Aircraft Engines

Abstract

The present invention discloses an aircraft engine, comprising a turbine, wherein the turbine comprises a rotor and a stator, and the aircraft engine further comprises a shaft failure speed limiting structure arranged between the rotor and the stator, wherein the shaft failure speed limiting structure comprises a first speed limiting part fixedly connected to the rotor and a second speed limiting part fixedly connected to the stator, and the shaft failure speed limiting structure is configured as follows: when the turbine is working normally, a gap is maintained between the first speed limiting part and the second speed limiting part; when a turbine shaft failure condition occurs in the turbine and the moving blades of the turbine move axially toward the downstream in the direction of gas flow, the first speed limiting part and the second speed limiting part approach each other and form an interference fit that becomes tighter the closer they are.

Description

Technical Field

[0001] The present invention relates to the field of aviation machinery, and in particular to an aviation engine. Background Art

[0002] In actual operation, turbine-driven engines may fail due to overtorque, resonance, fatigue, corrosion, material defects, manufacturing errors or other indirect events. Although the probability of failure is very small, once it occurs, it may lead to harmful consequences. For example, in aircraft engines, when the turbine shaft fails, the turbine rotor is decoupled from the front load (compressor), and driven by the high-energy gas discharged from the combustion chamber, the speed rises instantly or enters an overspeed state. When the speed rises to a certain level, the wheel stress reaches a critical level and breaks. The high-energy fragments of the rupture have the risk of penetrating the engine. Therefore, limiting the turbine overspeed after the turbine shaft fails is a very important consideration in the design of turbine engines.

[0003] In known aircraft engines, the rotor speed is usually directly monitored or converted by installing a speed sensor. The speed sensor is usually installed at the front end of the engine, which makes it difficult to monitor the increase in the rear turbine speed caused by the failure of the turbine shaft. Even if a sensor is added to the turbine end, for large civil aircraft engines, the control system takes a long time from monitoring and determining the occurrence of a turbine shaft failure to responding to oil cutting, and the response is slow and not timely.

[0004] It is known that adding honeycombs or friction-resistant devices to the low-pressure turbine stator structure (mainly the guide blades and the low-pressure turbine rear bearing casing), or designing axially swept low-pressure turbine guide blades, can limit the rotor speed by causing the rotor to move backward and collide with the stator or get stuck after a shaft failure occurs. However, after the low-pressure turbine rotor moves backward and collides with the stator structure, a collision force opposite to the axial force is generated, and the rotor may rebound, and the friction brake cannot be continued, further affecting the low-pressure turbine rotor speed limiting effect after the low-pressure shaft fails. Summary of the invention

[0005] The object of the present invention is to provide an aircraft engine which can respond quickly and effectively limit the increase of turbine rotation speed when a turbine shaft failure occurs.

[0006] The present invention discloses an aircraft engine, comprising a turbine, wherein the turbine comprises a rotor and a stator, and the aircraft engine further comprises a shaft failure speed limiting structure arranged between the rotor and the stator, wherein the shaft failure speed limiting structure comprises a first speed limiting part fixedly connected to the rotor and a second speed limiting part fixedly connected to the stator, and the shaft failure speed limiting structure is configured as follows: when the turbine is working normally, a gap is maintained between the first speed limiting part and the second speed limiting part; when a turbine shaft failure condition occurs in the turbine and the moving blades of the turbine move axially toward the downstream in the direction of gas flow, the first speed limiting part and the second speed limiting part approach each other and form an interference fit that becomes tighter the closer they are.

[0007] In some embodiments,

[0008] One of the first speed limiting part and the second speed limiting part comprises an annular plate, and the other comprises an annular groove corresponding to the annular plate, the annular plate, the annular groove and the turbine are coaxial, and when the turbine shaft fails and the turbine moves axially downstream in the direction of gas flow, the annular plate is inserted into the annular groove to cooperate with the annular groove to form the interference fit; and / or

[0009] One of the first speed limiting part and the second speed limiting part includes a speed limiting shaft, and the other includes a speed limiting hole corresponding to the annular plate. The speed limiting shaft, the speed limiting hole and the turbine are coaxial. When a turbine shaft failure occurs in the turbine and the turbine moves axially downstream in the direction of gas flow, the speed limiting shaft is inserted into the speed limiting hole to cooperate with the speed limiting hole to form the interference fit.

[0010] In some embodiments, the inlet end of the annular groove is gradually expanded, and the gradually expanded inlet end is used to guide the insertion of the annular plate, or the inlet end of the speed limiting hole is gradually expanded, and the gradually expanded inlet end is used to guide the insertion of the speed limiting shaft.

[0011] In some embodiments, the turbine is a low pressure turbine.

[0012] In some embodiments,

[0013] The rotor includes a low-pressure turbine blade, the stator includes a low-pressure turbine guide vane, the first speed limiter is provided at a trailing edge of the low-pressure turbine blade, and the second speed limiter is provided at a leading edge of the low-pressure turbine guide vane adjacent to the low-pressure turbine blade and located downstream of the low-pressure turbine blade; or

[0014] The rotor comprises a low-pressure turbine support cone wall for connecting the low-pressure turbine rotor blades and the low-pressure turbine shaft, the stator comprises a stator sealing ring having one end fixedly connected to the turbine interstage load-bearing casing (142) of the aircraft engine, the low-pressure turbine support cone wall is provided with a sealing portion for sealing with the stator sealing ring, the first speed limiting portion is provided on the sealing portion, and the second speed limiting portion is provided on the stator sealing ring; or

[0015] The rotor includes the end of the low-pressure turbine shaft, the stator includes a bearing seat connected to the rear bearing casing of the aircraft engine and connected to the end of the low-pressure turbine shaft through a bearing, the first speed limiting part is arranged on the end of the low-pressure turbine shaft, and the second speed limiting part is arranged on the bearing seat.

[0016] Based on the aircraft engine provided by the present invention, by setting a shaft failure speed limiting structure including a first speed limiting part and a second speed limiting part, when the turbine shaft fails, the moving blades of the turbine move axially downstream in the direction of the gas flow, the first speed limiting part and the second speed limiting part approach each other, and form an interference fit that is closer and tighter, and the friction force brought by the interference fit of the first speed limiting part and the second speed limiting part can limit the separation of the first speed limiting part and the second speed limiting part when the turbine shaft fails, and quickly reduce the speed of the turbine rotor when the turbine shaft fails. Since the first speed limiting part and the second speed limiting part can form an interference fit that is closer and tighter, when the turbine shaft fails, the rotor of the turbine moves downstream under the action of the gas and collides with the stator of the turbine, the greater the collision force, the tighter the interference fit between the first speed limiting part and the second speed limiting part, which can effectively prevent the turbine rotor from rebounding.

[0017] Further features and advantages of the present invention will become apparent from the following detailed description of exemplary embodiments of the present invention with reference to the attached drawings. BRIEF DESCRIPTION OF THE DRAWINGS

[0018] The drawings described herein are used to provide a further understanding of the present invention and constitute a part of this application. The exemplary embodiments of the present invention and their descriptions are used to explain the present invention and do not constitute an improper limitation of the present invention. In the drawings:

[0019] Figure 1 A schematic diagram of the structure of an aircraft engine according to an embodiment of the present invention;

[0020] Figure 2 for Figure 1 The schematic diagram of the structure of the shaft failure speed limiting structure of the aircraft engine shown;

[0021] Figure 3 A schematic cross-sectional view of a shaft failure speed limiting structure for an aircraft engine according to another embodiment;

[0022] Figure 4 for Figure 3 A schematic cross-sectional view of the second speed limiting portion of the shaft failure speed limiting structure shown;

[0023] Figure 5 A schematic structural diagram of a shaft failure speed limiting structure for an aircraft engine according to another embodiment;

[0024] Figure 6 A schematic diagram of the arrangement position of a shaft failure speed limiting structure of an aircraft engine according to another embodiment;

[0025] Figure 7 The present invention is a schematic diagram of the installation position of a shaft failure speed limiting structure of an aircraft engine according to another embodiment. DETAILED DESCRIPTION

[0026] The following will be combined with the drawings in the embodiments of the present invention to clearly and completely describe the technical solutions in the embodiments of the present invention. Obviously, the described embodiments are only part of the embodiments of the present invention, rather than all the embodiments. The following description of at least one exemplary embodiment is actually only illustrative and is by no means intended to limit the present invention and its application or use. Based on the embodiments of the present invention, all other embodiments obtained by ordinary technicians in this field without creative work are within the scope of protection of the present invention.

[0027] Unless otherwise specifically stated, the relative arrangement of the parts and steps described in these embodiments, numerical expressions and numerical values ​​do not limit the scope of the present invention. At the same time, it should be understood that, for ease of description, the sizes of the various parts shown in the accompanying drawings are not drawn according to the actual proportional relationship. The technology, method and equipment known to ordinary technicians in the relevant field may not be discussed in detail, but in appropriate cases, the technology, method and equipment should be regarded as a part of the authorization specification. In all examples shown and discussed here, any specific value should be interpreted as being merely exemplary, rather than as a limitation. Therefore, other examples of exemplary embodiments may have different values. It should be noted that similar reference numerals and letters represent similar items in the following drawings, so once a certain item is defined in an accompanying drawing, it does not need to be further discussed in subsequent drawings.

[0028] For ease of description, spatially relative terms such as "above", "above", "on the upper surface of", "above", etc. may be used here to describe the spatial positional relationship between a device or feature and other devices or features as shown in the figure. It should be understood that spatially relative terms are intended to include different orientations of the device in use or operation in addition to the orientation described in the figure. For example, if the device in the accompanying drawings is inverted, the device described as "above other devices or structures" or "above other devices or structures" will be positioned as "below other devices or structures" or "below other devices or structures". Thus, the exemplary term "above" can include both "above" and "below". The device can also be positioned in other different ways (rotated 90 degrees or in other orientations), and the spatially relative descriptions used here are interpreted accordingly.

[0029] like Figures 1 to 7 As shown, the aircraft engine of this embodiment includes a turbine, the turbine includes a rotor and a stator, the aircraft engine also includes a shaft failure speed limiting structure arranged between the rotor and the stator, the shaft failure speed limiting structure includes a first speed limiting portion 51 fixedly connected to the rotor and a second speed limiting portion 52 fixedly connected to the stator, and the shaft failure speed limiting structure is configured as follows: when the turbine is working normally, the first speed limiting portion 51 and the second speed limiting portion 52 maintain a gap; when the turbine shaft fails in the turbine, when the turbine blades move axially toward the downstream in the direction of gas flow, the first speed limiting portion 51 and the second speed limiting portion 52 approach each other and form an interference fit that becomes tighter the closer they are.

[0030] The shaft failure speed limiting structure can be set on the rotor and stator of the high-pressure turbine 20, or on the rotor and stator of the low-pressure turbine 30. The first speed limiting part 51 and the second speed limiting part 52 are arranged opposite to each other. When the turbine is working normally, the first speed limiting part 51 rotates with the rotor, and the rotation of the first speed limiting part 51 is not interfered by the second speed limiting part 52. When the turbine shaft fails, part of the turbine rotor (including the turbine blades) is separated from the turbine shaft and moves toward the downstream direction of the gas under the action of the gas. Driven by the moving rotor of the turbine, the first speed limiting part 51 maintains a gap with the second speed limiting part 52 at the beginning, then gradually approaches the second speed limiting part 52, and finally contacts the second speed limiting part 52 to form an interference fit. When the first speed limiting part 51 continues to move driven by the rotor of the turbine that continues to move downstream and the fit with the second speed limiting part 52 becomes deeper (that is, the first speed limiting part 51 and the second speed limiting part 52 are closer), the interference fit between the first speed limiting part 51 and the second speed limiting part 52 becomes tighter, that is, the first speed limiting part 51 and the second speed limiting part 52 are pressed tighter.

[0031] The above-mentioned interference fit between the first speed limiting part 51 and the second speed limiting part 52 can be achieved by setting a cone and a tapered hole. For example, the first speed limiting part 51 includes a cone, and the second speed limiting part 52 includes a tapered hole for inserting the cone corresponding to the outer surface shape of the cone. When the first speed limiting part 51 is close to the second speed limiting part 52, the cone is inserted into the tapered hole to form an interference fit. When the first speed limiting part 51 is closer to the second speed limiting part 52, that is, the cone is inserted deeper into the tapered hole, the tighter the interference fit, the more difficult it is to pull the cone out of the tapered hole. The second speed limiting part 52 may also include a cone, and the first speed limiting part 51 may also include a tapered hole corresponding to the outer surface shape of the cone.

[0032] The aircraft engine of this embodiment is provided with a shaft failure speed limiting structure including a first speed limiting portion 51 and a second speed limiting portion 52. When a turbine shaft failure occurs in the turbine and the moving blades of the turbine move axially downstream in the direction of the flow of the combustion gas, the first speed limiting portion 51 and the second speed limiting portion 52 approach each other and form an interference fit that becomes tighter the closer they get. The friction force caused by the interference fit between the first speed limiting portion 51 and the second speed limiting portion 52 can limit the separation of the first speed limiting portion 51 and the second speed limiting portion 52 when a turbine shaft failure occurs in the turbine, thereby quickly reducing the rotation speed of the turbine rotor when the turbine shaft fails. Since the first speed limiting part 51 and the second speed limiting part 52 can form an interference fit that becomes closer and tighter, when the turbine shaft fails, when the turbine rotor moves downstream under the action of the combustion gas and collides with the stator of the turbine, the greater the collision force, the tighter the interference fit between the first speed limiting part 51 and the second speed limiting part 52 will be, thereby effectively preventing the turbine rotor from rebounding, and the dynamic friction force of the second speed limiting part 52 on the first speed limiting part 51 will also be greater. The rotor of the turbine with a failed turbine shaft can quickly reduce its speed under the action of the dynamic friction force of the first speed limiting part 51.

[0033] In some embodiments, Figure 5 As shown, one of the first speed limiting part 51 and the second speed limiting part 52 includes an annular plate, and the other includes an annular groove corresponding to the annular plate, that is, the first speed limiting part 51 may include an annular plate, and the second speed limiting part 52 includes an annular groove, or the first speed limiting part 51 includes an annular groove, and the second speed limiting part 52 includes an annular plate. The annular plate, the annular groove and the turbine are coaxial. When the turbine shaft fails and the turbine moves axially downstream in the direction of gas flow, the annular plate is inserted into the annular groove to form an interference fit with the annular groove. In the embodiment shown in the figure, the first speed limiting part 51 includes an annular plate, and the second speed limiting part 52 includes an annular groove. Figure 5 In the embodiment shown, the annular plate is provided on the radially inner end of the moving blade, and the annular groove is provided on the radially inner end of the guide vane corresponding to the end of the moving blade (in Figure 5The structure of the guide vane is not drawn in the figure, only the partial structure of the annular groove on the guide vane is illustrated). In this embodiment, an annular plate and an annular groove structure are provided. When the annular plate is inserted into the annular groove, the radial inner surface and the radial outer surface of the annular plate are simultaneously matched with the wall surface of the annular groove, and the dynamic friction force is large when an interference fit is formed. The second speed limiting part 52 can effectively reduce the speed of the first speed limiting part 51 through the friction force.

[0034] In some embodiments, Figure 2 , Figure 3 and Figure 4 As shown, one of the first speed limiting part 51 and the second speed limiting part 52 includes a speed limiting shaft, and the other includes a speed limiting hole corresponding to the annular plate. In the figure, the first speed limiting part 51 includes the speed limiting shaft, and the second speed limiting part 52 includes the speed limiting hole. In some embodiments not shown in the figure, the first speed limiting part 51 may also include a speed limiting hole, and the second speed limiting part 52 may also include a speed limiting shaft. Figures 2 to 4 In the figure, only the structures of the first speed limiting part 51 and the second speed limiting part 52 are shown, and the rotor structure to which the first speed limiting part 51 is fixedly connected and the stator structure to which the second speed limiting part 52 is fixedly connected are not shown. The speed limiting shaft, the speed limiting hole and the turbine are coaxial. When the turbine shaft fails and the turbine moves axially downstream in the direction of gas flow, the speed limiting shaft is inserted into the speed limiting hole to form an interference fit with the speed limiting hole.

[0035] In some embodiments, Figure 5 As shown, the inlet end of the annular groove is gradually expanded, and the gradually expanded inlet end is used to guide the insertion of the annular plate. The annular groove includes a gradually expanded inlet end 521 and an interference fit section 522. The gradually expanded inlet end 521 gradually expands in the opposite direction of the axial movement direction of the first speed limiter 51 when the turbine shaft fails. The gradually expanded inlet end 521 is used to guide the annular plate when the annular plate is inserted. Under the guidance of the gradually expanded inlet end 521, the annular plate can be more easily aligned with the annular groove and enter the interference fit section 522 to achieve interference fit therewith.

[0036] In some embodiments, Figure 3 and Figure 4 As shown, the inlet end of the speed limiting hole is gradually expanded, and the gradually expanded inlet end is used to guide the insertion of the speed limiting shaft. The speed limiting hole includes a gradually expanded inlet end 521 and an interference fit section 522.

[0037] In some embodiments, the turbine is a low-pressure turbine. Shaft fracture failure of aircraft engines often occurs on the low-pressure turbine shaft 33, such as Figure 1 As shown, when the turbine shaft fails in the turbine, the low-pressure turbine shaft breaks at the shaft break point 331. The shaft speed limiting structure is arranged on the rotor and stator of the low-pressure turbine 30, which can more specifically provide speed limiting protection for the turbine shaft failure of the aircraft engine.

[0038] In some embodiments, Figure 1 , Figure 5 and Figure 6 As shown, the rotor includes a low-pressure turbine blade 32, the stator includes a low-pressure turbine guide vane 31, the first speed limiter 51 is arranged at the trailing edge of the low-pressure turbine blade 32, and the second speed limiter 52 is arranged on the leading edge of the low-pressure turbine guide vane 31 adjacent to the low-pressure turbine blade 32 and located downstream of the low-pressure turbine blade 32. The trailing edge of the low-pressure turbine blade 32 refers to the downstream side of the low-pressure turbine blade 32, and the leading edge refers to the upstream side. As shown in the figure, the shaft failure speed limiter structure can be arranged at B1, that is, the radially outer end of the low-pressure turbine blade 32 corresponds to the radially outer end of the low-pressure turbine guide vane 31. It can also be arranged at B2, that is, the radially inner end of the low-pressure turbine blade 32 corresponds to the radially inner end of the low-pressure turbine guide vane 31.

[0039] In some embodiments, Figure 1 and Figure 7 As shown, the rotor includes a low-pressure turbine support cone wall 332 for connecting the low-pressure turbine rotor blades 32 and the low-pressure turbine shaft 33, and the stator includes a stator sealing ring 1421 fixedly connected to the turbine interstage load-bearing casing (142) 142 of the aircraft engine at one end, and a sealing portion for sealing with the stator sealing ring 1421 is provided on the low-pressure turbine support cone wall 332, as shown in FIG. Figure 1 and Figure 7 As shown, the sealing part includes a rotor sealing ring 45 matched with a stator sealing ring 1421, a first speed limiting part 51 is arranged on the sealing part, and a second speed limiting part 52 is arranged on the stator sealing ring 1421. A first mounting part extending in the upstream direction can be arranged at the upstream end of the stator sealing ring 1421, and then the first speed limiting part 51 can be arranged on the mounting part, and the first speed limiting part 51 faces the downstream direction. A second mounting part is arranged at the downstream end of the stator sealing ring 1421, and the second mounting part is located downstream of the first mounting part. A second speed limiting part 52 facing the upstream direction is arranged on the second mounting part, and the first speed limiting part 51 is located upstream of the second speed limiting part 52 and opposite to it, that is, at Figure 1 and Figure 7 The C part shown can be provided with a shaft failure speed limiting structure.

[0040] exist Figure 1 In the illustrated embodiment, the rotor includes the end of the low-pressure turbine shaft 33, i.e., the end located downstream, the stator includes a bearing seat connected to the rear bearing casing 143 of the aircraft engine and connected to the end of the low-pressure turbine shaft 33 through a bearing, the first speed limiting part 51 is fixedly connected to the end of the low-pressure turbine shaft, and the second speed limiting part 52 is fixedly connected to the bearing seat.

[0041] In some embodiments, Figure 1As shown, an aircraft engine is also disclosed, the aircraft engine includes a fan 110, a low-pressure compressor 120, a high-pressure compressor 130, a combustion chamber 150, a high-pressure turbine 20, and a low-pressure turbine 30. The low-pressure compressor 120 includes a low-pressure compressor stator 121 and a low-pressure compressor rotor 122, the high-pressure compressor 130 includes a high-pressure compressor stator 131 and a high-pressure compressor rotor 132, the high-pressure turbine 20 includes a high-pressure turbine guide vane 21 and a high-pressure turbine moving blade 22, and the low-pressure turbine 30 includes a low-pressure turbine guide vane 31 and a low-pressure turbine moving blade 32. The fan 110 and the low-pressure compressor rotor 122 are driven by the low-pressure turbine 30, the low-pressure turbine shaft 33 is connected to the low-pressure turbine guide vane 32 through the low-pressure turbine support cone wall 332, the low-pressure turbine shaft 33 and the low-pressure turbine support cone wall 332 are generally connected by bolts, and the high-pressure compressor 130 is driven by the high-pressure turbine 20 and connected through the high-pressure turbine shaft 23. After being discharged from the combustion chamber 150, the high-temperature and high-energy combustion gas passes through the high-pressure turbine 20 and the low-pressure turbine 30 in turn, driving the high-pressure turbine blades 22 and the low-pressure turbine blades 32 to rotate. The high-pressure turbine blades 22 drive the high-pressure compressor rotor 132 at the front end to rotate, and the low-pressure turbine blades 32 drive the low-pressure compressor rotor 122 and the fan 110 to rotate. The compressor connection end of the low-pressure turbine shaft 33 connected to the low-pressure compressor is supported by the first roller bearing 101 and the first ball bearing 102, the turbine connection end of the low-pressure turbine shaft 33 connected to the low-pressure turbine support cone wall 332 is supported by the second roller bearing 105, the compressor connection end of the high-pressure turbine shaft 23 connected to the high-pressure compressor is supported by the second ball bearing 103, and the turbine connection end of the high-pressure turbine shaft 23 connected to the high-pressure turbine is supported by the third roller bearing 104. Roller bearings are mainly used to transmit radial forces, and ball bearings can transmit axial forces and radial forces at the same time. The axial force or radial force on the first roller bearing 101, the first ball bearing 102 and the second ball bearing 103 is mainly transmitted outward through the front bearing casing 141 inside the outer casing 140, and the force on the third roller bearing 104 and the second roller bearing 105 is respectively transmitted outward through the turbine interstage bearing casing (142) 142 and the rear bearing casing 143. The fan casing 111 is outside the fan 110, and the rear end is the guide support plate 112.

[0042] Finally, it should be noted that the above embodiments are only used to illustrate the technical solution of the present invention rather than to limit it. Although the present invention has been described in detail with reference to the preferred embodiments, ordinary technicians in the field should understand that the specific implementation methods of the present invention can still be modified or some technical features can be replaced by equivalents without departing from the spirit of the technical solution of the present invention, which should be included in the scope of the technical solution for protection of the present invention.

Claims

1. An aircraft engine, comprising a turbine, wherein the turbine comprises a rotor and a stator, wherein: The aircraft engine also includes a shaft failure speed limiting structure arranged between the rotor and the stator, the shaft failure speed limiting structure including a first speed limiting portion (51) fixedly connected to the rotor and a second speed limiting portion (52) fixedly connected to the stator, the shaft failure speed limiting structure being configured as follows: when the turbine is operating normally, a gap is maintained between the first speed limiting portion (51) and the second speed limiting portion (52); when a turbine shaft failure condition occurs in the turbine and the turbine blades move axially toward the downstream in the direction of gas flow, the first speed limiting portion (51) and the second speed limiting portion (52) approach each other and form an interference fit that becomes tighter the closer they are, the rotor includes low-pressure turbine blades, the stator includes low-pressure turbine guide vanes, the first speed limiting portion is arranged at the trailing edge of the low-pressure turbine blades, and the second speed limiting portion is arranged at an annular groove on the leading edge of the low-pressure turbine guide vanes adjacent to the low-pressure turbine blades and located downstream of the low-pressure turbine blades.

2. The aircraft engine according to claim 1, characterized in that: One of the first speed limiting part (51) and the second speed limiting part (52) comprises an annular plate, and the other comprises an annular groove corresponding to the annular plate, the annular plate, the annular groove and the turbine are coaxial, and when a turbine shaft failure occurs in the turbine and the moving blades of the turbine move axially downstream in the direction of gas flow, the annular plate is inserted into the annular groove to cooperate with the annular groove to form the interference fit; and / or One of the first speed limiting part (51) and the second speed limiting part (52) includes a speed limiting shaft, and the other includes a speed limiting hole corresponding to the annular plate. The speed limiting shaft, the speed limiting hole and the turbine are coaxial. When the turbine shaft fails and the moving blades of the turbine move axially downstream in the direction of gas flow, the speed limiting shaft is inserted into the speed limiting hole to cooperate with the speed limiting hole to form the interference fit.

3. The aircraft engine according to claim 2, characterized in that: The inlet end of the annular groove is gradually expanded, and the gradually expanded inlet end is used to guide the insertion of the annular plate, or the inlet end of the speed limiting hole is gradually expanded, and the gradually expanded inlet end is used to guide the insertion of the speed limiting shaft.

4. The aircraft engine according to any one of claims 1 to 3, characterized in that: The turbine is a low-pressure turbine (30).

5. The aircraft engine according to claim 4, characterized in that: The rotor comprises a low-pressure turbine blade (32), the stator comprises a low-pressure turbine guide vane (31), the first speed limiting portion (51) is arranged at a trailing edge of the low-pressure turbine blade (32), and the second speed limiting portion (52) is arranged at a leading edge of the low-pressure turbine guide vane (31) adjacent to the low-pressure turbine blade (32) and located downstream of the low-pressure turbine blade (32); or The rotor comprises a low-pressure turbine support cone wall (332) for connecting a low-pressure turbine rotor blade (32) and a low-pressure turbine shaft (33); the stator comprises a stator sealing ring (1421) having one end fixedly connected to a turbine interstage load-bearing casing (142) of the aircraft engine; a sealing portion for sealingly cooperating with the stator sealing ring (1421) is provided on the low-pressure turbine support cone wall (332); the first speed limiting portion (51) is provided on the sealing portion; and the second speed limiting portion (52) is provided on the stator sealing ring (1421); or The rotor includes the end of the low-pressure turbine shaft (33), the stator includes a bearing seat connected to the rear bearing casing (143) of the aircraft engine and connected to the end of the low-pressure turbine shaft (33) through a bearing, the first speed limiting part (51) is arranged on the end of the low-pressure turbine shaft (33), and the second speed limiting part (52) is arranged on the bearing seat.

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