Thermal deformation coordination structure of aero-engine turbine casing assembly

Abstract

The invention discloses an aero-engine turbine casing assembly thermal deformation coordination structure which comprises a turbine outer ring located on the outer side of a turbine blade, a turbine casing and a rear bearing casing which are axially connected, and a guider outer ring radially connected with the turbine casing, and the turbine outer ring is arranged between the guider outer ring and the turbine casing in a limited mode. Annular elastic parts are arranged on the outer sides of the two ends of the turbine outer ring and correspond to the turbine case and the rear bearing case correspondingly, and the annular elastic parts are used for achieving the radial positioning effect on the guider outer ring, providing space for radial thermal expansion of the turbine outer ring and reducing thermal stress of the turbine outer ring and the turbine case. The turbine outer ring is flexibly installed and positioned through the annular elastic parts, space is provided for free expansion of the turbine outer ring and the guider outer ring, the thermal stress of the turbine outer ring and the guider outer ring is reduced, the turbine outer ring cannot generate large plastic deformation, and the turbine outer ring and turbine blades can keep a proper radial gap.

Description

Technical Field

[0001] This invention relates to the field of turbine engine technology, specifically to a thermal deformation coordination structure for an aero-engine turbine casing assembly. Background Technology

[0002] Typically, a power turbine casing assembly consists of two layers: an outer ring-shaped casing for load-bearing, and an inner segmented outer ring to withstand the scouring of high-temperature airflow and to create rotor-stator clearance with the turbine blades. With the continuous improvement of aero-engine performance, turbine inlet temperatures are increasing, making the turbine outer ring, which directly contacts the high-temperature combustion gases, one of the most demanding components in the engine. Due to the circumferential and axial non-uniformity of the combustion gas temperature field, the turbine outer ring also experiences uneven heating. Furthermore, the differences in size and material compared to adjacent components lead to varying deformations. The rigid connection between the casing and the outer ring will cause significant thermal stress and uneven deformation in the turbine outer ring, affecting rotor-stator clearance and turbine engine efficiency, and potentially causing rubbing between the turbine blades and the outer ring. Summary of the Invention

[0003] To reduce the thermal stress and uneven deformation of the turbine casing and turbine outer ring, this invention provides a thermal deformation coordination structure for aero-engine turbine casing assemblies, in order to solve the problems of incomplete thermal stress release and thermal deformation mismatch in existing turbine casing assembly connection structures.

[0004] The present invention solves the above problems through the following technical solution:

[0005] A thermal deformation coordination structure for an aero-engine turbine casing assembly includes: a turbine outer ring located outside the turbine blades, an axially connected turbine casing and rear bearing casing, and a guide outer ring radially connected to the turbine casing. The turbine outer ring is positioned between the guide outer ring and the turbine casing, and annular elastic parts are provided on the outer sides of both ends of the turbine outer ring corresponding to the turbine casing and the rear bearing casing, respectively. The annular elastic parts are used to radially position the guide outer ring and provide space for the radial thermal expansion of the turbine outer ring, thereby reducing the thermal stress on the turbine outer ring and the turbine casing.

[0006] As a further improvement of the present invention, the annular elastic component includes at least a second annular elastic component and a fourth annular elastic component disposed at both ends of the turbine outer ring, for buffering the radial deformation of the turbine outer ring and isolating the high-temperature gas to reduce the conduction and radiation of high-temperature gas heat to the turbine casing.

[0007] As a further improvement of the present invention, one end of the second annular elastic part is welded to the outer side of the turbine casing near the combustion chamber casing, and the other end is welded to the top of the turbine outer ring.

[0008] The two ends of the fourth annular elastic component are welded to the lower middle part of the turbine casing and abut against the end of the turbine outer ring away from the combustion chamber casing.

[0009] As a further improvement of the present invention, the annular elastic component further includes a first annular elastic component, the two ends of which are welded to the upper part of the turbine outer ring, and the outer side of which abuts against the inside of the rear bearing housing.

[0010] As a further improvement of the present invention, the guide outer ring and the rear bearing housing are respectively formed with a second step plane and a third step plane at both ends of the turbine outer ring, which are used to fit into the two ends of the turbine outer ring to form axial positioning.

[0011] As a further improvement of the present invention, the annular elastic component further includes a third annular elastic component, the two ends of which are welded to the lower end of the turbine casing away from the combustion chamber casing, and the outer side abuts against the outer ring of the guide.

[0012] As a further improvement of the present invention, the outer ring of the guide is connected to the end of the turbine casing away from the rear bearing casing by a number of screws to achieve circumferential positioning.

[0013] As a further improvement of the present invention, the outer ring of the guide is connected to the turbine casing with a first stepped plane, which is used to fit into the turbine casing to form axial positioning.

[0014] As a further improvement of the present invention, one end of the screw passes through the second nut, the second stop washer, the turbine casing and the third annular elastic part in sequence and is then threadedly connected to the outer ring of the guide, and the turbine casing can move up and down relative to the screw.

[0015] As a further improvement of the present invention, the turbine casing has several ventilation holes for cooling its internal components;

[0016] A combustion chamber casing is provided between the turbine casing and the rear bearing casing. The turbine casing, the rear bearing casing, and the combustion chamber casing are connected by a number of locking washers, bolts, mounting nuts, and a first nut. The combustion chamber casing is embedded with a mounting nut. One end of the bolt passes through the turbine casing, the combustion chamber casing, the rear bearing casing, and the locking washer in sequence and is connected to the first nut. After the connection is completed, the locking lug of the locking washer is bent to the outside of the first nut to prevent loosening.

[0017] Compared with the prior art, the present invention has the following advantages and beneficial effects:

[0018] (1) The present invention employs four annular elastic parts to flexibly position the turbine outer ring between the guide outer ring, the turbine casing, and the rear bearing casing. When the turbine outer ring and the guide outer ring are subjected to thermal expansion and deformation by the combustion gas, the annular elastic parts are compressed, providing space for the turbine outer ring and the guide outer ring to expand freely, reducing the thermal stress of the turbine outer ring and the guide outer ring, preventing large plastic deformation of the turbine outer ring, and ensuring that the turbine outer ring and the turbine blades maintain a suitable radial clearance.

[0019] (2) The annular elastic part and the turbine outer ring cooperate to prevent the high temperature gas heat from radiating to the combustion chamber casing, turbine casing and rear bearing casing connecting flange, so that the working temperature at the three casing flanges is close and the deformation is coordinated. Attached Figure Description

[0020] Figure 1 This is a schematic diagram of a thermal deformation coordination structure for an aero-engine turbine casing assembly according to the present invention.

[0021] Reference numerals: 1. Combustion chamber casing; 2. Mounting locating nut; 3. First retaining washer; 4. Bolt; 5. First nut; 6. Rear bearing casing; 61. Third stepped plane; 7. First annular elastic component; 8. Second annular elastic component; 9. Turbine outer ring; 10. Turbine blade; 11. Guide vane; 12. Guide outer ring; 121. First stepped plane; 122. Second stepped plane; 13. Third annular elastic component; 14. Second retaining washer; 15. Screw; 16. Second nut; 17. Fourth annular elastic component; 18. Turbine casing. Detailed Implementation

[0022] The technical solutions of the embodiments of the present invention will be clearly and completely described below with reference to the accompanying drawings. Obviously, the described embodiments are only some embodiments of the present invention, and not all embodiments. Based on the embodiments of the present invention, all other embodiments obtained by those skilled in the art without creative effort are within the scope of protection of the present invention.

[0023] Example:

[0024] See Figure 1A thermal deformation coordination structure for an aero-engine turbine casing assembly includes: a turbine outer ring 9 located outside the turbine blade 10, an axially connected turbine casing 18 and rear bearing casing 6, and a guide outer ring 12 radially connected to the turbine casing 18. The turbine outer ring 9 is positioned between the guide outer ring 12 and the turbine casing 18, and annular elastic parts are respectively provided on the outer ends of the turbine outer ring 9 corresponding to the turbine casing 18 and the rear bearing casing 6 to provide radial positioning for the guide outer ring 12 and to provide space for radial thermal expansion of the guide outer ring 12 through the compressive deformation of the annular elastic parts. When the turbine outer ring 9 is thermally expanded by high-temperature combustion gases, the radial deformation compresses the annular elastic parts to provide space for radial deformation of the turbine outer ring 9, reducing the thermal stress on the turbine outer ring 9 and the turbine casing 18.

[0025] Specifically, a thermal deformation coordination structure for an aero-engine turbine casing assembly, mainly used in turbine engines, includes: a combustion chamber casing 1, a mounting nut 2, a first retaining washer 3, a bolt 4, a first nut 5, a rear bearing casing 6, a first annular elastic component 7, a second annular elastic component 8, a turbine outer ring 9, a turbine blade 10, a guide vane 11, a guide outer ring 12, a third annular elastic component 13, a second retaining washer 14, a screw 15, a second nut 16, a fourth annular elastic component 17, and a turbine casing 18, etc.

[0026] The turbine casing 18, the rear bearing casing 6, and the combustion chamber casing 1 are connected by several first locking washers 3, bolts 4, and first nuts 5. The combustion chamber casing 1 has an embedded mounting nut 2 for threaded engagement with the bolts 4. One end of the bolt 4 passes sequentially through the turbine casing 18, the combustion chamber casing 1, the rear bearing casing 6, and the first locking washers 3 before connecting to the first nut 5. After connection, the locking lug of the first locking washer 3 is bent outwards to prevent loosening. During assembly, the bolt 4 is first passed through the flange of the turbine casing 18, and the mounting nut 2 is used to fix the bolt 4 to the turbine casing 18. Then, the bolt 4 is passed through the combustion chamber casing 1 and the rear bearing casing 6, and finally, the first nut 5 and the first locking washer 3 are used to secure the bolt 4.

[0027] The annular elastic components include at least a second annular elastic component 8 and a fourth annular elastic component 17 disposed at both ends of the turbine outer ring 9. These components buffer the radial deformation of the turbine outer ring 9 and isolate the high-temperature combustion gas to reduce the conduction and radiation of heat from the high-temperature combustion gas to the turbine casing 18. Specifically, one end of the second annular elastic component 8 is welded to the outer side of the turbine casing 18 near the combustion chamber casing 1, and the other end is welded to the top of the turbine outer ring 9. Both ends of the fourth annular elastic component 17 are welded to the lower middle part of the turbine casing 18 and abut against the end of the turbine outer ring 9 away from the combustion chamber casing 1. Of course, other forms of the second annular elastic component 8 and the fourth annular elastic component 17 can also be used, as long as they can buffer the radial deformation of the turbine outer ring 9 and isolate the high-temperature combustion gas to reduce the conduction and radiation of heat from the high-temperature combustion gas to the turbine casing 18.

[0028] Of course, it may also include a first annular elastic part 7 and a third annular elastic part 13. The two ends of the first annular elastic part 7 are welded to the upper part of the turbine outer ring 9, and its outer side abuts against the inner part of the rear bearing housing 6. The two ends of the third annular elastic part 13 are welded to the lower part of the turbine housing 18 away from the combustion chamber housing 1, and its outer side abuts against the guide outer ring 12. The guide outer ring and the turbine housing are connected by several screws, which serve as circumferential positioning. A third annular elastic part 13 is provided between the guide outer ring and the turbine housing. This annular elastic part has several holes for the screws to pass through, and this annular elastic part has a specific stiffness, which can play a centering role when the guide outer ring is installed into the turbine housing, and can be compressed when the guide outer ring thermally expands. A second annular elastic part 8 and a fourth annular elastic part 17 are provided between the turbine outer ring and the turbine housing to buffer the radial deformation of the turbine outer ring and isolate the high-temperature combustion gas, reducing the conduction and radiation of the high-temperature combustion gas heat to the turbine housing. A first annular elastic component 7 is installed on the turbine outer ring near the rear bearing housing to isolate the high-temperature combustion gas and reduce the conduction and radiation of heat from the high-temperature combustion gas to the rear bearing housing. The first annular elastic component 7, the second annular elastic component 8, and the fourth annular elastic component 17 have specific stiffness. When the turbine outer ring undergoes thermal expansion under the action of high-temperature combustion gas, the three annular elastic components try to keep the turbine outer ring in a cylindrical shape.

[0029] In this embodiment, the two ends of the third annular elastic member 13 are welded to the lower part of the turbine casing 18 away from the combustion chamber casing 1; the two ends of the fourth annular elastic member 17 are welded to the lower part of the middle of the turbine casing 18; the A end of the second annular elastic member 8 is welded to the outer side of the turbine casing 18 near the combustion chamber casing 1, and the B end is welded to the upper part of the turbine outer ring 9; the two ends of the first annular elastic member 7 are welded to the upper part of the turbine outer ring 9 and are located outside the B end of the second annular elastic member 8. Preferably, the turbine casing 18 is provided with an inwardly recessed compression space at one end of the second annular elastic member 8 to increase the compression space of the second annular elastic member 8.

[0030] First, the first annular elastic component 7 is welded to the outer side of the turbine outer ring 9 near the rear bearing housing 6, i.e., above the turbine outer ring 9 between the rear bearing housing 6 and the turbine outer ring 9. Then, the turbine outer ring 9 is passed through the fourth annular elastic component 17, centered by the cylindrical surface of the guide outer ring 12, and axially positioned by the stepped plane. Finally, the B end of the second annular elastic component 8 is welded to the turbine outer ring 9. When the turbine outer ring 9 is thermally expanded by the high-temperature combustion gas, it radially deforms and compresses the first annular elastic component 7, the second annular elastic component 8, and the fourth annular elastic component 17, reducing the force on the turbine housing 18. The turbine outer ring 9, along with the first, second, and fourth annular elastic components, prevents the conduction and radiation of heat from the high-temperature combustion gas to the turbine housing 18, reducing the thermal deformation and thermal stress of the turbine housing 18.

[0031] Preferably, after the guide vane 11 is welded to the guide outer ring 12, the guide outer ring 12 is installed into the third annular elastic part 13 and the fourth annular elastic part 17, and the guide outer ring 12 is pressed tightly against the end face of the turbine casing 18, which serves to provide axial positioning and transmit axial force. The guide outer ring 12, the third annular elastic part 13, and the turbine casing 18 are connected by the second locking washer 14, the screw 15, and the second nut 16, which circumferentially positions these three parts. The turbine casing 18 can move up and down relative to the screw 15.

[0032] In this embodiment, a first stepped plane 121 is formed at the connection between the guide outer ring 12 and the turbine casing 18 for axial positioning by fitting with the turbine casing 18. A second stepped plane 122 is formed on the outer side of the guide outer ring 12 relative to the fourth annular elastic member 17 for axial positioning by fitting with one end of the turbine outer ring 9. Correspondingly, a third stepped plane 61 is formed on the other end of the turbine outer ring 9 for axial positioning with the other end of the turbine outer ring 9. Preferably, the length of the turbine outer ring 9 is less than the distance between the indentation of the second and third stepped planes to provide space for deformation of the turbine outer ring 9. The second annular elastic member 8 serves as a reinforcement to further enhance the elastic force between the turbine casing 18 and the turbine outer ring 9.

[0033] The turbine outer ring 9 and the second, third, and fourth annular elastic components of this invention prevent the conduction and radiation of heat from the high-temperature combustion gas to the turbine casing 18, keeping the connecting flanges of the combustion chamber casing, turbine casing, and rear bearing casing in a lower temperature range and with similar operating temperatures. This achieves coordinated thermal deformation of the three casings at the connecting flanges. In this embodiment, the first annular elastic component 7 has a specific stiffness, which acts as a centering element when the guide outer ring 12 is installed inside the turbine casing 18, and can be compressed during thermal expansion of the guide outer ring. The second, third, and fourth annular elastic components also have specific stiffness, and when the turbine outer ring 9 undergoes thermal expansion under the action of high-temperature combustion gas, the second, third, and fourth annular elastic components maintain the turbine outer ring's cylindrical shape as much as possible. The turbine casing 18 has several vent holes for cooling its internal elastic elements, turbine outer ring, etc.

[0034] Although the present invention has been described herein with reference to illustrative embodiments, the above embodiments are merely preferred embodiments of the present invention, and the implementation of the present invention is not limited to the above embodiments. It should be understood that those skilled in the art can devise many other modifications and implementations, which will fall within the scope and spirit of the principles disclosed in this application.

Claims

1. A thermal deformation coordination structure for an aero-engine turbine casing assembly, characterized in that, include: The turbine outer ring (9) is located outside the turbine blade (10), the turbine casing (18) and the rear bearing casing (6) are axially connected, and the guide ring (12) is radially connected to the turbine casing (18). The turbine outer ring (9) is limited between the guide ring (12) and the turbine casing (18), and the outer ends of the turbine outer ring (9) are respectively provided with annular elastic parts corresponding to the turbine casing (18) and the rear bearing casing (6). The annular elastic parts are used to radially position the guide ring (12) and provide space for the radial thermal expansion of the turbine outer ring (9), thereby reducing the thermal stress of the turbine outer ring (9) and the turbine casing (18).

2. The thermal deformation coordination structure for an aero-engine turbine casing assembly according to claim 1, characterized in that, The annular elastic component includes at least a second annular elastic component (8) and a fourth annular elastic component (17) disposed at both ends of the turbine outer ring (9) to buffer the radial deformation of the turbine outer ring (9) and isolate the high-temperature gas to reduce the conduction and radiation of the high-temperature gas heat to the turbine casing (18).

3. The thermal deformation coordination structure for an aero-engine turbine casing assembly according to claim 2, characterized in that, One end of the second annular elastic part (8) is welded to the outer side of the turbine casing (18) near the combustion chamber casing (1), and the other end is welded to the top of the turbine outer ring (9); The fourth annular elastic part (17) is welded at both ends to the lower part of the middle of the turbine casing (18) and abuts against the end of the turbine outer ring (9) away from the combustion chamber casing (1).

4. The thermal deformation coordination structure for an aero-engine turbine casing assembly according to claim 2, characterized in that, The annular elastic component also includes a first annular elastic component (7), the two ends of which are welded to the top of the turbine outer ring (9), and the outer side abuts against the inside of the rear bearing housing (6).

5. The thermal deformation coordination structure for an aero-engine turbine casing assembly according to claim 4, characterized in that, The guide outer ring (12) and the rear bearing housing (6) are respectively formed with a second step plane (122) and a third step plane (61) at both ends of the turbine outer ring (9), which are used to fit into the two ends of the turbine outer ring (9) to form axial positioning.

6. The thermal deformation coordination structure for an aero-engine turbine casing assembly according to claim 2, characterized in that, The annular elastic component also includes a third annular elastic component (13), the two ends of which are welded to the lower end of the turbine casing (18) away from the combustion chamber casing (1), and the outer side abuts against the outer ring (12) of the guide.

7. The thermal deformation coordination structure for an aero-engine turbine casing assembly according to claim 6, characterized in that, The outer ring (12) of the guide is connected to the end of the turbine casing (18) away from the rear bearing casing (6) by a number of screws (15) to achieve circumferential positioning.

8. The thermal deformation coordination structure for an aero-engine turbine casing assembly according to claim 7, characterized in that, The outer ring (12) of the guide is connected to the turbine casing (18) with a first stepped plane (121) for axial positioning by fitting with the turbine casing (18).

9. The thermal deformation coordination structure for an aero-engine turbine casing assembly according to claim 7, characterized in that, One end of the screw (15) passes through the second nut (16), the second stop washer (14), the turbine casing (18) and the third annular elastic part (13) in sequence and is threaded to the outer ring (12) of the guide. The turbine casing (18) can move up and down relative to the screw (15).

10. A thermal deformation coordination structure for an aero-engine turbine casing assembly according to any one of claims 1-9, characterized in that, The turbine casing (18) has several ventilation holes for cooling its internal components; A combustion chamber casing (1) is provided between the turbine casing (18) and the rear bearing casing (6). The turbine casing (18), the rear bearing casing (6) and the combustion chamber casing (1) are connected by several locking washers (3), bolts (4), mounting positioning nuts (2) and a first nut (5). The combustion chamber casing (1) is embedded with a mounting positioning nut (2). One end of the bolt (4) passes through the turbine casing (18), the combustion chamber casing (1), the rear bearing casing (6) and the locking washer (3) in sequence and is connected to the first nut (5). After the connection is completed, the locking lug of the locking washer (3) is bent to the outside of the first nut (5) to prevent loosening.

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