CN114837748A - Aircraft engine - Google Patents
Aircraft engine Download PDFInfo
- Publication number
- CN114837748A CN114837748A CN202110142455.7A CN202110142455A CN114837748A CN 114837748 A CN114837748 A CN 114837748A CN 202110142455 A CN202110142455 A CN 202110142455A CN 114837748 A CN114837748 A CN 114837748A
- Authority
- CN
- China
- Prior art keywords
- speed
- turbine
- speed limiting
- limiting part
- low
- Prior art date
- Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
- Pending
Links
- 230000000670 limiting effect Effects 0.000 claims abstract description 138
- 238000007789 sealing Methods 0.000 claims description 24
- 238000003780 insertion Methods 0.000 claims description 5
- 230000037431 insertion Effects 0.000 claims description 5
- 239000007789 gas Substances 0.000 description 12
- 230000009471 action Effects 0.000 description 5
- 238000011144 upstream manufacturing Methods 0.000 description 5
- 238000010586 diagram Methods 0.000 description 4
- 239000000567 combustion gas Substances 0.000 description 3
- 238000002485 combustion reaction Methods 0.000 description 3
- 238000000034 method Methods 0.000 description 3
- 230000004044 response Effects 0.000 description 3
- 238000013461 design Methods 0.000 description 2
- 238000000926 separation method Methods 0.000 description 2
- 238000013459 approach Methods 0.000 description 1
- 230000007797 corrosion Effects 0.000 description 1
- 238000005260 corrosion Methods 0.000 description 1
- 230000007547 defect Effects 0.000 description 1
- 230000002939 deleterious effect Effects 0.000 description 1
- 239000012634 fragment Substances 0.000 description 1
- 230000014509 gene expression Effects 0.000 description 1
- 238000004519 manufacturing process Methods 0.000 description 1
- 239000000463 material Substances 0.000 description 1
- 238000012986 modification Methods 0.000 description 1
- 230000004048 modification Effects 0.000 description 1
- 238000012544 monitoring process Methods 0.000 description 1
- 230000000149 penetrating effect Effects 0.000 description 1
- 230000008569 process Effects 0.000 description 1
- 238000006467 substitution reaction Methods 0.000 description 1
Images
Classifications
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
- F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
- F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
- F01D5/02—Blade-carrying members, e.g. rotors
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
- F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
- F01D11/00—Preventing or minimising internal leakage of working-fluid, e.g. between stages
- F01D11/003—Preventing or minimising internal leakage of working-fluid, e.g. between stages by packing rings; Mechanical seals
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
- F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
- F01D25/00—Component parts, details, or accessories, not provided for in, or of interest apart from, other groups
- F01D25/16—Arrangement of bearings; Supporting or mounting bearings in casings
- F01D25/162—Bearing supports
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
- F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
- F01D25/00—Component parts, details, or accessories, not provided for in, or of interest apart from, other groups
- F01D25/24—Casings; Casing parts, e.g. diaphragms, casing fastenings
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
- F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
- F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
- F01D5/12—Blades
- F01D5/14—Form or construction
- F01D5/147—Construction, i.e. structural features, e.g. of weight-saving hollow blades
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
- F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
- F01D9/00—Stators
- F01D9/02—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
- F02C—GAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
- F02C9/00—Controlling gas-turbine plants; Controlling fuel supply in air- breathing jet-propulsion plants
Landscapes
- Engineering & Computer Science (AREA)
- Mechanical Engineering (AREA)
- General Engineering & Computer Science (AREA)
- Chemical & Material Sciences (AREA)
- Combustion & Propulsion (AREA)
- Architecture (AREA)
- Turbine Rotor Nozzle Sealing (AREA)
Abstract
The invention discloses an aircraft engine, which comprises a turbine, wherein the turbine comprises a rotor and a stator, the aircraft engine also comprises a shaft failure speed-limiting structure arranged between the rotor and the stator, 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 works normally, a gap is kept between the first speed limiting part and the second speed limiting part; when the turbine shaft of the turbine fails and the movable blades of the turbine move downstream along the axial direction in the gas flow direction, the first speed limiting part and the second speed limiting part are close to each other and form closer and closer interference fit.
Description
Technical Field
The invention relates to the field of aeronautical machinery, in particular to an aeroengine.
Background
In actual operation of a turbine-driven engine, turbine shaft failure may occur due to over-torquing, resonance, fatigue, corrosion, material defects and manufacturing errors or other indirect events, and although the probability of turbine shaft failure is small, once it occurs, it may have deleterious consequences. For example, in an aircraft engine, when a turbine shaft fails, a rotor of a turbine is decoupled from a front-end load (a gas compressor), and simultaneously, under the driving of high-energy gas exhausted from a combustion chamber, the rotating speed is instantly increased or enters an overspeed rotation state, when the rotating speed is increased to a certain degree, the stress of a wheel disc reaches the critical point and is cracked, and cracked high-energy fragments have the risk of penetrating through the engine, so that the limitation of turbine overspeed after the turbine shaft fails is an important consideration factor in the design of the turbine engine.
In known aircraft engines, the rotation speed of a rotor is generally directly monitored or converted to obtain the rotation speed of the rotor by installing a rotation speed sensor, and the rotation speed sensor is generally installed at the front end of an engine and is difficult to monitor the rotation speed increase of a rear-end turbine caused by failure of a turbine shaft. Even if a sensor is added at the turbine end, for a large civil aircraft engine, the control system judges that the failure event of the turbine shaft occurs to the oil-cut response from monitoring, the duration of the whole process is long, the response is slow, and the response is not timely enough.
It is known to add honeycomb or friction resistant devices to the stator structure of the low pressure turbine (mainly directing to the blades and the rear bearing casing of the low pressure turbine), or to axially sweep the guide vane design of the low pressure turbine, so as to limit the rotor speed by the collision friction or the clamping stagnation of the rotor moving backward and the stator after the shaft failure event occurs. However, after the low-pressure turbine rotor moves backwards and collides with the stator structure, collision force opposite to axial force is generated, the rotor or the rotor rebounds, the collision and friction brake cannot be continuously carried out, and the rotating speed limiting effect of the low-pressure turbine rotor after the low-pressure shaft fails is further influenced.
Disclosure of Invention
The invention aims to provide an aircraft engine which can quickly respond and effectively limit the increase of the rotating speed of a turbine when a turbine shaft fails.
The invention discloses an aircraft engine, which comprises a turbine, wherein the turbine comprises a rotor and a stator, the aircraft engine also comprises a shaft failure speed-limiting structure arranged between the rotor and the stator, 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 works normally, a gap is kept between the first speed limiting part and the second speed limiting part; when the turbine shaft of the turbine fails and the movable blades of the turbine move downstream along the axial direction in the gas flow direction, the first speed limiting part and the second speed limiting part are close to each other and form closer and closer interference fit.
In some embodiments of the present invention, the,
one of the first speed limiting part and the second speed limiting part comprises an annular plate, the other one of the first speed limiting part and the second speed limiting part comprises an annular groove corresponding to the annular plate, the annular groove and the turbine are coaxial, and when the turbine shaft fails and the turbine moves downstream along the gas flow direction along the axial direction, the annular plate is inserted into the annular groove to be matched with the annular groove to form interference fit; and/or
One of the first speed limiting part and the second speed limiting part comprises a speed limiting shaft, the other one of the first speed limiting part and the second speed limiting part comprises a speed limiting hole corresponding to the annular plate, the speed limiting shaft, the speed limiting hole and the turbine are coaxial, and when the turbine shaft fails and the turbine moves towards the downstream of the gas flow direction along the axial direction, the speed limiting shaft is inserted into the speed limiting hole to be matched with the speed limiting hole to form interference fit.
In some embodiments, the inlet end of the annular groove is divergent, and the divergent inlet end is used for inserting and guiding the annular plate, or the inlet end of the speed limiting hole is divergent, and the divergent inlet end is used for inserting and guiding the speed limiting shaft.
In some embodiments, the turbine is a low pressure turbine.
In some embodiments of the present invention, the,
the rotor comprises a low-pressure turbine movable blade, the stator comprises a low-pressure turbine guide blade, the first speed limiting part is arranged on the rear edge of the low-pressure turbine movable blade, and the second speed limiting part is arranged on the front edge of the low-pressure turbine guide blade which is adjacent to the low-pressure turbine movable blade and is positioned at the downstream of the low-pressure turbine movable blade; or
The rotor comprises a low-pressure turbine supporting conical wall used for connecting a low-pressure turbine movable blade and a low-pressure turbine shaft, the stator comprises a stator sealing ring with one end fixedly connected with a turbine interstage force bearing casing (142) of the aircraft engine, a sealing part in sealing fit with the stator sealing ring is arranged on the low-pressure turbine supporting conical wall, the first speed limiting part is arranged on the sealing part, and the second speed limiting part is arranged on the stator sealing ring; or
The rotor comprises the tail end of the low-pressure turbine shaft, the stator comprises a bearing seat which is connected with a rear bearing box of the aircraft engine and is connected with the tail end of the low-pressure turbine shaft through a bearing, the first speed limiting part is arranged at the tail end of the low-pressure turbine shaft, and the second speed limiting part is arranged on the bearing seat.
Based on the aircraft engine provided by the invention, by arranging the shaft failure speed-limiting structure comprising the first speed-limiting part and the second speed-limiting part, when a turbine shaft failure working condition occurs in the turbine, and a movable blade of the turbine moves to the downstream of the gas flow direction along the axial direction, the first speed-limiting part and the second speed-limiting part are close to each other and form closer interference fit, and the friction force caused 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 failure working condition occurs in the turbine, so that the rotating speed of a rotor of the turbine under the turbine shaft failure working condition is quickly reduced. Because the closer and tighter interference fit can be formed between the first speed limiting part and the second speed limiting part, when the turbine shaft fails, the rotor of the turbine moves downwards under the action of gas and collides with the stator of the turbine, the larger the collision force is, the tighter the interference fit between the first speed limiting part and the second speed limiting part is, and the rebound of the rotor of the turbine can be effectively prevented.
Other features of the present invention and advantages thereof will become apparent from the following detailed description of exemplary embodiments thereof, which proceeds with reference to the accompanying drawings.
Drawings
The accompanying drawings, which are included to provide a further understanding of the invention and are incorporated in and constitute a part of this application, illustrate embodiment(s) of the invention and together with the description serve to explain the invention without limiting the invention. In the drawings:
FIG. 1 is a schematic structural diagram of an aircraft engine according to an embodiment of the invention;
FIG. 2 is a schematic structural view of a shaft failure speed limiting structure of the aircraft engine shown in FIG. 1;
FIG. 3 is a schematic cross-sectional structural view of an aircraft engine shaft failure speed limiting structure according to another embodiment;
FIG. 4 is a schematic cross-sectional view of a second governor portion of the axle failure governor structure shown in FIG. 3;
FIG. 5 is a schematic structural diagram of a shaft failure speed limiting structure of an aircraft engine according to yet another embodiment;
FIG. 6 is a schematic diagram of the arrangement position of a shaft failure speed limiting structure of an aircraft engine according to yet another embodiment;
fig. 7 is a schematic diagram of the position of a shaft failure speed limiting structure of an aircraft engine according to yet another embodiment.
Detailed Description
The technical solutions in the embodiments of the present invention will be clearly and completely described below with reference to the drawings in the embodiments of the present invention, and it is obvious that the described embodiments are only a part of the embodiments of the present invention, and not all of the embodiments. The following description of at least one exemplary embodiment is merely illustrative in nature and is in no way intended to limit the invention, its application, or uses. All other embodiments, which can be derived by a person skilled in the art from the embodiments given herein without making any creative effort, shall fall within the protection scope of the present invention.
The relative arrangement of the components and steps, the numerical expressions and numerical values set forth in these embodiments do not limit the scope of the present invention unless specifically stated otherwise. Meanwhile, it should be understood that the sizes of the respective portions shown in the drawings are not drawn in an actual proportional relationship for the convenience of description. Techniques, methods, and apparatus known to those of ordinary skill in the relevant art may not be discussed in detail but are intended to be part of the specification where appropriate. In all examples shown and discussed herein, any particular value should be construed as merely illustrative, and not limiting. Thus, other examples of the exemplary embodiments may have different values. It should be noted that: like reference numbers and letters refer to like items in the following figures, and thus, once an item is defined in one figure, further discussion thereof is not required in subsequent figures.
Spatially relative terms, such as "above … …," "above … …," "above … … surface," "above," and the like, may be used herein for ease of description to describe one device or feature's spatial relationship to another device or feature as illustrated in the figures. It will be understood that the spatially relative terms are intended to encompass different orientations of the device in use or operation in addition to the orientation depicted in the figures. For example, if a device in the figures is turned over, devices described as "above" or "on" other devices or configurations would then be oriented "below" or "under" the other devices or configurations. Thus, the exemplary term "above … …" can include both an orientation of "above … …" and "below … …". The device may be otherwise variously oriented (rotated 90 degrees or at other orientations) and the spatially relative descriptors used herein interpreted accordingly.
As shown in fig. 1 to 7, the aircraft engine of the present embodiment includes a turbine, the turbine includes a rotor and a stator, the aircraft engine further includes a shaft failure speed limiting structure disposed 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 to: when the turbine normally works, the first speed limiting part 51 and the second speed limiting part 52 keep a gap; when the turbine shaft failure condition occurs in the turbine and the movable blades of the turbine move downstream along the axial direction in the flow direction of the combustion gas, the first speed limiting portion 51 and the second speed limiting portion 52 are close to each other and form an interference fit which is closer and closer.
The shaft failure speed limiting structure may be disposed on the rotor and stator of the high pressure turbine 20 or may be disposed on the rotor and stator of the low pressure turbine 30. The first speed limiting part 51 is arranged opposite to the second speed limiting part 52, when the turbine normally works, the first speed limiting part 51 rotates along 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 rotor of the turbine (including the blades of the turbine) is separated from the constraint of the turbine shaft and moves to the downstream direction of the combustion gas under the action of the combustion gas. The first speed limiting part 51 is driven by the rotor of the moving turbine to keep a gap with the second speed limiting part 52 from the beginning to gradually approach the second speed limiting part 52 and to contact with the second speed limiting part 52 and form interference fit, when the first speed limiting part 51 is driven by the rotor of the turbine moving downstream continuously and moves continuously and is matched with the second speed limiting part 52 more deeply (namely, when the first speed limiting part 51 is closer to the second speed limiting part 52), the interference fit between the first speed limiting part 51 and the second speed limiting part 52 is tighter, namely, the first speed limiting part 51 and the second speed limiting part 52 are pressed more tightly.
The interference fit between the first speed-limiting portion 51 and the second speed-limiting portion 52 can be achieved by providing a tapered body and a tapered hole. For example, the first speed-limiting portion 51 includes a tapered body, the second speed-limiting portion 52 includes a tapered hole for insertion of the tapered body corresponding to the outer surface shape of the tapered body, the tapered body is inserted into the tapered hole to form an interference fit when the first speed-limiting portion 51 is close to the second speed-limiting portion 52, and the interference fit is tighter when the first speed-limiting portion 51 is closer to the second speed-limiting portion 52, that is, the position where the tapered body is inserted into the tapered hole is deeper, the tapered body is more difficult to be pulled out from the tapered hole. The second speed-limiting portion 52 may include a tapered body, and the first speed-limiting portion 51 may include a tapered hole corresponding to the shape of the outer surface of the tapered body.
According to the aircraft engine of the embodiment, by arranging the shaft failure speed-limiting structure comprising the first speed-limiting part 51 and the second speed-limiting part 52, when a turbine shaft failure working condition occurs in the turbine, and a movable blade of the turbine moves to the downstream of the gas flow direction along the axial direction, the first speed-limiting part 51 and the second speed-limiting part 52 are close to each other and form closer interference fit, and friction force caused by the interference fit of the first speed-limiting part 51 and the second speed-limiting part 52 can limit separation of the first speed-limiting part 51 and the second speed-limiting part 52 when the turbine shaft failure working condition occurs in the turbine, so that the rotating speed of a rotor of the turbine under the turbine shaft failure working condition is rapidly reduced. Because the closer and tighter interference fit can be formed between the first speed limiting part 51 and the second speed limiting part 52, when the turbine shaft fails, and when the rotor of the turbine moves downstream under the action of 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, so that the rebound of the rotor of the turbine can be effectively prevented, the greater the dynamic friction force of the second speed limiting part 52 on the first speed limiting part 51 will be, and the rotational speed of the rotor of the turbine shaft that fails can be rapidly reduced under the action of the dynamic friction force received by the first speed limiting part 51.
In some embodiments, as shown in fig. 5, one of the first speed limiting part 51 and the second speed limiting part 52 includes an annular plate, and the other one 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, and when the turbine shaft fails and the turbine moves towards the downstream of the gas flow direction along the axial direction under the working condition that the turbine shaft fails, the annular plate is inserted into the annular groove to form interference fit with the annular groove. In the embodiment shown in the drawings, the first speed limit portion 51 includes an annular plate, and the second speed limit portion 52 includes an annular groove. In the embodiment shown in fig. 5, the annular plate is provided on the end portion of the radially inner side of the movable blade, and the annular groove is provided on the end portion of the radially inner side of the guide blade corresponding to the end portion of the movable blade (the structure of the guide blade is not shown in fig. 5, and only the partial structure of the guide blade on which the annular groove is provided is illustrated).
In some embodiments, as shown in fig. 2, 3 and 4, one of the first speed limiting portion 51 and the second speed limiting portion 52 includes a speed limiting shaft, and the other includes a speed limiting hole corresponding to the annular plate, in the figures, the first speed limiting portion 51 includes the speed limiting shaft, the second speed limiting portion 52 includes the speed limiting hole, in some embodiments not shown in the figures, the first speed limiting portion 51 may also include the speed limiting hole, and the second speed limiting portion 52 may also include the speed limiting shaft. In fig. 2 to 4, only the structures of the first speed limiting portion 51 and the second speed limiting portion 52 are shown, and the rotor structure fixedly connected to the first speed limiting portion 51 and the stator structure fixedly connected to the second speed limiting portion 52 are not illustrated. The speed-limiting shaft, the speed-limiting hole and the turbine are coaxial, and when the turbine shaft fails and the turbine moves towards the downstream of the gas flowing direction along the axial direction, the speed-limiting shaft is inserted into the speed-limiting hole to form interference fit with the speed-limiting hole.
In some embodiments, as shown in FIG. 5, the inlet end of the annular groove is divergent, and the divergent inlet end is used to guide the insertion of the annular plate. The annular groove comprises a gradually-expanding inlet end 521 and an interference fit section 522, the gradually-expanding inlet end 521 is gradually expanded along the direction opposite to the axial moving direction of the first speed limiting part 51 under the condition that the turbine shaft of the turbine fails, the gradually-expanding inlet end 521 is used for guiding the annular plate when the annular plate is inserted, and the annular plate can be more easily aligned to the annular groove under the guiding action of the gradually-expanding inlet end 521 and enters the interference fit section 522 to realize interference fit with the annular groove.
In some embodiments, as shown in FIGS. 3 and 4, the inlet end of the speed-limiting orifice is flared, and the flared inlet end is used to guide the insertion of the speed-limiting shaft. The speed limiting orifice includes a diverging inlet end 521 and an interference fit section 522.
In some embodiments, the turbine is a low pressure turbine. Shaft fracture failure of the aircraft engine mostly occurs on the low-pressure turbine shaft 33, as shown in fig. 1, when the turbine shaft failure working condition occurs in the turbine, the low-pressure turbine shaft fractures 331 at the shaft, the shaft speed limiting structure is arranged on a rotor and a stator of the low-pressure turbine 30, and speed limiting protection can be performed on the turbine shaft failure of the aircraft engine in a more targeted manner.
In some embodiments, as shown in fig. 1, 5 and 6, the rotor comprises a low pressure turbine blade 32, the stator comprises a low pressure turbine vane 31, the first speed limiter 51 is provided at a trailing edge of the low pressure turbine blade 32, and the second speed limiter 52 is provided on a leading edge of the low pressure turbine vane 31 adjacent to the low pressure turbine blade 32 and downstream of the low pressure turbine blade 32. The trailing edge of the low pressure turbine bucket 32 refers to the downstream side of the low pressure turbine bucket 32 and the leading edge refers to the upstream side. As shown, a shaft failure rate limiting structure may be provided at B1, i.e., the radially outer end of the low pressure turbine bucket 32 and the radially outer end of the low pressure turbine vane 31 correspond. It may be provided at B2, that is, the radially inner end of the low-pressure turbine bucket 32 and the radially inner end of the low-pressure turbine vane 31 correspond to each other.
In some embodiments, as shown in fig. 1 and 7, the rotor includes a low-pressure turbine supporting conical wall 332 for connecting the low-pressure turbine rotor blade 32 and the low-pressure turbine shaft 33, the stator includes a stator sealing ring 1421 having one end fixedly connected to a turbine stage bearing casing (142)142 of the aircraft engine, a sealing portion for sealing the stator sealing ring 1421 is provided on the low-pressure turbine supporting conical wall 332, as shown in fig. 1 and 7, the sealing portion includes a rotor sealing ring 45 for sealing the stator sealing ring 1421, 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. A first mounting portion extending in an upstream direction may be provided at an upstream end of the stator sealing ring 1421, and then the first speed limiting portion 51 may be provided on the mounting portion, with the first speed limiting portion 51 facing in a downstream direction. A second mounting part is arranged at the downstream end of the stator sealing ring 1421, the second mounting part is positioned at the 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 positioned at the upstream of the second speed limiting part 52 and is opposite to the second speed limiting part, namely, a shaft failure speed limiting structure can be arranged at the part C shown in fig. 1 and 7.
In the embodiment shown in fig. 1, the rotor comprises the end of the low-pressure turbine shaft 33, i.e. the end located downstream, the stator comprises a bearing seat connected to the rear bearing case 143 of the aircraft engine and to the end of the low-pressure turbine shaft 33 via a bearing, the first governor part 51 is fixedly connected to the end of the low-pressure turbine shaft, and the second governor part 52 is fixedly connected to the bearing seat.
In some embodiments, as shown in fig. 1, an aircraft engine is further 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 high-pressure turbine guide vanes 21 and high-pressure turbine blades 22, and the low-pressure turbine 30 includes low-pressure turbine guide vanes 31 and low-pressure turbine blades 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 vanes 32 via a low-pressure turbine support cone wall 332, the low-pressure turbine shaft 33 and the low-pressure turbine support cone wall 332 are connected, typically by bolts, and the high-pressure compressor 130 is driven by the high-pressure turbine 20 and connected via a high-pressure turbine shaft 23. After being discharged from the combustion chamber 150, the high-temperature high-energy gas passes through the high-pressure turbine 20 and the low-pressure turbine 30 in sequence, the high-pressure turbine rotor blades 22 and the low-pressure turbine rotor blades 32 are driven to rotate, the high-pressure turbine rotor blades 22 drive the front-end high-pressure compressor rotor 132 to rotate, and the low-pressure turbine rotor blades 32 drive the low-pressure compressor rotor 122 and the fan 110 to rotate. The compressor connecting end of the low-pressure turbine shaft 33 connected with the low-pressure compressor is supported by a first roller bearing 101 and a first ball bearing 102, the turbine connecting end of the low-pressure turbine shaft 33 connected with the low-pressure turbine supporting conical wall 332 is supported by a second roller bearing 105, the compressor connecting end of the high-pressure turbine shaft 23 connected with the high-pressure compressor is supported by a second ball bearing 103, and the turbine connecting end of the high-pressure turbine shaft 23 connected with the high-pressure turbine is supported by a third roller bearing 104. The roller bearing is mainly used for transmitting radial force, and the ball bearing can simultaneously transmit axial force and radial force. Axial or radial forces on the first roller bearing 101, the first ball bearing 102 and the second ball bearing 103 are mainly transmitted outwards through a front bearing case 141 inside an outer case 140, and the stress of the third roller bearing 104 and the second roller bearing 105 is transmitted outwards through a turbine interstage bearing case (142)142 and a rear bearing case 143 respectively. The fan 110 has a fan casing 111 at its outer side and a flow guiding plate 112 at its rear end.
Finally, it should be noted that: the above examples are only intended to illustrate the technical solution of the present invention and not to limit it; although the present invention has been described in detail with reference to preferred embodiments, those skilled in the art will understand that: modifications to the specific embodiments of the invention or equivalent substitutions for parts of the technical features may be made; without departing from the spirit of the present invention, it is intended to cover all aspects of the invention as defined by the appended claims.
Claims (5)
1. An aircraft engine, comprising a turbine, the turbine including a rotor and a stator, characterized in that the aircraft engine further includes a shaft failure speed-limiting structure disposed 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 to: when the turbine normally works, the first speed limiting part (51) and the second speed limiting part (52) keep a gap; when the turbine shaft failure condition occurs in the turbine and the movable blades of the turbine move downstream along the axial direction in the gas flow direction, the first speed limiting part (51) and the second speed limiting part (52) are close to each other and form an interference fit which is closer and closer.
2. The aircraft engine of claim 1,
one of the first speed limiting part (51) and the second speed limiting part (52) comprises an annular plate, the other one of the first speed limiting part and the second speed limiting part comprises an annular groove corresponding to the annular plate, the annular groove and the turbine are coaxial, and when a turbine shaft failure working condition occurs in the turbine and a movable blade of the turbine moves downstream along the gas flow direction along the axial direction, the annular plate is inserted into the annular groove to be matched 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) comprises a speed limiting shaft, the other one of the first speed limiting part and the second speed limiting part comprises 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 movable blade of the turbine moves towards the downstream of the gas flowing direction along the axial direction, the speed limiting shaft is inserted into the speed limiting hole to form interference fit with the speed limiting hole.
3. The aircraft engine according to claim 2, wherein the inlet end of the annular groove is divergent and the divergent inlet end is used for insertion guidance of the annular plate, or the inlet end of the speed-limiting hole is divergent and the divergent inlet end is used for insertion guidance of the speed-limiting shaft.
4. An aircraft engine according to any one of claims 1 to 3, characterised in that the turbine is a low pressure turbine (30).
5. The aircraft engine of claim 4,
the rotor comprises a low-pressure turbine movable blade (32), the stator comprises a low-pressure turbine guide blade (31), the first speed limiting part (51) is arranged at the rear edge of the low-pressure turbine movable blade (32), and the second speed limiting part (52) is arranged on the front edge of the low-pressure turbine guide blade (31) which is adjacent to the low-pressure turbine movable blade (32) and is located at the downstream of the low-pressure turbine movable blade (32); or
The rotor comprises a low-pressure turbine supporting conical wall (332) used for connecting a low-pressure turbine moving blade (32) and a low-pressure turbine shaft (33), the stator comprises a stator sealing ring (1421) with one end fixedly connected with a turbine interstage force bearing casing (142) of the aircraft engine, a sealing part in sealing fit with the stator sealing ring (1421) is arranged on the low-pressure turbine supporting conical wall (332), the first speed limiting part (51) is arranged on the sealing part, and the second speed limiting part (52) is arranged on the stator sealing ring (1421); or
The rotor comprises the tail end of the low-pressure turbine shaft (33), the stator comprises a bearing seat which is connected with a rear bearing casing (143) of the aircraft engine and is connected with the tail end of the low-pressure turbine shaft (33) through a bearing, the first speed limiting part (51) is arranged on the tail end of the low-pressure turbine shaft (33), and the second speed limiting part (52) is arranged on the bearing seat.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202110142455.7A CN114837748A (en) | 2021-02-02 | 2021-02-02 | Aircraft engine |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202110142455.7A CN114837748A (en) | 2021-02-02 | 2021-02-02 | Aircraft engine |
Publications (1)
| Publication Number | Publication Date |
|---|---|
| CN114837748A true CN114837748A (en) | 2022-08-02 |
Family
ID=82561043
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CN202110142455.7A Pending CN114837748A (en) | 2021-02-02 | 2021-02-02 | Aircraft engine |
Country Status (1)
| Country | Link |
|---|---|
| CN (1) | CN114837748A (en) |
Citations (6)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US5029439A (en) * | 1988-12-15 | 1991-07-09 | Societe National D'etude Et De Construction De Moteurs D'aviation "S.N.E.C.M.A." | Gas turbine engine including a turbine braking device |
| CN1755063A (en) * | 2004-09-28 | 2006-04-05 | 斯奈克玛公司 | Turbine overspeed limiting device |
| CN1950590A (en) * | 2004-04-07 | 2007-04-18 | 西门子公司 | Turbo-machine and rotor therefor |
| CN101737087A (en) * | 2008-11-12 | 2010-06-16 | 阿特拉斯·科普柯能源有限公司 | Rotor of turbine |
| CN104379289A (en) * | 2012-03-29 | 2015-02-25 | 大陆汽车有限公司 | Turbine rotor for an exhaust-gas turbine and method for producing the turbine rotor |
| CN105339589A (en) * | 2013-07-03 | 2016-02-17 | 大陆汽车有限公司 | Rotor for a turbocharger device, turbocharger device having a rotor, and shaft for a rotor of said type |
-
2021
- 2021-02-02 CN CN202110142455.7A patent/CN114837748A/en active Pending
Patent Citations (6)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US5029439A (en) * | 1988-12-15 | 1991-07-09 | Societe National D'etude Et De Construction De Moteurs D'aviation "S.N.E.C.M.A." | Gas turbine engine including a turbine braking device |
| CN1950590A (en) * | 2004-04-07 | 2007-04-18 | 西门子公司 | Turbo-machine and rotor therefor |
| CN1755063A (en) * | 2004-09-28 | 2006-04-05 | 斯奈克玛公司 | Turbine overspeed limiting device |
| CN101737087A (en) * | 2008-11-12 | 2010-06-16 | 阿特拉斯·科普柯能源有限公司 | Rotor of turbine |
| CN104379289A (en) * | 2012-03-29 | 2015-02-25 | 大陆汽车有限公司 | Turbine rotor for an exhaust-gas turbine and method for producing the turbine rotor |
| CN105339589A (en) * | 2013-07-03 | 2016-02-17 | 大陆汽车有限公司 | Rotor for a turbocharger device, turbocharger device having a rotor, and shaft for a rotor of said type |
Similar Documents
| Publication | Publication Date | Title |
|---|---|---|
| EP3205840B1 (en) | Rotor support system with shape memory alloy components for a gas turbine engine | |
| US9909451B2 (en) | Bearing assembly for supporting a rotor shaft of a gas turbine engine | |
| US8166746B2 (en) | Rotor containment element with frangible connections | |
| CN106988799B (en) | Metal attachment system integrated into a composite structure | |
| US20180363665A1 (en) | Geared turbofan with overspeed protection | |
| US11624295B2 (en) | Gas turbine engine | |
| EP3686561B1 (en) | Shaft monitoring system | |
| US11561235B2 (en) | Shaft monitoring system | |
| EP2460976A2 (en) | Gas turbine rotor containment | |
| US10975729B2 (en) | Gas turbine engine | |
| US11643944B2 (en) | Turbine shaft of a turbomachine and method for protecting against overspeed of said shaft | |
| CN114837749A (en) | Aircraft engine | |
| US20210095602A1 (en) | Gas turbine engine | |
| US11060419B2 (en) | Planetary gear drive and aircraft gas turbine with a planetary gear drive | |
| CN114837748A (en) | Aircraft engine | |
| CN217558429U (en) | Aircraft engine | |
| US10677089B2 (en) | Mitigation of effects of fan blade off in a gas turbine engine | |
| US11542828B2 (en) | Gas turbine engine | |
| CN114837752A (en) | Aircraft engine | |
| CN213598075U (en) | Turbine engine and shaft failure detection system thereof | |
| EP3309364B1 (en) | System of an engine | |
| CN115126545A (en) | Aircraft engine | |
| CN116357404A (en) | Braking structure for preventing turbine from flying after broken shaft of aero-engine |
Legal Events
| Date | Code | Title | Description |
|---|---|---|---|
| PB01 | Publication | ||
| PB01 | Publication | ||
| SE01 | Entry into force of request for substantive examination | ||
| SE01 | Entry into force of request for substantive examination |