Turbine rotor structure for preventing blades from moving backwards and aero-engine

Abstract

The invention discloses a turbine rotor structure capable of preventing blades from moving backwards and an aero-engine, the turbine rotor structure comprises turbine blades, a turbine disc, a labyrinth sealing ring, a clamping ring and a vibration reduction piece, the turbine blades and the turbine disc are connected through tenon teeth and mortises, and the tenon teeth and the mortises are each of a two-tooth structure; the vibration reduction pieces are installed between the turbine blades and the turbine disc. The labyrinth sealing ring is assembled on the turbine disc and axially limited through the clamping ring. The turbine blade is in a bent and twisted shape; the problem that the blade moves backwards can be solved, and therefore the safety and reliability of the aero-engine are guaranteed.

Description

Technical Field

[0001] The present invention belongs to the technical field of aeroengines, and in particular, relates to a turbine rotor structure for preventing blades from moving backward. The present invention also relates to an aeroengine. Background Art

[0002] The axial position of aircraft engine turbine blades is fixed by locking plates, which are usually manually hammered to bend the two ends of the locking plates and riveted to both sides of the turbine disk. Due to the small size of the locking plates, there is a certain dispersion in the hammering force, hammering location, and angle during the manual hammering and riveting process. Therefore, there is a certain dispersion in the riveting quality of the locking plates. The structural reliability of the locking plates to axially fix the blades is low, and there is a risk that the locking plates cannot constrain the axial position of the blades, causing the blades to move backward. If the turbine blades move backward, the rear end face of the turbine blade edge plate may axially rub against the inner ring of the turbine guide, causing the gas turbine rotor to get stuck, affecting the normal use of the engine in the field.

[0003] The existing turbine rotor consists of turbine blades, turbine discs, comb-tooth sealing rings, and locking plates. The blades and discs are connected by a three-tooth fir-tree-shaped tenon, and the comb-tooth sealing ring is fixed to the turbine disc by interference fit. The turbine blades are matched with the locking plate bosses through the locking plate grooves at the bottom of the tenons, and the two ends of the locking plates are bent and riveted to the turbine disc, as shown in the figure. The axial position of the blades is fixed by the locking plates. The locking plates are riveted downward on both sides of the turbine disc. During operation, the bent parts on both sides of the locking plates are subject to the centrifugal force and tend to loosen upwards; and the locking plates are made of wire with a thickness of 0.8, and the locking strength is low. Under the action of loads such as blade vibration and aerodynamic impact, they are prone to yielding, being straightened and warped, resulting in failure of the locking function and the risk of blades moving backward.

[0004] The existing patent publication number CN110439628A discloses a connection structure between gas turbine blades and turbine discs, including a turbine disc, a plurality of blades matched with the turbine disc, and a clamping ring and a buckle for axially limiting the blades. It can minimize the probability of damage to the turbine disc due to stress concentration under high load and high thermal stress conditions; it has high reliability, a reduced number of parts, convenient and quick assembly and disassembly, and can improve assembly efficiency. It does not require the use of special fixtures during installation, and parts can be reused. This patent assembles the clamping ring through the front end grooves of the blades and the turbine disc, and then screws on the locking plate to limit the movement of the blades. However, drilling holes in the turbine disc will affect the strength and service life of the disc; using pins to fix the buckles, there is a risk of the pins falling off, resulting in locking failure. Summary of the invention

[0005] The present invention is mainly aimed at the problem that the structural reliability of the locking plate for axially fixing the blades in the prior art is low, and there is a hidden danger that the locking plate cannot constrain the axial position of the blades, resulting in the blades moving backwards. If the turbine blades move backwards, the rear end surface of the turbine blade edge plate may axially rub against the inner ring of the turbine guide, causing the gas turbine rotor to get stuck, affecting the normal use of the engine in the field. The present invention provides a turbine rotor structure and an aircraft engine that can prevent the blades from moving backwards.

[0006] In view of the above technical problems, the technical solutions of the present invention are as follows:

[0007] A turbine rotor structure and an aircraft engine for preventing blades from moving backwards, comprising turbine blades, a turbine disk, a comb-tooth sealing ring, a clamping ring and a vibration-damping plate, wherein the turbine blades and the turbine disk are connected via tenons and tenons, and the tenons and tenons are both two-tooth structures; the vibration-damping plate is installed between the turbine blades and the turbine disk; the comb-tooth sealing ring is assembled on the turbine disk and axially limited by the clamping ring; the turbine blades are in a bent and twisted shape.

[0008] Furthermore, a turbine blade boss is provided on the lower side of the turbine blade edge plate.

[0009] Furthermore, the comb teeth sealing ring is installed on the turbine disc with an interference fit.

[0010] Furthermore, the rear end surface of the comb teeth sealing ring presses the front side of the blade tenon.

[0011] Furthermore, a turbine disk boss is provided on the tenon edge of the turbine disk, and a turbine disk groove is provided on the side surface of the turbine disk boss.

[0012] Furthermore, the side surface of the turbine disk groove is a slope extending downward, and the slope is connected to the turbine disk boss through a continuous curved surface.

[0013] Furthermore, when assembling the turbine blade and the turbine disk, the turbine blade boss is fitted with the bottom of the turbine disk groove, and the turbine disk boss is fitted with the end surface of the lower side of the turbine blade edge plate.

[0014] Furthermore, the upper end surface and the side surface of the turbine disk boss are connected through an arc surface, and the thickness of the boss is 5 to 8 mm.

[0015] Furthermore, a groove is formed at each of the matching positions of the comb teeth sealing ring and the turbine disk journal, and a clamping ring in an open state is installed at the groove.

[0016] Furthermore, the vibration reduction sheet is a damping sheet material or a resin-based material.

[0017] Furthermore, an aircraft engine adopts the above-mentioned turbine rotor structure.

[0018] Compared with the prior art, the present invention has the following beneficial effects:

[0019] The present invention comprises a turbine blade, a turbine disc, a comb-tooth sealing ring, a clamping ring and a vibration-damping plate, wherein the turbine blade is connected to the turbine disc through a tenon and a tenon groove, and both the tenon and the tenon groove are two-tooth structures; the vibration-damping plate is installed between the turbine blade and the turbine disc; the comb-tooth sealing ring is assembled on the turbine disc and axially limited by the clamping ring; the turbine blade is bent and twisted; the improved turbine rotor structure can prevent the turbine blade from moving backward, and the boss on the outer edge of the turbine disc is added to fit with the end face of the lower side of the blade edge plate to limit the backward movement of the blade, and the rear end face of the comb-tooth sealing ring presses the front side of the blade tenon, and the axial movement of the comb-tooth sealing ring relative to the turbine disc is limited by adding the clamping ring to prevent the blade from moving forward; thereby solving the problem of turbine blade backward movement and the rotor jamming failure caused by it, greatly reducing the damage to the turbine blade and the turbine disc, and improving the working reliability of the aircraft engine. BRIEF DESCRIPTION OF THE DRAWINGS

[0020] Figure 1 It is a schematic diagram of the turbine rotor structure of the present invention;

[0021] Figure 2 It is a schematic diagram of the structure of the turbine disk of the present invention;

[0022] Figure 3 It is a schematic diagram of the structure of the turbine blade of the present invention;

[0023] Figure 4 It is a diagram showing the assembly relationship between the turbine blade and the turbine disk of the present invention;

[0024] Figure 5 It is a schematic diagram of the assembly of turbine blades and guide vanes;

[0025] In the above figure, 1. turbine blade; 2. turbine disk; 3. comb tooth sealing ring; 4. retaining ring; 5. vibration damping plate; 6. edge plate; 7. turbine disk groove; 8. turbine disk boss; 9. turbine blade boss; 10. turbine blade tenon; 11. turbine guide. DETAILED DESCRIPTION

[0026] In order to clearly illustrate the technical features of the application scheme of the present invention, the present invention is described in detail below through specific implementation methods and in conjunction with the accompanying drawings.

[0027] In the following description, many specific details are set forth to facilitate a full understanding of the present application. However, the present application may also be implemented in other ways different from those described herein. Therefore, the scope of protection of the present application is not limited to the specific embodiments disclosed below.

[0028] In addition, in the description of the present application, it should be understood that the terms "center", "upper", "lower", "front", "back", "left", "right", "vertical", "horizontal", "top", "bottom", "inner", "outer", "axial", "radial", "circumferential" and the like indicate positions or positional relationships based on the positions or positional relationships shown in the accompanying drawings, and are only for the convenience of describing the present application and simplifying the description, rather than indicating or implying that the device or element referred to must have a specific orientation, be constructed and operated in a specific orientation, and therefore cannot be understood as a limitation on the present application. In addition, the terms "first" and "second" are used only for descriptive purposes, and cannot be understood as indicating or implying relative importance or implicitly indicating the number of technical features indicated. Thus, the features defined as "first" and "second" may explicitly or implicitly include one or more of such features. In the description of the present application, the meaning of "multiple" is two or more, unless otherwise clearly and specifically defined.

[0029] In this application, unless otherwise clearly specified and limited, the terms "installed", "connected", "connected", "fixed" and the like should be understood in a broad sense, for example, it can be a fixed connection, a detachable connection, or an integral connection; it can be a mechanical connection, an electrical connection, or a communication; it can be a direct connection, or an indirect connection through an intermediate medium, it can be the internal connection of two elements or the interaction relationship between two elements. For ordinary technicians in this field, the specific meanings of the above terms in this application can be understood according to specific circumstances.

[0030] In the present application, unless otherwise clearly specified and limited, a first feature "above" or "below" a second feature may be that the first and second features are in direct contact, or the first and second features are in indirect contact through an intermediate medium. In the description of this specification, reference terms; "one embodiment", "some embodiments", "example", "specific example", or "some examples" and the like are intended to mean that the specific features, structures, materials or characteristics described in conjunction with the embodiment or example are included in at least one embodiment or example of the present application. In this specification, the schematic representations of the above terms do not necessarily refer to the same embodiment or example. Moreover, the specific features, structures, materials or characteristics described may be combined in an appropriate manner in any one or more embodiments or examples.

[0031] Example 1

[0032] like Figure 1As shown, a turbine rotor structure and an aircraft engine for preventing blades from moving backward include a turbine blade 1, a turbine disk 2, a comb-tooth sealing ring 3, a clamping ring 4 and a vibration-damping plate 5. The turbine blade 1 and the turbine disk 2 are connected by tenons and tenons, and the tenons and tenons are both two-tooth structures; the vibration-damping plate 5 is installed between the turbine blade 1 and the turbine disk 2; the comb-tooth sealing ring 3 is assembled on the turbine disk 2 and is axially limited by the clamping ring 4; the turbine blade 1 is bent and twisted.

[0033] like Figure 2 and Figure 3 As shown, in this embodiment, by improving the structure of the mortise and tenon parts of the turbine blade 1 and the turbine disk 2, the original 3-tooth structure is changed to a 2-tooth structure, and a turbine disk groove 7 and a turbine disk boss 8 are added to the outer edge of the turbine disk 2; wherein, the side surface of the turbine disk groove 7 is a downwardly extending inclined surface, the inclined surface gradually extends downward and is transitionally connected to the root of the turbine disk boss 8 through a continuous curved surface; a turbine blade boss 9 is provided on the end surface of the lower side of the turbine blade edge plate.

[0034] like Figure 4 As shown, when assembling the turbine blade 1 and the turbine disk 2, firstly, multiple turbine blades 1 are sequentially installed into the turbine disk 2, ensuring that the turbine blade boss 9 fits and fits with the bottom of the turbine disk groove 7, that is, the rear end of the turbine blade boss 9 fits and fits with the front end of the turbine disk groove 7, and the turbine disk boss 8 on the outer edge of the turbine disk 2 fits with the lower end face of the edge plate 6 of the turbine blade 1, so as to limit the backward movement of the turbine blade 1, and then the clamping ring 4 is installed into the groove at the front end of the turbine disk 2, and the diameter of the clamping ring 4 is limited by the assembly tool, and then the comb tooth sealing ring 3 is guided by the assembly tool to cooperate with the shaft neck of the turbine disk 2, and its rear end face presses the front side of the turbine blade tenon 10, and finally the assembly tool is withdrawn, and the clamping ring 4 is in an open state, and the diameter increases. The maximum state is limited by the inner diameter of the clamping groove of the comb tooth sealing ring 3. At this time, the clamping ring 4 cooperates with the comb tooth sealing ring 3 and the groove of the turbine disk 2, and the axial movement of the comb tooth sealing ring 3 relative to the turbine disk 2 is limited by adding the clamping ring 4 to prevent the blade from moving forward.

[0035] The present invention effectively limits the axial movement of the turbine blade 1, thereby solving the problem of backward movement of the turbine blade 1 and the resulting turbine rotor jamming failure, greatly reducing damage to the turbine blade 1 and the turbine disk 2, and improving the working reliability of the aircraft engine.

[0036] This embodiment also provides an aircraft engine, which adopts the turbine rotor structure described in this embodiment.

[0037] Example 2

[0038] like Figure 1As shown, a turbine rotor structure and an aircraft engine for preventing blades from moving backward include a turbine blade 1, a turbine disk 2, a comb-tooth sealing ring 3, a clamping ring 4 and a vibration-damping plate 5. The turbine blade 1 and the turbine disk 2 are connected by tenons and tenons, and the tenons and tenons are both two-tooth structures; the vibration-damping plate 5 is installed between the turbine blade 1 and the turbine disk 2; the comb-tooth sealing ring 3 is assembled on the turbine disk 2 and is axially limited by the clamping ring 4; the turbine blade 1 is bent and twisted.

[0039] like Figure 4 As shown, the improved turbine rotor structure consists of a turbine blade 1, a turbine disc 2, a comb-tooth sealing ring 3, a clamping ring 4, and a vibration damping plate 5. Compared with the traditional turbine rotor structure, the addition of the clamping ring 4 and the vibration damping plate 5 can effectively prevent the turbine blade 1 from moving and ensure the stability of the turbine engine. Specifically, a groove is processed at the joint between the comb-tooth sealing ring 3 and the journal of the turbine disc 2, and a clamping ring 4 is added to the groove. The clamping ring 4 is an open annular structure, which limits the axial movement of the comb-tooth sealing ring 3 relative to the turbine disc 2; the outer diameter of the rear end face of the comb-tooth sealing ring 3 is increased to press the front side of the turbine blade tenon 10, and the vibration damping plate 5 is added, which can reduce the vibration stress of the turbine blade 1 and increase the service life of the turbine blade 1.

[0040] The clamp ring 4 limits the axial movement of the comb teeth sealing ring 3 relative to the turbine disk 2. The rear end face of the comb teeth sealing ring 3 presses the front side of the turbine blade tenon 10 to prevent the blade from moving forward. The turbine disk boss 8 on the outer edge of the turbine disk 2 fits with the end face under the edge plate 6 of the turbine blade 1 to limit the backward movement of the blade, thereby ensuring the stability of the turbine blade 1. Figure 5 As shown, the improved turbine rotor axial fixing structure has high reliability and can prevent the turbine blades 1 from moving and being scraped by the turbine guide vane 11 .

[0041] Example 3

[0042] like Figure 1 As shown, a turbine rotor structure and an aircraft engine for preventing blades from moving backward include a turbine blade 1, a turbine disk 2, a comb-tooth sealing ring 3, a clamping ring 4 and a vibration-damping plate 5. The turbine blade 1 and the turbine disk 2 are connected by tenons and tenons, and the tenons and tenons are both two-tooth structures; the vibration-damping plate 5 is installed between the turbine blade 1 and the turbine disk 2; the comb-tooth sealing ring 3 is assembled on the turbine disk 2 and is axially limited by the clamping ring 4; the turbine blade 1 is bent and twisted.

[0043] like Figure 4 As shown, the tenons and tenons of the turbine rotor of the present invention are asymmetric structures, and the turbine disk boss 8 on the outer edge of the turbine disk 2 is fitted with the end surface of the lower side of the edge plate 6 of the turbine blade 1; specifically, the thickness of the turbine disk boss 8 is 6 mm, and the turbine disk 2 is an irregular three-dimensional structure with a curved side surface, which can effectively limit the backward movement of the blades and also prevent cold air leakage.

[0044] In this embodiment, if Figure 3 As shown, a vibration damping plate 5 is provided between the lower end surface of the edge plate 6 of the turbine blade 1 and the upper end surface of the tenon of the turbine disk 2. A damping plate material can be selected for vibration reduction, and the vibration amplitude of the blade can be reduced through its elastic structure. The damping plate is usually made of a material with elastic and damping properties, such as a specific alloy; resin-based or rubber shock-absorbing materials can also be selected to effectively absorb the impact energy of vibration.

[0045] Obviously, the above-described embodiments of the present invention are merely examples for clearly illustrating the present invention, and are not intended to limit the embodiments of the present invention. For those skilled in the art, other different forms of changes or modifications can be made based on the above description. It is not necessary and impossible to list all the embodiments here. Any modifications, equivalent substitutions, improvements, etc. made within the spirit and principles of the present invention should be included in the protection scope of the claims of the present invention.

Claims

1. A turbine rotor structure for preventing blades from moving backward, characterized in that: It includes turbine blades, turbine disc, comb teeth sealing ring, clamping ring and vibration damping plate, the turbine blades and turbine disc are connected by tenons and tenons, and the tenons and tenons are two-tooth structures; the vibration damping plate is installed between the turbine blades and the turbine disc; the comb teeth sealing ring is assembled on the turbine disc and axially limited by the clamping ring; the turbine blades are bent and twisted.

2. A turbine rotor structure for preventing blades from moving backward according to claim 1, characterized in that: A turbine blade boss is provided on the lower side of the turbine blade edge plate.

3. A turbine rotor structure for preventing blades from moving backward according to claim 1, characterized in that: The comb teeth sealing ring is installed on the turbine disk by interference fit.

4. A turbine rotor structure for preventing blades from moving backward according to claim 3, characterized in that: The rear end surface of the comb teeth sealing ring presses the front side of the blade tenon.

5. A turbine rotor structure for preventing blades from moving backward according to claim 1, characterized in that: A turbine disc boss is provided on the tenon tooth edge of the turbine disc, and a turbine disc groove is provided on the side surface of the turbine disc boss.

6. A turbine rotor structure for preventing blades from moving backward according to claim 5, characterized in that: The side surface of the turbine disc groove is a slope extending downward, and the slope is connected to the turbine disc boss through a continuous curved surface.

7. A turbine rotor structure for preventing blades from moving backward according to claim 5, characterized in that: When assembling the turbine blade and the turbine disk, the turbine blade boss is fitted with the bottom of the turbine disk groove, and the turbine disk boss is fitted with the end surface of the lower side of the turbine blade edge plate.

8. A turbine rotor structure for preventing blades from moving backward according to claim 7, characterized in that: The upper end surface and the side surface of the turbine disk boss are connected through an arc surface, and the thickness of the boss is 5 to 8 mm.

9. A turbine rotor structure for preventing blades from moving backward according to claim 1, characterized in that: A groove is respectively opened at the matching position between the comb tooth sealing ring and the turbine disk journal, and a clamping ring in an open state is installed at the groove.

10. An aircraft engine, characterized in that: A turbine rotor structure as claimed in any one of claims 1 to 9 is adopted.

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