Gas turbine rotor structure, design method, assembly method and aero-engine

Abstract

The invention discloses a gas turbine rotor structure, a design method, an assembly method and an aero-engine. The gas turbine rotor structure comprises a turbine disc and a flow guide disc, a thickness reduction structure is arranged on a radial plate part of the turbine disc, and a weight reduction groove is formed in a disc center low-stress area of the turbine disc; the second surface of the flow guide disc is constructed into a first straight structure from the first seam allowance to the radial outer end area and constructed into a second straight structure from the first seam allowance to the radial inner end, and the second straight structure is obliquely arranged at a preset angle; the first surface is sequentially constructed into a first cambered surface structure, a third straight structure and a second cambered surface structure in the main bearing area, and the second surface is sequentially constructed into a third cambered surface structure, a fourth straight structure and a fourth cambered surface structure in the main bearing area; the circle centers of the first cambered surface structure, the third cambered surface structure and the fourth cambered surface structure are located on the side away from the axial direction, and the circle center of the second cambered surface structure is located on the side close to the axial direction. And a segmented straight structure is arranged in a low-stress area of the radial inner end wall surface of the flow guide disc.

Description

Technical Field

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

[0002] In order to improve thermal efficiency and power output, small and medium-sized aircraft engines are constantly improving turbine efficiency and gas turbine rotor inlet temperature (RIT). When the gas temperature at the turbine inlet of small and medium-sized aircraft engines exceeds 1500K, in order to improve the efficiency of the gas turbine and reduce the rotor stage load, measures such as optimizing the turbine rotor and stator blade profile, adopting a multi-stage turbine configuration, and increasing the rotor mid-diameter tangential speed are generally adopted. The first two measures are commonly used in the current aerodynamic design of small and medium-sized aircraft engines. Under the premise of ensuring the aerodynamic efficiency of the turbine, the turbine life level can be well achieved. Increasing the turbine rotor speed (i.e., the turbine rotor mid-diameter tangential speed) under the same design conditions can also greatly improve the turbine efficiency, while reducing the number of turbine stages and the complexity of structural design. However, it will greatly increase the difficulty of turbine structural strength design. It is known that the mid-diameter tangential speed of the single-stage turbine rotor of small and medium-sized aircraft engines at home and abroad is below 480m / s, and the two-stage is even lower, and the mid-diameter tangential speed is less than 350m / s. Optimizing the rotor and stator blade shape 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, weight, etc., which is not conducive to improving the engine's power-to-weight ratio, while 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 of high difficulty in designing the structural strength of gas turbine rotors for small and medium-sized aircraft engines in the prior art.

[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 plate 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 plate, 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 at 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 tenon 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 disk; the turbine disk 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 disk to the inner cavity of the rotor cooling blades and out to the rear end face of the turbine disk.

[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 to limit the axial position of 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 radial plate of the turbine disk based on finite element analysis;

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

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

[0019] S4. Weight reduction design of 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, and the assembly method comprises:

[0021] A1. Place each rotor cooling blade in sequence from the mortise and tenon position at the front end of the turbine disc, and axially move the rotor cooling blade to the boss structure position;

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

[0023] A3. Transfer the guide plate from the front end journal of the turbine disk to the guide plate, and move the guide plate axially until the guide plate fits the front end surface of the rotor cooling blade;

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

[0025] According to another aspect of the present invention, there is also provided an aircraft engine comprising 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 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-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 in a water droplet shape, so as to improve 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, 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 water droplet-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 improving the overall strength and life of the turbine rotor under high temperature and high tangential speed.

[0028] In addition to the above-described purposes, features and advantages, the present invention has other purposes, features and advantages. The present invention will be further described in detail with reference to the accompanying drawings. BRIEF DESCRIPTION OF THE DRAWINGS

[0029] 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:

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

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

[0032] Figure 3 The structure of the preferred embodiment of the present invention is shown in FIG. Figure 1 ;

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

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

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

[0036] Figure 7 It 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 The structure of the preferred embodiment of the present invention is shown in FIG. Figure 2 ;

[0038] Fig. 9 is a schematic diagram of cooling air flow direction in a preferred embodiment of the present invention;

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

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

[0041] Fig.12 is a schematic diagram of the installation of rotor cooling blades in a preferred embodiment of the present invention;

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

[0043] Fig.14 It is a schematic diagram of the installation of the nut and the locking plate of the preferred embodiment of the present invention;

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

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

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

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

[0048] Fig.19 It is an equivalent stress distribution diagram of the turbine disk design thickness reduction structure and weight reduction groove of the preferred embodiment of the present invention;

[0049] Fig. 20 It is a circumferential stress distribution diagram of the turbine disk design thickness reduction structure and the weight reduction groove of the preferred embodiment of the present invention;

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

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

[0052] Fig.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] Fig.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] Fig.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] Fig.26 It is a circumferential 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.

[0056] Legend:

[0057] 1. Boss structure 2. Radial 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 first stop and 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 radial plate 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 radial plate direction, and a weight-reducing groove R1 is provided in the low stress area of ​​the center of the turbine disk; the reduced thickness structure has a first inclined section L1 and a second inclined section L2 respectively located on the first surface and the second surface, and 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 radial outer end area of ​​the first stop 4 and is constructed as a second straight structure L5 in the radial 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 arc surface structure R2, a third straight structure L9 and a second arc 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 arc surface structure R3, a fourth straight structure L8 and a fourth arc surface structure R4 in the main load-bearing area from the first stop 4 to the second stop 5, the center 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;

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

[0064] In a specific embodiment, reference Fig.15 and Fig.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 was that the addition of the inclined surface at the spoke plate of the turbine disk 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 showed that the rupture margin of the turbine disk was increased by more than 2.1%. The reasonable setting of the inclined surface greatly improved the life of the turbine disk while increasing the rupture margin, which better solved the problem of high strength life of the single-stage turbine rotor at high tangential speed; further, the weight reduction design was carried out in the low stress area of ​​the turbine disk wheel center to further reduce the weight of the turbine disk. At the same time, the load-bearing capacity of the disk center is improved. The design is further optimized based on the thickened structural design, and the weight can be reduced by 37g in the end, 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 increased 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 increased by more than 3.1%. Therefore, by designing the thickened 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 bearing area of ​​the conventional guide plate adopts a rounded corner design, and the rest of the structure is a straight structure design; this embodiment optimizes the design of the main bearing area, adjusts the second straight structure L5 of the upper straight segment of the main bearing area to be inclined at a small angle, and optimizes the first surface to be sequentially constructed in the main bearing area as a first arc surface structure R2, a third straight structure L9 and a second arc surface structure, and the second surface is sequentially constructed in the main bearing area from the first stop 4 to the second stop 5 as a third arc surface structure R3, a fourth straight structure L8 and a fourth arc surface structure R4, the center 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 bearing area is an inclined teardrop-shaped configuration as a whole, so as to improve the bearing capacity. Fig.21 and Fig. 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. Figure 23 to Figure 26, the guide plate structure of this invention is 19MPa and 45Mpa high respectively. By increasing the width of the main bearing area of ​​the rotating plate, 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 is increased by 227g, and the rupture margin is increased by 1.5%. Similarly, in order to reduce the weight increase caused by the water drop-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, in this embodiment, the fifth straight structure L6 with different radial positions is designed. and the sixth straight structure L7, the guide disc bearing capacity is maximized, the weight is reduced by 53g, the maximum equivalent stress and circumferential stress of the disc center are increased by 7MPa and 16MPa respectively, and the rupture margin is still increased by 1% compared with the conventional structure; in summary, the guide disc structure of this embodiment increases in weight by 174g, but the maximum equivalent stress and radial stress are reduced by 12MPa and 19MPa respectively, that is, the maximum stress is reduced by more than 1.2%, and the rupture margin is increased by 1%, which improves the rupture resistance and life of the guide disc, and makes an important contribution to the improvement of the overall strength and life of the turbine rotor at high temperature and high tangential speed;

[0066] The single-stage gas turbine rotor obtained by the structural design of this embodiment has a significantly improved mid-diameter tangential speed of the gas turbine rotor (above 530 m / s) and has a strength and life level comparable to that of a single-stage or two-stage gas turbine rotor with a low 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 thickened 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 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 to enhance the load-bearing capacity of the inner diameter of the guide disk, the weight increase is controlled based on the straight structure design between the arc structures, and the weight reduction 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.

[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 at 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 based on the 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 disc is provided with vent holes 6 evenly distributed along the circumferential direction, 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 along the circumferential direction at the position of the first stop 4, 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 surface of the turbine disc, so as to realize cooling air induction;

[0071] Further, 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 the conventional design, the gap between the tenon of the rotor cooling blade and the tenon groove of the gas turbine disk 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 disk, a guide plate needs to be added to the rear end of the gas turbine disk in the prior art; the tenon profile of the rotor cooling blade of this embodiment and the tenon groove profile of the turbine disk are designed with a small gap (i.e., 0.03-0.05 mm), 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 cooling blades of the single-stage gas turbine rotor 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 of the 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 of the 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 disk, 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, and 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 to limit the axial position of 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, the radial plate of the turbine disk is designed with a reduced thickness structure. By reducing the volume of the radial plate area in the upper part of the gas turbine disk, the volume of the low radius and low stress level area of ​​the wheel center is reduced. Under the premise of not increasing the weight of the turbine disk, the radial stress level of the radial plate area is appropriately increased, and the maximum stress of the turbine disk wheel center is significantly reduced;

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

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

[0080] S4. Weight reduction design of low stress area of ​​guide plate based on finite element analysis.

[0081] The final structural design is 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 in sequence from the mortise and tenon position at the front end of the turbine disk, and axially move the rotor cooling blade to the position of boss structure 1;

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

[0085] A3. Transfer the guide plate from the front end journal of the turbine disk to the guide plate, and move the guide plate axially until the guide plate fits the front end surface of the rotor cooling blade;

[0086] A4.Insert the locking plates and screw the nuts in order 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 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. 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 plate 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 plate, 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.

2. The gas turbine rotor structure according to claim 1, characterized in that: The center of gravity of the comb teeth at the radial 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, characterized in that: 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 disk is provided with ventilation holes evenly distributed along the circumference for introducing external cooling air into the gap between the guide disk and the turbine disk; the turbine disk is provided with ventilation grooves evenly distributed circumferentially at the first stop position for guiding the cooling air in the gap between the guide disk and the turbine disk to the inner cavity of the rotor cooling blades and out to the rear end face of the turbine disk.

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 0.05-0.06mm. The first stop of the guide plate is fitted with a small clearance, and the radial clearance is 0.05-0.07mm.

6. The gas turbine rotor structure according to claim 1, characterized in that: The straightness of the second straight structure is greater than 0 and less than 3.5°.

7. The gas turbine rotor structure according to any one of claims 1 to 6, 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.

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

9. An assembly method, characterized in that: For assembling the gas turbine rotor structure according to claim 7, the gas turbine rotor structure further comprising a sealing wire, a nut and a locking plate, the assembling method comprising: A1. Place each rotor cooling blade in sequence from the mortise and tenon position at the front end of the turbine disc, and axially move the rotor cooling blade to the boss structure position; A2. Heat the guide plate to above 200℃ and put the sealing wire into the installation groove of the guide plate; A3. Transfer the guide plate from the front end journal of the turbine disk to the guide plate, and move the guide plate axially until the guide plate fits the front end surface of the rotor cooling blade; A4.Insert the locking plates and screw the nuts in order until the guide plate reaches the preset compression amount to complete the assembly.

10. An aircraft engine, characterized in that: A gas turbine rotor structure according to any one of claims 1 to 7 is used.

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