Blisk for high thrust ratio gas turbine engine and turbine engine

Abstract

The invention discloses an integral blade ring for a high thrust ratio gas turbine engine and the turbine engine, the integral blade ring comprises a first-stage combined rotor, a second-stage combined rotor and a third-stage combined rotor, the first-stage combined rotor comprises a first inner ring, a first outer ring and a first compressor blade; the second-stage combined rotor comprises a second inner ring, a second outer ring, a second compressor blade and a first turbine blade; the three-stage combined rotor comprises a third inner ring, a third outer ring, a third compressor blade and a second turbine blade; the first outer ring, the second outer ring and the third outer ring are sequentially connected in the axial direction to form a gas compressor flow channel, and a first bearing seat is arranged on the first inner ring; the centrifugal force of the blade is borne by the rotor drum, so that the power capability of the blade at a high tangential speed can be effectively utilized, and the problem of performance reduction caused by air leakage of a blade tip can be solved; a connecting shaft for transmitting work between the gas compressor and the turbine is omitted, the number of bearings and bearing cavities can be reduced, the overall weight of the engine is reduced, and the thrust-weight ratio of the engine is increased.

Description

Technical Field

[0001] The present invention relates to the technical field of aeroengines, and in particular, to an integral blade ring for a high thrust-to-weight ratio gas turbine engine. In addition, the present invention also relates to a turbine engine comprising the integral blade ring for a high thrust-to-weight ratio gas turbine engine. Background Art

[0002] like Figure 1 As shown, the structure of the existing aviation gas turbine engine is that the rotating compressor and turbine are connected on the same shaft, and a heat source (combustion chamber) is installed between the two. Air is continuously sucked into the compressor, and after being compressed and pressurized therein, it enters the combustion chamber and is sprayed and burned to become high-temperature and high-pressure combustion gas, and then enters the turbine to expand and do work.

[0003] Since the compressor, combustion chamber and turbine are arranged linearly along the axial direction, the force transmission path is too long, and at least three bearings and three bearing cavities need to be arranged in front of the compressor, after the compressor, in front of the turbine, and after the turbine. This results in a complex structure of the lubricating oil system, which not only makes the entire engine larger in size, but also increases the weight of the engine, affecting the engine's thrust-to-weight ratio. In addition, the span between the compressor flow path and the turbine flow path is large, and the design of the air induction system is complex. Summary of the invention

[0004] The present invention provides an integral blade ring and a turbine engine for a high thrust-to-weight ratio gas turbine engine, so as to solve the technical problems that the existing aviation gas turbine engine has a too long force transmission path, requires multiple bearing fulcrums and multiple bearing cavities, the entire engine is large in size and heavy, which affects the engine thrust-to-weight ratio, and the lubricating oil system and the air bleed system are complex in design.

[0005] According to one aspect of the present invention, an integral blade ring for a high thrust-ratio gas turbine engine is provided, comprising a first-stage combined rotor, a second-stage combined rotor and a third-stage combined rotor, the first-stage combined rotor comprising a first inner ring, a first outer ring and a first compressor blade arranged between the first inner ring and the first outer ring; the second-stage combined rotor comprising a second inner ring, a second outer ring, a second compressor blade arranged between the second inner ring and the second outer ring, and a first turbine blade arranged on the outer side of the second outer ring; the third-stage combined rotor comprising a third inner ring, a third outer ring, a third compressor blade arranged between the third inner ring and the third outer ring, and a second turbine blade arranged on the outer side of the third outer ring; the first outer ring, the second outer ring and the third outer ring are sequentially connected along the axial direction to form a compressor flow channel, and a first bearing seat is arranged on the first inner ring.

[0006] Furthermore, reinforcement rings are arranged on the outer sides of the second outer ring and the third outer ring respectively, the second outer ring is connected to the first turbine blades via the reinforcement rings, and the third outer ring is connected to the second turbine blades via the reinforcement rings.

[0007] Furthermore, a first air inlet hole facing the root of the first turbine blade is formed on the second outer ring, and a second air inlet hole facing the root of the second turbine blade is formed on the third outer ring.

[0008] Furthermore, a first limiting structure for limiting the axial position of the first turbine blade is arranged on the second outer ring, and 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 chamber, and the first cooling chamber and the compressor flow passage pass through the first air duct hole; a second limiting structure for limiting the axial position of the second turbine blade is arranged on the third outer ring, and 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 chamber, and the second cooling chamber and the compressor flow passage pass through the second air duct hole.

[0009] Furthermore, comb teeth are arranged on the first limiting structure and the second limiting structure.

[0010] According to another aspect of the present invention, a turbine engine is also provided, which includes the above-mentioned integral blade ring for a high thrust-ratio gas turbine engine, and also includes an outer casing, a center shaft, a diffuser, a combustion chamber, a turbine flow channel and a tail nozzle; the outer casing is rotatably sealed and connected to the first outer ring, the diffuser is rotatably sealed and connected to the third outer ring, and the two ends of the center shaft are rotatably connected to the first bearing seat and the third inner ring through bearings respectively; the first turbine blade extends into the turbine flow channel, and the comb teeth on the first limiting structure are rotatably sealed and connected to the outer wall surface of the turbine flow channel; the second turbine blade extends into the tail nozzle, and is arranged at one end of the tail nozzle close to the turbine flow channel, and the comb teeth on the second limiting structure are rotatably sealed and connected to the outer wall surface of the turbine flow channel.

[0011] Furthermore, an inlet guide vane is arranged on the outer casing, a first lubricating oil channel is opened on the inlet guide vane, a second lubricating oil channel for introducing lubricating oil into the bearing is opened on the central shaft, and the first lubricating oil channel is connected to the second lubricating oil channel; lubricating oil sealing structures are respectively arranged on the first bearing seat and the third inner ring, and the lubricating oil sealing structures are arranged at both ends of the bearing.

[0012] Furthermore, the first lubricating oil passage includes an oil inlet passage and an oil return passage.

[0013] Furthermore, a first-stage stator blade and a second-stage stator blade are arranged on the central axis, the first-stage stator blade is arranged between the first compressor blade and the second compressor blade, and the second-stage stator blade is arranged between the second compressor blade and the third compressor blade.

[0014] Furthermore, annular grooves respectively matched with the first inner ring, the second inner ring and the third inner ring are formed on the central axis, so that the first inner ring, the second inner ring, the third inner ring and the central axis are flush with the cylindrical surface of the outer casing.

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

[0016] The high thrust ratio gas turbine engine of the present invention uses an integral blade ring, wherein 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 channel is formed inside the rotor drum, and the first compressor blade, the second compressor blade, and the third compressor blade are arranged on the inner circumferential 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 all wide at the part connected to the rotor drum, and narrow at the part close to the central axis. The centrifugal force of the first compressor blade, the second compressor blade, and the third compressor blade is borne by the rotor drum, which can not only effectively utilize the working capacity of the blade at a high tangential speed, but also solve the problem of tip leakage ( The first turbine blades and the second turbine blades are connected to the outer circumferential surface of the rotor drum. When working, the rotor drum drives the first compressor blades, the second compressor blades, the third compressor blades, the first turbine blades and the second turbine blades to rotate synchronously. Compared with the traditional structure, the interchangeable rotor and stator structure improves the work capacity of the compressor blades and makes full use of the space in the vertical axial direction. The connecting shaft for transmitting work between the compressor and the turbine is eliminated. The two ends of the rotor drum are connected to the bearings through the first bearing seat and the third inner ring respectively, which can reduce the number of bearings and bearing cavities, thereby simplifying the structure, shortening the overall axial length of the engine, reducing the overall weight of the engine, and improving the thrust-to-weight ratio of the engine.

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

[0018] The drawings constituting a part of this application are used to provide a further understanding of the present invention. The exemplary embodiments of the present invention and their descriptions are used to explain the present invention and do not constitute an improper limitation of the present invention. In the drawings:

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

[0020] Figure 2 2 is a schematic structural diagram of an integral blade ring for a high thrust-to-weight ratio gas turbine engine according to a preferred embodiment of the present invention;

[0021] Figure 3 is a structural schematic diagram of a two-stage combined rotor according to a preferred embodiment of the present invention;

[0022] Figure 4 It is one of the structural schematic diagrams of the existing compressor rotor;

[0023] Figure 5 This is the second structural schematic diagram of the existing compressor rotor;

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

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

[0026] Figure 8 is a schematic structural diagram of a turbine engine according to a preferred embodiment of the present invention;

[0027] Fig. 9 It is a schematic structural diagram of the central axis of a preferred embodiment of the present invention.

[0028] Legend:

[0029] 1. Integral blade ring; 11. First stage 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 stage combined rotor; 121. Second inner ring; 122. Second outer ring; 123. Second compressor blade; 124. First turbine blade; 13. Third stage combined rotor; 131. Third inner ring; 132. Third outer ring; 133. Third compressor blade; 134. Second turbine blade; 1 4. Reinforcement ring; 15. First air inlet hole; 16. Second air inlet hole; 17. First limiting structure; 18. Second limiting structure; 19. Combustion ring; 2. Outer casing; 21. Inlet guide vane; 211. First lubricating oil channel; 3. Center axis; 31. Second lubricating oil channel; 32. First stage stator blade; 33. Second stage stator blade; 34. Annular groove; 4. Diffuser; 5. Combustion chamber; 6. Turbine flow channel; 7. Tail nozzle; 8. Bearing; 100. Wheel; 101. Compressor blade; 102. Blade crown. DETAILED DESCRIPTION

[0030] The embodiments of the present invention are described in detail below with reference to the accompanying drawings. However, the present invention can be implemented in many different ways as defined and covered below.

[0031] like Figure 2and Figure 3 As shown, the integrated blade ring for the high thrust-to-weight ratio gas turbine engine of 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 arranged 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 arranged between the second inner ring 121 and the second outer ring 122. and first turbine blades 124 arranged on the outer side of the second outer ring 122; the three-stage combined rotor 13 includes a third inner ring 131, a third outer ring 132, third compressor blades 133 arranged between the third inner ring 131 and the third outer ring 132, and second turbine blades 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 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 wheel disc 100, and the centrifugal force of the compressor blades 101 is borne by the wheel disc 100 at the center. Due to the existence of centrifugal force, the compressor blades 101 must be designed to be sharp at the top and wide at the bottom, so as to ensure that the stress levels at the tip and root of the blades are the same and the material is effectively utilized; however, the high tangential speed of the blade tip is not effectively utilized. Figure 5 As shown, in order to effectively utilize the high tangential speed of the blade tip, a blade crown 102 is arranged at the blade tip of the compressor blade 101. Although the blade crown 102 can reduce the leakage (secondary flow) of the blade tip of the compressor blade 101 and improve the work efficiency, the centrifugal force is greatly increased due to the increased weight of the blade tip. The centrifugal force of the blade structure with a crown is still borne by the central wheel 100. To ensure that the wheel 100 meets the use requirements, the wheel 100 needs to be made very heavy, which will not only increase the overall weight of the engine and affect the engine thrust-to-weight ratio, but also occupy a large amount of space.

[0033] The high thrust ratio gas turbine engine of this embodiment uses an integral blade ring, and the first outer ring 112 and the second outer ring 122, and the second outer ring 122 and the third outer ring 132 are respectively connected by bolts to form a rotor drum, and a compressor flow channel is formed inside the rotor drum. The first compressor blades 113, the second compressor blades 123, and the third compressor blades 133 are arranged on the inner circumferential surface of the rotor drum. In order to ensure the same stress level, the first compressor blades 113, the second compressor blades 123, and the third compressor blades 133 are all wide at the parts connected to the rotor drum, and narrow at the parts away from the rotor drum. The centrifugal force of the first compressor blades 113, the second compressor blades 123, and the third compressor blades 133 is borne by the rotor drum. Since the tangential speed at the outer ring is high and the working area is large, not only can the working capacity of the blades at the high tangential speed be effectively utilized, but also the reduction of performance due to tip leakage (secondary flow loss) can be solved; the first turbine blades 12 4. The second turbine blades 134 are connected to the outer circumferential surface of the rotor drum. When working, the rotor drum drives the first compressor blades 113, the second compressor blades 123, the third compressor blades 133, the first turbine blades 124 and the second turbine blades 134 to rotate synchronously. Compared with the traditional structure, the interchangeable rotor-stator structure improves the work capacity of the compressor blades and makes full use of the space in the vertical axial direction. The connecting shaft for transferring work between the compressor and the turbine is eliminated. The two ends of the rotor drum are connected to the bearings through the first bearing seat 114 and the third inner ring 131 respectively, which can shorten the axial length of the turbine engine by more than one time, and at the same time can reduce the number of bearings and bearing cavities, thereby simplifying the structure, reducing the overall weight of the engine, and improving the thrust-to-weight ratio of the engine; it can be understood that the number of the first-stage combined rotor 11, the second-stage combined rotor 12 or the third-stage combined rotor 13 can be adjusted and combined to form an integral blade ring to meet the use requirements of different scenarios.

[0034] like Figure 2As shown, in this embodiment, reinforcement rings 14 are respectively arranged on the outer sides 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 reinforcement ring 14, the reinforcement 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 reinforcement ring 14, the reinforcement ring 14 on the third outer ring 132, the third compressor blade 133 and the third inner ring 131 are integrally formed, the reinforcement ring 14 can increase the strength of the rotor drum, the tongue and groove structure for assembling the turbine blade is processed on the reinforcement ring 14, the centrifugal force of the blade is borne by the reinforcement ring 14, and deformation of the rotor drum can be avoided. Optionally, a reinforcement ring 14 is arranged on the outer surface of the first outer ring 112. The reinforcement ring 14 can increase the strength of the first outer ring 112 to prevent the first outer ring 112 from deforming, and can also act as a counterweight to improve the stability of the rotor drum. Optionally, the reinforcement ring 14 is a hollow structure to reduce the weight of the rotor drum.

[0035] like Figure 6 As shown in the figure, since the compressor, combustion chamber and turbine are arranged linearly along the axial direction, the span of the compressor flow channel and the turbine flow channel is large, the bleed air path of the cooling gas is long, and the bleed air system design is complex. Figure 2 and Figure 7 As shown, since the compressor decelerates and increases the pressure of the airflow, the cross-sectional area of ​​the compressor flow path gradually decreases from the inlet to the outlet; the turbine increases and decreases the pressure of the airflow, and the cross-sectional area of ​​the turbine flow path 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 circumferential surface of the rotor drum, in order to maximize the use of the space perpendicular to the axial direction, the structural characteristics of the compressor flow path and the turbine flow path are integrated, and the turbine and the compressor are placed side by side. The second outer ring 122 is provided with a first air inlet hole 15 facing the root of the first turbine blade 124. The cooling gas flows to the root of the first turbine blade 124 through the first air inlet hole 15, and the connection between the first turbine blade 124 and the reinforcement ring 14 is cooled. The cooling gas can enter the first turbine blade 124 to cool the blade body of the first turbine blade 124; the third outer ring 132 is provided with a second air inlet hole 16 facing the root of the second turbine blade 134. The cooling gas flows to the root of the second turbine blade 134 through the second air inlet hole 16 to cool the connection between the second turbine blade 134 and the reinforcement ring 14, and can enter the second turbine blade 134 to cool the blade body of the second turbine blade 134; the air inlet path of the cooling gas can be greatly reduced, and the design of the air inlet system can be simplified.

[0036] like Figure 2As shown, in this embodiment, a first limiting structure 17 for limiting the axial position of the first turbine blade 124 is arranged on the second outer ring 122. 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 and the compressor flow passage pass through the first air inlet hole 15. The first limiting structure 17 is arranged on both sides of the first turbine blade 124, thereby limiting the axial position of the first turbine blade 124 and ensuring the stability of the first turbine blade 124. The first cooling chamber can prevent the cooling gas from quickly entering the first turbine blade 124, thereby ensuring the cooling effect; a first limiting structure 17 for limiting the second turbine blade 124 is arranged on the third outer ring 132. The second limiting structure 18 at the axial position of the 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 reinforcement ring 14 form a second cooling chamber. The second cooling chamber and the compressor flow channel are connected through the second air inlet hole 16. The reinforcement ring 14 is provided with a tenon 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 tenon on the reinforcement ring 14, thereby limiting the second turbine blade 134 and ensuring the stability of the second turbine blade 134. The second cooling chamber can prevent the cooling gas from quickly entering the second turbine blade 134 and ensure the cooling effect.

[0037] like Figure 2 As shown, in this embodiment, comb teeth 19 are arranged on the first limiting structure 17 and the second limiting structure 18, and the comb teeth 19 and the outer circumferential surface of the turbine flow channel 6 form a rotating sealing structure to prevent the airflow of the turbine flow channel 6 from leaking out.

[0038] like Figure 8 and Fig. 9As shown, a turbine engine includes the above-mentioned integral blade ring 1 for a high thrust-to-weight ratio gas turbine engine, and also includes 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 rotatably sealed and connected to the first outer ring 112, the diffuser 4 is rotatably sealed and connected to the third outer ring 132, and both ends of the central shaft 3 are rotatably connected to the first bearing seat 114 and the third inner ring 131 through bearings 8 respectively; the first turbine blade 124 extends into the turbine flow channel 6, and the comb teeth 19 on the first limiting structure 17 are rotatably sealed and connected to the outer wall surface of the turbine flow channel 6; the second turbine blade 134 extends into the tail nozzle 7, and Arranged at one end of the tail nozzle 7 close to the turbine flow channel 6, the comb teeth 19 on the second limiting structure 18 are rotatably sealed and connected to the outer wall surface of the turbine flow channel 6; it makes full use of the space in the vertical axial direction, eliminates the connecting shaft for transmitting work between the compressor and the turbine, and the two ends of the integral blade ring 1 are rotatably connected to the central axis 3 through bearings 8 respectively. The first bearing seat 114 and the third inner ring 131 are located on the same stator component, which improves the concentricity of the front and rear bearings, thereby simplifying the structure, reducing the difficulty of processing, reducing the number of bearings and bearing cavities, shortening the overall axial length of the engine, reducing the overall weight of the engine, and improving the thrust-to-weight ratio of the engine.

[0039] like Figure 8 and Fig. 9 As shown, in this embodiment, an inlet guide vane 21 is arranged on the outer casing 2, a first lubricating oil channel 211 is opened on the inlet guide vane 21, a second lubricating oil channel 31 for introducing lubricating oil into the bearing 8 is opened on the central shaft 3, and the first lubricating oil channel 211 is communicated with the second lubricating oil channel 31; lubricating oil sealing structures 116 are arranged on the first bearing seat 114 and the third inner ring 131 respectively, and the lubricating oil sealing structures 116 are arranged at both ends of the bearing 8, and the first lubricating oil channel 211 is integrated on the inlet guide vane 21, and the lubricating oil enters the bearing 8 at the front and rear ends of the central shaft 3 respectively through the first lubricating oil channel 211 and the second lubricating oil channel 31. The structure is simple and can maximize the use of the internal space of the inlet guide vane 21. On the one hand, it can reduce the weight of the inlet guide vane 21, thereby reducing the weight of the entire engine. On the other hand, it can avoid the influence of the additional first lubricating oil channel 211 on the flow field.

[0040] In this embodiment, the first lubricating oil channel 211 includes an oil inlet channel and an oil return channel, which can form a lubricating oil circulation loop to ensure the lubrication effect of the bearing 8. Since the second lubricating oil channel 31 of the center shaft 3 is used for oil inlet and return, the design of the oil circuit is greatly simplified. Moreover, since the number of bearing cavities is reduced, the number of engine stators is reduced, which can reduce the overall assembly difficulty of the engine and improve the disassembly and assembly efficiency.

[0041] like Figure 8 and Fig. 9As shown, in this embodiment, the first-stage stator blades 32 and the second-stage stator blades 33 are arranged on the central shaft 3. The first-stage stator blades 32 are arranged between the first compressor blades 113 and the second compressor blades 123, and the second-stage stator blades 33 are arranged between the second compressor blades 123 and the third compressor blades 133, which can decelerate and pressurize the airflow and improve the work efficiency of the compressor.

[0042] like Figure 8 and Fig. 9 As shown, in this embodiment, an annular groove 34 is opened on the central shaft 3, which is respectively 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 facing the outer casing 2 to avoid affecting the flow field in the compressor flow channel.

[0043] The above description is only a preferred embodiment of the present invention and is not intended to limit the present invention. For those skilled in the art, the present invention may have various modifications and variations. Any modification, equivalent replacement, improvement, etc. made within the spirit and principle of the present invention shall be included in the protection scope of the present invention.

Claims

1. An integral blade ring for a high thrust-to-weight ratio gas turbine engine, characterized in that: The invention comprises a first-stage combined rotor (11), a second-stage combined rotor (12) and a third-stage combined rotor (13), wherein the first-stage combined rotor (11) comprises 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-stage combined rotor (12) comprises 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 third compressor blade (123) arranged on the second outer ring (122). The invention relates to a first turbine blade (124) on the outer surface; the three-stage combined rotor (13) comprises 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 sequentially connected along the axial direction to form a compressor flow channel, and a first bearing seat (114) is arranged on the first inner ring (111).

2. The integrated blade ring for a high thrust ratio gas turbine engine according to claim 1, characterized in that: Reinforcement rings (14) are respectively arranged on the outer sides of the second outer ring (122) and the third outer ring (132); the second outer ring (122) is connected to the first turbine blades (124) via the reinforcement ring (14); and the third outer ring (132) is connected to the second turbine blades (134) via the reinforcement ring (14).

3. The integrated blade ring for a high thrust ratio gas turbine engine according to claim 2, characterized in that: The second outer ring (122) is provided with a first air inlet hole (15) facing the root of the first turbine blade (124), and the third outer ring (132) is provided with a second air inlet hole (16) facing the root of the second turbine blade (134).

4. The integrated blade ring for a high thrust ratio gas turbine engine according to claim 3, characterized in that: 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, and the first cooling chamber and the compressor flow passage pass through the first air inlet 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 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, and the second cooling chamber and the compressor flow passage pass through the second air inlet hole (16).

5. The integrated blade ring for a high thrust ratio gas turbine engine according to claim 4, characterized in that: Comb teeth (19) are arranged on the first limiting structure (17) and the second limiting structure (18).

6. A turbine engine, characterized in that: The invention comprises an integral blade ring for a high thrust-to-weight ratio gas turbine engine as claimed in claim 5, 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 rotatably sealedly connected to the first outer ring (112), the diffuser (4) is rotatably sealedly connected to the third outer ring (132), and both ends of the central shaft (3) are respectively connected to the first bearing seat (114) and the third inner ring (132) through bearings (8). (131) is rotatably connected; the first turbine blade (124) extends into the turbine flow channel (6), and the comb teeth (19) on the first limiting structure (17) are rotatably sealed and connected to 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), and the comb teeth (19) on the second limiting structure (18) are rotatably sealed and connected to the outer wall surface of the turbine flow channel (6).

7. The turbine engine according to claim 6, characterized in that The outer casing (2) is provided with an inlet guide vane (21), and the inlet guide vane (21) is provided with a first lubricating oil passage (211); the central shaft (3) is provided with a second lubricating oil passage (31) for introducing lubricating oil into the bearing (8), and the first lubricating oil passage (211) is communicated with the second lubricating oil passage (31); and the first bearing seat (114) and the third inner ring (131) are respectively provided with lubricating oil sealing structures (116), and the lubricating oil sealing structures (116) are arranged at both ends of the bearing (8).

8. The turbine engine according to claim 7, characterized in that The first lubricating oil passage (211) comprises an oil inlet passage and an oil return passage.

9. The turbine engine according to claim 6, characterized in that A first-stage stator blade (32) and a second-stage stator blade (33) are arranged on the central shaft (3), wherein 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 according to claim 6, characterized in that The central shaft (3) is provided with annular grooves (34) respectively 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).

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