High-pressure turbine working blade and aero-engine

Abstract

The invention provides a high-pressure turbine working blade and an aero-engine, and belongs to the technical field of aero-engines. The high-pressure turbine working blade comprises a blade body, a margin plate and a tenon; a basin side damping groove and a back side damping groove which are used for being matched with adjacent high-pressure turbine working blades to form a damping block containing space are formed in the basin side and the back side of the lower surface of the margin plate respectively, and an axial limiting face is arranged at the axial front end of the basin side damping groove. The axial limiting face is narrower in the axial length compared with the axial front end edge of the back side damping groove so that the axial limiting face can be in matched contact with the front end surface of the damper in the axial direction. According to the high-pressure turbine working blade provided by the invention, an unspecific point contact matching mode of the damper in the axial direction is adjusted into a surface contact matching mode, so that the single-side axial positioning of the damper is ensured, the contact matching of the damper and the blade is more stable, and the damping effect is improved.

Description

Technical Field

[0001] The present application relates to the field of aero-engines, and in particular to a high-pressure turbine blade and an aero-engine. Background Art

[0002] The high-pressure turbine rotor of an aircraft engine operates in a high-temperature, high-pressure, and high-speed environment. Under these conditions, the high-pressure turbine blades experience vibration stresses of a specific order. Resonance can occur at certain speeds, endangering the blades and even the entire engine. For high-pressure turbine blades without shrouds, dampers are often used when vibration occurs during operation. Friction between the damper and the blade creates structural damping, reducing vibration stress.

[0003] The fit between the damper and the blade—if it deviates from the design requirements, the damping effect will be lower than expected, resulting in the blade being unable to withstand the increased vibration stress and being damaged. Therefore, a reasonable structural design is required to ensure that the damper can achieve the damping effect on the blade.

[0004] Existing dampers often feature a clearance fit in the radial and tangential directions, meaning the damper is inserted between the blade flanges. This ensures that the damper fits smoothly into the turbine disk and does not fall out of the cavity when the blade is assembled. However, lack of clear axial control can lead to deviations from design requirements during operation between the damper and the flanges, resulting in poor contact between the damper and the blade flanges, a significant reduction in contact area, and unsatisfactory damping performance. This can increase blade vibration stress and even lead to high-cycle fatigue fracture.

[0005] In addition, the current axial positioning of the high-pressure turbine blades and the damper assembly is point contact, the mating surface angles are different, the contact surface is unstable, and it is difficult to ensure the radial contact of the damper at the blade edge position, resulting in the damper not being able to play the expected damping effect and insufficient damping effect. The damper and the blades are in poor contact, resulting in local uneven force under the action of centrifugal force during operation, leading to failure. The existing damper also has a gas sealing function. The damper and the blades are in poor contact, and under the action of centrifugal force during operation, the sealing effect is insufficient, resulting in excessive leakage of cold air into the mainstream gas channel, causing performance loss of the entire machine. Summary of the Invention

[0006] The purpose of the present application is to provide a high-pressure turbine blade and an aircraft engine to solve or alleviate at least one problem in the background technology.

[0007] The technical solution of the present application is: a high-pressure turbine working blade, the high-pressure turbine working blade comprising a blade body, an edge plate and a tenon, the basin side and the back side of the lower surface of the edge plate are respectively provided with a basin side damping groove and a back side damping groove for cooperating with adjacent high-pressure turbine working blades to form a damping block accommodating space, wherein the axial front end of the basin side damping groove is provided with an axial limiting surface, and the axial limiting surface is narrower in axial length than the axial front end edge of the back side damping groove, so as to cooperate and contact with the front end surface of the damper in the axial direction.

[0008] In at least one embodiment of the present application, one or more bosses are provided in the basin-side damping groove and the back-side damping groove along the tangential direction, for limiting the gap of the damper.

[0009] In at least one embodiment of the present application, when there are multiple bosses, the multiple bosses are distributed along the axial direction, and the length of the boss in the last section of the axial distribution is shorter than the lengths of the other bosses.

[0010] In at least one embodiment of the present application, the bosses in the basin-side damping groove and the back-side damping groove are collinearly arranged in the tangential direction.

[0011] In at least one embodiment of the present application, the thickness of the boss is 1 mm to 5 mm.

[0012] In at least one embodiment of the present application, the boss and the axial limit surface are both cast in the low stress area of ​​the high-pressure turbine working blade, and the sharp edges of the boss and the axial limit surface are rounded to reduce stress concentration.

[0013] On the other hand, the present application provides an aircraft engine, comprising any of the high-pressure turbine blades described above.

[0014] The high-pressure turbine blades and aircraft engines provided by this application have the following advantages:

[0015] 1) Ensures more effective damping effect

[0016] The blades limit the damper in all directions and reasonably constrain the position of the damper in the blades, ensuring that the damper has good contact and friction with the blades and that the damping effect is effectively exerted;

[0017] 2) Improve the reliability of the damper

[0018] By increasing the axial contact area between the blade and the damper, the damper's stress area is effectively reduced, the stress concentration effect is reduced, and the damper's reliability is improved;

[0019] 3) Ensure the sealing effect of the damper

[0020] Reasonable positioning of the damper ensures full contact between the damper and the blades, reduces cold air leakage, and ensures a tight sealing effect. BRIEF DESCRIPTION OF THE DRAWINGS

[0021] In order to more clearly illustrate the technical solutions provided by this application, the following is a brief introduction to the accompanying drawings. Obviously, the accompanying drawings described below are only some embodiments of this application.

[0022] Figure 1 This is an enlarged schematic diagram of the overall layout of the high-pressure turbine working blades of this application.

[0023] Figure 2 This is a schematic diagram of the separation state of the high-pressure turbine working blades and the damper in this application.

[0024] Figure 3 This is a schematic diagram of the assembly state of the high-pressure turbine working blades and damper in this application.

[0025] Reference numerals:

[0026] 10-High-pressure turbine blades

[0027] 11-Leaf

[0028] 12-Edge Board

[0029] 121- Basin side damping groove

[0030] 122-Axial limit surface

[0031] 123-Boss

[0032] 124-Back side damping groove

[0033] 125-Axial front edge

[0034] 13-tenon

[0035] 20-Damper

[0036] 21-Front end surface DETAILED DESCRIPTION

[0037] In order to make the purpose, technical solutions and advantages of the implementation of this application clearer, the technical solutions in the embodiments of this application will be described in more detail below in conjunction with the drawings in the embodiments of this application.

[0038] In order to overcome the problems existing in the damper and high-pressure turbine blades in the prior art, the present application provides a high-pressure turbine blade and an aircraft engine.

[0039] like Figures 1 to 3As shown, the high-pressure turbine blade 10 provided in this application includes a blade body 11, a flange 12, and a tenon 13. The blade body 11 and the tenon 13 are respectively located on the upper and lower sides of the flange 12. The tenon 13 is used to connect with the turbine disk to achieve the fixation of the high-pressure turbine blade 10. For example, the tenon 13 can have a dovetail shape, a longitudinal tree shape, or a sawtooth shape.

[0040] The lower surface of the edge plate 12 is provided with a basin-side damping groove 121 and a back-side damping groove 124, respectively, for mating with adjacent high-pressure turbine blades 10', thereby forming a mounting space for the damper 20. An axial limiting surface 122 is provided at the axial front portion (i.e., the front portion in the airflow direction) of the basin-side damping groove 121, for mating with the front end surface 21 of the damper 20 in the axial direction (i.e., the X-direction). The axial limiting surface 122 is narrower in axial length than the axial front end edge 125 of the back-side damping groove 124.

[0041] by Figure 2 and Figure 3 Taking the high-pressure turbine blade 10 and the adjacent high-pressure turbine blade 10' shown as an example, the basin side of the edge plate 12 of the high-pressure turbine blade 10 is provided with a basin side damping groove 121, and the axial front portion thereof is provided with an axial limiting surface 122, and the back side of the edge plate 12 is provided with a back side damping groove 124, and the axial front portion thereof is provided with an axial front end edge 125. The back side of the edge plate 12' of the adjacent high-pressure turbine blade 10' has a back side damping groove 124' that is compatible with the basin side damping groove 121 of the high-pressure turbine blade 10, and the axial front portion of the back damping groove 124' is provided with an axial front end edge 125'. From Figure 2 It can be seen that the axial front end edge 125' of the adjacent high-pressure turbine working blade 10' is not aligned with the axial limit surface 122 of the high-pressure turbine working blade 10, and the axial length of the latter is smaller than the axial length of the former, that is, the axial limit surface 122 of the high-pressure turbine working blade 10 is narrower in the axial direction than the axial front end edge 125' of the adjacent high-pressure turbine working blade 10'.

[0042] In the working state, under the influence of centrifugal force, the damper 20 will move toward the axial limit surface 122 of the high-pressure turbine working blade 10, and the axial limit surface 122 can axially limit the front end surface 21 of the damper, thereby ensuring that the damper 20 has sufficient surface contact friction with the high-pressure turbine working blade 10 and the lower surface of the edge plate of the adjacent high-pressure turbine working blade 10' to achieve a damping effect.

[0043] In the present application, a certain space is provided at the axial rear portion of the pot side and the back side of the high-pressure turbine rotor blade 10 to ensure the assembly of the damper 20 .

[0044] One or more bosses 123 extending in the tangential (Y-direction) direction are provided within the basin-side damping groove 121 and the back-side damping groove 124 of the high-pressure turbine blade 10 to provide clearance limiting, ensuring that the damper 20 does not fall off from the assembly position of the basin-side damping groove 121 and the back-side damping groove 124 when the engine is not operating. The bosses 123 have a certain thickness in the radial (Z-direction) direction, which can be designed as needed. For example, the thickness of the bosses 123 can be set to 1 mm to 5 mm.

[0045] In a preferred embodiment of the present application, there are three bosses 123, which are distributed front and back in the axial direction, and the tangential length of the rear boss 123 is smaller than that of the middle and front bosses 123, so that the rear boss 123 that cooperates with the high-pressure turbine working blade 10 and the adjacent high-pressure turbine working blade 10' only limits the gap of the damper 20 in the tangential direction to ensure the assembly of the damper 20.

[0046] Furthermore, the bosses 123 in the basin-side damping groove 121 and the back-side damping groove 124 are substantially collinear in the tangential direction.

[0047] In the present application, the boss 123 and the axial limit surface 122 in the high-pressure turbine blade 10 that contact the damper 20 can be set in the low-stress area of ​​the high-pressure turbine blade 10 by casting. The low-stress area can be obtained through simulation. At the same time, the boss 123 and the axial limit surface 122 are provided with appropriate sharp edge rounding and other structures to avoid local stress concentration. At the same time, it will not have a significant impact on production and processing, and can ensure sufficient casting feasibility. For example, the shape of the boss 123 can be rectangular, runway-shaped, or long strip, etc. The boss 123 shown in the figure of this application is a long strip with a round head.

[0048] In some embodiments of the present application, the damping block 20 is a metal damping block, a composite material damping block, or a damping structure with a metal shell and an internal filling of fiber or rubber blocks.

[0049] Compared with the existing large clearance matching method used in the matching structure between the high-pressure turbine blades and the damper, the radial friction contact is generated by the centrifugal force of the damping structure when the engine is working, and there is no special constraint in the axial direction, so that the damping structure will produce axial stringing during operation. On the one hand, it accelerates the failure and damage of the damping structure, and on the other hand, it reduces the expected damping effect. The high-pressure turbine blade provided by this application adjusts the matching method of the damper's unspecific axial direction and point contact to a surface contact matching method, ensuring the unilateral axial positioning of the damper and making the contact and matching between the damper and the blade more stable. This structure comprehensively considers multiple factors such as blade cooling, casting processability, sealing, weight and damping caused by friction, and uses the smallest possible changes and adjustments to achieve the desired matching function.

[0050] Finally, the present application also provides an aircraft engine, which includes the above-mentioned high-pressure turbine working blades.

[0051] The high-pressure turbine blades and aircraft engines provided by this application have the following advantages:

[0052] 1) Ensures more effective damping effect

[0053] The blades limit the damper in all directions and reasonably constrain the position of the damper in the blades, ensuring that the damper has good contact and friction with the blades and that the damping effect is effectively exerted;

[0054] 2) Improve the reliability of the damper

[0055] By increasing the axial contact area between the blade and the damper, the damper's stress area is effectively reduced, the stress concentration effect is reduced, and the damper's reliability is improved;

[0056] 3) Ensure the sealing effect of the damper

[0057] Reasonable positioning of the damper ensures full contact between the damper and the blades, reduces cold air leakage, and ensures a tight sealing effect.

[0058] The above description is merely a specific embodiment of the present application, but the scope of protection of the present application is not limited thereto. Any changes or substitutions that can be easily conceived by a person skilled in the art within the technical scope disclosed in this application should be included in the scope of protection of the present application. Therefore, the scope of protection of the present application should be based on the scope of protection of the claims.

Claims

1. A high-pressure turbine blade, characterized in that: The high-pressure turbine working blade includes a blade body, an edge plate and a tenon. The basin side and the back side of the lower surface of the edge plate are respectively provided with a basin side damping groove and a back side damping groove for cooperating with adjacent high-pressure turbine working blades to form a damping block accommodating space, wherein the axial front end of the basin side damping groove is provided with an axial limiting surface, and the axial limiting surface is narrower in axial length than the axial front end edge of the back side damping groove, so as to cooperate and contact with the front end surface of the damper in the axial direction.

2. The high-pressure turbine blade according to claim 1, wherein: One or more bosses are provided in the basin-side damping groove and the back-side damping groove along the tangential direction, for limiting the gap of the damper.

3. The high-pressure turbine blade according to claim 2, wherein: When there are multiple bosses, the bosses are distributed along the axial direction, and the length of the boss in the last section of the axial distribution is shorter than the lengths of the other bosses.

4. The high-pressure turbine rotor blade according to claim 2 or 3, characterized in that: The bosses in the basin-side damping groove and the back-side damping groove are collinearly arranged in the tangential direction.

5. The high-pressure turbine rotor blade according to claim 4, characterized in that: The thickness of the boss is 1 mm to 5 mm.

6. The high-pressure turbine rotor blade according to claim 5, characterized in that: The boss and the axial limiting surface are both arranged in the low stress area of ​​the high-pressure turbine working blade by casting, and the sharp edges of the boss and the axial limiting surface are rounded to reduce stress concentration.

7. An aircraft engine, characterized in that: The aircraft engine comprises the high-pressure turbine blade according to any one of claims 1 to 6.

Images