Gas turbine rotor structure, design method, assembly method and aircraft engine

Abstract

The present invention discloses a gas turbine rotor structure, a design method, an assembly method and an aircraft engine, comprising a turbine disk and a guide disk. A thickening structure is provided at the spoke portion of the turbine disk, and a weight-reducing groove is provided in a low-stress area at the center of the turbine disk. The second surface of the guide disk is constructed as a first straight structure in the radial outer end area from the first stop and as a second straight structure in the radial inner end area from the first stop, and the second straight structure is inclined at a preset angle. The first surface is constructed as a first arc surface structure, a third straight structure and a second arc surface structure in sequence in the main load-bearing area, and the second surface is constructed as a third arc surface structure, a fourth straight structure and a fourth arc surface structure in sequence in the main load-bearing area. The centers of the first arc surface structure, the third arc surface structure and the fourth arc surface structure are located on the side away from the axial direction, and the center of the second arc surface structure is located on the side close to the axial direction. A segmented straight structure is provided in the low-stress area of ​​the radial inner end wall of the guide disk.

Description

Technical Field

[0001] The present invention relates to the technical field of aero-engines, and in particular to a gas turbine rotor structure, a design method, an assembly method, and an aero-engine. Background Art

[0002] To improve thermal efficiency and power output, small and medium-sized aircraft engines are constantly striving to increase turbine efficiency and turbine rotor inlet temperature (RIT). When the turbine inlet gas temperature exceeds 1500K, current measures to improve turbine efficiency and reduce rotor stage loads are typically implemented through optimizing turbine rotor and stator blade profiles, adopting a multi-stage turbine configuration, and increasing the rotor mid-diameter tangential speed. The first two measures are commonly used in turbine aerodynamic design for small and medium-sized aircraft engines. While maintaining turbine aerodynamic efficiency, they can effectively achieve turbine life. Under the same design conditions, increasing the turbine rotor speed (i.e., the turbine rotor mid-diameter tangential speed) can also significantly improve turbine efficiency while reducing the number of turbine stages and structural design complexity. However, this significantly increases the difficulty of turbine structural strength design. The mid-diameter tangential speed of single-stage turbines for small and medium-sized aircraft engines, both domestically and internationally, is below 480 m / s. For two-stage turbines, the speed is even lower, at less than 350 m / s. Optimizing the rotor and stator blade profile will increase the difficulty of designing and casting the cooling structure of the hollow air-cooled blades, thereby increasing costs; adopting a two-stage turbine design will increase the turbine's axial length, number of parts, and weight, which is not conducive to improving the engine's power-to-weight ratio, while also increasing assembly complexity. Summary of the Invention

[0003] The present invention provides a gas turbine rotor structure, a design method, an assembly method and an aircraft engine to solve the technical problem in the prior art of high difficulty in designing the structural strength of gas turbine rotors for small and medium-sized aircraft engines with increased speed.

[0004] According to one aspect of the present invention, a gas turbine rotor structure is provided, comprising a turbine disk, a guide disk, and rotor cooling blades.

[0005] The outer wall of the spoke portion of the turbine disk is provided with a reduced thickness structure, the thickness of which gradually decreases from the end of the turbine disk toward the spoke, and a weight-reducing groove is provided in a low-stress area at the center of the turbine disk;

[0006] The side of the guide plate away from the turbine disk is a first surface, and the side facing the turbine disk is a second surface.

[0007] The second surface of the guide plate is constructed as a first straight structure in the radial outer end area from the first stop and is constructed as a second straight structure in the radial inner end area from the first stop, and the second straight structure is inclined at a preset angle; the first surface of the guide plate is constructed as a first curved surface structure, a third straight structure and a second curved surface structure in the main load-bearing area from the first stop to the second stop, and the second surface of the guide plate is constructed as a third curved surface structure, a fourth straight structure and a fourth curved surface structure in the main load-bearing area from the first stop to the second stop, the centers of the first curved surface structure, the third curved surface structure and the fourth curved surface structure are located on the side away from the axial direction, and the center of the second curved surface structure is located on the side close to the axial direction;

[0008] A segmented straight structure is provided in the low stress area of ​​the radial inner end wall of the guide plate.

[0009] As a further improvement of the above technical solution, the center of gravity of the comb teeth on the radial outer end portion of the first surface of the guide plate is deviated from the axial direction and the first stop.

[0010] As a further improvement of the above technical solution, the gap between the tenon profile of the rotor cooling blade and the tongue profile of the turbine disk is 0.03-0.05 mm.

[0011] As a further improvement of the above technical solution, the axially extended portion of the guide plate is provided with vent holes evenly distributed along the circumference, for introducing external cooling air into the gap between the guide plate and the turbine plate; the turbine plate is provided with vent grooves evenly distributed circumferentially at the first stop position, for guiding the cooling air in the gap between the guide plate and the turbine plate to the inner cavity of the rotor cooling blades and out to the rear end face of the turbine plate.

[0012] As a further improvement of the above technical solution, the guide plate is pre-tightened and the second stop of the guide plate has a radial interference of 0.05-0.06mm, and the first stop of the guide plate adopts a small clearance fit and the radial fit clearance is 0.05-0.07mm.

[0013] As a further improvement of the above technical solution, the straightness of the second straight structure is greater than 0 and less than 3.5°.

[0014] As a further improvement of the above technical solution, a boss structure is provided on one side of the top of the turbine disk for axially limiting the rotor cooling blades.

[0015] According to another aspect of the present invention, a design method is provided for designing the above-mentioned gas turbine rotor structure, the design method comprising:

[0016] S1. Design a reduced-thickness structure for the turbine disc's spokes based on finite element analysis;

[0017] S2. Design a weight-reducing slot in the low-stress area of ​​the turbine disk center based on finite element analysis;

[0018] S3. Optimize the design of the main load-bearing parts of the guide plate based on finite element analysis;

[0019] S4. Design of weight reduction for low stress area of ​​guide plate based on finite element analysis.

[0020] According to another aspect of the present invention, there is also provided an assembly method for assembling the above-mentioned gas turbine rotor structure, wherein the gas turbine rotor structure further comprises a sealing wire, a nut, and a locking plate. The assembly method comprises:

[0021] A1. Place each rotor cooling blade sequentially into the mortise and tenon groove at the front end of the turbine disk, and axially move the rotor cooling blade to the boss structure.

[0022] A2. Heat the guide plate to above 200°C and place the sealing wire into the installation groove of the guide plate;

[0023] A3. Attach the guide plate to the turbine disk from the front journal. Move the guide plate axially until it contacts the front surface of the rotor cooling blades.

[0024] A4.Insert the locking plates and screw the nuts in place until the guide plate reaches the preset compression amount to complete the assembly.

[0025] According to another aspect of the present invention, an aircraft engine is provided, which includes the above-mentioned gas turbine rotor structure.

[0026] The present invention has the following beneficial effects:

[0027] Compared with the conventional turbine disk design, the turbine disk of this gas turbine rotor structure is designed with a reduced thickness structure to reduce the volume of the upper spoke area in the turbine disk, and at the same time, the weight-reducing groove is designed to reduce the volume of the low stress area in the wheel center, so as to appropriately increase the radial stress level in the spoke area without increasing the weight of the turbine disk, and significantly reduce the maximum stress in the center of the turbine disk; by optimizing the main load-bearing area of ​​the guide disk, the main load-bearing area is overall inclined teardrop-shaped, so as to achieve an improvement in the load-bearing capacity, and by increasing the width of the main load-bearing area of ​​the rotating disk body, the load-bearing capacity of the inner diameter of the guide disk is enhanced, and the weight increase is controlled based on the straight structure design between the arc structures, and the weight-reducing design of the segmented straight structure in the low stress area of ​​the guide disk is combined to reduce the influence of the weight increase caused by the teardrop-shaped structure in the main load-bearing area, thereby greatly improving the rupture resistance and life of the guide disk, and making an important contribution to achieving the improvement of the overall strength and life of the turbine rotor under high temperature and high tangential speed.

[0028] 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 below with reference to the accompanying drawings. BRIEF DESCRIPTION OF THE DRAWINGS

[0029] The accompanying drawings, which constitute part of this application, are intended to provide a further understanding of the present invention. The exemplary embodiments of the present invention and their descriptions are intended to explain the present invention and do not constitute an undue limitation of the present invention. In the accompanying drawings:

[0030] Figure 1 It is a schematic diagram of the structure of a turbine disc in the prior art;

[0031] Figure 2 This is a schematic diagram of the structure of a guide plate in the prior art;

[0032] Figure 3 This is a schematic diagram of the structure of a preferred embodiment of the present invention Figure 1 ;

[0033] Figure 4 is a schematic structural diagram of a turbine disc according to a preferred embodiment of the present invention;

[0034] Figure 5 This is a schematic structural diagram of a guide plate according to a preferred embodiment of the present invention;

[0035] Figure 6 2. It is a schematic structural diagram of a rotor cooling blade according to a preferred embodiment of the present invention;

[0036] Figure 7 This is a schematic diagram of the structure of the rotor cooling blades and the turbine disk in the preferred embodiment of the present invention;

[0037] Figure 8 This is a schematic diagram of the structure of a preferred embodiment of the present invention Figure 2 ;

[0038] Figure 9 Schematic diagram of cooling air flow direction in a preferred embodiment of the present invention;

[0039] Figure 10 The ventilation groove structure of the preferred embodiment of the present invention is shown in FIG. Figure 1 ;

[0040] Figure 11 The ventilation groove structure of the preferred embodiment of the present invention is shown in FIG. Figure 2 ;

[0041] Figure 12 This is a schematic diagram of the installation of rotor cooling blades according to a preferred embodiment of the present invention;

[0042] Figure 13 This is a schematic diagram of the installation of the guide plate and the sealing wire in a preferred embodiment of the present invention;

[0043] Figure 14 This is a schematic diagram of the installation of a nut and a locking plate according to a preferred embodiment of the present invention;

[0044] Figure 15 This is a turbine disk equivalent stress distribution diagram of the prior art;

[0045] Figure 16 This is a circumferential stress distribution diagram of a turbine disk in the prior art;

[0046] Figure 17 This is the equivalent stress distribution diagram of the turbine disk thickness reduction structure of the preferred embodiment of the present invention.

[0047] Figure 18 This is a circumferential stress distribution diagram of a turbine disk with a reduced thickness design according to a preferred embodiment of the present invention;

[0048] Figure 19 This is an equivalent stress distribution diagram of the turbine disk design thickness reduction structure and weight reduction groove in a preferred embodiment of the present invention;

[0049] Figure 20 This is a diagram showing the circumferential stress distribution of the turbine disk design with reduced thickness and weight reduction grooves according to a preferred embodiment of the present invention;

[0050] Figure 21 This is a diagram of equivalent stress distribution of a guide plate in the prior art;

[0051] Figure 22 This is a circumferential stress distribution diagram of a guide plate in the prior art;

[0052] Figure 23 This is the equivalent stress distribution diagram of the optimized design of the main bearing area of ​​the guide plate of the preferred embodiment of the present invention

[0053] Figure 24 This is a circumferential stress distribution diagram of the optimized design of the main bearing area of ​​the guide plate of the preferred embodiment of the present invention;

[0054] Figure 25 This is an equivalent stress distribution diagram of the optimized design and weight reduction design of the main bearing area of ​​the guide plate of the preferred embodiment of the present invention;

[0055] Figure 26 This is a circumferential stress distribution diagram of the optimized design and weight reduction design of the main load-bearing area of ​​the guide plate in the preferred embodiment of the present invention.

[0056] Legend:

[0057] 1. Boss structure 2. Spoke plate position 3. Comb teeth 4. First stop 5. Second stop 6. Vent L1. First inclined section L2. Second inclined section L3. Reduced thickness section L4. First straight structure L5. Second straight structure L6. Fifth straight structure L7. Sixth straight structure L8. Fourth straight structure L9. Third straight structure L10. Axial distance R1 between the first stop and the second stop. Weight reduction groove R2. First arc surface structure R3. Third arc surface structure R4. Fourth arc surface structure. DETAILED DESCRIPTION

[0058] 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.

[0059] like Figures 3 to 14 As shown, the gas turbine rotor structure of this embodiment includes a turbine disk, a guide disk, and rotor cooling blades.

[0060] The outer wall of the turbine disk's spokes is designed with a reduced thickness structure, with the thickness gradually decreasing from the ends of the turbine disk toward the spokes. A weight-reducing groove R1 is provided in the low-stress area at the center of the turbine disk. The reduced thickness structure includes a first inclined section L1 and a second inclined section L2 located on the first and second surfaces, respectively. The minimum thickness after the reduction is the reduced thickness section L3.

[0061] The side of the guide plate away from the turbine disk is the first surface, and the side facing the turbine disk is the second surface.

[0062] The second surface of the guide plate is constructed as a first straight structure L4 in the radially outer end area of ​​the first stop 4 and is constructed as a second straight structure L5 in the radially inner end area of ​​the first stop 4, and the second straight structure L5 is inclined at a preset angle; the first surface of the guide plate is constructed as a first curved surface structure R2, a third straight structure L9 and a second curved surface structure in the main load-bearing area from the first stop 4 to the second stop 5, and the second surface of the guide plate is constructed as a third curved surface structure R3, a fourth straight structure L8 and a fourth curved surface structure R4 in the main load-bearing area from the first stop 4 to the second stop 5, the centers of the first curved surface structure R2, the third curved surface structure R3 and the fourth curved surface structure R4 are located on the side away from the axial direction, and the center of the second curved surface structure is located on the side close to the axial direction;

[0063] A segmented straight structure is provided in a low stress area of ​​the radial inner end wall of the guide plate.

[0064] In one embodiment, reference Figure 15 and Figure 16 , indicating that the maximum equivalent stress and circumferential stress of the wheel center will reach 1665MPa and 1384MPa according to the conventional turbine disk design. Figures 17 to 20, the results of setting the reduced thickness structure at the spoke plate position 2 were 52MPa and 39MPa higher respectively. The main reason is that the addition of the inclined surface to the turbine disc spoke plate in the high tangential speed design reduced the mass of the spoke plate position 2 by 144g, but the radial stress at the spoke plate position 2 was only increased by 14MPa, which still met the strength and life requirements of the spoke plate position 2; the results show that the rupture margin of the turbine disc is improved by more than 2.1%. The reasonable setting of the inclined surface greatly improves the life of the turbine disc while also greatly increasing the rupture margin, which better solves the problem of high strength and life of the single-stage turbine rotor at high tangential speed; further, a weight reduction design is carried out in the low stress area of ​​the turbine disc wheel center to further reduce the weight of the turbine disc At the same time, the load-bearing capacity of the disk center is improved. Based on the design of the reduced thickness structure and further optimization of the design, the weight can be reduced by 37g, so that the equivalent stress and circumferential stress of the turbine disk wheel center are further reduced by 23MPa and 22MPa respectively, and the rupture margin is further improved by 1%. In summary, the maximum equivalent stress and circumferential stress of the turbine disk wheel center of the final configuration are reduced by 75MPa and 61MPa respectively compared with the conventional turbine disk, and the rupture margin is improved by more than 3.1%. Therefore, by designing the reduced thickness structure and the weight-reducing groove R1 in the low-stress area, the maximum stress is reduced by 4.5%, and the rupture margin is increased by more than 3%, breaking through the bottleneck of significantly improving the strength and life of the turbine disk under high temperature and high tangential speed.

[0065] refer to Figure 2 , the main load-bearing area of ​​the conventional guide plate adopts a rounded corner design, and the rest of the part is a straight structure design; this embodiment optimizes the design of the main load-bearing area, adjusts the second straight structure L5 of the upper straight section of the main load-bearing area to be inclined at a small angle, and optimizes the first surface in the main load-bearing area to sequentially construct the first arc surface structure R2, the third straight structure L9 and the second arc surface structure; the second surface in the main load-bearing area is sequentially constructed from the first stop 4 to the second stop 5 to form the third arc surface structure R3, the fourth straight structure L8 and the fourth arc surface structure R4; the centers of the first arc surface structure R2, the third arc surface structure R3 and the fourth arc surface structure R4 are located on the side away from the axial direction, and the center of the second arc surface structure is located on the side close to the axial direction. The main load-bearing area as a whole has an inclined teardrop-shaped configuration, thereby improving the load-bearing capacity. Figure 21 and Figure 22 According to the conventional guide plate design, the maximum equivalent stress and circumferential stress of the wheel center will reach 1027MPa and 1123MPa respectively. Figures 23 to 26, the guide plate structure is 19MPa and 45Mpa high respectively. By increasing the width of the main bearing area of ​​the rotating disc, that is, the radius of the first arc surface structure and the length of the third straight structure L9, the bearing capacity of the inner diameter of the guide plate is enhanced. The weight increase is controlled based on the straight structure design between the arc structures. In this embodiment, the weight increases by 227g, and the rupture margin increases by 1.5%. Similarly, in order to reduce the weight increase caused by the teardrop-shaped structure in the main bearing area, the weight reduction design of the segmented straight structure in the low stress area of ​​the guide plate is combined. For example, this embodiment designs a fifth straight structure L6 with different radial positions. The sixth straight structure L7 maximizes the guide disc's load-bearing capacity, reduces weight by 53g, and increases the maximum equivalent stress and circumferential stress at the disc center by 7MPa and 16MPa, respectively. Compared to conventional structures, the rupture margin is still 1% higher. In summary, the guide disc structure of this embodiment increases weight by 174g, but the maximum equivalent stress and radial stress decrease by 12MPa and 19MPa, respectively, meaning the maximum stress decreases by more than 1.2%. At the same time, the rupture margin is increased by 1%, significantly improving the guide disc's rupture resistance and lifespan, making a significant contribution to improving the overall strength and lifespan of turbine rotors at high temperatures and high tangential speeds.

[0066] The single-stage gas turbine rotor obtained by the structural design of this embodiment has a significantly increased mid-diameter tangential speed of the gas turbine rotor (above 530 m / s) while having a strength and lifespan comparable to that of a single-stage or two-stage gas turbine rotor with a lower mid-diameter tangential speed (below 480 m / s).

[0067] In summary, compared with the conventional turbine disk design, the turbine disk of the gas turbine rotor structure is designed with a reduced thickness structure to reduce the volume of the upper spoke area in the turbine disk, and at the same time, the weight reduction groove R1 is designed to reduce the volume of the low stress area of ​​the wheel center, so as to appropriately increase the radial stress level of the spoke area without increasing the weight of the turbine disk, and significantly reduce the maximum stress of the turbine disk center; by optimizing the main load-bearing area of ​​the guide disk, the main load-bearing area is overall in an inclined teardrop shape, so as to achieve an improvement in the load-bearing capacity, and by increasing the width of the main load-bearing area of ​​the rotating disk body to enhance the load-bearing capacity of the inner diameter of the guide disk body, based on the straight structure design between the arc structures to control the weight increase, combined with the weight reduction design of the segmented straight structure in the low stress area of ​​the guide disk to reduce the impact of the weight increase caused by the teardrop structure in the main load-bearing area, thereby greatly improving the rupture resistance and life of the guide disk, and making an important contribution to achieving the improvement of the overall strength and life of the turbine rotor under high temperature and high tangential speed.

[0068] In this embodiment, the straightness of the second straight structure L5 is greater than 0 and less than 3.5° to match the large-radius arc surface structure design of the main bearing area.

[0069] In one embodiment, the center of gravity of the comb teeth 3 on the radial outer end portion of the first surface of the guide disc is deviated from the axial direction and the first stop 4, thereby forming a pendulum effect under the action of centrifugal load during operation, ensuring that the end face of the guide disc rim fits the upper end face of the turbine disc, thereby enhancing its load-bearing capacity.

[0070] In one embodiment, the axially extended portion of the guide plate is provided with vent holes 6 evenly distributed along the circumference for introducing external cooling air into the gap between the guide plate and the turbine disk; the turbine disk is provided with vent grooves evenly distributed along the circumference at the position of the first stop 4 for guiding the cooling air in the gap between the guide plate and the turbine disk into the inner cavity of the rotor cooling blades and out to the rear end face of the turbine disk, thereby achieving cooling air induction;

[0071] Furthermore, the gap between the tenon profile of the rotor cooling blade and the tenon profile of the turbine disk is 0.03-0.05 mm;

[0072] It should be noted that the single-stage design can reduce the axial length, number of parts, weight, etc. of the turbine, further improve the engine power-to-weight ratio, and the complexity of rotor assembly; while in conventional designs, the gap between the rotor cooling blade tenon and the gas turbine disc tenon is more than 0.2 mm. In order to prevent a large amount of cold air from leaking from the rear end of the gas turbine disc, a guide plate needs to be added to the rear end of the gas turbine disc in the prior art; the rotor cooling blade tenon profile and the turbine disc tenon profile are designed with a small gap (i.e., 0.03-0.05 mm) in this embodiment, which greatly limits the leakage of cold air when the gas turbine rotor has no rear baffle or rear guide plate, thereby achieving precise control of the air supply to the single-stage gas turbine rotor cooling blade and reducing the complexity of the structure.

[0073] In this embodiment, the first arc surface structure R2, the third arc surface structure R3 and the fourth arc surface structure R4 all have a large radius of 10-15mm; the large radius first arc surface structure R2 is combined with the addition of the third straight structure L9 to enhance the bearing capacity of the guide plate body; the large radius third arc surface structure R3 and the fourth arc surface structure R4 are used to enhance the bearing capacity while controlling the cold air flow gap between the guide plate and the turbine plate, further realizing precise control of the air supply to the cooling blades.

[0074] In one embodiment, the guide plate is pre-tightened (after pre-tightening, the axial distance L10 between the first stop and the second stop is shortened by 0.3-0.4 mm) and the second stop 5 of the guide plate has a radial interference of 0.05-0.06 mm. The first stop 4 of the guide plate adopts a small clearance fit and the radial fit clearance is 0.05-0.07 mm, thereby improving the centering reliability of the gas turbine rotor at high speed.

[0075] In one embodiment, a boss structure 1 is provided on one side of the top of the turbine disk for axially limiting the rotor cooling blades.

[0076] The design method of this embodiment is used to design the above-mentioned gas turbine rotor structure, and the design method includes:

[0077] S1. Based on finite element analysis, a reduced-thickness design was implemented for the turbine disk's spoke area. By reducing the volume of the upper spoke area in the gas turbine disk, the volume of the low-radius, low-stress zone at the wheel center was reduced. This approach allowed for a moderate increase in radial stress in the spoke area without increasing the weight of the turbine disk, significantly reducing the maximum stress at the turbine disk's wheel center.

[0078] S2. Design a weight-reducing slot in the low-stress area of ​​the turbine disk's center based on finite element analysis; optimize the iterative thickness reduction structure and the design of the weight-reducing slot R1 based on finite element analysis;

[0079] S3. Optimize the design of the main load-bearing parts of the guide plate based on finite element analysis;

[0080] S4. Design of weight reduction for low stress area of ​​guide plate based on finite element analysis.

[0081] The final structural design was obtained by comparing the gas turbine rotor obtained by finite element analysis with the existing design, achieving a significant improvement in strength and life at high tangential speeds.

[0082] The assembly method of this embodiment is used to assemble the above-mentioned gas turbine rotor structure, and the assembly method includes:

[0083] A1. Place each rotor cooling blade sequentially into the mortise and tenon groove at the front end of the turbine disk, and axially move the rotor cooling blade to the boss structure 1 position;

[0084] A2. Heat the guide plate to above 200°C and place the sealing wire into the installation groove of the guide plate;

[0085] A3. Attach the guide plate to the turbine disk from the front journal. Move the guide plate axially until it contacts the front surface of the rotor cooling blades.

[0086] A4.Insert the locking plates and screw the nuts in place until the guide plate reaches the preset compression amount to complete the assembly.

[0087] On the other hand, an aircraft engine is also provided, which is equipped with the above-mentioned gas turbine rotor.

[0088] The foregoing description is merely a preferred embodiment of the present invention and is not intended to limit the present invention. Those skilled in the art will readily appreciate that various modifications and variations of the present invention are possible. Any modifications, equivalent substitutions, or improvements made within the spirit and principles of the present invention are intended to be within the scope of protection of the present invention.

Claims

1. A gas turbine rotor structure, comprising a turbine disk, a guide disk, and rotor cooling blades, characterized in that: The outer wall of the spoke portion of the turbine disk is provided with a reduced thickness structure, the thickness of which gradually decreases from the end of the turbine disk toward the spoke, and a weight-reducing groove is provided in a low-stress area at the center of the turbine disk; The side of the guide plate away from the turbine disk is a first surface, and the side facing the turbine disk is a second surface. The second surface of the guide plate is constructed as a first straight structure in the radial outer end area from the first stop and is constructed as a second straight structure in the radial inner end area from the first stop, and the second straight structure is inclined at a preset angle; the first surface of the guide plate is constructed as a first curved surface structure, a third straight structure and a second curved surface structure in the main load-bearing area from the first stop to the second stop, and the second surface of the guide plate is constructed as a third curved surface structure, a fourth straight structure and a fourth curved surface structure in the main load-bearing area from the first stop to the second stop, the centers of the first curved surface structure, the third curved surface structure and the fourth curved surface structure are located on the side away from the axial direction, and the center of the second curved surface structure is located on the side close to the axial direction; A segmented straight structure is provided in the low stress area of ​​the radial inner end wall of the guide plate, and the segmented straight structure includes a fifth straight structure and a sixth straight structure at different radial positions.

2. The gas turbine rotor structure according to claim 1, characterized in that: The center of gravity of the grate teeth on the radially outer end portion of the first surface of the guide plate is deviated from the axial direction and the first stop.

3. The gas turbine rotor structure according to claim 1, wherein: The gap between the tenon profile of the rotor cooling blade and the tenon profile of the turbine disk is 0.03-0.05 mm.

4. The gas turbine rotor structure according to claim 3, characterized in that: The axially extended portion of the guide disc is provided with vent holes evenly distributed along the circumference for introducing external cooling air into the gap between the guide disc and the turbine disc; the turbine disc is provided with vent grooves evenly distributed circumferentially at the first stop position for guiding the cooling air in the gap between the guide disc and the turbine disc to the inner cavity of the rotor cooling blades and out to the rear end face of the turbine disc.

5. The gas turbine rotor structure according to claim 1, characterized in that: The guide plate is pre-tightened and assembled, and the second stop of the guide plate has a radial interference of 0.05-0.06 mm. The first stop of the guide plate adopts a small clearance fit, and the radial fit clearance is 0.05-0.07 mm.

6. The gas turbine rotor structure according to any one of claims 1 to 5, characterized in that: A boss structure is provided on one side of the top of the turbine disk for axially limiting the rotor cooling blades.

7. A design method, characterized in that: For designing the gas turbine rotor structure according to any one of claims 1 to 6, the design method comprises: S1. Design a reduced-thickness structure for the turbine disc's spokes based on finite element analysis; S2. Design a weight-reducing slot in the low-stress area of ​​the turbine disk center based on finite element analysis; S3. Optimize the design of the main load-bearing parts of the guide plate based on finite element analysis; S4. Design of weight reduction for low stress area of ​​guide plate based on finite element analysis.

8. An assembly method, characterized in that: For assembling the gas turbine rotor structure according to claim 6, the gas turbine rotor structure further comprising a sealing wire, a nut and a locking plate, the assembly method comprising: A1. Place each rotor cooling blade sequentially into the mortise and tenon groove at the front end of the turbine disk, and axially move the rotor cooling blade to the boss structure. A2. Heat the guide plate to above 200°C and place the sealing wire into the installation groove of the guide plate; A3. Attach the guide plate to the turbine disk from the front journal. Move the guide plate axially until it contacts the front surface of the rotor cooling blades. A4.Insert the locking plates and screw the nuts in place until the guide plate reaches the preset compression amount to complete the assembly.

9. An aircraft engine, characterized in that: The gas turbine rotor structure according to any one of claims 1 to 6 is applied.

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