Blisk for high thrust ratio gas turbine engines and turbine engine

Abstract

The application discloses a whole blade ring for a high thrust ratio gas turbine engine and a turbine engine, which comprises a first combined rotor, a second combined rotor and a third combined rotor, the first combined rotor comprises a first inner ring, a first outer ring and first compressor blades, the second combined rotor comprises a second inner ring, a second outer ring, second compressor blades and first turbine blades, and the third combined rotor comprises a third inner ring, a third outer ring, third compressor blades and second turbine blades; the first outer ring, the second outer ring and the third outer ring are connected in sequence along the axial direction and form a compressor flow channel, and the first inner ring is provided with a first bearing seat; the centrifugal force of the blades is borne by a rotor drum, the work capacity of the blades at a high tangential speed can be effectively utilized, the reduction of the performance caused by the blade tip leakage can be solved, the connecting shaft for transmitting power between the compressor and the turbine is cancelled, the number of bearings and bearing cavities can be reduced, the overall weight of the engine can be reduced, and the thrust-to-weight ratio of the engine can be improved.

Description

TECHNICAL FIELD

[0001] The present application relates to the technical field of aero-engine, in particular to a whole blade ring for high thrust ratio gas turbine engine. In addition, the present application also relates to a turbine engine comprising the whole blade ring for high thrust ratio gas turbine engine. BACKGROUND

[0002] As shown in the prior art, the rotating compressor and turbine are connected on the same shaft, and a heat source (combustion chamber) is arranged between the compressor and the turbine. Air is continuously sucked into the compressor, compressed and pressurized, and then enters the combustion chamber to be injected and burned into high-temperature and high-pressure gas, which then enters the turbine to expand and do work. Figure 1

[0003] Since the compressor, combustion chamber and turbine are linearly arranged along the axial direction, the transmission route is too long, and at least three bearings and three bearing cavities need to be arranged before the compressor, after the compressor, before the turbine and after the turbine, which leads to a complex structure of the oil system, a large size of the whole engine, an increase in the weight of the engine, and an impact on the thrust-to-weight ratio of the engine. In addition, the span of the compressor flow passage and the turbine flow passage is large, and the design of the air induction system is complex. SUMMARY

[0004] The present application provides a whole blade ring for high thrust ratio gas turbine engine and a turbine engine to solve the technical problems of the prior art, such as the long transmission route of the aero gas turbine engine, the need for multiple bearing supports and multiple bearing cavities, the large size and weight of the whole engine, the impact on the thrust-to-weight ratio of the engine, and the complex design of the oil system and the air induction system.

[0005] According to one aspect of the present application, a whole blade ring for high thrust ratio gas turbine engine is provided, comprising a first combined rotor, a second combined rotor and a third combined rotor. The first combined rotor comprises a first inner ring, a first outer ring and first compressor blades arranged between the first inner ring and the first outer ring. The second combined rotor comprises a second inner ring, a second outer ring, second compressor blades arranged between the second inner ring and the second outer ring, and first turbine blades arranged on the outer side of the second outer ring. The third combined rotor comprises a third inner ring, a third outer ring, third compressor blades arranged between the third inner ring and the third outer ring, and second turbine blades arranged on the outer side of the third outer ring. The first outer ring, the second outer ring and the third outer ring are connected in sequence along the axial direction and form a compressor flow passage. The first inner ring is provided with a first bearing seat.

[0006] ​Further, the outer side surface of the second outer ring and the third outer ring is respectively provided with a reinforcing ring, the second outer ring is connected with the first turbine blade through the reinforcing ring, and the third outer ring is connected with the second turbine blade through the reinforcing ring.

[0007] Further, the second outer ring is provided with a first bleed hole towards the root of the first turbine blade, and the third outer ring is provided with a second bleed hole towards the root of the second turbine blade.

[0008] Further, the second outer ring is provided with a first limiting structure for limiting the axial position of the first turbine blade, the first limiting structure, the outer side surface of the second outer ring, the root of the first turbine blade and the reinforcing ring form a first cooling cavity, and the first cooling cavity is communicated with the compressor flow passage through the first bleed hole; the third outer ring is provided with a second limiting structure for limiting the axial position of the second turbine blade, the second limiting structure, the outer side surface of the third outer ring, the root of the second turbine blade and the reinforcing ring form a second cooling cavity, and the second cooling cavity is communicated with the compressor flow passage through the second bleed hole.

[0009] Further, the first limiting structure and the second limiting structure are provided with a labyrinth.

[0010] According to another aspect of the present application, a turbine engine is also provided, which comprises the integral blade ring for high thrust ratio gas turbine engine and an outer casing, a center shaft, a diffuser, a combustion chamber, a turbine flow passage and an exhaust nozzle; the outer casing is rotatably and sealingly connected with the first outer ring, the diffuser is rotatably and sealingly connected with the third outer ring, and the two ends of the center shaft are rotatably connected with the first bearing seat and the third inner ring through bearings; the first turbine blade extends into the turbine flow passage, and the labyrinth on the first limiting structure is rotatably and sealingly connected with the outer wall surface of the turbine flow passage; the second turbine blade extends into the exhaust nozzle and is arranged at one end of the exhaust nozzle close to the turbine flow passage, and the labyrinth on the second limiting structure is rotatably and sealingly connected with the outer wall surface of the turbine flow passage.

[0011] Further, the outer casing is provided with an inlet guide vane, the inlet guide vane is provided with a first lubricating oil passage, the center shaft is provided with a second lubricating oil passage for supplying lubricating oil to the bearing, and the first lubricating oil passage is communicated with the second lubricating oil passage; the first bearing seat and the third inner ring are respectively provided with a lubricating oil sealing structure, and the lubricating oil sealing structure is arranged at the two ends of the bearing.

[0012] Further, the first lubricating oil passage comprises an oil inlet passage and an oil return passage.

[0013] Further, a first stator blade and a second stator blade are arranged on the central shaft, the first stator blade is arranged between the first compressor blade and the second compressor blade, and the second stator blade is arranged between the second compressor blade and the third compressor blade.

[0014] Further, annular grooves adapted to the first inner ring, the second inner ring and the third inner ring are formed on the central shaft, so that the first inner ring, the second inner ring, the third inner ring and the central shaft are flush with the cylindrical surface of the outer casing.

[0015] The present application has the following beneficial effects:

[0016] The integral blade ring for the high thrust ratio gas turbine engine of the present application is characterized in that the first outer ring and the second outer ring, and the second outer ring and the third outer ring are connected by bolts to form a rotor drum, a compressor flow passage is formed in the rotor drum, the first compressor blade, the second compressor blade and the third compressor blade are arranged on the inner cylindrical surface of the rotor drum, in order to ensure the same stress level, the first compressor blade, the second compressor blade and the third compressor blade are wide at the part connected to the rotor drum, and narrow at the part close to the central shaft, the centrifugal force of the first compressor blade, the second compressor blade and the third compressor blade is borne by the rotor drum, which not only effectively utilizes the work function of the blade at the high tangential speed, but also solves the performance reduction caused by the tip leakage (secondary flow); the first turbine blade and the second turbine blade are connected to the outer cylindrical surface of the rotor drum, during operation, the rotor drum drives the first compressor blade, the second compressor blade, the third compressor blade, the first turbine blade and the second turbine blade to rotate synchronously, compared with the traditional structure, the interchanging of the rotor stator structure improves the work function of the compressor blade, and fully utilizes the space in the vertical axial direction, cancels the connecting shaft for transmitting power between the compressor and the turbine, the two ends of the rotor drum are connected to the bearing through the first bearing seat and the third inner ring, which can reduce the number of bearings and bearing cavities, thereby simplifying the structure, shortening the axial length of the whole engine, reducing the overall weight of the engine, and improving the thrust-to-weight ratio of the engine.

[0017] In addition to the objects, features and advantages described above, the present application has other objects, features and advantages. The present application will be further described in detail below with reference to the drawings. BRIEF DESCRIPTION OF DRAWINGS

[0018] The accompanying drawings, which form a part of this application, are included to provide a further understanding of the application, illustrate preferred embodiments of the present application, and assist in the explanation of the present application. In the drawings:

[0019] Figure 1 is a structural schematic diagram of the existing aviation gas turbine engine;

[0020] Figure 2 is a structure diagram of the whole blade ring of the high thrust ratio gas turbine engine of the preferred embodiment of the present application;

[0021] Figure 3 is a structure diagram of the two-stage combined rotor of the preferred embodiment of the present application;

[0022] Figure 4 is one of the structure diagrams of the existing compressor rotor;

[0023] Figure 5 is the second structure diagram of the existing compressor rotor;

[0024] Figure 6 is a cold air bleed diagram of the existing aviation gas turbine engine;

[0025] Figure 7 is a cold air bleed diagram of the preferred embodiment of the present application;

[0026] Figure 8 is a structure diagram of the turbine engine of the preferred embodiment of the present application;

[0027] Figure 9 is a structure diagram of the center shaft of the preferred embodiment of the present application.

[0028] Legend:

[0029] 1, whole blade ring; 11, first combined rotor; 111, first inner ring; 112, first outer ring; 113, first compressor blade; 114, first bearing seat; 116, lubricating oil sealing structure; 12, second combined rotor; 121, second inner ring; 122, second outer ring; 123, second compressor blade; 124, first turbine blade; 13, third combined rotor; 131, third inner ring; 132, third outer ring; 133, third compressor blade; 134, second turbine blade; 14, reinforcing ring; 15, first bleed hole; 16, second bleed hole; 17, first limiting structure; 18, second limiting structure; 19, grid; 2, outer casing; 21, inlet guide vane; 211, first lubricating oil channel; 3, center shaft; 31, second lubricating oil channel; 32, first stator blade; 33, second stator blade; 34, annular groove; 4, diffuser; 5, combustion chamber; 6, turbine flow passage; 7, tail nozzle; 8, bearing; 100, wheel disc; 101, compressor blade; 102, shroud. DETAILED DESCRIPTION

[0030] The embodiments of the present application are described in detail below with reference to the accompanying drawings, but the present application can be implemented in various different ways as defined and covered by the following.

[0031] As Figure 2and Figure 3 As shown, the integral blade ring for the high thrust-to-weight ratio gas turbine engine in this embodiment includes a first-stage combined rotor 11, a second-stage combined rotor 12, and a third-stage combined rotor 13. The first-stage combined rotor 11 includes a first inner ring 111, a first outer ring 112, and a first compressor blade 113 disposed between the first inner ring 111 and the first outer ring 112. The second-stage combined rotor 12 includes a second inner ring 121, a second outer ring 122, and a second compressor blade 123 disposed between the second inner ring 121 and the second outer ring 122. The first turbine blade 124 is arranged on the outer surface of the second outer ring 122; the three-stage combined rotor 13 includes a third inner ring 131, a third outer ring 132, a third compressor blade 133 arranged between the third inner ring 131 and the third outer ring 132, and a second turbine blade 134 arranged on the outer surface of the third outer ring 132; the first outer ring 112, the second outer ring 122 and the third outer ring 132 are connected in sequence along the axial direction to form a compressor flow channel, and a first bearing seat 114 is arranged on the first inner ring 111.

[0032] like Figure 4 As shown, conventional compressor blades 101 are arranged on the outer circumference of the disk 100. The centrifugal force of the compressor blades 101 is borne by the central disk 100. Due to the presence of centrifugal force, the compressor blades 101 must be designed with a pointed top and a wide bottom to ensure that the stress levels at the blade tip and root are the same, thus making effective use of the material; however, the high tangential velocity at the blade tip is not effectively utilized. Figure 5 As shown, in order to effectively utilize the high tangential velocity at the blade tip, a blade crown 102 is installed at the blade tip of the compressor blade 101. Although the blade crown 102 can reduce air leakage (secondary flow) at the blade tip of the compressor blade 101 and improve the working efficiency, it also increases the centrifugal force significantly due to the increased blade tip weight. The centrifugal force of the crowned blade structure is still borne by the central impeller 100. To ensure that the impeller 100 meets the usage requirements, the impeller 100 needs to be made very bulky, which will not only increase the overall weight of the engine and affect the engine thrust-to-weight ratio, but also occupy a lot of space.

[0033] The integral blade ring of the high thrust ratio gas turbine engine of the embodiment is formed by bolt connection between the first outer ring 112 and the second outer ring 122 and between the second outer ring 122 and the third outer ring 132 to form a rotor drum, and a compressor flow passage is formed in the rotor drum. The first compressor blade 113, the second compressor blade 123 and the third compressor blade 133 are arranged on the inner circular surface of the rotor drum. In order to ensure the same stress level, the first compressor blade 113, the second compressor blade 123 and the third compressor blade 133 are wide at the part connected to the rotor drum and narrow at the part away from the rotor drum. The centrifugal force of the first compressor blade 113, the second compressor blade 123 and the third compressor blade 133 is borne by the rotor drum. Since the tangential velocity is high and the working area is large at the outer ring, not only the working capacity of the blade at the high tangential velocity can be effectively utilized, but also the reduction of performance caused by tip leakage (secondary flow loss) can be solved. The first turbine blade 124 and the second turbine blade 134 are connected to the outer circular surface of the rotor drum. During operation, the rotor drum drives the first compressor blade 113, the second compressor blade 123, the third compressor blade 133, the first turbine blade 124 and the second turbine blade 134 to rotate synchronously. Compared with the traditional structure, the interchange of the rotor and stator structures improves the working capacity of the compressor blade, fully utilizes the space in the vertical axial direction, cancels the connecting shaft for transmitting power between the compressor and the turbine, and connects the bearings through the first bearing seat 114 and the third inner ring 131 at both ends of the rotor drum. The axial length of the turbine engine can be shortened by more than one time, the number of bearings and bearing cavities can be reduced, the structure is simplified, the overall weight of the engine is reduced, and the thrust-to-weight ratio of the engine is improved. It can be understood that the number of the first combined rotor 11, the second combined rotor 12 or the third combined rotor 13 can be adjusted and combined to form the integral blade ring to meet the use requirements in different scenarios.

[0034] As Figure 2As shown, in this embodiment, reinforcing rings 14 are respectively arranged on the outer surfaces of the second outer ring 122 and the third outer ring 132. The second outer ring 122 is connected to the first turbine blade 124 through the reinforcing ring 14. The reinforcing ring 14 on the second outer ring 122, the second outer ring 122, the second compressor blade 123, and the second inner ring 121 are integrally formed. The third outer ring 132 is connected to the second turbine blade 134 through the reinforcing ring 14. The reinforcing ring 14 on the third outer ring 132, the third outer ring 132, the third compressor blade 133, and the third inner ring 131 are integrally formed. The reinforcing ring 14 can increase the strength of the rotor drum. The tenon and groove structure for assembling the turbine blade is machined on the reinforcing ring 14. The centrifugal force of the blade is borne by the reinforcing ring 14, which can prevent the rotor drum from deforming. Optionally, a reinforcing ring 14 is provided on the outer surface of the first outer ring 112. The reinforcing ring 14 can increase the strength of the first outer ring 112 and prevent deformation of the first outer ring 112. On the other hand, it can also act as a counterweight to improve the stability of the rotor drum. Optionally, the reinforcing ring 14 has a hollow structure to reduce the weight of the rotor drum.

[0035] like Figure 6 As shown, because the compressor, combustion chamber, and turbine are arranged linearly along the axial direction, the span between the compressor flow channel and the turbine flow channel is large, the bleed gas traction path is long, and the bleed gas system design is complex. For example... Figure 2 and Figure 7 As shown, because the compressor decelerates and pressurizes the airflow, the cross-sectional area of ​​the compressor flow channel gradually decreases from the inlet to the outlet; while the turbine accelerates and depressurizes the airflow, the cross-sectional area of ​​the turbine flow channel gradually increases from the inlet to the outlet. In this embodiment, since the first turbine blade 124 and the second turbine blade 134 are connected to the outer circumference of the rotor drum, in order to maximize the use of the space perpendicular to the axial direction, the structural features of the compressor flow channel and the turbine flow channel are integrated, and the turbine and compressor are placed side by side. A first air intake hole 15 is opened on the second outer ring 122 facing the root of the first turbine blade 124. Cooling gas flows to the root of the first turbine blade 124 through the first air intake hole 15 to cool the connection between the first turbine blade 124 and the reinforcing ring 14. However, it can also enter the first turbine blade 124 to cool the blade body of the first turbine blade 124; the third outer ring 132 has a second air duct 16 facing the root of the second turbine blade 134, through which the cooling gas flows to the root of the second turbine blade 134 to cool the connection between the second turbine blade 134 and the reinforcing ring 14, and can also enter the second turbine blade 134 to cool the blade body of the second turbine blade 134; this can greatly reduce the air duct path of the cooling gas and simplify the air duct system design.

[0036] like Figure 2As shown, in the embodiment, the second outer ring 122 is provided with a first limiting structure 17 for limiting the axial position of the first turbine blade 124, the first limiting structure 17, the outer side surface of the second outer ring 122, the root of the first turbine blade 124 and the reinforcing ring 14 form a first cooling chamber, the first cooling chamber is communicated with the compressor flow passage through the first air guiding hole 15, the first turbine blade 124 is provided with the first limiting structure 17 on both sides, thereby limiting the axial position of the first turbine blade 124, ensuring the stability of the first turbine blade 124, the first cooling chamber can avoid the cooling gas from rapidly entering the inside of the first turbine blade 124, ensuring the cooling effect; the third outer ring 132 is provided with a second limiting structure 18 for limiting the axial position of the second turbine blade 134, the second limiting structure 18, the outer side surface of the third outer ring 132, the root of the second turbine blade 134 and the reinforcing ring 14 form a second cooling chamber, the second cooling chamber is communicated with the compressor flow passage through the second air guiding hole 16, the reinforcing ring 14 is provided with a mortise for limiting the right side position of the second turbine blade 134, the second limiting structure 18 is arranged on the left side of the second turbine blade 134 and cooperates with the mortise on the reinforcing ring 14, thereby limiting the second turbine blade 134 of the second turbine blade 134, ensuring the stability of the second turbine blade 134, the second cooling chamber can avoid the cooling gas from rapidly entering the inside of the second turbine blade 134, ensuring the cooling effect.

[0037] As Figure 2 shown, in the embodiment, the first limiting structure 17 and the second limiting structure 18 are provided with a grid tooth 19, the grid tooth 19 and the outer circular surface of the turbine flow passage 6 form a rotary sealing structure, avoiding the air flow of the turbine flow passage 6 from leaking out.

[0038] As Figure 8 and Figure 9As shown in the figure, a turbine engine comprises the integral blade ring 1 for high thrust ratio gas turbine engine, and further comprises an outer casing 2, a central shaft 3, a diffuser 4, a combustion chamber 5, a turbine flow passage 6 and a tail nozzle 7; the outer casing 2 is in rotational sealing connection with the first outer ring 112, the diffuser 4 is in rotational sealing connection with the third outer ring 132, the two ends of the central shaft 3 are respectively in rotational connection with the first bearing seat 114 and the third inner ring 131 through bearings 8; the first turbine blade 124 extends into the turbine flow passage 6, the louver 19 on the first limiting structure 17 is in rotational sealing connection with the outer wall surface of the turbine flow passage 6; the second turbine blade 134 extends into the tail nozzle 7 and is arranged at one end of the tail nozzle 7 close to the turbine flow passage 6, the louver 19 on the second limiting structure 18 is in rotational sealing connection with the outer wall surface of the turbine flow passage 6; it fully utilizes the space in the vertical axial direction, cancels the connecting shaft for transmitting power between the compressor and the turbine, the two ends of the integral blade ring 1 are respectively in rotational connection with the central shaft 3 through bearings 8, the first bearing seat 114 and the third inner ring 131 are located on the same stator, the concentricity of the front and rear bearings is improved, thereby simplifying the structure, reducing the machining difficulty, reducing the number of bearings and bearing cavities, shortening the axial length of the whole engine, reducing the overall weight of the engine and improving the thrust-to-weight ratio of the engine.

[0039] As shown in the figure, Figure 8 and Figure 9 In this embodiment, the inlet guide vane 21 is arranged on the outer casing 2, the first oil passage 211 is formed on the inlet guide vane 21, the second oil passage 31 is formed on the central shaft 3 for feeding oil to the bearing 8, and the first oil passage 211 is in communication with the second oil passage 31; the oil sealing structure 116 is arranged on the first bearing seat 114 and the third inner ring 131 respectively, the oil sealing structure 116 is arranged at the two ends of the bearing 8, the first oil passage 211 is integrated on the inlet guide vane 21, and the oil enters the bearings 8 at the front and rear ends of the central shaft 3 through the first oil passage 211 and the second oil passage 31 respectively; the structure is simple, the internal space of the inlet guide vane 21 can be maximized, on the one hand, the weight of the inlet guide vane 21 can be reduced, thereby reducing the overall weight of the engine, on the other hand, the influence of the first oil passage 211 on the flow field can be avoided.

[0040] In this embodiment, the first oil passage 211 comprises an oil inlet passage and an oil return passage, which can form an oil circulation loop to ensure the lubrication effect of the bearing 8; since the second oil passage 31 of the central shaft 3 is used for oil feeding and oil return, the oil passage design is greatly simplified, and since the number of bearing cavities is reduced, the number of engine stators is reduced, the overall assembly difficulty of the engine is reduced, and the disassembly and assembly efficiency is improved.

[0041] As shown in the figure, Figure 8 and Figure 9As shown, in the embodiment, the first-stage stator blade 32 and the second-stage stator blade 33 are arranged on the center shaft 3, the first-stage stator blade 32 is arranged between the first compressor blade 113 and the second compressor blade 123, and the second-stage stator blade 33 is arranged between the second compressor blade 123 and the third compressor blade 133, so that the airflow can be decelerated and pressurized, and the working efficiency of the compressor can be improved.

[0042] As shown in Figure 8 and Figure 9 In the embodiment, the annular groove 34 adapted to the first inner ring 111, the second inner ring 121 and the third inner ring 131 is arranged on the center shaft 3, so that the first inner ring 111, the second inner ring 121, the third inner ring 131 and the center shaft 3 are flush with the cylindrical surface of the outer casing 2, so as to avoid affecting the flow field in the compressor flow passage.

[0043] The above only describes the preferred embodiments of the present application and is not used to limit the present application. For those skilled in the art, the present application can have various modifications and changes. Any modification, equivalent replacement, improvement, etc. within the spirit and principle of the present application shall be included in the protection scope of the present application.

Claims

1. A blisk for a high thrust ratio gas turbine engine, characterized by, The combined rotor includes a first combined rotor (11), a second combined rotor (12) and a third combined rotor (13), the first combined rotor (11) includes a first inner ring (111), a first outer ring (112) and a first compressor blade (113) arranged between the first inner ring (111) and the first outer ring (112); the second combined rotor (12) includes a second inner ring (121), a second outer ring (122), a second compressor blade (123) arranged between the second inner ring (121) and the second outer ring (122) and a first turbine blade (124) arranged on the outer side of the second outer ring (122); the third combined rotor (13) includes a third inner ring (131), a third outer ring (132), a third compressor blade (133) arranged between the third inner ring (131) and the third outer ring (132) and a second turbine blade (134) arranged on the outer side of the third outer ring (132); the first outer ring (112), the second outer ring (122) and the third outer ring (132) are connected in sequence along the axial direction and form a compressor flow passage, and the first inner ring (111) is provided with a first bearing seat (114).

2. The blisk for a high thrust ratio gas turbine engine of claim 1, wherein, The outer side of the second outer ring (122) and the third outer ring (132) is respectively provided with a reinforcing ring (14), the second outer ring (122) is connected with the first turbine blade (124) through the reinforcing ring (14), and the third outer ring (132) is connected with the second turbine blade (134) through the reinforcing ring (14).

3. The bladed ring for a high thrust ratio gas turbine engine of claim 2, wherein, The second outer ring (122) is provided with a first bleed hole (15) towards the root of the first turbine blade (124), and the third outer ring (132) is provided with a second bleed hole (16) towards the root of the second turbine blade (134).

4. The bladed ring for a high thrust ratio gas turbine engine of claim 3, wherein, The second outer ring (122) is provided with a first limiting structure (17) for limiting the axial position of the first turbine blade (124), the first limiting structure (17), the outer side of the second outer ring (122), the root of the first turbine blade (124) and the reinforcing ring (14) form a first cooling chamber, and the first cooling chamber is connected with the compressor flow passage through the first bleed hole (15); the third outer ring (132) is provided with a second limiting structure (18) for limiting the axial position of the second turbine blade (134), the second limiting structure (18), the outer side of the third outer ring (132), the root of the second turbine blade (134) and the reinforcing ring (14) form a second cooling chamber, and the second cooling chamber is connected with the compressor flow passage through the second bleed hole (16).

5. The bladed ring for a high thrust ratio gas turbine engine of claim 4, wherein, The first limiting structure (17) and the second limiting structure (18) are provided with a grid (19).

6. A turbine engine characterized by, The integral blade ring for high thrust ratio gas turbine engine as claimed in claim 5 further comprises an outer casing (2), a central shaft (3), a diffuser (4), a combustion chamber (5), a turbine flow channel (6) and a tail nozzle (7); the outer casing (2) is in rotational sealing connection with the first outer ring (112), the diffuser (4) is in rotational sealing connection with the third outer ring (132), both ends of the central shaft (3) are in rotational connection with the first bearing seat (114) and the third inner ring (131) through bearings (8); the first turbine blade (124) extends into the turbine flow channel (6), the labyrinth (19) on the first limiting structure (17) is in rotational sealing connection with the outer wall surface of the turbine flow channel (6); the second turbine blade (134) extends into the tail nozzle (7) and is arranged at one end of the tail nozzle (7) close to the turbine flow channel (6), the labyrinth (19) on the second limiting structure (18) is in rotational sealing connection with the outer wall surface of the turbine flow channel (6).

7. The turbine engine of claim 6, wherein, An inlet guide vane (21) is arranged on the outer casing (2), a first oil passage (211) is formed on the inlet guide vane (21), a second oil passage (31) for supplying oil to the bearing (8) is formed on the central shaft (3), and the first oil passage (211) is in communication with the second oil passage (31); the first bearing seat (114) and the third inner ring (131) are respectively provided with oil sealing structures (116), and the oil sealing structures (116) are arranged at both ends of the bearing (8).

8. The turbine engine of claim 7, wherein, The first oil passage (211) comprises an oil inlet passage and an oil return passage.

9. The turbine engine of claim 6, wherein, A first-stage stator blade (32) and a second-stage stator blade (33) are arranged on the central shaft (3), the first-stage stator blade (32) is arranged between the first compressor blade (113) and the second compressor blade (123), and the second-stage stator blade (33) is arranged between the second compressor blade (123) and the third compressor blade (133).

10. The turbine engine of claim 6, wherein, An annular groove (34) is formed on the central shaft (3) and is adapted to the first inner ring (111), the second inner ring (121) and the third inner ring (131), so that the first inner ring (111), the second inner ring (121), the third inner ring (131) and the central shaft (3) are flush with the cylindrical surface of the outer casing (2).

Images