a sleeve-like structure;

[0023] The plate provides a mating surface for mating with one of the two edge plates of the turbine guide vane, and a connection structure for connecting with the pull plate;

[0024] The sleeve-like structure integrally extends from the plate toward another of the brackets;

[0025] The sleeve-shaped structure comprises a force transmission platform which contacts and cooperates with the inner wall surface of the airfoil portion of the turbine guide vane.

[0026] In one embodiment, the other of the two brackets comprises another plate;

[0027] The other plate provides a mating surface that cooperates with the other of the two edge plates of the turbine guide vane, and also provides a force transmission platform that contacts and cooperates with the inner wall surface of the airfoil portion of the turbine guide vane, and a connection structure connected to the pull plate.

[0028] According to an embodiment of the present invention, there is a gap in the radial direction at the connection of the two brackets, and the gap is sufficient to absorb the thermal mismatch of the two brackets relative to the turbine guide vane caused by thermal expansion. At the same time, a pull plate is introduced to radially limit the two brackets to reduce the gap between the bracket and the turbine guide vane in the working state. The axial and circumferential loads on the turbine guide vane are transmitted to the upstream and downstream components of the turbine guide vane through the two brackets, and the radial force on the pull plate is small, so the gap between the guide vane and the turbine guide vane under high temperature can be reduced while maintaining small stress. BRIEF DESCRIPTION OF THE DRAWINGS

[0029] 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:

[0030] Figure 1 It is a schematic diagram of the turbine guide vane assembly structure;

[0031] Figure 2 It is a schematic exploded view of the turbine guide vane assembly structure;

[0032] Figure 3 is a schematic diagram of a turbine guide vane;

[0033] Figure 4 is a schematic diagram of the upper bracket;

[0034] Figure 5 is a schematic diagram of the lower bracket;

[0035] Figure 6 It is a schematic diagram of the force transmission platform abutting against the inner wall of the turbine guide vane;

[0036] Figure 7 It is a schematic diagram of the pull plate;

[0037] Figure 8 is a schematic diagram of another embodiment of a pull plate;

[0038] Fig. 9 It is an axial cross-sectional view of the turbine guide vane assembly structure;

[0039] Fig.10 yes Fig. 9 A partial enlarged view of the upper middle part;

[0040] Fig.11 It is a radial cross-sectional view of the turbine guide vane assembly structure;

[0041] Fig.12 yes Fig.11 A partial enlarged view of the lower middle part;

[0042] Fig.13 yes Fig. 9 A partial enlargement of the lower middle part. DETAILED DESCRIPTION

[0043] 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.

[0044] As used in this article, the term "turbine vane" is a component in the high-temperature turbine of a gas turbine engine (aircraft engine or gas turbine) that rectifies the high-temperature and high-pressure gas from upstream and outputs it to the downstream. It is installed on the casing and together with the casing forms a stator.

[0045] Ceramic matrix composites (Ceramics matrix composites, CMC) is a kind of continuous fiber reinforced composite material with ceramic as the matrix. Its main materials currently include carbon / silicon carbide, silicon carbide / silicon carbide, zirconium boride / silicon carbide and hafnium boride / silicon carbide, etc. It has the characteristics of high operating temperature, good oxidation resistance and microcrack resistance, light weight, high strength and stiffness.

[0046] Directional terms such as "axial", "circumferential" and "radial" are based on the turbine guide vanes or the nozzle formed by the turbine guide vanes when they are installed on the turbine. Figure 1 Shown in.

[0047] In the embodiments described below, the turbine guide vanes and the pull plates are both made of CMC materials, and the two brackets are made of metal materials, such as nickel-based metal materials. The relevant descriptions of these embodiments are also applicable to other embodiments in which the thermal expansion coefficient of the bracket material is much larger than the thermal expansion coefficient of the turbine guide vane material.

[0048] Figure 1 and Figure 2 The assembly structure of the turbine guide vane and its exploded view. The assembly structure mainly includes the CMC guide vane 1 and the support frame 2. The support frame 2 mainly includes the metal upper bracket 21, the metal lower bracket 22 and the CMC material pull plate 23. The CMC pull plate 23 only plays the role of transmitting radial force and radial positioning. At least one of the metal upper and lower brackets has a structure connected to other upstream and downstream components.

[0049] Figure 3The CMC guide vane 1 shown in the figure includes an upper edge plate 12, a lower edge plate 13 and a blade body 11, and the blade body is divided into a pressure side and a suction side. The upper and lower edge plates are provided with assembly surfaces 121, 131 for assembly with a metal upper bracket 21 and a metal lower bracket 22.

[0050] Figure 4 and Figure 5 The metal upper bracket 21 and the metal lower bracket 22 in the support frame 2 are provided, wherein the metal upper bracket 21 includes a plate and a sleeve-like structure 211 attached to the plate. The sleeve-like structure 211 can withstand certain stress and can be used as an impact cooling sleeve after being punched. The sleeve-like structure 211 can be rigidly connected to one side of the upper and lower brackets by bolt connection, riveting, pin connection, integral connection, etc. In the embodiment shown in the figure, the sleeve 211 is integrally connected to the plate of the upper bracket 21.

[0051] The metal lower bracket 22 includes a plate and a columnar structure 223 with two pin holes 227 extending from the plate, on which the force transmission platform 221 is disposed. In addition, the metal lower bracket 22 also has a connection structure for connecting with the CMC pull plate 23, an example of which includes an ear piece 224, an ear piece hole 226 and a through hole 225.

[0052] There are at least one force transmission platform 212, 221 on each of the upper and lower brackets 21, 22, and at least three force transmission platforms 212, 221 in total. The metal upper bracket 21 is matched with the CMC upper edge plate 12 through the matching surfaces 121, 213, and the metal lower bracket 22 is matched with the CMC lower edge plate 13 through the matching surfaces 131, 222. A connection structure for connecting the pull plate 23 is reserved on the metal upper bracket 21, and the assembly structure includes a threaded hole 215. Two pin holes 214 are reserved at the bottom of the sleeve-like structure 211.

[0053] Figure 6 The figure shows that in the assembly state, the force transmission platforms 212 and 221 are inserted into the inner cavity of the CMC guide vane 1 and abut against the pressure side of the inner cavity wall to transmit the aerodynamic force exerted on the CMC guide vane 1 and perform circumferential positioning. The upper force transmission platform 212 is inserted into the inner cavity from the upper side for positioning, and the lower force transmission platform 221 is inserted into the inner cavity from the lower side for positioning, which facilitates assembly.

[0054] Figure 7 One embodiment of the CMC pull plate 23 is a single plate with holes 231, 232 processed at both ends for inserting pins or bolts 25a, 25c to connect the pull plate 23 to the metal component. The pull plate 23 can be a flat plate with continuous fibers or a filler. Its shape can also have a certain arc surface to ensure that it can be inserted into the guide vane 1. Fig. 9As shown, the upper end of the pull plate 23 is connected to the upper bracket 21, and the lower end is connected to the lower bracket 22, and a tensioning force is applied to the two brackets 21, 22 to press the two brackets 21, 22 on the two edge plates 12, 13 of the turbine guide vane 1. Since the thermal expansion coefficient of the material of the pull plate 23 is close enough to the thermal expansion coefficient of the turbine guide vane 1, the two brackets 21, 22 can be kept pressed on the two edge plates of the turbine guide vane even when heated. In one embodiment, the thermal expansion coefficient of the pull plate 23 is the same as the thermal expansion coefficient of the turbine guide vane 1. In another embodiment, the thermal expansion coefficients of the two are slightly different. Due to the different internal and external temperatures of the turbine guide vane, the thermal expansion coefficient of the pull plate is slightly larger than that of the guide vane. The thermal expansion coefficient of the pull plate generally does not exceed 150% of the thermal expansion coefficient of the guide vane. The thermal expansion coefficient of the pull plate is close enough to the thermal expansion coefficient of the turbine guide vane, and therefore is much smaller than the thermal expansion coefficient of the bracket. This is particularly obvious when the pull plate and the turbine guide vane are formed by CMC and the bracket is formed by metal.

[0055] Figure 8 It is another embodiment of the CMC pull plate 23. The CMC continuous fiber 232 is wound into a long racetrack shape. The CMC continuous fiber is, for example, SiC. The bends are used to insert pins or bolts 25a, 25c or other fasteners. The direction of the fiber 232 in this form can more effectively bear the radial force, and the cavity in the middle can also be used for gas circulation and weight reduction.

[0056] according to Fig. 9 In one embodiment, the CMC pull plate 23 is connected to the upper and lower metal brackets 21 and 22 respectively, wherein the connection structure between the CMC pull plate 23 and the metal upper bracket 21 includes a connector, and in one embodiment of the connector, a metal spring 24. The connection structure between the CMC pull plate 23 and the lower bracket 22 includes a bolt 25c, a nut 26b, and an ear hole 226. The CMC pull plate 23 passes through the through hole 225 and is connected to the ear 224 through the bolt 25c and the nut 26b passing through the ear hole 226. In another embodiment, the connection method of the CMC pull plate 23 and the metal upper and lower brackets 21 and 22 is not limited to the above method, and the connection can also be achieved by pins, slots, etc.

[0057] Fig.10 yes Fig. 9A partial enlarged view of the upper end. The CMC pull plate 23 is fixed to the metal spring 24 by bolts 25a and nuts 26b. In another embodiment, the fixing method is replaced by pin connection. The size of the two light holes 231 and 232 on the CMC pull plate 23 is slightly larger than the outer diameter of the bolt 25a to avoid the generation of thermal stress. The connection structure between one end of the pull plate 23 and the bracket 21 provides a pre-tightening force applied to the pull plate 23. For example, the two ends of the metal spring 24 are fixed to the threaded holes 215 on the upper metal bracket 21 by bolts 25b. Elastic gaskets 27 are arranged on both sides of the connection position between the metal spring 24 and the metal upper bracket 21 and the bolts 25b. One embodiment of the elastic gasket is a disc spring, which is used to adjust the radial position of the metal spring 24 and provide a certain pre-tightening force, thereby adjusting the radial assembly position of the CMC pull plate 23, so that the upper edge of the upper hole 231 of the CMC pull plate 23 and the lower edge of the lower hole 232 in the assembled state are in contact with the upper and lower bolts 25a, 25c, and a certain initial tension is generated on the CMC pull plate 23. Through the matching surfaces 121, 131, 213, 222 of the CMC pull plate 23 and the upper and lower cover plates 21, 22 with the CMC upper edge plate 12 and the lower edge plate 13, the CMC blade 1 is clamped. The width direction of the pull plate 23 can be along the chord direction of the blade to provide sufficient assembly space for the pull plate 23, or it can extend from the pressure surface to the suction surface. In one embodiment, the width direction of the pull plate 23 extends from the pressure surface side to the suction surface side, or the fasteners or bolts or pins used to install the pull plate 23 extend roughly in the chord direction of the turbine guide vane. If the lower bracket is deformed, the pull plate 23 can swing with the position connected to the upper bracket, such as the fasteners or bolts or pins, as a fulcrum, to buffer the in-plane force on the pull plate 23, which can reduce the axial and circumferential forces on the pull plate 23, and further strengthen the function of the pull plate 23 to only transmit radial forces. In one embodiment, the CMC pull plate 23 constitutes a part of the spoiler structure or the guide vane cooling air path, such as forming a part of the sleeve wall or the sleeve cavity partition, and is sealed with other metal parts by setting a sealing sheet.

[0058] Fig.11 and Fig.12It is a cross-sectional view of the assembly at another angle and a partial enlarged view of the lower part. The pin hole 227 on the lower metal bracket 22 is interference fit with the two pins 28, and the pin 28 is clearance fit with the pin hole 214 on the upper bracket 21, and a gap is left in the radial direction. The size of the gap can offset the mismatch caused by thermal expansion, that is, it is sufficient to absorb the thermal mismatch of the two brackets relative to the turbine guide vane due to thermal expansion. The thermal mismatch refers to the amount of deformation relative to the normal assembly position. In one embodiment, only radial relative movement is allowed between the pin 28 and the pin hole 214, and only axial and circumferential force transmission is allowed. In the working state, since the thermal expansion coefficient of the metal is greater than that of the CMC, the length of the metal sleeve 211 increases and extends downward, and the metal lower bracket 22 is radially positioned by the CMC pull plate 23 through the assembly structure, and the radial position change is relatively small. At this time, the sleeve 211 and the lower bracket 22 undergo radial relative movement, and no gap is generated between the CMC blade 1 and the metal upper and lower brackets 21, 22.

[0059] Fig.13 for Fig.12 A cross-sectional view from another angle. During operation, the components of aerodynamic force generated in the circumferential and axial directions are mainly transmitted through the force transmission platforms 212, 221 on the metal brackets 21, 22 on both sides, wherein the load on the lower bracket 22 is finally transmitted to the upstream and downstream components through the pin 28, the sleeve 211 and the upper metal cover plate 21, without passing through the CMC pull plate 23. The radial force with a smaller load component is transmitted to the upper metal cover plate 21 through the lower cover plate 22 and the CMC pull plate 23. The CMC pull plate 23 only transmits radial force. Under working conditions, due to the high external temperature, the actual thermal expansion of the blade 1 part will be greater than that of the CMC pull plate 23. The difference in thermal expansion is shared by the deformation of the lower cover plate 22 itself and / or the strain of the pull plate 23 and / or the elastic gasket 27, so that the force on each component is not too large, and the preload force between the blade 1 and the support frame 2 is increased in the working state, and the connection is tighter.

[0060] Although the present invention is disclosed as above with preferred 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. Therefore, any modification, equivalent change and modification made to the above embodiments according to the technical essence of the present invention without departing from the content of the technical solution of the present invention shall fall within the protection scope defined by the claims of the present invention.

Claims

1. A turbine guide vane assembly structure, comprising a turbine guide vane and two brackets, wherein the thermal expansion coefficient of the material of the two brackets is greater than the thermal expansion coefficient of the material of the turbine guide vane, the two brackets are respectively matched with two edge plates of the turbine guide vane, the two brackets are connected at the hollow part of the turbine guide vane, and the turbine guide vane is clamped between the two brackets, so that the axial and circumferential loads on the turbine guide vane are transmitted to the upstream and downstream components of the turbine guide vane through the two brackets, characterized in that: There is a gap in the radial direction at the connection between the two brackets, and the gap is sufficient to absorb the thermal mismatch between the two brackets and the turbine guide vane caused by thermal expansion; The turbine guide vane assembly structure also includes a pull plate, one end of which is connected to one of the two brackets, and the other end of which is connected to the other of the two brackets. A tensioning force is applied to the two brackets to press the two brackets against the two edge plates of the turbine guide vane, and the thermal expansion coefficient of the material of the pull plate is sufficiently close to the thermal expansion coefficient of the turbine guide vane, so that the two brackets can be kept pressed against the two edge plates of the turbine guide vane even when heated.

2. The turbine guide vane assembly structure according to claim 1, characterized in that: The connection structure between one end of the pull plate and one of the two brackets provides a pre-tightening force applied by the pull plate.

3. The turbine guide vane assembly structure according to claim 2, characterized in that: The connection structure comprises a connection piece with two ends connected to the bracket and the middle connected to the pull plate, and an elastic gasket for providing the pre-tightening force is arranged between the two ends of the connection piece and the bracket.

4. The turbine guide vane assembly structure according to claim 3, characterized in that: The connecting piece is a metal spring.

5. The turbine guide vane assembly structure according to any one of claims 1 to 4, characterized in that: At least one of the two brackets can be elastically deformed under the force of the pulling plate when it expands due to heat.

6. The turbine guide vane assembly structure according to claim 1, characterized in that: The turbine guide vane and the pull plate are made of CMC material.

7. The turbine guide vane assembly structure according to claim 6, characterized in that: The fiber preform of the pull plate is formed into a long runway shape by winding CMC continuous fibers, and the bending parts are used for inserting fasteners.

8. The turbine guide vane assembly structure according to claim 1, characterized in that: At the connection point of the two brackets, one of them provides a pin and the other provides a pin hole. The pin and the pin hole are slidably matched. The depth of the pin hole is greater than the matching length of the pin to provide the gap. Only radial relative movement is allowed between the pin and the pin hole, and only axial and circumferential force transmission is allowed.

9. The turbine guide vane assembly structure according to claim 1, characterized in that: One of the two brackets includes a plate and a sleeve-like structure; The plate provides a mating surface for mating with one of the two edge plates of the turbine guide vane, and a connection structure for connecting with the pull plate; The sleeve-like structure integrally extends from the plate toward another of the brackets; The sleeve-shaped structure comprises a force transmission platform which contacts and cooperates with the inner wall surface of the airfoil portion of the turbine guide vane.

10. The turbine guide vane assembly structure according to claim 9, characterized in that: The other of the two brackets includes another plate; The other plate provides a mating surface that cooperates with the other of the two edge plates of the turbine guide vane, and also provides a force transmission platform that contacts and cooperates with the inner wall surface of the airfoil portion of the turbine guide vane, and a connection structure connected to the pull plate.

11. The turbine guide vane assembly structure according to claim 9 or 10, characterized in that: The pull plate forms a turbulent structure or a ventilation cooling structure inside the sleeve, becoming a part of the cooling air path of the turbine guide vane.

12. The turbine guide vane assembly structure according to claim 1, characterized in that: The thermal expansion coefficient of the material of the pull plate is greater than the thermal expansion coefficient of the turbine guide vane, and is at most 150% of the thermal expansion coefficient of the turbine guide vane.

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