CN110206646B - Aeroengine bearing support subassembly and aeroengine - Google Patents

Aeroengine bearing support subassembly and aeroengine Download PDF

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Publication number
CN110206646B
CN110206646B CN201810166235.6A CN201810166235A CN110206646B CN 110206646 B CN110206646 B CN 110206646B CN 201810166235 A CN201810166235 A CN 201810166235A CN 110206646 B CN110206646 B CN 110206646B
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ring
bearing
aircraft engine
limiting ring
annular
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CN110206646A (en
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赵芝梅
宋会英
郑李鹏
唐振南
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AECC Commercial Aircraft Engine Co Ltd
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AECC Commercial Aircraft Engine Co Ltd
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    • FMECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02CGAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
    • F02C7/00Features, components parts, details or accessories, not provided for in, or of interest apart form groups F02C1/00 - F02C6/00; Air intakes for jet-propulsion plants
    • F02C7/06Arrangements of bearings; Lubricating

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  • Engineering & Computer Science (AREA)
  • Chemical & Material Sciences (AREA)
  • Combustion & Propulsion (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Shafts, Cranks, Connecting Bars, And Related Bearings (AREA)
  • Support Of The Bearing (AREA)

Abstract

The invention aims to provide an aeroengine bearing support assembly and an aeroengine. The invention provides an aeroengine bearing support assembly, which comprises a mounting seat, a bearing seat and a bearing seat, wherein the mounting seat is provided with a mounting surface which is a part of a spherical surface; the aeroengine bearing support assembly further comprises a positioning ring, a connecting piece, a first limiting ring and a second limiting ring; the first limiting ring and the second limiting ring are respectively contacted with the spherical surface of the mounting surface and are used for respectively supporting the bearing; the positioning ring is fixedly arranged on the mounting surface; the connecting piece is used for fixedly connecting the first limiting ring, the second limiting ring and the positioning ring into a whole. The positioning ring is used for bearing overload force to separate from the mounting surface. The positioning ring is clamped tightly, so that the positioning ring cannot fly off, and the safety is high. The aero-engine provided by the invention comprises the aero-engine bearing support assembly, so that the aero-engine bearing support assembly has high safety.

Description

Aeroengine bearing support subassembly and aeroengine
Technical Field
The invention relates to an aeroengine bearing assembly and an aeroengine.
Background
According to the requirements of airworthiness regulations (FAR33.74, FAR33.94), commercial aircraft engines must ensure that the occurrence of an FBO event (fan blade fly-out) does not result in catastrophic consequences. In order to ensure the safety of the aircraft engine after an FBO event occurs, the conventional method is to increase the structural strength of each component on the force transmission path of the engine, such as a rotor support structure, a bearing casing, an installation system and the like, and increase the stress reserve margin of the engine so as to ensure that key components cannot be damaged under the FBO load. However, this results in an increased weight of the engine, which is disadvantageous for weight-saving design.
In recent years, as engine thrust increases, fan blade diameters have increased, and imbalance loads from FBO events have also increased. To ensure that FBO events do not have catastrophic consequences without significantly increasing engine weight, the concept of a fused design is applied to engine design. A common fuse design is to place a primary fuse component, typically a mechanically weak structure such as a thinned section, necked-down bolt, etc., near bearing No. 1, which is closest to the fan, and which can fail under a predetermined load (fuse threshold). After the FBO event occurs, the first fusing component near the No. 1 bearing fails, the fan rotor and stator supporting structures such as an intermediate casing are decoupled, unbalanced loads transmitted to the intermediate casing through the bearing, a support and the like are reduced, and other key components of the engine are prevented from being damaged.
However, after the failure of the primary fuse element, the constraint of the fan shaft at bearing number 1 is released and only at bearing number 2, the swing radius is larger than before the failure of the primary fuse element. This can lead to local bending deformation and stress concentration of the fan shaft at the bearing No. 2 due to the constraint, and can even lead to the breakage of the fan shaft and the flying-off of the fan rotor, thus endangering the safety of the airplane. Therefore, a secondary fusing design is required to be performed near the No. 2 bearing so as to avoid excessive stress concentration generated by the fan shaft and ensure the safety of the engine.
As shown in fig. 5, US6491497B1 discloses an assembly for supporting a rotor in case of unbalance, in which there is an extension shaft, a mounting race 350 protruding from the tip of the extension shaft, which cooperates with a concave surface 364 of an inner race of a bearing, and cages 354, 355 located on both sides of the mounting race 350 keep the concave surface 364 of the inner race of the bearing stationary relative to the mounting race 350 in a normal operating state. When a shift in the center of gravity of the rotor occurs beyond a set limit, one of the cages 354, 355 fractures, allowing relative sliding between the spherical surface of the mount race 350 and the concave surface 364. Although the retainers 354, 355 are fused by being broken, the portions broken off from the retainers 354, 355 still fly off, causing secondary damage to the engine.
Disclosure of Invention
The invention aims to provide an aeroengine bearing support assembly which cannot cause secondary damage to an engine when fused.
It is also an object of the present invention to provide an aircraft engine including an aircraft engine bearing support assembly as described above, and therefore having a high degree of safety.
To achieve the object, an aircraft engine bearing support assembly for supporting a bearing comprises a mounting block having a mounting face; the mounting surface is a part of a spherical surface;
the aeroengine bearing support assembly further comprises a positioning ring, a connecting piece, a first limiting ring and a second limiting ring;
the first limiting ring and the second limiting ring are in spherical contact with the mounting surface respectively and are used for supporting the inner ring of the bearing respectively; the positioning ring is connected with the mounting surface and is positioned between the first limiting ring and the second limiting ring; the connecting piece fixedly connects the first limiting ring, the second limiting ring and the positioning ring into a whole;
the positioning ring is arranged to be separated from the mounting surface under the condition of bearing overload force, and allows the first limiting ring, the second limiting ring and the positioning ring which are fixedly connected into a whole by the connecting piece to slide on the mounting surface.
The aircraft engine bearing support assembly is further characterized in that the connecting piece is a long bolt, one end of the long bolt penetrates through one of the first limiting ring and the second limiting ring, then penetrates through the positioning ring, and then is connected with the other one of the first limiting ring and the second limiting ring, so that the first limiting ring, the second limiting ring and the positioning ring are fixedly connected into a whole.
The aero-engine bearing support assembly is further characterized in that a first contact surface is arranged on the radial inner side of the first limiting ring, a second contact surface is arranged on the radial inner side of the second limiting ring, and the first contact surface and the second contact surface are respectively in spherical contact with the mounting surface;
the radial outer side of the first limiting ring and the radial outer side of the second limiting ring are respectively arranged to axially abut against two axial end faces of the inner ring of the bearing so as to axially limit the bearing, and radially abut against two convex shoulders of the inner ring of the bearing so as to radially limit the bearing.
The aero-engine bearing support assembly further characterized in that a radially outer side of the first retainer ring has a first annular face and a first annular flange projecting radially outwardly from the first annular face; the outer side of the second spacing ring is provided with a second annular surface and a second annular flange, and the second annular flange is protruded outwards along the radial direction from the second annular surface;
the first annular flange and the second annular flange are arranged at intervals in the axial direction and are respectively abutted against two end faces of the inner ring of the bearing in the axial direction and two convex shoulders of the bearing in the radial direction, so that the inner ring of the bearing is supported, and the inner ring, the first annular flange, the second annular flange, the first annular surface and the second annular surface jointly limit to form a closed annular cavity.
The aeroengine bearing support assembly is further characterized in that the aeroengine bearing support assembly also comprises a damping energy-absorbing body made of damping energy-absorbing materials; the annular cavity is filled with the damping energy-absorbing material so as to form the damping energy-absorbing body in the annular cavity.
The aeroengine bearing support assembly is further characterized in that the damping energy-absorbing material is metal rubber or foamed aluminum.
To achieve the object, an aircraft engine comprises a bearing, the bearing comprises an inner ring, a roller and an outer ring, the aircraft engine further comprises an aircraft engine bearing support assembly as described above, the aircraft engine bearing support assembly comprises a mounting seat, and the mounting seat is provided with a mounting surface; the mounting surface is a part of a spherical surface;
the aeroengine bearing support assembly further comprises a positioning ring, a connecting piece, a first limiting ring and a second limiting ring;
the first limiting ring and the second limiting ring are in spherical contact with the mounting surface respectively and are used for supporting the inner ring of the bearing respectively; the positioning ring is connected with the mounting surface and is positioned between the first limiting ring and the second limiting ring; the connecting piece is used for fixedly connecting the first limiting ring, the second limiting ring and the positioning ring into a whole;
the positioning ring is arranged to be separated from the mounting surface under the condition of bearing overload force, so that the first limiting ring, the second limiting ring and the positioning ring which are fixedly connected into a whole by the connecting piece can slide on the mounting surface.
The aircraft engine is further characterized in that the aircraft engine further comprises an intermediate casing and a fan shaft;
the outer ring is fixedly arranged on the intermediate casing; the mounting seat is fixedly mounted on the fan shaft and is positioned on the inner side of the inner ring.
The aircraft engine is further characterized in that the bearing is a rear support bearing of the fan shaft of the aircraft engine.
The aircraft engine is further characterized in that the inner ring is provided with two convex shoulders along the axial direction, and the two convex shoulders are respectively arranged on two end surfaces of the inner ring in the axial direction in a protruding mode.
The positive progress effects of the invention are as follows: the invention provides an aeroengine bearing support assembly, which is used for supporting a bearing and comprises a mounting seat, a bearing seat and a bearing seat, wherein the mounting seat is provided with a mounting surface; the mounting surface is a part of a spherical surface; the aeroengine bearing support assembly further comprises a positioning ring, a connecting piece, a first limiting ring and a second limiting ring; the first limiting ring and the second limiting ring are in spherical contact with the mounting surface respectively and are used for supporting the bearing respectively; the positioning ring is fixedly arranged on the mounting surface and is positioned between the first limiting ring and the second limiting ring; the connecting piece is used for fixedly connecting the first limiting ring, the second limiting ring and the positioning ring into a whole. The positioning ring is used for bearing overload force and separating from the mounting surface, so that the first limiting ring, the second limiting ring and the positioning ring which are fixedly connected into a whole by the connecting piece can slide on the mounting surface.
Because the positioning ring is still clamped by the first limiting ring and the second limiting ring after being disconnected with the mounting surface, the positioning ring cannot fly off, and secondary damage to the engine cannot be caused. The aero-engine provided by the invention comprises the aero-engine bearing support assembly, so that the aero-engine bearing support assembly has high safety.
Drawings
The above and other features, properties and advantages of the present invention will become more apparent from the following description of the embodiments with reference to the accompanying drawings, in which:
FIG. 1 is a cross-sectional view of an aircraft engine according to the present invention;
FIG. 2 is an enlarged view taken at B of FIG. 1 showing the bearing and an aircraft engine bearing support assembly;
FIG. 3 is an enlarged view at C of FIG. 2;
FIG. 4 is a schematic view of the first stop collar, the second stop collar and the positioning ring sliding on the mounting surface;
FIG. 5 is a schematic view of a prior art assembly for supporting a rotor in case of imbalance.
Detailed Description
The present invention is further described in the following description with reference to specific embodiments and the accompanying drawings, wherein the details are set forth in order to provide a thorough understanding of the present invention, but it is apparent that the present invention can be embodied in many other forms different from those described herein, and it will be readily appreciated by those skilled in the art that the present invention can be implemented in many different forms without departing from the spirit and scope of the invention.
In the description of the present invention, it is to be understood that the terms "length", "width", "upper", "lower", "front", "rear", "left", "right", "vertical", "horizontal", "top", "bottom", "inner" and "outer" etc. indicate orientations or positional relationships based on those shown in the drawings, and are only for convenience of description and simplicity of description, and do not indicate or imply that the referenced devices or elements must have a particular orientation, be constructed and operated in a particular orientation, and thus, are not to be construed as limiting the present invention.
Furthermore, the terms "first", "second" and "first" are used for descriptive purposes only and are not to be construed as indicating or implying relative importance or implicitly indicating the number of technical features indicated. Thus, a feature defined as "first" or "second" may explicitly or implicitly include one or more of that feature. In the description of the present invention, "a plurality" means two or more unless specifically defined otherwise.
In the present invention, unless otherwise expressly stated or limited, the terms "mounted," "connected," "secured," and the like are to be construed broadly and can, for example, be fixedly connected, detachably connected, or integrally formed; can be mechanically or electrically connected; either directly or indirectly through intervening media, either internally or in any other relationship. The specific meanings of the above terms in the present invention can be understood by those skilled in the art according to specific situations.
It should be noted that fig. 1-5 are exemplary only, are not drawn to scale, and should not be construed as limiting the scope of the invention as actually claimed. The terms "axial," "radial," and circumferential in the present disclosure are used with reference to the orientation of aircraft engine 100.
FIG. 1 illustrates an embodiment of an aircraft engine 100 of the present invention, aircraft engine 100 having a central axis A-A, and only one half of a cross-sectional view of aircraft engine 100 is shown, with portions not shown being entirely symmetrical with portions shown about central axis A-A. The axis surrounded by the "ring" in the present invention is the central axis a-a.
With continued reference to fig. 1, aircraft engine 100 includes a fan rotor 3, an intermediate case 4, bearings 1 and aircraft engine bearing support assembly 2, a fan shaft 5 and an extension shaft 6. Bearing 1 is in the embodiment shown in fig. 1 a bearing number 2 of an aircraft engine 100. In further embodiments, the aero engine bearing support assembly 2 may also be used to support bearings in other locations of an aero engine, such as bearing number 1 7.
Embodiments of the present invention will now be described in detail, taking as an example the application of an aircraft engine bearing support assembly 2 to bearing number 2.
In this embodiment, the bearing 1, i.e. the No. 2 bearing, comprises an inner ring 10, rollers 11 and an outer ring 12, the outer ring 12 being connected to the intermediate housing 4, the inner ring 10 being fusably connected to the extension shaft 6 of the fan shaft 5 via the aero engine bearing support assembly 2, i.e. the inner ring 10 is supported by the aero engine bearing support assembly 2. The front end of the fan shaft 5 is supported by a bearing 7 No. 1, a supporting conical wall 8 connects the bearing 7 No. 1 to the intermediate casing 4, and the supporting conical wall 8 is an important path for transferring the load of the fan rotor to the intermediate casing. In other embodiments, the aero engine bearing support assembly 2 may support other portions of the bearing 1, such as the outer race 12, instead of the inner race 10.
In the embodiment shown in fig. 1, the fan shaft 5 is supported by both bearing No. 1 and bearing No. 2, wherein bearing No. 1 is a rolling rod bearing, providing radial constraint to the fan shaft 5; the bearing number 2 is a ball bearing that provides both axial and radial restraint to the extension shaft 6 of the fan shaft 5.
In order to reduce the load transmitted from the fan rotor 3 to the intermediate casing 4 after an FBO event, a primary fuse element 9, which may be embodied as a thinned section, is provided on the supporting cone wall 8 of the bearing No. 1 bearing 7. The extension shaft 6 of the fan shaft 5 is connected to bearing No. 2 through the aero engine bearing assembly 2 and further to the intermediate case 4. The fusing design at the No. 1 bearing 7 is called a primary fusing design, in order to further reduce the stress of a fan shaft and improve the safety, the fusing design at the No. 2 bearing is also required, and the fusing design at the No. 2 bearing is called a secondary fusing design.
The design of the fuse at bearing number 2 may be achieved by the specific construction of the aircraft engine bearing support assembly 2.
Referring to fig. 1, 2 and 3, an aircraft engine bearing support assembly 2 includes a mount 20, the mount 20 having a mounting surface 20 a. The mounting surface 20a is located inside the inner ring 10 of the bearing 1 in order to support the inner ring 10 of the bearing 1. The mounting surface 20a is designed concentrically with the inner ring 10 of the bearing 1 and is disposed opposite to the inner surface of the inner ring 10 in the radial direction. The mounting surface 20a is a portion of a spherical surface, which may be a convex spherical surface or a concave spherical surface. The mounting seat 20 is annular, and correspondingly the mounting surface 20a is also annular. The mounting 20 is concentric with the bearing 1 about a central axis a-a.
In order to support the inner ring 10, the aeroengine bearing support assembly 2 further comprises a positioning ring 21, a connecting piece 22, a first limit ring 23 and a second limit ring 24; the first limiting ring 23 and the second limiting ring 24 are respectively sleeved on the mounting surface 20a and are in spherical contact with the mounting surface 20a, and respectively support the inner ring 10 of the bearing 1; in addition, the first retainer ring 23 and the second retainer ring 24 can be in spherical hinge connection with the mounting surface 20 a. The positioning ring 21, the first stop collar 23 and the second stop collar 24 are all arranged around the central axis a-a.
The positioning ring 21 is connected with the mounting surface 20a and is positioned between the first limiting ring 23 and the second limiting ring 24; the connecting piece 22 is used for fixedly connecting the first spacing ring 23, the second spacing ring 24 and the positioning ring 21 into a whole, and enabling the first spacing ring 23 and the second spacing ring 24 to clamp the positioning ring 21. The form of the connector 22 is numerous and will be discussed later.
The joint of the positioning ring 21 and the mounting surface 20a is designed to be fused, for example, a thinned section may be used for connection, so as to ensure that the positioning ring 21 is separated from the mounting surface 20a under the condition of bearing an overload force, and allow the first stop collar 23, the second stop collar 24 and the positioning ring 21 which are fixedly connected into a whole by the connecting member 22 to slide on the mounting surface 20a, thereby achieving the purpose of fused design.
In a more specific embodiment, the retaining ring 21 has a connection surface that is a portion of a spherical surface that is fused (which may be by an adhesive) to the mounting surface 20 a. The attachment surface is capable of sliding on the attachment surface 20a when the attachment surface is disconnected from the attachment surface 20a (i.e., the adhesive fails).
In another embodiment, the positioning ring 21 may be integrally manufactured with the mounting seat 20 such that the positioning ring 21 protrudes outward in the radial direction from the mounting surface 20a, and in this embodiment, a crack or a groove for fusing may be provided at the intersection of the positioning ring 21 and the mounting surface 20a along the arc of the mounting surface 20a, so that when the positioning ring 21 is disconnected from the mounting surface 20a, the positioning ring 21 forms an arc-shaped cross section, and the cross section can also slide on the mounting surface 20 a.
As shown in fig. 4, due to the connecting member 22, the positioning ring 21 is still clamped by the first stop collar 23 and the second stop collar 24 after being disconnected from the mounting surface 20a, so that the positioning ring 21 does not fly off, and secondary damage to the engine is avoided.
In one embodiment, the connecting member 22 may be a long bolt, and one end of the connecting member 22 is configured to pass through the positioning ring 21 after passing through one of the first stop collar 23 and the second stop collar 24, and then to be connected with the other one of the first stop collar 23 and the second stop collar 24, so as to fixedly connect the first stop collar 23, the second stop collar 24 and the positioning ring 21 into a whole. Correspondingly, one of the first limiting ring 23 and the second limiting ring 24 is provided with a through hole, and the other is provided with a connecting hole; the positioning ring 21 is provided with a through hole, and the connecting piece 22 penetrates through the through hole, so that the positioning ring 21 is always sleeved on the connecting piece 22, and the positioning ring 21 is further prevented from flying off after being broken.
The specific structure of the first retainer ring 23 and the second retainer ring 24 will be discussed below.
With continued reference to fig. 3 and 4, the radially inner side of the first retainer ring 23 has a first contact surface 23a, the radially inner side of the second retainer ring 24 has a second contact surface 24a, and the first contact surface 23a and the second contact surface 24a are configured to be in spherical contact with the mounting surface 20a, respectively; the radially outer side of the first retainer ring 23 and the radially outer side of the second retainer ring 24 are arranged to respectively abut against both end faces of the inner ring 10 of the bearing 1 in the axial direction to axially retain the bearing 1, and to respectively abut against both convex shoulders 10a of the inner ring 10 of the bearing 1 in the radial direction to radially retain the bearing 1. The first contact surface 23a and the second contact surface 24a are portions of spherical surfaces, and may be concave spherical surfaces or convex spherical surfaces.
The radially outer side of the first retainer ring 23 has a first annular face 23b and a first annular flange 23c, the first annular flange 23c projecting radially outward from the first annular face 23 b; the outer side of the second stop collar 24 has a second annular surface 24b and a second annular flange 24c, the second annular flange 24c projecting radially outwardly from the second annular surface 24 b;
the first annular flange 23c and the second annular flange 24c are arranged at intervals in the axial direction and are arranged to respectively abut against two end faces of the inner ring 10 of the bearing 1 in the axial direction and two shoulders 10a of the bearing 1 in the radial direction, so that the inner ring 10 of the bearing 1 is supported, and the inner ring 10, the first annular flange 23c, the second annular flange 24c, the first annular surface 23b and the second annular surface 24b jointly define a closed annular cavity 2 a.
In order to further relieve the radial load in an unbalanced state, the aeroengine bearing support assembly 2 further comprises a damping energy-absorbing body 25 made of a damping energy-absorbing material; the annular cavity 2a is filled with damping and energy-absorbing material to form a damping and energy-absorbing body 25 in the annular cavity 2 a. The damping energy-absorbing material can be metal rubber or foamed aluminum.
As shown in fig. 4, after an FBO event occurs, the fan rotor 3 generates a huge unbalanced load, the positioning ring 21 fails at the mounting surface 20a in a preset manner, so that the first and second limiting rings 23 and 24 form a spherical hinge connection with the mounting surface 20a, the pitching stiffness of the fan shaft at the bearing No. 2 is released, and the situation that the swing radius of the fan rotor 3 is increased and the fan shaft 5 generates an excessive stress concentration at the bearing No. 2 after the primary fusing component at the bearing No. 1 fails is avoided. Meanwhile, the shoulder 10a on the inner ring 10 is broken, so that the inner ring 10 and the first and second spacing rings 23 and 24 can generate relative displacement in a limited range along the radial direction, and the damping energy absorber 25 is squeezed, thereby further consuming the energy of the FBO and reducing the load transmitted to key components such as the intermediate casing 4 and the like.
Although the present invention has been described with reference to the preferred embodiments, it is not intended to limit the present invention, and those skilled in the art can make modifications and variations without departing from the spirit and scope of the present invention.

Claims (10)

1. An aircraft engine bearing support assembly for supporting a bearing (1), comprising a mount (20), the mount (20) having a mounting face (20 a); the mounting surface (20a) is a part of a spherical surface;
the aeroengine bearing support assembly (2) is characterized by further comprising a positioning ring (21), a connecting piece (22), a first limiting ring (23) and a second limiting ring (24);
the first limiting ring (23) and the second limiting ring (24) are in spherical contact with the mounting surface (20a) respectively and are used for supporting the inner ring (10) of the bearing (1) respectively; the positioning ring (21) is connected with the mounting surface (20a) and is positioned between the first limiting ring (23) and the second limiting ring (24); the connecting piece (22) fixedly connects the first spacing ring (23), the second spacing ring (24) and the positioning ring (21) into a whole;
the positioning ring (21) is arranged to be separated from the mounting surface (20a) under the condition of bearing overload force, and allows the first limiting ring (23), the second limiting ring (24) and the positioning ring (21) which are fixedly connected into a whole by the connecting piece (22) to slide on the mounting surface (20 a).
2. The aircraft engine bearing assembly according to claim 1, wherein the connecting member (22) is a long bolt, one end of which is arranged to pass through the positioning ring (21) after passing through one of the first retainer ring (23) and the second retainer ring (24) and then to be connected with the other one of the first retainer ring (23) and the second retainer ring (24), thereby fixedly connecting the first retainer ring (23), the second retainer ring (24) and the positioning ring (21) into a whole.
3. The aircraft engine bearing assembly according to claim 1, wherein a radially inner side of the first retainer ring (23) has a first contact surface (23a), and a radially inner side of the second retainer ring (24) has a second contact surface (24a), the first contact surface (23a) and the second contact surface (24a) being for spherical contact with the mounting surface (20a), respectively;
the radial outer side of the first limiting ring (23) and the radial outer side of the second limiting ring (24) are respectively arranged to axially abut against two axial end faces of the inner ring (10) of the bearing (1) so as to axially limit the bearing (1), and radially abut against two convex shoulders (10a) of the inner ring (10) of the bearing (1) so as to radially limit the bearing (1).
4. The aircraft engine bearing assembly as defined in claim 3, wherein a radially outer side of the first retainer ring (23) has a first annular face (23b) and a first annular flange (23c), the first annular flange (23c) projecting radially outwardly from the first annular face (23 b); the outer side of the second limiting ring (24) is provided with a second annular surface (24b) and a second annular flange (24c), and the second annular flange (24c) protrudes outwards from the second annular surface (24b) in the radial direction;
the first annular flange (23c) and the second annular flange (24c) are arranged at intervals in the axial direction and are arranged to respectively abut against two end faces of the inner ring (10) of the bearing (1) in the axial direction and abut against two shoulders (10a) of the bearing (1) in the radial direction, so that the inner ring (10) of the bearing (1) is supported, and the inner ring (10), the first annular flange (23c), the second annular flange (24c), the first annular surface (23b) and the second annular surface (24b) jointly define a closed annular cavity (2 a).
5. The aeroengine bearing support assembly according to claim 4, wherein said aeroengine bearing support assembly (2) further comprises an energy damping body (25) comprised of an energy damping material; the annular cavity (2a) is filled with the damping energy-absorbing material, so that the damping energy-absorbing body (25) is formed in the annular cavity (2 a).
6. The aero engine bearing assembly according to claim 5 wherein said damping energy absorbing material is metal rubber or foamed aluminum.
7. An aircraft engine comprising a bearing (1), the bearing (1) comprising an inner race (10), rollers (11) and an outer race (12), characterised in that the aircraft engine (100) further comprises an aircraft engine bearing assembly (2) according to any one of claims 1 to 6.
8. An aircraft engine according to claim 7, characterised in that the aircraft engine (100) further comprises an intermediate casing (4) and a fan shaft (5);
the outer ring (12) is fixedly arranged on the intermediate casing (4); the mounting seat (20) is fixedly mounted on the fan shaft (5) and is positioned on the inner side of the inner ring (10).
9. An aircraft engine according to claim 8, characterised in that the bearing (1) is a rear support bearing for the fan shaft (5) of the aircraft engine.
10. An aircraft engine according to claim 7, characterized in that the inner ring (10) has two shoulders (10a) in the axial direction, and the two shoulders (10a) are respectively provided in a protruding manner on both end faces of the inner ring (10) in the axial direction.
CN201810166235.6A 2018-02-28 2018-02-28 Aeroengine bearing support subassembly and aeroengine Active CN110206646B (en)

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CN112498708B (en) * 2020-06-01 2022-02-08 重庆宗申航空发动机制造有限公司 Aviation unmanned aerial vehicle and aeroengine installing support
CN112983651B (en) * 2021-04-26 2023-07-28 黄锴 Small aviation double-rotor unmanned aerial vehicle engine

Citations (6)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP1900910A2 (en) * 2006-09-08 2008-03-19 Pratt & Whitney Canada Corp. Thrust bearing housing for a gas turbine engine
EP1596038A3 (en) * 2004-05-12 2009-01-14 United Technologies Corporation Gas turbine engine bearing support
CN103119254A (en) * 2010-10-13 2013-05-22 斯奈克玛 Connecting module between a drive shaft of an engine fan and a rolling element bearing
CN105510044A (en) * 2015-12-31 2016-04-20 苏州东菱科技有限公司 High-speed rotor blade flying-off test device and test method
CN105822366A (en) * 2016-04-20 2016-08-03 中国科学院工程热物理研究所 Low-pressure rotor supporting structure of engine with fusing design
CN107237655A (en) * 2016-03-28 2017-10-10 中国航发商用航空发动机有限责任公司 Aero-engine and its fan blade fly off blowout method under load

Family Cites Families (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US9777592B2 (en) * 2013-12-23 2017-10-03 Pratt & Whitney Canada Corp. Post FBO windmilling bumper

Patent Citations (6)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP1596038A3 (en) * 2004-05-12 2009-01-14 United Technologies Corporation Gas turbine engine bearing support
EP1900910A2 (en) * 2006-09-08 2008-03-19 Pratt & Whitney Canada Corp. Thrust bearing housing for a gas turbine engine
CN103119254A (en) * 2010-10-13 2013-05-22 斯奈克玛 Connecting module between a drive shaft of an engine fan and a rolling element bearing
CN105510044A (en) * 2015-12-31 2016-04-20 苏州东菱科技有限公司 High-speed rotor blade flying-off test device and test method
CN107237655A (en) * 2016-03-28 2017-10-10 中国航发商用航空发动机有限责任公司 Aero-engine and its fan blade fly off blowout method under load
CN105822366A (en) * 2016-04-20 2016-08-03 中国科学院工程热物理研究所 Low-pressure rotor supporting structure of engine with fusing design

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