CN112443360B

CN112443360B - Aeroengine blade and aeroengine

Info

Publication number: CN112443360B
Application number: CN201910807736.2A
Authority: CN
Inventors: 陈杰; 李秋胜
Original assignee: City University of Hong Kong CityU
Current assignee: City University of Hong Kong CityU
Priority date: 2019-08-29
Filing date: 2019-08-29
Publication date: 2022-09-27
Anticipated expiration: 2039-08-29
Also published as: CN112443360A

Abstract

The invention discloses an aircraft engine blade and an aircraft engine, wherein the aircraft engine blade (2) comprises a blade body (21), a vibration reduction cavity (3) arranged along the diameter direction of the aircraft engine is arranged in the blade body (21), and a vibration reduction plug (4) capable of reducing the vibration of the engine blade is filled in the vibration reduction cavity (3). A damping cavity is arranged in the damping cavity of the blade of the aero-engine, and a damping plug used for preventing the blade of the aero-engine from vibrating is filled in the damping cavity. The vibration damping plug has obvious effects of buffering and absorbing the vibration energy of the blade and the like, greatly reduces the vibration heating of the engine blade, further solves the vibration problem of the wheel disc-blade during high-speed rotation, prevents the blade fatigue problem and prolongs the service life of the engine blade.

Description

Aeroengine blade and aeroengine

Technical Field

The invention relates to the field of engines, in particular to an aircraft engine blade and an aircraft engine comprising the same.

Background

The aircraft engine is the heart of the aircraft and is of great importance to the safety of the aircraft. The aircraft engine structure is unique, mainly in that the aircraft engine pursues the thrust-weight ratio, the working efficiency and the economy as high as possible. For a long time, the problems of operational reliability and structural integrity have been the bottleneck restricting the development of high performance engines, and one of the core problems is the structural vibration of the turbine wheel disc-blade under severe working conditions of high pressure and high rotating speed. Statistics show that in an aircraft engine, blade fatigue failure and accidents caused by disk structure vibration account for about 70% of total failure and accidents. It should be noted that, since the occurrence and development of a failure of the wheel disc structure are characterized by rapidity and great destructiveness, once the failure occurs, the failure often causes catastrophic results. Thus, the dynamic and vibration characteristics of the wheel disk assembly have a decisive influence on the fatigue life and performance of the aircraft engine.

Disclosure of Invention

In order to solve the problem that the existing engine blade is easy to fatigue and damage, the invention provides an aircraft engine blade and an aircraft engine. The vibration damping plug has obvious effects of buffering and absorbing the vibration energy of the blade and the like, greatly reduces the vibration heating of the engine blade, further solves the vibration problem of the wheel disc-blade during high-speed rotation, prevents the blade fatigue problem and prolongs the service life of the engine blade.

The technical scheme adopted by the invention for solving the technical problems is as follows: the utility model provides an aeroengine blade, includes the blade body, this internal damping chamber that sets up along aeroengine's diameter direction that is equipped with of blade, and the damping intracavity is filled has the damping stopper that can weaken this engine blade vibration.

The inner end of the vibration reduction cavity is of a closed structure, the outer end of the vibration reduction cavity is of an open structure, a gland capable of sealing the vibration reduction cavity is arranged at the outer end of the vibration reduction cavity, and the gland is detachably connected with the blade body.

The gland is spliced with the blade body, and the outer end of blade body is equipped with the gland groove that is used for matcing the installation gland, and the gland groove is seted up along aeroengine's axis direction, and the installation direction of gland is the same with aeroengine's axis direction, the shape closure between gland and the blade body.

The both sides of gland are equipped with a plurality of interlock beads, and the interlock bead sets up along aeroengine's axis direction, and the both sides of gland groove are equipped with the interlock recess with interlock bead assorted.

The cooperation of gland and blade body is tight fit, gland and blade body sealing connection, and the surface of interlock recess is equipped with the graphite sealing layer.

The section shape of the vibration reduction cavity is the equal proportion reduction of the section shape of the blade body, and the side wall thickness of the blade body at the position of the vibration reduction cavity is uniform.

The internal surface in the damping chamber is equipped with the locking sand grip that a plurality of diameter directions along aeroengine set up, and the circumference interval arrangement in the damping chamber is followed to a plurality of locking sand grips, and the damping is filled completely in the damping intracavity in the damping stopper, and the material of damping stopper is high temperature resistant flexible damping material.

The damping intracavity is equipped with middle strengthening rib, and the damping chamber is divided into a plurality of cellular time damping chambeies by middle strengthening rib, and the damping stopper is filled in each time damping intracavity.

The aero-engine comprises a wheel disc and a plurality of engine blades, wherein the engine blades are arranged at intervals along the circumferential direction of the wheel disc, and the engine blades are the engine blades.

The aeroengine blade further comprises a connecting part, the connecting part is connected with the inner end of the blade body, the connecting part is in joggle joint with the wheel disc, and the installation direction of the engine blade is the same as the axis direction of the aeroengine.

The invention has the beneficial effects that: through set up semi-closed damping chamber in engine blade, it has the damping stopper that is used for preventing engine blade vibration to fill in the damping chamber, can dismantle to be equipped with the gland corresponding to the open end in damping chamber on engine blade. The vibration damping plug has the function of buffering and absorbing vibration energy, and buffers and absorbs the vibration energy of the whole engine blade, so that the vibration of the blade in the high-speed rotation process is prevented from influencing the structure of the whole blade; the gland is used for sealing the vibration reduction cavity, the vibration reduction plug is prevented from throwing out the blades along the vibration reduction cavity to cause safety accidents, and meanwhile, external gas can be prevented from entering the vibration reduction cavity to cause vibration of the vibration reduction cavity to influence vibration of the blades and influence the service life of the blades. The middle reinforcing rib strengthens the connection of blade structures on two sides of the cavity, has a structure strengthening effect, prevents the blades from deforming, increases the contact area of the damping plug and the engine blades, and is more favorable for buffering and absorbing vibration energy. The secondary vibration damping cavity is arranged in a honeycomb shape, so that the contact area between the vibration damping plug and the blade is obviously increased, the vibration damping plug has more remarkable effects on the buffer absorption of the vibration energy of the blade and the like, the vibration of the engine blade is greatly reduced, the vibration problem of the wheel disc-blade during high-speed rotation is further solved, the problem of blade fatigue is prevented, and the service life of the engine blade is prolonged. Compared with the traditional engine blade without a damping cavity and a damping plug, the engine blade provided by the invention has the advantage that the fatigue life is increased by 15-20% through comparison of a vibration fatigue test.

Drawings

The accompanying drawings, which are incorporated in and constitute a part of this application, illustrate embodiments of the invention and, together with the description, serve to explain the invention and not to limit the invention.

FIG. 1 is a schematic view of the overall structure of an aircraft engine blade according to the invention in one embodiment.

Fig. 2 is a schematic structural view of a blade body and a connecting portion of an aircraft engine blade according to a first embodiment of the invention.

FIG. 3 is a top view of an aircraft engine blade according to the present invention in one embodiment.

Fig. 4 is a schematic sectional view of a gland in the first embodiment.

Fig. 5 is a plan view of the gland in the first embodiment.

Fig. 6 is a sectional view taken along a-a in fig. 1.

Fig. 7 is a schematic view of the overall structure of an aircraft engine blade according to a second embodiment of the invention.

Fig. 8 is a sectional view taken along the direction B-B in fig. 7.

Fig. 9 is a schematic view of the overall structure of an aircraft engine blade according to the third embodiment of the invention.

Fig. 10 is a sectional view taken along the direction C-C in fig. 9.

1. A wheel disc; 2. an engine blade; 3. a vibration damping cavity; 4. a damping plug; 5. a gland; 6. a gland groove; 7. a stop convex strip; 8. a middle reinforcing rib;

21. a blade body; 22. a connecting portion;

31. a secondary damping chamber;

51. an occlusal bead;

61. and (4) meshing the grooves.

Detailed Description

It should be noted that, in the present application, the embodiments and features of the embodiments may be combined with each other without conflict. The present invention will be described in detail below with reference to the embodiments with reference to the attached drawings.

Example one

An aircraft engine blade 2 comprises a blade body 21, a vibration damping cavity 3 arranged along the diameter direction of an aircraft engine is arranged in the blade body 21, and a vibration damping plug 4 capable of weakening the vibration of the engine blade is filled in the vibration damping cavity 3, as shown in figure 1.

The vibration damping plug 4 has obvious effects of buffering and absorbing the vibration energy of the engine blade 2 and the like, greatly reduces the vibration heating of the engine blade, further solves the vibration problem of the wheel disc-blade during high-speed rotation, prevents the blade fatigue problem and prolongs the service life of the engine blade.

In this embodiment, the inner end of the vibration damping cavity 3 is a closed structure, the outer end of the vibration damping cavity 3 is an open structure, the outer end of the vibration damping cavity 3 is provided with a gland 5 capable of closing the vibration damping cavity 3, the gland 5 is located at the outer end of the blade body 21, the gland 5 is detachably connected with the blade body 21, and the gland 5 enables the vibration damping cavity 3 to become a sealed cavity. The gland 5 is used for sealing the damping cavity 3, preventing the damping plug 4 from throwing away the blade along the damping cavity 3, and simultaneously preventing the external gas from entering the damping cavity 3 to cause the vibration of the damping cavity 3 to influence the vibration of the blade and influence the service life of the blade.

The inner end of the damper chamber 3 is the end of the damper chamber 3 that faces the disk 1 to be described below, i.e., the lower end of the damper chamber 3 in fig. 1. The outer end of the damping cavity 3 is the end of the damping cavity 3 facing away from the wheel disc 1, i.e. the upper end of the damping cavity 3 in fig. 1. The inner end of the blade body 21 is the end of the blade body 21 facing the disk 1, i.e. the lower end of the blade body 21 in fig. 1, and the outer end of the blade body 21 is the end of the blade body 21 facing away from the disk 1, i.e. the upper end of the blade body 21 in fig. 1. The detachable connection of the gland 5 and the blade body 21 can be realized by screw connection or plug connection. The mounting direction of the gland 5 may be the same as the axial direction of the aircraft engine, and may be the same as the diameter direction of the aircraft engine.

Preferably, the gland 5 is inserted into the blade body 21, the outer end of the blade body 21 is provided with a gland groove 6 for matching and installing the gland 5, the gland groove 6 is arranged along the axial direction of the aircraft engine, the installation direction of the gland 5 is the same as the axial direction of the aircraft engine, and the gland 5 and the blade body 21 are locked in shape. The axis of the aircraft engine is perpendicular to the paper of figure 1, and in figure 1, the gland 5 can be mounted in the gland groove 6 or the gland 5 can be dismounted from the gland groove 6 by moving the gland 5 in a direction perpendicular to the paper of figure 1.

In order to further improve the combination between the gland 5 and the vane body 21, the size and shape of the gland 5 are matched with those of the gland groove 6, a plurality of meshing convex ribs 51 are arranged on two sides of the gland 5, the meshing convex ribs 51 are arranged along the axial direction of the aircraft engine, meshing grooves 61 matched with the meshing convex ribs 51 are arranged on two sides of the gland groove 6, and after assembly, the meshing convex ribs 51 are matched and inserted with the meshing grooves 61, as shown in fig. 1 to 5.

For the convenience of assembly and disassembly, the gland groove 6 penetrates through the outer end of the blade body 21 along the axis of the aircraft engine, as shown in fig. 3, during assembly and disassembly, only the gland 5 needs to slide into the gland groove 6 of the blade body 21 along the axial direction of the aircraft engine (i.e. the direction perpendicular to the paper surface of fig. 1), when the wheel disc 1 rotates to drive the blade to rotate, the gland 5 is tightly limited on the engine blade 2 under the centrifugal action, and the gland 5 is not easy to slide from one side end surface of the engine blade 2.

The gland 5 and the gland groove 6 can be mutually matched through the concave-convex structures such as the occlusion convex ribs 51 and the occlusion grooves 61, so that the connection between the gland 5 and the blade body 21 is more stable. The shape of the gland 5 is adapted to the shape of the engine blade 2, which is beneficial to the consistency of the two stresses.

In order to prevent the high-speed rotation process, the vibration reduction cavity 3 is communicated with the outside to cause airflow impact, the blade structure and the airflow compression function of the blade are damaged, and the gland 5 completely seals the open end (outer end) of the vibration reduction cavity 3 to prevent the vibration reduction cavity 3 from being communicated with the outside. Further, the blade body 21 and the gland 5 are arranged in a sealing mode, the sealing arrangement can adopt a graphite sealing mode, a graphite sealing layer (not shown in the figure) can be arranged on the surface of the gland groove 6 and is arranged along with the shape of the gland groove 6, the graphite sealing layer can enable the gland 5 and the blade to have buffering and prevent abrasion, the engine blade 2 and the gland 5 can be matched more tightly, the gland 5 and the blade also have buffering and prevent abrasion, and meanwhile, the air tightness is achieved, and air flow is prevented from being introduced into the vibration reduction cavity 3.

In the present embodiment, the fitting of the gland 5 to the vane body 21 is a tight fit. The top of the damping plug 4 is pushed against the gland 5 and fills the damping chamber 3 to improve damping performance and also prevent external air from entering the damping chamber 3 to cause impact.

In this embodiment, in order to make the whole stress of the damping cavity 3 uniform, the damping cavity 3 is arranged in the middle of the engine blade 2, and the contour shape of the damping cavity 3 changes along with the contour shape of the engine blade 2, so that the damping plugs 4 in the damping cavity 3 are uniformly stressed in all directions and have uniform buffering and absorbing effects on vibration energy, the uniform stress change of the blade structure is facilitated, and the service life of the blade is prolonged. That is, the sectional shape of the damper cavity 3 is reduced in an equal proportion to the sectional shape of the blade body 21, and the thickness of the side wall of the blade body 21 at the damper cavity 3 portion is uniform.

In the present embodiment, the inner surface of the damping cavity 3 is provided with a plurality of stop ribs 7 arranged along the diameter direction of the aircraft engine, the stop ribs 7 are arranged at intervals along the circumferential direction of the damping cavity 3, and the damping plug 4 is completely filled in the damping cavity 3, as shown in fig. 6. The stop convex strip 7 has a buckling effect on the vibration damping plug 4, can enhance the binding force between the vibration damping plug 4 and the vibration damping cavity 3 and prevent the vibration damping plug 4 from loosening, and is favorable for the connection and the combination of the vibration damping plug 4 and the vibration damping cavity 3. The cross section of the stop rib 7 may be arc-shaped, but may have other shapes.

In this embodiment, the damping plug 4 has the function of buffering and absorbing vibration energy, so as to buffer and absorb the vibration energy of the whole engine blade 2 and prevent the vibration of the blade in the high-speed rotation process from affecting the whole blade structure, and in the blade rotation process, the temperature of the blade is high, the damping plug 4 is made of a high-temperature-resistant flexible damping material, and preferably, the damping plug 4 is made of an existing aerogel material, preferably, an existing silicon aerogel or carbon aerogel; the aerogel has the advantages of low density, high temperature resistance, strong pressure resistance, porous structure and good heat insulation performance, and has the functions of sound absorption, noise reduction, vibration buffering, absorption and the like.

In this embodiment, aeroengine blade 2 still includes connecting portion 22, and connecting portion 22 is connected with the inner of blade body 21, and connecting portion 22 is connected with blade body 21 and forms the engine blade 2 of integral type, and connecting portion 22 is joggled with rim plate 1, and the installation direction of engine blade 2 is the same with this aeroengine's axis direction, and connecting portion 22 is equivalent to the tenon, and rim plate 1 is corresponding to be equipped with mortise or tongue-and-groove. In fig. 1, the movement of the engine blade 2 in a direction perpendicular to the paper of fig. 1 makes it possible to mount the attachment portion 22 of the engine blade 2 in the mortise or mortise of the disk 1 or to dismount the attachment portion 22 from the mortise or mortise of the disk 1.

In addition, in order to improve the vibration damping effect of the connecting part 22, the inner end of the vibration damping cavity 3 extends into the connecting part 22, as shown in fig. 1 and 2, when the rotary engine rotates, the motion vibration between the connecting part 22 and the wheel disc 1 absorbs a part of vibration energy through the vibration damping plug 4, the vibration problem of the wheel disc 1-blade during high-speed rotation is solved, the problem of blade fatigue is effectively prevented, and the service life of the engine blade 2 is prolonged.

An aircraft engine is described below, which includes a disk 1 and a plurality of engine blades, which are uniformly spaced along the circumferential direction of the disk 1, and the engine blades are the engine blades 2 described above, as shown in fig. 1.

Only a part of the structure of the wheel disc 1 is shown in fig. 1, the wheel disc 1 is circular, the circular ring corresponding to the wheel disc 1 is parallel to the paper surface of fig. 1, the engine blade 2 is connected with the wheel disc 1 through the connecting part 22 and realizes the rotation of the whole engine blade 2, in the embodiment, the connecting part 22 is of a joggle structure and is joggled with the wheel disc 1, and the wheel disc 1 is stably rotated under the centrifugal action.

When the engine blade 2 rotates to generate vibration, the engine blade is easy to generate heat, usually, the vibration is mostly concentrated on the central axial position of the blade, the heat of the natural blade is also positioned at the position, the vibration damping cavity 3 is arranged at the middle position of the blade, the vibration damping plug 4 is arranged in the vibration damping cavity, the vibration energy can be buffered and absorbed at the main vibration generating position, the vibration generation is reduced, the whole engine blade 2 is less influenced by the vibration, the vibration problem of the wheel disc 1-the blade during high-speed rotation is more effectively solved, the blade fatigue problem is prevented, and the service life of the engine blade 2 is prolonged. The aerogel is used as the material of the vibration damping plug 4, and is applied to the engine blade 2 due to the structural performance of the aerogel, the effect is obvious, the porous structure, the density and the heat conduction performance of the aerogel obviously reduce vibration and heat, and the vibration damping effect is better.

Example two

The present embodiment is an improvement of the first embodiment, and the main difference between the present embodiment and the first embodiment is that an intermediate rib 8 is added in the damping chamber 3, the damping chamber 3 is divided into a plurality of secondary damping chambers 31 by the intermediate rib 8, and the sections of the secondary damping chambers 31 are fan-shaped, as shown in fig. 7 and 8.

Through set up middle strengthening rib 8 in damping chamber 3, its middle strengthening rib 8 separates damping chamber 3 for a plurality of time damping chamber 31, and damping stopper 4 corresponds fills in a plurality of times damping chamber 31, and middle strengthening rib 8 strengthens the contact of cavity both sides blade structure, has the structure additional strengthening effect, prevents that the blade from warping, and damping stopper 4 is favorable to absorbing the buffering of vibrational energy more with the area of contact increase of engine blade 2 simultaneously.

In this embodiment, the intermediate reinforcing rib 8 is provided in the thickness direction of the damper chamber 3. The shape of the middle reinforcing rib 8 is not limited, and may be a linear type, a V-shaped, a W-shaped, and a plurality of middle reinforcing ribs 8 may be provided, as long as the vibration damping chamber 3 is divided and connected to the blade structures at other positions, preferably, the middle reinforcing rib 8 is a linear type, the bottom of the middle reinforcing rib is connected to the bottom of the vibration damping chamber 3, and five middle reinforcing ribs 8 are provided.

Other technical features of the present embodiment are the same as those of the first embodiment, and the present embodiment will not be described in detail for the sake of brevity.

EXAMPLE III

The present embodiment is an improvement of the first embodiment, and the main difference between the present embodiment and the first embodiment is that an intermediate reinforcing rib 8 is added in the damping chamber 3, the damping chamber 3 is divided into a plurality of secondary damping chambers 31 by the intermediate reinforcing rib 8, and the plurality of secondary damping chambers 31 form a honeycomb structure, that is, the section of the damping chamber 3 has a honeycomb structure, as shown in fig. 9 and 10. The cross section of the pore channel in the honeycomb structure can be regular hexagon or round.

The cross section of the secondary damping cavity 31 is honeycomb-shaped, so that the contact area can be increased, and the damping plugs 4 are filled in a plurality of honeycomb-shaped secondary damping cavities 31. The secondary damping cavity 31 is arranged to be honeycomb-shaped, so that the contact area between the damping plug 4 and the blade is obviously increased, the damping plug 4 has more remarkable effects on the damping absorption of the vibration energy of the blade and the like, the vibration heating of the engine blade 2 is greatly reduced, the vibration problem of the wheel disc 1-blade during high-speed rotation is solved, the blade fatigue problem is prevented, and the service life of the engine blade 2 is prolonged. In addition, the honeycomb structure strengthens the structural stability of the whole blade due to the cross connection between the middle reinforcing ribs 8, so that the blade is more impact-resistant.

The above description is only exemplary of the invention and should not be taken as limiting the scope of the invention, so that the invention is intended to cover all modifications and equivalents of the embodiments described herein. In addition, the technical features and the technical schemes, and the technical schemes can be freely combined and used.

Claims

1. The blade of the aero-engine is characterized in that the blade (2) of the aero-engine comprises a blade body (21), a vibration reduction cavity (3) arranged along the diameter direction of the aero-engine is arranged in the blade body (21), and a vibration reduction plug (4) capable of reducing vibration of the blade of the aero-engine is filled in the vibration reduction cavity (3);

the inner end of the vibration reduction cavity (3) is of a closed structure, the outer end of the vibration reduction cavity (3) is of an open structure, and the outer end of the vibration reduction cavity (3) is provided with a gland (5) capable of closing the vibration reduction cavity (3);

the gland (5) is inserted into the blade body (21), a gland groove (6) used for installing the gland (5) in a matched mode is formed in the outer end of the blade body (21), the gland groove (6) is formed in the axial direction of the aero-engine, the installation direction of the gland (5) is the same as the axial direction of the aero-engine, and the gland (5) and the blade body (21) are locked in a shape mode;

a plurality of occlusion ribs (51) are arranged on two sides of the gland (5), the occlusion ribs (51) are arranged along the axial direction of the aircraft engine, and occlusion grooves (61) matched with the occlusion ribs (51) are arranged on two sides of the gland groove (6);

the gland (5) is matched with the blade body (21) in a tight fit mode, the gland (5) is connected with the blade body (21) in a sealing mode, and a graphite sealing layer is arranged on the surface of the meshing groove (61);

the inner surface in the vibration reduction cavity (3) is provided with a plurality of stop convex strips (7) arranged along the diameter direction of the aircraft engine, the stop convex strips (7) are arranged at intervals along the circumferential direction of the vibration reduction cavity (3), and the vibration reduction plug (4) is completely filled in the vibration reduction cavity (3);

the damping cavity (3) is internally provided with a middle reinforcing rib (8), the damping cavity (3) is divided into a plurality of honeycomb-shaped secondary damping cavities (31) by the middle reinforcing rib (8), and a damping plug (4) is filled in each secondary damping cavity (31);

the aeroengine blade (2) further comprises a connecting part (22), the connecting part (22) is connected with the inner end of the blade body (21), and the inner end of the vibration reduction cavity (3) extends into the connecting part (22).

2. The aircraft engine blade according to claim 1, characterised in that the cross-sectional shape of the damping chamber (3) is an isometric reduction of the cross-sectional shape of the blade body (21), the side wall thickness of the blade body (21) being uniform in the region of the damping chamber (3).

3. An aircraft engine, characterized in that it comprises a wheel disc (1) and a plurality of engine blades, spaced along the circumferential direction of the wheel disc (1), said engine blades being the engine blades (2) according to claim 1.

4. An aircraft engine according to claim 3, characterised in that the attachment portion (22) is a dovetail connection with the wheel disc (1), the mounting direction of the engine blade (2) being the same as the axial direction of the aircraft engine.