Sealing structure and aero-engine

Abstract

The invention relates to the technical field of aero-engines, and discloses a sealing structure and an aero-engine, the sealing structure comprises: a turbine shaft, the outer wall of which is provided with a first conveying hole; the turbine disc comprises a disc shaft and a final-stage disc, the disc shaft is fixedly arranged on the outer side of the turbine shaft in a sleeving mode, a second conveying hole is formed in the disc shaft, and the first conveying channel communicates with the cooling space and the second conveying channel at the same time through the first conveying hole; the sealing cover is arranged on the outer side of the disc shaft in a sleeving manner; the first sealing assembly is provided with a third conveying hole, a second conveying channel, a second conveying hole, a third conveying hole and a sealing cavity which are sequentially communicated, a conveying path is directly formed through the first conveying hole, the second conveying channel, the second conveying hole and the third conveying hole in a hole opening mode, airflow in the turbine shaft can be directly guided into the sealing cavity, and therefore the sealing effect of the turbine shaft is improved. The flow of airflow flowing into the sealing cavity can be adjusted by adjusting the flow areas of the first conveying hole, the second conveying hole and the third conveying hole, and multi-stage adjustment of the airflow of the sealing cavity is achieved.

Description

Technical Field

[0001] The present invention relates to the technical field of aeroengines, and in particular to a sealing structure and an aeroengine. Background Art

[0002] Aircraft engines are highly complex and sophisticated thermal machines. When designing the power turbine bearing cavity of an aircraft engine, in order to avoid oil leakage from the bearing cavity, it is usually necessary to draw a stream of high-pressure gas from the high-pressure compressor to seal the bearing cavity. During the sealing process, there is a certain pressure difference between the sealing position and the bearing cavity. When the pressure difference is too small, the sealing is likely to be unreliable and cause partial lubricating oil leakage. When the pressure difference is too large, too much sealing gas will flow into the bearing cavity. Therefore, in order to ensure the sealing effect of the power turbine bearing cavity, the pressure difference between the sealing position and the bearing cavity needs to be controlled within a reasonable range.

[0003] At present, the position of the existing power turbine bearing cavity is usually located between the power turbine final stage disk and the sealing cover sleeved on the outside. The sealing form is that a first sealing component is arranged between the sealing cover and the power turbine final stage disk, and a sealing cavity is formed between the sealing cover and the first sealing component. An air hole is opened on the sealing cover and a separate ventilation pipeline is used to connect to the air hole to introduce high-pressure gas into the sealing cavity. In this sealing form, the pressure difference between the sealing cavity and the bearing cavity is usually difficult to effectively adjust, and there is a problem that the sealing requirements cannot be met. Summary of the invention

[0004] In view of this, the present invention provides a sealing structure and an aircraft engine to solve the problem that the sealing form of the existing power turbine bearing cavity is difficult to meet the sealing requirements.

[0005] In the first aspect, the present invention provides a sealing structure, including: a turbine shaft, a first delivery hole is opened on the outer wall, and a first delivery channel for delivering sealing gas is provided inside; a turbine disk, including a disk shaft and a final disk, the disk shaft is fixedly sleeved on the outer side of the turbine shaft, one end of the disk shaft along the axial direction forms a terminal, the disk shaft is fixedly sleeved on the outer side of the terminal, and the space away from the disk shaft along the axial direction forms a cooling space, the space between the inner wall of the terminal and the outer wall of the turbine shaft forms a second delivery channel, a second delivery hole is provided on the disk shaft, and the first delivery channel is connected with the cooling space and the second delivery channel at the same time through the first delivery hole; a sealing cover is sleeved on the outer side of the disk shaft and forms a bearing cavity with the space between the disk shaft and the first sealing component, the inner wall of the sealing cover is sealed with the outer wall of the disk shaft through the first sealing component, the first sealing component has a third delivery hole, the first sealing component cooperates with the sealing cover and encloses to form a sealing cavity, the second delivery channel, the second delivery hole, the third delivery hole and the sealing cavity are connected in sequence.

[0006] Beneficial effect: A delivery path is directly formed by opening holes through the first delivery hole, the second delivery channel, the second delivery hole and the third delivery hole. The airflow inside the turbine shaft can be directly led to the sealing cavity without setting a separate pipeline outside the sealing cover. Since the airflow flowing out of the first delivery hole will flow into the cooling space and the second delivery channel at the same time, the airflow flow at the second delivery channel can be adjusted by adjusting the number, position and size of the first delivery holes. Since the airflow in the second delivery channel will also flow through the second delivery hole and the third delivery hole in turn, the flow of the second delivery hole and the third delivery hole can also be adjusted. The area can also adjust the air flow rate flowing into the sealing cavity. In addition, the flow rate of the air flow discharged outward from the sealing cavity can be adjusted by adjusting the specifications of the first sealing component to further adjust the pressure of the sealing cavity. Through a simple and effective structure, the function of introducing airflow from the turbine axial sealing cavity is realized, and at the same time, a multi-stage adjustment of the air flow in the sealing cavity can be achieved through a variety of adjustment methods, so that the air pressure in the sealing cavity can be adjusted within a larger range. Therefore, the flexibility of adjusting the pressure difference between the sealing cavity and the bearing cavity can be improved, which effectively solves the problem that the sealing form of the existing power turbine bearing cavity is difficult to meet the sealing requirements.

[0007] In an optional embodiment, the space between the sealing cover and the final disk forms an exhaust space, the sealing cavity is located between the bearing cavity and the exhaust space, the first sealing component also has a fourth delivery hole, and the second delivery hole is connected to the exhaust space through the fourth delivery hole.

[0008] Beneficial effect: This type of first sealing component can adjust the flow rate of the airflow delivered from the second delivery hole to the third delivery hole by adjusting the specifications of the fourth delivery hole. On the basis of the diversion adjustment at the first delivery hole, a second diversion adjustment process is added, which further improves the ability to adjust the airflow entering the sealing cavity, and greatly improves the flexibility of adjusting the pressure at the sealing cavity.

[0009] In an optional embodiment, the first sealing component includes a supporting part and two sealing parts, the supporting part is sleeved on the outer wall of the disk shaft, the two sealing parts are located between the outer wall of the supporting part and the inner wall of the sealing cover and are arranged at intervals along the axial direction, the outer wall of the supporting part, the side walls of the two sealing parts and the inner wall of the sealing cover together form a sealing cavity, and the two sealing parts are located between the exhaust space and the bearing cavity.

[0010] Beneficial effect: This type of sealing cavity can further adjust the air pressure in the sealing cavity by adjusting the size of the sealing portion away from the bearing cavity. The third flow diversion adjustment process can be performed here, greatly improving the adjustability of the air pressure in the sealing cavity.

[0011] In an optional embodiment, a diversion cavity is provided inside the support portion, the third delivery hole is opened on the outer peripheral surface of the support portion, the fourth delivery hole is opened on the end surface of the support portion away from the bearing cavity, and the second delivery hole is connected to the third delivery hole and the fourth delivery hole at the same time through the diversion cavity.

[0012] Beneficial effects: This form of support portion is easier to manufacture and has stronger bending resistance. In addition, the diversion cavity can buffer the airflow entering through the second delivery hole, and then discharge it through the third delivery hole and the fourth delivery hole, which can improve the uniformity of the airflow discharged from the third delivery hole and the fourth delivery hole.

[0013] In an optional embodiment, the support portion has an annular groove opening toward the disk shaft, and the support portion is surrounded by the annular groove and the outer peripheral surface of the disk shaft to form a diversion cavity.

[0014] Beneficial effects: This form of diversion chamber structure is simpler and more reliable, and the annular groove is easy to process and manufacture.

[0015] In an optional embodiment, the sealing structure also includes a second sealing component, which is arranged between the sealing cover and the final stage disk and is located in the exhaust space, the space between the second sealing component and the first sealing component in the exhaust space forms a first exhaust cavity, and the space on the other side of the second sealing component in the exhaust space forms a second exhaust cavity.

[0016] Beneficial effect: The second sealing component blocks the airflow flowing from the first exhaust chamber to the second exhaust chamber, increases the pressure in the first exhaust chamber, and prevents excessive airflow in the diversion chamber from flowing into the exhaust space.

[0017] In an optional embodiment, an axial extension line of the second delivery hole intersects with a radial extension line of the disk shaft, an axial extension line of the fourth delivery hole intersects with an axis of the support portion, and the third delivery hole extends radially along the support portion; and / or, the number of the first delivery holes is multiple and they are arranged at circumferential intervals along the outer wall of the turbine shaft, and the first delivery holes are strip holes extending axially.

[0018] Beneficial effect: The second delivery hole and the fourth delivery hole arranged obliquely can increase the heat exchange area between the airflow and the hole walls of the second delivery hole and the fourth delivery hole, thereby improving the cooling effect of the airflow. In addition, this type of first delivery hole has little effect on the pressure of the airflow discharged from the first delivery channel, and also has little effect on the structural strength of the turbine shaft.

[0019] In an optional embodiment, the pressure difference between the sealing chamber and the bearing chamber is greater than 3.5 kPa, the ratio of the cross-sectional area of ​​the second delivery hole to the cross-sectional area of ​​the third delivery hole is in the range of 0.87 to 0.88, the ratio of the cross-sectional area of ​​the fourth delivery hole to the cross-sectional area of ​​the third delivery hole is in the range of 0.093 to 0.095, and the ratio of the flow rate of the second delivery hole to the flow rate of the third delivery hole is in the range of 0.3 to 0.8.

[0020] Beneficial effect: The conveying hole under this proportional relationship can achieve efficient sealing.

[0021] In an optional embodiment, an annular groove opening toward the turbine shaft is formed on the inner wall of the disk shaft, and a second delivery hole is formed on the inner wall of the annular groove away from the turbine shaft. The space inside the annular groove forms a connecting cavity, and the second delivery channel is connected to the second delivery hole through the connecting cavity. The ratio of the pressure at the connecting cavity to the pressure of the second exhaust cavity is in the range of 2.8 to 3.2, and the ratio of the pressure at the connecting cavity to the pressure of the bearing cavity is in the range of 1.8 to 2.2.

[0022] Beneficial effects: The structure is simple and reliable, and the connecting cavity can buffer the airflow flowing out of the second conveying channel.

[0023] In a second aspect, the present invention further provides an aircraft engine, which comprises: the above-mentioned sealing structure. BRIEF DESCRIPTION OF THE DRAWINGS

[0024] In order to more clearly illustrate the specific implementation methods of the present invention or the technical solutions in the prior art, the drawings required for use in the specific implementation methods or the description of the prior art will be briefly introduced below. Obviously, the drawings described below are some implementation methods of the present invention. For ordinary technicians in this field, other drawings can be obtained based on these drawings without paying creative work.

[0025] Figure 1 A schematic diagram of the arrangement of a sealing structure in a cutaway state according to an embodiment of the present invention;

[0026] Figure 2 for Figure 1 Schematic diagram of the airflow path of the sealing structure shown.

[0027] Description of reference numerals:

[0028] 1. turbine shaft; 101. delivery hole; 102. first delivery channel;

[0029] 2. turbine disk; 201. disk shaft; 202. final disk; 203. cooling space; 204. second delivery channel; 205. second delivery hole; 206. connecting cavity;

[0030] 3. Sealing cover; 301. Bearing cavity;

[0031] 4. First sealing component; 401. Third delivery hole; 402. Sealing cavity; 403. Exhaust space; 4031. First exhaust cavity; 4032. Second exhaust cavity; 404. Fourth delivery hole; 405. Support part; 406. Sealing part; 407. Diversion cavity;

[0032] 5. The second sealing component. DETAILED DESCRIPTION

[0033] In order to make the purpose, technical solution and advantages of the embodiments of the present invention clearer, the technical solution in the embodiments of the present invention will be clearly and completely described below in conjunction with the drawings in the embodiments of the present invention. Obviously, the described embodiments are 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 those skilled in the art without creative work are within the scope of protection of the present invention.

[0034] Combine the following Figure 1 to Figure 2 , describing an embodiment of the present invention.

[0035] In the related art, due to the limited space between the sealing cover and the power turbine final stage disk, the selectable specifications and sizes of the first sealing component are also limited, and the pressure at the sealing cavity can only be adjusted by changing the specifications and sizes of the first sealing component. When the selectable specifications and sizes of the first sealing component are limited, the adjustable range of the pressure difference between the sealing cavity and the bearing cavity will be limited to a smaller range, resulting in the pressure difference being difficult to flexibly and effectively adjust to the required range, and thus difficult to match the requirements of the engine running at different speeds. It is often easy for the pressure difference to be too large or too small, making the design of the aircraft engine difficult.

[0036] According to an embodiment of the present invention, on the one hand, a sealing structure is provided, comprising: a turbine shaft 1, a turbine disc 2, a sealing cover 3 and a first sealing assembly 4, wherein a first delivery hole 101 is provided on the outer wall of the turbine shaft 1, and a first delivery channel 102 for delivering sealing gas is provided inside; the turbine disc 2 comprises a disc shaft 201 and a final disc 202, the disc shaft 201 is fixedly sleeved on the outer side of the turbine shaft 1, one end of the disc shaft 201 along the axial direction forms a terminal, the disc shaft 201 is fixedly sleeved on the outer side of the terminal, and a space away from the disc shaft 201 along the axial direction forms a cooling space 203, and the space between the inner wall of the terminal and the outer wall of the turbine shaft 1 forms a second delivery channel 204, a second delivery hole 205 is provided on the disk shaft 201, and the first delivery channel 102 is connected with the cooling space 203 and the second delivery channel 204 at the same time through the first delivery hole 101; the sealing cover 3 is sleeved on the outer side of the disk shaft 201, and forms a bearing cavity 301 with the space between the sealing cover 3 and the disk shaft 201; the inner wall of the sealing cover 3 is sealed with the outer wall of the disk shaft 201 through the first sealing component 4, and the first sealing component 4 has a third delivery hole 401, and the first sealing component 4 cooperates with the sealing cover 3 to enclose a sealing cavity 402, and the second delivery channel 204, the second delivery hole 205, the third delivery hole 401 and the sealing cavity 402 are connected in sequence.

[0037] By using the sealing structure of the present embodiment, a conveying path is directly formed by opening holes through the first conveying hole 101, the second conveying channel 204, the second conveying hole 205 and the third conveying hole 401. Without setting a separate pipeline outside the sealing cover 3, the airflow inside the turbine shaft 1 can be directly led to the sealing cavity 402. Since the airflow flowing out of the first conveying hole 101 will flow into the cooling space 203 and the second conveying channel 204 at the same time, the airflow flow at the second conveying channel 204 can be adjusted by adjusting the number, position and size of the first conveying holes 101. Since the airflow in the second conveying channel 204 will also flow through the second conveying hole 205 and the third conveying hole 401 in sequence, the airflow in the second conveying channel 204 can also be adjusted by adjusting the second conveying hole 101. The flow area of ​​205 and the third delivery hole 401 can also adjust the air flow rate flowing into the sealing chamber 402. In addition, the size of the first sealing component 4 can be adjusted to adjust the flow rate of the airflow discharged outward from the sealing chamber 402, so as to further adjust the pressure of the sealing chamber 402. Through a simple and effective structure, the function of introducing airflow from the turbine shaft 1 to the sealing chamber 402 is realized, and at the same time, a multi-stage adjustment of the air flow in the sealing chamber 402 can be achieved through a variety of adjustment methods, so that the air pressure in the sealing chamber 402 can be adjusted within a larger range. Therefore, the flexibility of adjusting the pressure difference between the sealing chamber 402 and the bearing chamber can be improved, and the problem that the sealing form of the existing power turbine bearing chamber is difficult to meet the sealing requirements is effectively solved.

[0038] Among them, it should be noted that the bearing cavity 301 formed between the sealing cover 3 and the disk shaft 201 is the bearing cavity, and the first sealing component 4 is the sealing component used to seal the opening of the bearing cavity 301; the final disk 202 is the disk body located at the axial end of the turbine disk 2, and the final disk 202 and the disk shaft 201 are an integral component; the cooling space 203 is not a completely open or completely closed space, and the power turbine also has a primary disk. The cooling space 203 is the space between the final disk 202 and the primary disk, and there is another sealing structure at the end of the cooling space 203. The airflow entering the cooling space 203 will move along the gap between the final disk 202 and the primary disk, and a reliable seal will be formed at the sealing structure here, and excess airflow will enter the mainstream combustion gas through the sealing structure flowing into this place.

[0039] Specifically, Figure 2 As shown, after the airflow inside the turbine shaft 1 flows out through the first delivery hole 101, a part of it enters the second delivery channel 204, and the other part moves radially along the disk surface of the final stage disk 202, thereby adjusting the gas flow entering the second delivery channel 204. At the same time, the airflow that does not enter the second delivery channel 204 can also exchange heat with the final stage disk 202 to cool the final stage disk 202, thereby playing a cooling role.

[0040] In a possible implementation, the space between the sealing cover 3 and the final disk 202 forms an exhaust space 403, the sealing chamber 402 is located between the bearing chamber 301 and the exhaust space 403, the first sealing component 4 also has a fourth delivery hole 404, the second delivery hole 205 is connected to the exhaust space 403 through the fourth delivery hole 404, in this form of the first sealing component 4, the second delivery hole 205 is simultaneously connected to the third delivery hole 401 and the fourth delivery hole 404, a portion of the airflow flowing through the second delivery hole 205 can enter the exhaust space 403 through the fourth delivery hole 404, by adjusting the size of the fourth delivery hole 404, the flow rate of the airflow delivered from the second delivery hole 205 to the third delivery hole 401 can be adjusted, on the basis of the diversion adjustment at the first delivery hole 101, a second diversion adjustment process is added, which further improves the ability to adjust the airflow entering the sealing chamber 402, and greatly improves the flexibility of adjusting the pressure at the sealing chamber 402.

[0041] The specific diameters and numbers of the second delivery holes 205, the third delivery holes 401, and the fourth delivery holes 404 are not limited. The second delivery holes 205, the third delivery holes 401, and the fourth delivery holes 404 are multiple in number and are arranged at intervals along the circumferential direction. The specific diameters and numbers can be determined according to the air system flow calculation results during the design, and will not be described in detail here.

[0042] In a possible implementation, Figure 1As shown, the first sealing component 4 includes a support portion 405 and two sealing portions 406. The support portion 405 is sleeved on the outer wall of the disc shaft 201. The two sealing portions 406 are located between the outer wall of the support portion 405 and the inner wall of the sealing cover 3 and are spaced apart along the axial direction. The outer wall of the support portion 405, the side walls of the two sealing portions 406 and the surrounding of the sealing cover 3 form a sealing cavity 402. The two sealing portions 406 are located between the exhaust space 403 and the bearing cavity 301. In this form, the sealing cavity 402 and the sealing cavity 40 2 cooperates with the sealing portion 406 on the side close to the bearing cavity 301 to form an air curtain to effectively seal the bearing cavity 301. The airflow in the sealing cavity 402 can enter the exhaust space 403 through the sealing portion 406 on the side away from the bearing cavity 301. By adjusting the size of the sealing portion 406 on the side away from the bearing cavity 301, the air pressure in the sealing cavity 402 can be further adjusted. The third diversion adjustment process can be performed here, which greatly improves the adjustability of the air pressure in the sealing cavity 402.

[0043] Specifically, there is no limitation on the specific types of the two sealing parts 406 , which may be a graphite sealing structure or a comb teeth sealing structure, and can be flexibly selected according to needs.

[0044] Furthermore, the first delivery hole 101, the second delivery hole 205, the third delivery hole 401 and the fourth delivery hole 404 are all throttling holes. The sealing structure of the present application does not completely rely on graphite seals and comb seals to control the sealing flow, but cleverly controls the flow of each branch through multiple throttling holes, thereby achieving efficient sealing and more flexibly adjusting the air pressure in the sealing cavity 402. In particular, the fourth delivery hole 404 can introduce more airflow into the exhaust space 403 to cool the other side of the final stage disk 202, thereby achieving simultaneous cooling of both sides of the final stage disk 202, and avoiding excessive airflow into the bearing cavity 301, resulting in excessive sealing flow in the bearing cavity of the power turbine and causing lubricating oil over-temperature failure.

[0045] If the power turbine bearing cavity seal is found to have oil leakage or overheating faults during the test run, it is only necessary to perform reprocessing on the throttle hole structures, increase the number of holes or block some of the throttle holes, and adjust the flow distribution of each branch to achieve rapid test verification of relevant improvement measures.

[0046] In a possible implementation, a shunt cavity 407 is provided inside the support portion 405, the third delivery hole 401 is provided on the outer peripheral surface of the support portion 405, the fourth delivery hole 404 is provided on the end surface of the support portion 405 away from the bearing cavity 301, and the second delivery hole 205 is connected to the third delivery hole 401 and the fourth delivery hole 404 through the shunt cavity 407. This form of the support portion 405 is easier to manufacture and has stronger bending resistance. In addition, the shunt cavity 407 can buffer the airflow entering through the second delivery hole 205, and then discharge it through the third delivery hole 401 and the fourth delivery hole 404, which can improve the uniformity of the airflow discharged from the third delivery hole 401 and the fourth delivery hole 404.

[0047] Specifically, there is no limitation on the specific form of the diversion cavity 407 , which may be a cavity formed by the inner wall of the support portion 405 , or a cavity formed by the support portion 405 and adjacent components.

[0048] In a possible implementation, the support portion 405 has an annular groove opening toward the disk shaft 201, and the support portion 405 and the outer peripheral surface of the disk shaft 201 are surrounded by the annular groove to form a diversion cavity 407. This type of diversion cavity 407 has a simpler and more reliable structure, and the annular groove is easy to process and manufacture.

[0049] Specifically, there is no limitation on the shape of the inner wall of the annular groove, and it can be a continuous curved surface or a wall formed by a combination of a plane and a curved surface, which can be flexibly selected according to usage requirements.

[0050] In a possible embodiment, the sealing structure also includes a second sealing component 5, which is arranged between the sealing cover 3 and the final stage disk 202 and is located in the exhaust space 403. The space between the second sealing component 5 and the first sealing component 4 in the exhaust space 403 forms a first exhaust cavity 4031, and the space on the other side of the exhaust space 403 on the other side of the second sealing component 5 forms a second exhaust cavity 4032. The second sealing component 5 blocks the airflow flowing from the first exhaust cavity 4031 to the second exhaust cavity 4032, thereby increasing the pressure in the first exhaust cavity 4031 and preventing excessive airflow in the diversion cavity 407 from flowing into the exhaust space 403.

[0051] The airflow entering the second exhaust cavity 4032 will flow into the mainstream fuel gas channel of the engine and be used to seal the rear stage of the final stage disk 202 .

[0052] Specifically, the sealing portion 406 and the second sealing assembly 5 are one of a graphite seal and a comb seal.

[0053] Preferably, the second sealing component 5 is a comb-teeth sealing structure, and the sealing portion 406 is a graphite sealing structure.

[0054] In a possible implementation, the axial extension line of the second delivery hole 205 intersects with the radial extension line of the disk shaft 201, the axial extension line of the fourth delivery hole 404 intersects with the axis of the support portion 405, and the third delivery hole 401 extends radially along the support portion 405. Such obliquely arranged second delivery holes 205 and fourth delivery holes 404 can increase the heat exchange area between the airflow and the hole walls of the second delivery holes 205 and the fourth delivery holes 404, thereby improving the cooling effect of the airflow.

[0055] It can be understood that, as an alternative embodiment, the axial extension line of the second delivery hole 205 is parallel to the radial extension line of the disk shaft 201 , and the axial extension line of the fourth delivery hole 404 is parallel to the axis of the support portion 405 .

[0056] In a possible embodiment, there are multiple first delivery holes 101 and they are arranged at circumferential intervals along the outer wall of the turbine shaft 1. The first delivery holes 101 are strip holes extending along the axial direction. This type of first delivery holes 101 has little effect on the pressure of the airflow discharged from the first delivery channel 102 and also has little effect on the structural strength of the turbine shaft 1.

[0057] It can be understood that, as an alternative implementation, the first delivery hole 101 can also be a circular hole or a hole of other shapes, which can be flexibly selected according to needs.

[0058] In a possible implementation, the pressure difference between the sealing chamber 402 and the bearing chamber 301 is greater than 3.5 kPa, the ratio of the cross-sectional area of ​​the second delivery hole 205 to the cross-sectional area of ​​the third delivery hole 401 is in the range of 0.87 to 0.88, the ratio of the cross-sectional area of ​​the fourth delivery hole 404 to the cross-sectional area of ​​the third delivery hole 401 is in the range of 0.093 to 0.095, and the ratio of the flow rate of the second delivery hole 205 to the flow rate of the third delivery hole 401 is in the range of 0.3 to 0.8. The delivery holes under this proportional relationship can achieve efficient sealing.

[0059] It should be noted that the cross-sectional area is the flow area of ​​the hole.

[0060] Preferably, the ratio of the cross-sectional areas of the second delivery hole 205, the third delivery hole 401 and the fourth delivery hole 404 is 56:64:6; the flow rate of the second delivery hole 205 is equal to the sum of the flow rate of the third delivery hole 401 and the flow rate of the fourth delivery hole 404, the flow rate of the second delivery hole 205 is less than the flow rate of the third delivery hole 401, and the ratio of the flow rate of the second delivery hole 205 to the flow rate of the third delivery hole 401 is 6:9.

[0061] Furthermore, the above ratio is the ratio when the number of the second delivery holes 205 , the third delivery holes 401 and the fourth delivery holes 404 are the same, or the above ratio is the ratio of the total number of the second delivery holes 205 , the third delivery holes 401 and the fourth delivery holes 404 .

[0062] In a possible embodiment, an annular groove opening toward the turbine shaft 1 is formed on the inner wall of the disk shaft 201, and the second delivery hole 205 is formed on the inner wall of the annular groove away from the turbine shaft 1. The space inside the annular groove forms a connecting chamber 206. The second delivery channel 204 is connected to the second delivery hole 205 through the connecting chamber 206. The ratio of the pressure at the connecting chamber 206 to the pressure of the second exhaust chamber 4032 is in the range of 2.8 to 3.2, and the ratio of the pressure at the connecting chamber 206 to the pressure of the bearing chamber 301 is in the range of 1.8 to 2.2. The structure is simple and reliable, and the connecting chamber 206 can buffer the airflow flowing out of the second delivery channel 204.

[0063] Preferably, the ratio of the pressure at the connecting chamber 206 to the pressure of the second exhaust chamber 4032 is 3, and the ratio of the pressure at the connecting chamber 206 to the pressure of the bearing chamber 301 is 2.

[0064] When the engine is in an emergency state, the pressure in the connecting cavity 206 is greater than twice the pressure in the bearing cavity 301 .

[0065] According to an embodiment of the present invention, on the other hand, there is provided an aircraft engine, which includes: the above-mentioned sealing structure.

[0066] Although the embodiments of the present invention have been described in conjunction with the accompanying drawings, those skilled in the art may make various modifications and variations without departing from the spirit and scope of the present invention, and such modifications and variations are all within the scope defined by the appended claims.

Claims

1. A sealing structure, characterized in that: include: The turbine shaft (1) has a first delivery hole (101) on its outer wall and a first delivery channel (102) inside for delivering sealing gas; A turbine disk (2), comprising a disk shaft (201) and a final disk (202), wherein the disk shaft (201) is fixedly sleeved on the outside of the turbine shaft (1), one end of the disk shaft (201) along the axial direction forms a terminal, the disk shaft (201) is fixedly sleeved on the outside of the terminal, and a space away from the disk shaft (201) along the axial direction forms a cooling space (203), a space between the inner wall of the terminal and the outer wall of the turbine shaft (1) forms a second delivery channel (204), a second delivery hole (205) is provided on the disk shaft (201), and the first delivery channel (102) is simultaneously connected to the cooling space (203) and the second delivery channel (204) through the first delivery hole (101); A sealing cover (3) is sleeved on the outer side of the disk shaft (201), and the space between the sealing cover and the disk shaft (201) forms a bearing cavity (301); A first sealing component (4), the inner wall of the sealing cover (3) is sealed with the outer wall of the disk shaft (201) through the first sealing component (4), the first sealing component (4) has a third delivery hole (401), the first sealing component (4) cooperates with the sealing cover (3) and encloses to form a sealing cavity (402), the second delivery channel (204), the second delivery hole (205), the third delivery hole (401) and the sealing cavity (402) are connected in sequence.

2. The sealing structure according to claim 1, characterized in that: The space between the sealing cover (3) and the final disk (202) forms an exhaust space (403); the sealing cavity (402) is located between the bearing cavity (301) and the exhaust space (403); the first sealing component (4) further has a fourth delivery hole (404); and the second delivery hole (205) is connected to the exhaust space (403) through the fourth delivery hole (404).

3. The sealing structure according to claim 2, characterized in that: The first sealing component (4) includes a supporting portion (405) and two sealing portions (406), wherein the supporting portion (405) is sleeved on the outer wall of the disk shaft (201), and the two sealing portions (406) are located between the outer wall of the supporting portion (405) and the inner wall of the sealing cover (3) and are spaced apart along the axial direction, and the outer wall of the supporting portion (405), the side walls of the two sealing portions (406) and the inner wall of the sealing cover (3) together form the sealing cavity (402), and the two sealing portions (406) are located between the exhaust space (403) and the bearing cavity (301).

4. The sealing structure according to claim 3, characterized in that: The support portion (405) has a shunt cavity (407) inside, the third delivery hole (401) is opened on the outer peripheral surface of the support portion (405), the fourth delivery hole (404) is opened on the end surface of the support portion (405) away from the bearing cavity (301), and the second delivery hole (205) is connected to the third delivery hole (401) and the fourth delivery hole (404) at the same time through the shunt cavity (407).

5. The sealing structure according to claim 4, characterized in that: The support portion (405) has an annular groove opening toward the disk shaft (201), and the support portion (405) is surrounded by the outer peripheral surface of the disk shaft (201) through the annular groove to form the diversion cavity (407).

6. The sealing structure according to any one of claims 2 to 5, characterized in that: The sealing structure also includes a second sealing component (5), which is arranged between the sealing cover (3) and the final disk (202) and is located in the exhaust space (403), and the space between the second sealing component (5) and the first sealing component (4) of the exhaust space (403) forms a first exhaust cavity (4031), and the space on the other side of the second sealing component (5) of the exhaust space (403) forms a second exhaust cavity (4032).

7. The sealing structure according to any one of claims 3 to 5, characterized in that: The axial extension line of the second delivery hole (205) intersects with the radial extension line of the disk shaft (201), the axial extension line of the fourth delivery hole (404) intersects with the axis of the support portion (405), and the third delivery hole (401) extends along the radial direction of the support portion (405); And / or, the first delivery holes (101) are multiple in number and are arranged at intervals along the circumferential direction of the outer wall of the turbine shaft (1), and the first delivery holes (101) are strip-shaped holes extending along the axial direction.

8. The sealing structure according to any one of claims 2 to 5, characterized in that: The pressure difference between the sealing cavity (402) and the bearing cavity (301) is greater than 3.5 kPa, the ratio of the cross-sectional area of ​​the second delivery hole (205) to the cross-sectional area of ​​the third delivery hole (401) is in the range of 0.87 to 0.88, the ratio of the cross-sectional area of ​​the fourth delivery hole (404) to the cross-sectional area of ​​the third delivery hole (401) is in the range of 0.093 to 0.095, and the ratio of the flow rate of the second delivery hole (205) to the flow rate of the third delivery hole (401) is in the range of 0.3 to 0.

8.

9. The sealing structure according to claim 6, characterized in that: The inner wall of the disk shaft (201) is provided with an annular groove opening toward the turbine shaft (1); the second delivery hole (205) is provided on the inner wall of the annular groove away from the turbine shaft (1); the space inside the annular groove forms a connecting chamber (206); the second delivery channel (204) is connected to the second delivery hole (205) through the connecting chamber (206); the ratio of the pressure at the connecting chamber (206) to the pressure of the second exhaust chamber (4032) is in the range of 2.8 to 3.2; and the ratio of the pressure at the connecting chamber (206) to the pressure of the bearing chamber (301) is in the range of 1.8 to 2.

2.

10. An aircraft engine, characterized in that: include: The sealing structure according to any one of claims 1 to 9.

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