An engine high- and low-pressure turbine cooling structure and method with interstage support plates

Abstract

The present invention discloses a high- and low-pressure turbine cooling structure and method for an engine with an interstage support plate, the cooling structure comprising: an internal pipeline arranged between the engine casing and the third bushing, for introducing cooling air into the second cavity outside the bearing seat; the interstage support plate is arranged between the first bushing, the second bushing and the casing, and forms a fifth cavity between the interstage support plate and the casing, and forms a fourth cavity between the interstage support plate and the first bushing and the second bushing; the present invention forms a two-way flow bleed air design between the internal channels of the support plate by cooperating and sealing the parts between the interstage support plate, the internal pipeline and the third bushing, making the internal cavity structure of the support plate more compact, greatly improving the utilization rate of the internal channels of the interstage support plate, and the exhaust gas through the internal channels of the support plate is first used to cool the multiple transverse ribs arranged inside the interstage support plate, and then is discharged directly from the casing 1 to the engine outer duct through the fifth cavity, thereby increasing thrust and improving the utilization efficiency of the engine bleed air.

Description

Technical Field

[0001] The present invention belongs to the technical field of turbine engines, and in particular relates to a cooling structure and method for a high- and low-pressure turbine of an engine with an interstage support plate. Background Art

[0002] Conventional engine high and low pressure bearing cooling and sealing bleed air solutions through interstage support plates are as follows Figure 1 As shown, part of the sealing air between the rear stages of the high-pressure turbine is used for sealing the rear of the high-pressure disc, and the other part is used for sealing the interstage bearing seat through the grate teeth, and is mixed with the bleed air from the interstage support plate through the grate teeth on the right side of the bearing seat, and is used for sealing the interstage in front of the low-pressure disc.

[0003] In conventional technical solutions, as the thermal cycle parameters of aircraft engines increase, especially when the compressor pressure ratio reaches 24, the sealing air behind the high-pressure turbine disc needs to have a higher bleed air pressure, about 0.8 MPa. The corresponding bleed air position temperature can reach 450°C, and the bleed air will have a temperature rise when passing through the turbine disc core and the compressor disc cavity. If this high-temperature air is used to directly seal the bearing cavity, the overall temperature level of the bearing seat will increase, which will bring a large thermal load to the bearing and affect the normal use of the bearing. At the same time, there is a risk of coking of the lubricating oil in the bearing cavity, affecting the normal operation of the engine. Summary of the Invention

[0004] In response to the above problems, on the one hand, the present invention discloses a cooling structure for a high- and low-pressure turbine of an engine with an interstage support plate, wherein the high- and low-pressure turbine of the engine comprises a first liner, a second liner, and a third liner, wherein a first cavity is formed between the first liner and the third liner, and a third cavity is formed between the second liner and the third liner, and the cooling structure comprises:

[0005] an internal pipeline, the internal pipeline being arranged between the casing of the engine and the third bushing and being used for passing cooling air into the second cavity outside the bearing seat;

[0006] an interstage support plate, disposed between the first liner, the second liner and the casing, and forming a fifth cavity between the interstage support plate and the casing, and forming a fourth cavity between the interstage support plate and the first liner and the second liner;

[0007] The interstage support plate includes an upper edge plate, a support plate, and a lower edge plate. A fifth cavity is formed between the upper edge plate and the casing, and a fourth cavity is formed between the lower edge plate and the first bushing and the second bushing. The support plate is disposed between the upper edge plate and the lower edge plate to support and fix the upper edge plate and the lower edge plate.

[0008] The internal pipeline passes through the casing, the upper edge plate, the lower edge plate and the third bushing in sequence and is inserted into the third cavity.

[0009] Furthermore, the bearing seat is arranged on a rotating base, and the side of the rotating base facing the high-pressure turbine disk is set as the high-pressure side, and the side facing the low-pressure turbine disk is set as the low-pressure side;

[0010] The two sides of the rotating base form a dynamic seal with the third bushing through the first grate teeth and the third grate teeth respectively;

[0011] The two sides of the bearing seat are respectively sealed with carbon seals and the second grate teeth to form dynamic seals with the rotating base;

[0012] The space formed between the top of the bearing seat, the rotating base and the third bushing is defined as a second cavity.

[0013] Furthermore, the second bushing and the third bushing on the high-pressure side are connected to the first bushing via a first bolt, and the second bushing and the third bushing on the low-pressure side are fixedly connected via a second bolt.

[0014] Furthermore, the end of the rotating base on the high-pressure side is provided with two side branch mounting plates, which respectively form dynamic seals with the first bushing and the third bushing through the brush seal and the first grate teeth;

[0015] A first cavity is formed between the two-way bypass mounting plate and the first and third bushings.

[0016] Furthermore, a sixth cavity is formed between the outer side of the first bushing on the high-pressure side and the high-pressure turbine disc. The sixth cavity is interconnected with the first cavity through a brush seal, and the first cavity and the fourth cavity are interconnected.

[0017] Furthermore, the rotating base on the low-pressure side is fixedly connected to the low-pressure turbine disc by a bolt assembly;

[0018] A dynamic seal is formed between the low-pressure turbine disc and the second bushing via the fourth grate teeth;

[0019] A seventh cavity is formed between the low-pressure turbine disk on the low-pressure side and the second liner and the third liner.

[0020] Furthermore, the middle portions of the upper edge plate and the lower edge plate are provided with openings, the top two sides of the upper edge plate are clamped to the casing, and the bottom two sides of the lower edge plate are respectively connected to the first bushing and the second bushing;

[0021] The support plate is an annular structure, and the first and second ends of the annular structure are fixedly connected to the upper edge plate and the lower edge plate respectively. A support plate cavity is formed between the support plate and the upper edge plate and the lower edge plate. The fourth cavity and the fifth cavity are connected to each other through the support plate cavity.

[0022] Furthermore, a plurality of transverse ribs are provided on the inner wall of the annular structure of the support plate.

[0023] Furthermore, the internal pipeline passes through the support plate cavity, and a mounting flange is provided on the top of the internal pipeline for connecting the internal pipeline with the outer wall of the casing, and an interface is provided on the top of the internal pipeline for receiving cooling air delivered by an external air source.

[0024] Furthermore, the top of the internal pipeline also forms a contact seal with the inner wall of the mounting hole of the casing;

[0025] The bottom of the internal pipeline is set to be conical, and cooperates with the third bushing to form a conical surface seal.

[0026] On the other hand, the present invention also discloses a method for cooling high- and low-pressure turbines of an engine with interstage support plates, the method comprising:

[0027] An external air source introduces cooling air into the internal pipeline through an interface;

[0028] The cooling air passes through the internal pipeline and enters the second cavity, and acts on the surface of the bearing seat to perform surface blowing cooling;

[0029] A portion of the cooling gas passes through the second cavity and acts on the bearing cavity to seal the bearing cavity;

[0030] Another part of the cooling gas passes through the second cavity and enters the seventh cavity to seal the interstage before the low-pressure turbine disk;

[0031] Another part of the cooling air also passes through the second cavity, enters the first cavity, mixes with the high-temperature airflow flowing into the first cavity from the sixth cavity, and is discharged to the engine duct through the fourth cavity and the fifth cavity in sequence.

[0032] Compared with the prior art, the present invention has the following beneficial effects:

[0033] 1) The interstage support plate, internal piping, and third bushing are sealed together to create a bidirectional air flow design within the internal channels of the support plate. This makes the internal structure of the support plate more compact and significantly improves the utilization rate of the internal channels of the interstage support plate.

[0034] 2) A separate stream of cooling air is introduced from the outside to meet the cooling and sealing requirements of the bearing seat below the hot end component, effectively reducing the thermal load on the bearing seat below the hot end component and ensuring the normal operation of the engine bearing. Furthermore, a pressure relief exhaust system is formed in the first chamber through the brush seal and the first grate teeth, making the pressure in the first chamber lower than that in the sixth and second chambers. This effectively prevents the high-temperature sealing air from the sixth chamber behind the high-pressure plate from directly entering the bearing chamber, which would otherwise cause risks such as high bearing thermal load, oil coking, and high oil temperature rise.

[0035] 3) The exhaust gas passing through the internal channel of the support plate (a mixture of a small amount of high-temperature air leaked from the sixth cavity and the cooling air from the first cavity) is first used to cool the interstage guide vane (a plurality of transverse ribs provided on the inner cavity wall of the support plate), and then is discharged directly from the casing to the engine outer duct through the fifth cavity, thereby increasing thrust and improving the utilization efficiency of the engine bleed air.

[0036] Other features and advantages of the present invention will be described in the following description, and in part will become apparent from the description, or will be understood by practicing the present invention. The purpose and other advantages of the present invention can be realized and obtained by the structures pointed out in the description, claims and drawings. BRIEF DESCRIPTION OF THE DRAWINGS

[0037] In order to more clearly illustrate the embodiments of the present invention or the technical solutions in the prior art, the following is a brief introduction to the drawings required for use in the embodiments or the description of the prior art. Obviously, the drawings described below are some embodiments of the present invention. For ordinary technicians in this field, other drawings can be obtained based on these drawings without paying any creative work.

[0038] Figure 1 The figure shows a high-pressure turbine and low-pressure turbine bearing cooling structure according to the prior art;

[0039] Figure 2 A schematic diagram of a bearing cooling structure between high-pressure and low-pressure turbines according to an embodiment of the present invention is shown;

[0040] Figure 3 A schematic diagram of an interstage support plate of a cooling structure according to an embodiment of the present invention is shown;

[0041] Figure 4 A schematic diagram of the airflow direction of the high- and low-pressure turbine inter-bearing cooling structure according to an embodiment of the present invention is shown.

[0042] In the accompanying drawings: 1. Casing; 2. Internal piping; 3. Upper edge plate; 4. Support plate; 5. Lower edge plate; 6. First bushing; 7. Second bushing; 8. Third bushing; 9. Low-pressure turbine disc; 10. High-pressure turbine disc; 11. Bearing seat; 12. Rotating base; 13. Carbon seal; 14. First grate teeth; 15. Second grate teeth; 16. Third grate teeth; 17. Fourth grate teeth; 18. Brush seal; 19. First chamber; 20. Second chamber; 21. Third chamber; 22. Fourth chamber; 23. Fifth chamber; 24. Sixth chamber; 25. Seventh chamber; 26. O-ring; 27. First bolt; 28. Second bolt; 29. ​​Interface. DETAILED DESCRIPTION

[0043] To make the objectives, technical solutions, and advantages of the embodiments of the present invention more clear, the technical solutions in the embodiments of the present invention will be clearly and completely described below in conjunction with the accompanying drawings in the embodiments of the present invention. Obviously, the described embodiments are only part of the embodiments of the present invention, not all of the embodiments. Based on the embodiments of the present invention, all other embodiments obtained by ordinary technicians in this field without making creative efforts shall fall within the scope of protection of the present invention.

[0044] The present invention proposes a high- and low-pressure turbine cooling structure of an engine with an interstage support plate, wherein the high- and low-pressure turbine of the engine includes a first bushing 6, a second bushing 7 and a third bushing 8, a first cavity 19 is formed between the first bushing 6 and the third bushing 8, and a third cavity 21 is formed between the second bushing 7 and the third bushing 8, and the cooling structure includes an internal pipeline 2 and an interstage support plate. The internal pipeline 2 is arranged between the casing 1 and the third bushing 8 of the engine, and is used to pass cooling air into the second cavity 20 outside the bearing seat 11; the interstage support plate is arranged between the first bushing 6, the second bushing 7 and the casing 1, and forms a fifth cavity 23 between the interstage support plate and the casing 1, and forms a fourth cavity 22 between the interstage support plate and the first bushing 6 and the second bushing 7; wherein, the interstage support plate includes an upper edge plate 3, a support plate 4 and a lower edge plate 5, a fifth cavity 23 is formed between the upper edge plate 3 and the casing 1, a fourth cavity 22 is formed between the lower edge plate 5 and the first bushing 6 and the second bushing 7, and the support plate 4 is arranged between the upper edge plate 3 and the lower edge plate 5, and is used to support and fix the upper edge plate 3 and the lower edge plate 5; the internal pipeline 2 passes through the casing 1, the upper edge plate 3, the lower edge plate 5 and the third bushing 8 in sequence, and is inserted into the third cavity 21.

[0045] refer to Figure 2 As shown, the bearing seat 11 is arranged on the rotating base 12, and the side of the rotating base 12 facing the high-pressure turbine disk 10 is set as the high-pressure side, and the side facing the low-pressure turbine disk 9 is set as the low-pressure side; the two sides of the rotating base 12 form a dynamic seal with the third bushing 8 through the first grate 14 and the third grate 16 respectively; the two sides of the bearing seat 11 form a dynamic seal with the rotating base 12 through the carbon seal 13 and the second grate 15 respectively; the space formed between the top of the bearing seat 11 and the rotating base 12 and the third bushing 8 is defined as the second cavity 20.

[0046] The second and third bushings 7 and 8 on the high-pressure side are connected to the first bushing 6 via first bolts 27, and the second and third bushings 7 and 8 on the low-pressure side are fixedly connected via second bolts 28. A two-way bypass mounting plate is provided at the end of the rotating base 12 on the high-pressure side, forming a dynamic seal with the first and third bushings 6 and 8 via a brush seal 18 and a first grate 14, respectively; a first cavity 19 is formed between the two-way bypass mounting plate and the first and third bushings 6 and 8.

[0047] A sixth chamber 24 is formed between the outside of the first high-pressure-side bushing 6 and the high-pressure turbine disk 10. This sixth chamber 24 communicates with the first chamber 19 via a brush seal 18, and the first chamber 19 communicates with the fourth chamber 22. The rotating base 12 on the low-pressure side is fixedly connected to the low-pressure turbine disk 9 via a bolt assembly; a dynamic seal is formed between the low-pressure turbine disk 9 and the second bushing 7 via the fourth grate 17; and a seventh chamber 25 is formed between the low-pressure turbine disk 9 on the low-pressure side and the second bushing 7 and the third bushing 8.

[0048] It is necessary to explain that the brush seal 18 can effectively reduce the leakage flow of high-temperature air from the sixth chamber 24 to the first chamber 19, reduce the temperature of the mixed gas, and create favorable conditions for cooling the support plate 4. At the same time, the air discharged to the outer duct mainly comes from a position with lower pressure, which reduces the impact on the performance of the whole machine. The first grate teeth 14 and the pressure relief exhaust ensure that the pressure of the second chamber 20 is higher than that of the first chamber 19, and the air in the first chamber 19 will not flow back into the second chamber 20. The design ensures that there is a pressure difference margin between them. In addition, the materials used for the carbon seal 13 include flexible graphite sealing materials, reinforced graphite sealing materials and carbon fiber composite materials. Based on the self-lubricating properties and heat resistance of the carbon sealing materials, carbon sealing materials are widely used in the aerospace field. For example, it replaces some asbestos materials and rubber sealing materials used in the aerospace field, and high-strength graphite sealing rings used in turbo pumps of aerospace engines.

[0049] refer to Figure 2 As shown, the brush seal 18 is fixedly mounted on the bottom of the first bushing 6 and is a stator component. The corresponding rotating base 12 is a rotating component, and a runway is provided on the side mounting plate of the rotating base 12 to form a dynamic seal with the brush seal 18. The first grate teeth 14, the second grate teeth 15 and the third grate teeth 16 are respectively mounted on the rotating base 12, and are used to cooperate with the structural parts corresponding to the first grate teeth 14, the second grate teeth 15 and the third grate teeth 16 to form a dynamic seal. Further, the low-pressure turbine disc 9 and the rotating base 12 are fixed by a bolt assembly, that is, the low-pressure turbine disc 9 and the rotating base 12 are both rotating components, and the fourth grate teeth 17 are mounted on the low-pressure turbine disc 9 to form a dynamic seal with the third bushing 8.

[0050] In one embodiment of the present invention, the present invention controls the sealing air volume between the forward rotating static stages of the low-pressure rotor (low-pressure turbine disk 9) through two third grate teeth 16 and fourth grate teeth 17 with different radial heights, and adopts a double grate tooth sealing design. Even if one of the grate teeth seals fails (the gap is very large), the other grate teeth can ensure that the sealing air volume between the front stages of the low-pressure rotor is at an appropriate level, thereby preventing the sealing air volume between the front stages of the low-pressure rotor from being too large due to grate tooth failure, and preventing the cavity pressure of the second cavity 20 from being greatly reduced, causing the high-temperature air behind the high-pressure turbine disk 10 to be poured into the bearing cavity.

[0051] refer to Figure 3 As shown, the upper and lower edge plates 3 and 5 have openings in their middle portions. The top and sides of the upper edge plate 3 engage the casing 1, while the bottom and sides of the lower edge plate 5 are connected to the first and second bushings 6 and 7, respectively. The support plate 4 is an annular structure, with the ends of the annular structure fixedly connected to the upper and lower edge plates 3 and 5, respectively. A support plate cavity is formed between the support plate 4, the upper and lower edge plates 3, and the lower edge plates 5. The fourth and fifth cavities 22 and 23 are connected to each other through the support plate cavity. Several transverse ribs are provided on the inner wall of the annular structure of the support plate 4.

[0052] In one embodiment of the present invention, the support plate 4, upper edge plate 3, and lower edge plate 5 are integrally formed using a monolithic casting process, effectively improving structural stability. The internal pipeline 2 runs through the support plate cavity, with cooling air being transported from the top to the bottom of the internal pipeline 2. During exhaust, a mixture of cooling air and high-temperature air is transported from the fourth cavity 22 to the fifth cavity 23 along the outer wall of the internal pipeline 2. The gas transport directions during intake and exhaust are opposite. This bidirectional air flow structure of the interstage support plate isolates the air intake and pressure relief exhaust within the support plate 4 from each other, increasing the compactness of the support plate 4 and significantly improving the utilization rate of the internal channels of the interstage support plate.

[0053] Specifically, during the cooling process, the pressure and temperature of the sixth chamber 24 are the highest, while the pressure and temperature of the second chamber 20 are lower than those of the sixth chamber 24. The pressure relief technology effectively reduces the pressure of the first chamber 19, making the pressure of the first chamber 19 lower than that of the sixth chamber 24 and the second chamber 20. This allows the air in the sixth chamber 24 and the second chamber 20 to mix in the first chamber 19 and be discharged to the engine duct through the support plate 4. Directly discharging the mixed air in the first chamber 19 to the duct through the channel between the inner wall of the interstage support plate and the bleed air line not only increases engine thrust but also reduces the pressure in the first chamber 19, preventing the high-temperature, high-pressure sealing air from the interstage after the high-pressure turbine disk 10 from directly contacting the bearing seat 11 and entering the bearing cavity. This allows a stream of low-temperature, low-pressure cooling air to cool the bearing seat 11 and seal the bearing cavity, and inhibits the high-temperature air from the high-pressure turbine disk 10 from directly entering the bearing cavity.

[0054] refer to Figure 2 As shown, the top of internal conduit 2 is provided with a mounting flange for matingly connecting it to the outer wall of casing 1. A port 29 is also provided at the top of internal conduit 2 for receiving cooling air from an external air source. The top of internal conduit 2 also forms a contact seal with the inner wall of the mounting hole in casing 1. The bottom of internal conduit 2 is tapered and forms a conical seal with third bushing 8.

[0055] In one embodiment of the present invention, a first groove is provided on internal conduit 2, 6 mm below the mounting flange, for receiving an O-ring 26. This groove forms a seal with casing 1, preventing air from leaking from fifth chamber 23. The bottom of internal conduit 2 is tapered, forming a conical seal with third bushing 8. A second groove is provided 6 mm above the bottom of the tapered surface, for receiving an O-ring 26. This seal forms a seal with third bushing 8, isolating air from fourth chamber 22 and third chamber 21. During assembly, internal conduit 2 is inserted directly through the pre-set mounting opening on casing 1 and into the bottom of the tapered seal hole in third bushing 8.

[0056] On the other hand, the present invention also provides a method for cooling high- and low-pressure turbines of an engine with interstage support plates, the method comprising:

[0057] The external air source introduces cooling air into the internal pipeline 2 through the interface 29;

[0058] The cooling air passes through the internal pipe 2 and enters the second cavity 20, and acts on the surface of the bearing seat 11 to perform surface blowing cooling;

[0059] A portion of the cooling gas passes through the second cavity 20 and acts on the bearing cavity to seal the bearing cavity;

[0060] Another part of the cooling gas passes through the second cavity 20 and enters the seventh cavity 25 to perform sealing between the front stages of the low-pressure turbine disk 9;

[0061] Another part of the cooling air also passes through the second cavity 20, enters the first cavity 19, mixes with the high-temperature airflow flowing into the first cavity 19 from the sixth cavity 24, and is discharged to the engine duct through the fourth cavity 22 and the fifth cavity 23 in sequence.

[0062] refer to Figure 4 As shown, specifically, low-temperature (about 200°C) cooling air flows through the internal pipe 2 to the second cavity 20 for blowing cooling on the surface of the bearing seat 11. A portion of the airflow passes through the carbon seal 13 and the second grate 15 for sealing the bearing cavity. The right side of the bearing seat 11 bleeds air through the third grate 16 and the fourth grate 17 to leak into the main channel for sealing between the low-pressure turbine forward static stage. The first cavity 19 and the fourth cavity 22 are connected, and the fourth cavity 22 and the fifth cavity 23 are connected through the channel formed between the pipeline and the inner wall of the support plate 4. On the one hand, the present invention increases thrust by discharging the air in the fifth cavity 23 to the engine outer duct, effectively reducing the pressure of the decompression first cavity 19, making the pressure of the first cavity 19 lower than that of the sixth cavity 24 and the second cavity 20, which can effectively prevent the high-temperature air (about 500°C) in the sixth cavity 24 behind the high-pressure turbine disk 10 from directly heating the bearing seat 11 and the sealed bearing cavity. On the other hand, the present invention arranges transverse ribs inside the support plate 4, which can enhance the heat exchange effect between the exhaust gas and the inner wall of the support plate 4, and has a better cooling effect on the support plate 4.

[0063] To achieve a sealed bearing cavity by introducing a separate stream of cooling air and prevent high-temperature air from entering the bearing cavity from behind the high-pressure turbine disc 10, the present invention separates a first chamber 19 from a second chamber 20 using a brush seal. Cooling air is then introduced from the outside into the internal pipeline 2, where it enters the second chamber 20 to cool and seal the bearing cavity. By depressurizing and exhausting the first chamber 19, the cavity pressure is reduced, preventing high-temperature air from the sixth chamber 24 from directly entering the bearing cavity. Simultaneously, the depressurized air from the first chamber 19 is passed through the inner cavity of the support plate 4 for enhanced flow and heat exchange, cooling the support plate 4. The air is then discharged directly to the engine duct through an exhaust port on the casing 1 to increase thrust.

[0064] The technical solution of the present invention has been shown through relevant one-dimensional calculation and analysis that, under the premise of ensuring that the sealing of the high-pressure turbine rear stage meets the requirements, a cooling air is introduced through the internal pipeline 2 arranged in the support plate 4 to cool the bearing seat 11, the sealing bearing cavity and the low-pressure front stage seal, and a pressure relief first cavity 19 is formed through the brush seal 18 and the first grate teeth 14, so that the sealing air and cooling air with higher temperature in the high-pressure turbine rear stage are discharged through the channel between the inner wall of the support plate 4 and the air bleed pipeline. The solution is reasonable and feasible, which not only improves the compactness of the engine interstage support plate and the utilization rate of the internal channel space, but also the mixed exhaust gas can increase the engine outer duct thrust. When the engine is working at the maximum state, the adoption of this solution can reduce the temperature of the bearing seat 11 by about 70°C, while reducing the temperature rise of the lubricating oil in the bearing cavity and the risk of lubricating oil coking.

[0065] Although the present invention has been described in detail with reference to the aforementioned embodiments, those skilled in the art should understand that they can still modify the technical solutions described in the aforementioned embodiments, or make equivalent replacements for some of the technical features therein; and these modifications or replacements do not cause the essence of the corresponding technical solutions to deviate from the spirit and scope of the technical solutions of the embodiments of the present invention.

Claims

1. A high-pressure and low-pressure turbine cooling structure of an engine with an interstage support plate, wherein the high-pressure and low-pressure turbine of the engine comprises a first bushing (6), a second bushing (7) and a third bushing (8), a first cavity (19) is formed between the first bushing (6) and the third bushing (8), and a third cavity (21) is formed between the second bushing (7) and the third bushing (8), characterized in that: The cooling structure comprises: An internal pipeline (2), the internal pipeline (2) being arranged between the casing (1) and the third bushing (8) of the engine and being used for passing cooling air into the second cavity (20) outside the bearing seat (11); an interstage support plate, disposed between the first bushing (6), the second bushing (7) and the casing (1), and forming a fifth cavity (23) between the interstage support plate and the casing (1), and forming a fourth cavity (22) between the interstage support plate and the first bushing (6) and the second bushing (7); The interstage support plate comprises an upper edge plate (3), a support plate (4) and a lower edge plate (5); a fifth cavity (23) is formed between the upper edge plate (3) and the casing (1); a fourth cavity (22) is formed between the lower edge plate (5) and the first bushing (6) and the second bushing (7); the support plate (4) is arranged between the upper edge plate (3) and the lower edge plate (5) to support and fix the upper edge plate (3) and the lower edge plate (5); The internal pipeline (2) passes through the casing (1), the upper edge plate (3), the lower edge plate (5) and the third bushing (8) in sequence and is inserted into the third cavity (21).

2. The cooling structure according to claim 1, characterized in that: The bearing seat (11) is arranged on a rotating base (12), and the side of the rotating base (12) facing the high-pressure turbine disc (10) is set as the high-pressure side, and the side facing the low-pressure turbine disc (9) is set as the low-pressure side; Both sides of the rotating base (12) form a dynamic seal with the third bushing (8) through the first grate teeth (14) and the third grate teeth (16); Dynamic seals are formed between the two sides of the bearing seat (11) and the rotating base (12) through the carbon seal (13) and the second grate teeth (15); The space formed between the top of the bearing seat (11), the rotating base (12) and the third bushing (8) is defined as a second cavity (20).

3. The cooling structure according to claim 2, characterized in that: The side of the second bushing (7) facing the high-pressure side and the third bushing (8) are connected to the first bushing (6) via a first bolt (27), and the second bushing (7) and the third bushing (8) on the low-pressure side are fixedly connected via a second bolt (28).

4. The cooling structure according to claim 2, characterized in that: The end of the rotating base (12) on the high-pressure side is provided with two side branch mounting plates, which respectively form dynamic seals with the first bushing (6) and the third bushing (8) through a brush seal (18) and a first grate tooth (14); A first cavity (19) is formed between the two-way bypass mounting plate and the first bushing (6) and the third bushing (8).

5. The cooling structure according to claim 2, characterized in that: A sixth cavity (24) is formed between the outer side of the first bushing (6) on the high-pressure side and the high-pressure turbine disc (10). The sixth cavity (24) is communicated with the first cavity (19) through a brush seal (18), and the first cavity (19) and the fourth cavity (22) are communicated with each other.

6. The cooling structure according to claim 2, characterized in that: The rotating base (12) on the low-pressure side is fixedly connected to the low-pressure turbine disc (9) via a bolt assembly; A dynamic seal is formed between the low-pressure turbine disc (9) and the second bushing (7) via a fourth grate tooth (17); A seventh cavity (25) is formed between the low-pressure turbine disc (9) on the low-pressure side and the second liner (7) and the third liner (8).

7. The cooling structure according to claim 1, characterized in that: The middle parts of the upper edge plate (3) and the lower edge plate (5) are provided with openings, the top two sides of the upper edge plate (3) are clamped with the casing (1), and the bottom two sides of the lower edge plate (5) are respectively connected to the first bushing (6) and the second bushing (7); The support plate (4) is an annular structure, and the first and second ends of the annular structure are fixedly connected to the upper edge plate (3) and the lower edge plate (5), respectively. A support plate cavity is formed between the support plate (4) and the upper edge plate (3) and the lower edge plate (5), and the fourth cavity (22) and the fifth cavity (23) are connected to each other through the support plate cavity.

8. The cooling structure according to claim 7, characterized in that: A plurality of transverse ribs are provided on the inner wall of the annular structure of the support plate (4).

9. The cooling structure according to claim 7, characterized in that: The internal pipeline (2) passes through the support plate cavity, and a mounting flange is provided at the top of the internal pipeline (2) for connecting the internal pipeline (2) with the outer wall surface of the casing (1); and an interface (29) is provided at the top of the internal pipeline (2) for receiving cooling air delivered by an external air source.

10. The cooling structure according to any one of claims 1 to 9, characterized in that: The top of the internal pipeline (2) also forms a contact seal with the inner wall of the mounting hole of the casing (1); The bottom of the internal pipeline (2) is configured to be conical, and cooperates with the third bushing (8) to form a conical surface seal.

11. A method for cooling high and low pressure turbines of an engine with interstage support plates, characterized in that: The method comprises: The external air source passes cooling air into the internal pipeline (2) through the interface (29); The cooling air passes through the internal pipeline (2) and enters the second cavity (20) outside the bearing seat (11), acts on the surface of the bearing seat (11), and performs surface blowing cooling; A portion of the cooling gas passes through the second cavity (20) and acts on the bearing cavity to seal the bearing cavity; Another portion of the cooling gas passes through the second cavity (20) and enters the seventh cavity (25) to perform sealing between the front stages of the low-pressure turbine disc (9); wherein the seventh cavity (25) is formed between the low-pressure turbine disc (9) on the low-pressure side and the second liner (7) and the third liner (8); Another part of the cooling air also passes through the second cavity (20) and enters the first cavity (19), and after being mixed with the high-temperature airflow flowing from the sixth cavity (24) into the first cavity (19), it passes through the fourth cavity (22) and the fifth cavity (23) in sequence and is discharged to the engine outer duct; wherein, the first cavity (19) is formed between the first bushing (6) and the third bushing (8), and the sixth cavity (24) is formed between the outer side of the first bushing (6) on the high-pressure side and the high-pressure turbine disc (10); the interstage support plate is arranged between the first bushing (6), the second bushing (7) and the casing (1), and forms a fifth cavity (23) between the interstage support plate and the casing (1), and forms a fourth cavity (22) between the interstage support plate and the first bushing (6) and the second bushing (7).

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