A turbine rotor blade of a ceramic matrix composite skeleton structure for an aeroengine

Abstract

The application discloses an aero-engine ceramic matrix composite framework structure turbine rotor blade and belongs to the technical field of aero-engines. The blade comprises a force transmission framework of a ceramic matrix composite material, a flow passage profile, a tip sealing block and a connecting pin. The force transmission framework has dovetail single-tooth tenon heads, an adapter section and at least three groups of branch force transmission beams. The flow passage profile is attached to the end face of the force transmission framework through a lug and is connected to the branch force transmission beams through the pin. Cooling air flows into the cavity between the flow passage profile and the force transmission framework from the bottom of the tenon head through the internal passage of the force transmission framework and is finally discharged from the air film holes on the flow passage profile. The blade fully utilizes the characteristics of high-temperature resistance, light weight and high strength of the ceramic matrix composite material, solves the force transmission and assembly problems caused by the anisotropy of the composite material through the optimized force transmission and connecting structure of the framework, improves the working temperature of the blade, reduces the cooling air consumption and the weight of the blade, and thus improves the thrust-to-weight ratio and working efficiency of the aero-engine.

Description

Technical Field

[0001] This invention belongs to the field of aero-engine technology, and in particular relates to a turbine rotor blade with a ceramic matrix composite skeleton structure for aero-engines. Background Technology

[0002] Turbine rotor blades are one of the core components of the hot-end parts of aero-engines. Their function is to convert the energy of high-temperature, high-pressure combustion gases into mechanical work to drive the compressor and engine accessories. As aero-engines pursue higher thrust-to-weight ratios and thermal efficiency, turbine inlet temperatures are constantly increasing, placing extremely stringent requirements on the temperature resistance, load-bearing capacity, and lightweighting of turbine rotor blades.

[0003] Currently, turbine rotor blades for high-performance aero engines are generally made of nickel-based or cobalt-based high-temperature alloys, and are precision-cast into hollow blades with complex internal cooling structures. These blades are typically connected to the turbine disk using multi-toothed "fir tree-shaped" tenons, and cooling air drawn from the compressor is introduced. The blades are further cooled through labyrinthine cooling channels, baffles, and film cooling holes inside the blades to ensure safe operation at combustion gas temperatures exceeding the material's melting point.

[0004] However, this traditional technological approach is gradually approaching its physical limits and faces the following prominent bottlenecks:

[0005] 1. Material temperature limit: Even with the most advanced single-crystal high-temperature alloys and thermal barrier coatings, the upper limit of its long-term operating temperature is difficult to exceed 1150℃, which restricts the further improvement of the turbine inlet temperature.

[0006] 2. High cost of cooling efficiency: In order to protect the blades, a large amount of high-pressure cooling air (about 15%-25% of the total flow) is required. This part of the gas does not fully participate in combustion and work, which directly leads to a decrease in engine cycle efficiency.

[0007] 3. The contradiction between structural complexity and weight: The complex internal cooling structure leads to significant stress concentration, often requiring increased local thickness to meet strength requirements, making the blades bulky. At the same time, the high density of high-temperature alloys (approximately 8.9 g / cm³) is not conducive to reducing engine weight.

[0008] 4. Material development is in a predicament: In order to improve high-temperature performance, more high-density refractory elements have to be added to the alloy, which further increases the material density, running counter to the goal of lightweighting.

[0009] Ceramic matrix composites (CMCs) are considered ideal materials for hot-end components of next-generation high-performance aero-engines due to their low density (approximately one-third that of high-temperature alloys), high specific strength, and excellent high-temperature oxidation and creep resistance. Using CMCs to manufacture turbine blades is expected to raise operating temperatures to over 1400°C and significantly reduce or even eliminate film cooling, thereby revolutionizing engine efficiency and thrust-to-weight ratio.

[0010] However, the inherent anisotropy (mechanical properties strongly depend on fiber orientation) and brittleness of CMC make it impractical to directly apply traditional integral metal blade structures and "fir tree" multi-tooth connection methods. The continuum mechanics approach based on isotropic materials and connection methods under complex stress states in metal blade design cannot adapt to the design principles of CMC, which emphasize fiber-guided force transmission and avoidance of interlaminar shear and stress concentration. Therefore, there is an urgent need for a novel turbine rotor blade structure specifically tailored to the characteristics of CMC to address key issues such as force transmission path design, component connections, cooling organization, and compatibility with the metal rotor disk. Summary of the Invention

[0011] Based on the problems existing in the above-mentioned background technology, the present invention aims to provide a turbine rotor blade with a ceramic matrix composite skeleton structure for aero-engines, which solves the limitations of existing high-temperature alloy turbine rotor blades in terms of temperature resistance, weight and cooling gas consumption.

[0012] The embodiments of the present invention are implemented as follows:

[0013] This invention provides a turbine rotor blade with a ceramic matrix composite skeleton structure for an aero-engine, which includes a force transmission skeleton made of ceramic matrix composite, a flow channel shape, a blade tip sealing block, and multiple pins.

[0014] The force transmission frame includes an integrally formed dovetail-shaped single-tooth tenon and a transition section; multiple branched force transmission beams are arranged sequentially from top to bottom on both side walls of the transition section, each branched force transmission beam extends outward from the side wall of the transition section and each branched force transmission beam is provided with a pin hole; a cold air channel is provided inside the force transmission frame, the bottom end of the cold air channel is connected to the bottom end of the dovetail-shaped single-tooth tenon, and the top end of the cold air channel extends to the transition section and is connected to one side wall of the transition section;

[0015] The flow channel includes a blade, a rim plate, and two lugs; the blade is located on the upper end face of the rim plate, and the two lugs are located on the lower end face of the rim plate; the leading edge of the blade has multiple air film holes communicating with its interior, and the blade has multiple pin holes for connection.

[0016] The leaf tip sealing block is assembled at the tip of the leaf body and matches the inner shape of the leaf body to form a closed cavity;

[0017] The transition section of the force transmission frame and the branched force transmission beams are located inside the blade. The dovetail-shaped single-tooth tenon is located below the edge plate. The inner sidewalls of the two lugs are fixedly connected to the two sidewalls of the dovetail-shaped single-tooth tenon respectively. Multiple branched force transmission beams are fixedly connected to the inner wall of the blade by passing pins through pin holes on the blade and pin holes on the branched force transmission beams.

[0018] After the prefabricated components of the force transmission frame, flow channel shape, blade tip sealing block, and multiple pins are assembled to form turbine rotor blades, they are then cured by an integrated matrix to form a continuous matrix.

[0019] This invention discloses a turbine rotor blade with a ceramic matrix composite skeleton structure for an aero-engine. The load-bearing skeleton is the main load-bearing structure of the blade, functioning similarly to the skeleton of the human body, responsible for transmitting the centrifugal force, aerodynamic loads, and other loads borne by the blade to the turbine disk. The flow channel shape constitutes the aerodynamic surface of the blade, its function being to form an efficient gas flow channel, converting gas energy into mechanical work. It is similar to a "skin" covering the skeleton. The blade tip sealing block is assembled inside the blade tip, its function being to seal the top of the internal cavity formed by the flow channel shape and the load-bearing skeleton, preventing cooling gas leakage from the blade tip and ensuring cooling effect. The mechanical connections between the various components mainly rely on pins.

[0020] The basic principle of this aero-engine ceramic matrix composite (CMC) skeleton structure turbine rotor blade is as follows: Cooling airflow enters from the air supply channel on the turbine disk into the air inlet at the bottom of the dovetail-shaped single-tooth tenon, flows through the pre-reserved cold air channel inside the force transmission skeleton, and is delivered to the hollow cavity between the blade body and the force transmission skeleton. The cooling air flows within this cavity, dissipating heat from the inner wall of the blade body (especially the leading edge and pressure surface bearing the highest heat load) through convection and heat exchange, and finally exits through film cooling holes located at the leading edge and other positions on the blade body, forming a low-temperature film on the outer surface of the blade to insulate against the high-temperature combustion gases. Because CMC itself is heat-resistant, the required cooling air volume is significantly reduced compared to metal blades, and the cooling structure is also simplified.

[0021] Furthermore, the matrix material for the force transmission frame, flow channel shape, blade tip seal, and pins is SiC, and the reinforcing fibers are all SiC fibers. All components in the turbine rotor blade of the aero-engine ceramic matrix composite frame structure are made of SiC fiber-reinforced SiC matrix composite material. This allows the turbine rotor blade to maintain excellent high-temperature strength, oxidation resistance, and creep resistance even at temperatures above 1400℃, while its density is only about one-third that of high-temperature alloys.

[0022] Furthermore, in the fiber preform of the force transmission skeleton, the fibers are oriented perpendicular to the engine axis. This design allows the force transmission skeleton to most effectively withstand the enormous centrifugal force generated by the high-speed rotation of the blades, aligning the fiber direction with the direction of the principal tensile stress and fully leveraging the advantages of CMC fiber reinforcement.

[0023] Furthermore, each branched force transmission beam is inclined upwards, and the angle between each branched force transmission beam and the main trunk of the transition section is 20°~40°. This angle range can ensure a smooth force transmission path, effectively transfer the load to the main trunk, and also conform to the composite material layup design principle, which is conducive to the continuous laying of fibers and avoids fiber bending and performance loss caused by sharp turns.

[0024] Furthermore, the thickness of the branched force transmission beam is 1.5mm to 3mm. This range is the result of careful consideration: if the thickness is less than 1.5mm, the beam's stiffness is insufficient, making it difficult to effectively fix the flow channel shape and transmit loads; if the thickness is greater than 3mm, the force transmission skeleton will be too thick overall, increasing weight, and in subsequent matrix densification processes such as chemical vapor infiltration, the excessively thick cross-section is prone to defects due to uneven gas permeation.

[0025] Furthermore, the multiple branched force transmission beams are divided into three groups of branched force transmission beams from top to bottom; the first group of branched force transmission beams includes a branched force transmission beam located at the tip of the leaf blade and its end is attached to the inner wall of the back side of the leaf blade.

[0026] The remaining groups of branched force transmission beams each include two branched force transmission beams located on the two side walls of the transition section, with the ends of the two branched force transmission beams respectively attached to the inner wall of the blade basin and the inner wall of the blade back. The above design enables multiple branched force transmission beams to support the flow channel shape at multiple points, ensuring the stability of the flow channel shape and providing it with multi-path load transfer.

[0027] Furthermore, each branched force transmission beam is connected to the blade by at least two pins.

[0028] Furthermore, the thickness of the force transmission frame gradually decreases from the root to the tip of the dovetail-shaped single-tooth tenon, achieving a uniform strength distribution of the material and minimizing the weight of the frame while ensuring safety.

[0029] Furthermore, the leaf wall thickness gradually increases from the leaf tip to the leaf root, with a minimum wall thickness of not less than 1.5 mm and a maximum wall thickness of not more than 3 mm. The root bears the greatest centrifugal force, so a thicker wall is required; the tip experiences less force, so the wall can be thinned to reduce weight. The wall thickness is controlled within a minimum of 1.5 mm and a maximum of 3 mm to ensure good process feasibility while withstanding aerodynamic and thermal loads.

[0030] Furthermore, the diameter of the pin is 0.1 mm larger than the diameter of the pin hole and pin bore it mates with. This creates a slight interference fit. A certain preload is generated during assembly, which helps eliminate microscopic gaps between components, improves connection stiffness and force transmission efficiency, and maintains a tight connection during thermal cycling.

[0031] Compared to existing pressure regulating valves, the advantages of this invention are:

[0032] 1. This invention relates to a turbine rotor blade with a ceramic matrix composite (CMC) skeleton structure for aero-engines, proposing a split-reintegration design concept of "CMC skeleton for main load-bearing + CMC skin for shaping". By centrally handling high loads through the load-transmitting skeleton, ensuring aerodynamics through the flow channel shape, and achieving overall integration through optimized pin connections and integrated curing, a novel structure has been successfully constructed that meets the extreme operating conditions of turbine blades while fully adapting to the anisotropy and brittleness of CMC materials. This opens up a practical and feasible technical path for the engineering application of CMC in turbine rotor blades.

[0033] 2. The turbine rotor blade of this invention, with its ceramic matrix composite skeleton structure, utilizes SiC fiber-reinforced SiC matrix composite material for all components. This allows the turbine rotor blade to maintain excellent high-temperature strength, oxidation resistance, and creep resistance even at temperatures above 1400℃, while maintaining a density only about one-third that of high-temperature alloys. The structural design of this invention fully leverages the advantages of the materials, providing a fundamental guarantee for achieving a leap in blade performance.

[0034] 3. The turbine rotor blade of the aero-engine ceramic matrix composite skeleton structure of the present invention has circumferentially oriented force-transmitting skeleton fibers that directly resist centrifugal force, resulting in extremely high material utilization. The dendritic force-transmitting beam efficiently transfers the load of the flow channel shape to the skeleton backbone at the optimal angle (20°~40°) and thickness (1.5mm~3mm), avoiding stress concentration and interlaminar shear. The variable thickness design of the force-transmitting skeleton and flow channel shape realizes the equal strength design of "material matching as needed" and eliminates redundant mass. Compared with traditional high-temperature alloy blades, the blade of the present invention can achieve a weight reduction of 50%~75%, which makes a significant contribution to improving the thrust-to-weight ratio of the engine.

[0035] 4. The turbine rotor blade with a ceramic matrix composite skeleton structure for aero-engines disclosed in this invention, due to the inherent high-temperature resistance of the ceramic matrix composite material, combined with effective internal convection and film cooling, allows the blade to operate reliably for extended periods at combustion gas temperatures of 1200℃~1400℃, an improvement of approximately 100℃~300℃ compared to advanced high-temperature alloy blades. This improved temperature resistance directly translates to a reduction in the required cooling air volume. The cooling air requirement of the blades in this invention can be reduced by 60%~80% compared to traditional blades, allowing more air to be used for combustion and significantly improving the engine's thermal efficiency and power output.

[0036] 5. The turbine rotor blade of the aero-engine ceramic matrix composite skeleton structure in this invention adopts a dovetail-shaped single-tooth tenon and a wheel disk. In the actual design of the dovetail-shaped single-tooth tenon, the magnitude and direction of the pressure on the contact surface between the blade and the wheel disk can be changed by adjusting the tenon contact width and the tenon contact angle. This can better adjust and manage the assembly stress and thermal stress caused by the difference in thermal expansion coefficient between the CMC blade and the metal turbine disk, and improve the matching between the blade and the metal wheel disk. Attached Figure Description

[0037] To more clearly illustrate the technical solutions in the embodiments of the present invention or the prior art, the drawings used in the embodiments will be briefly described below. Obviously, the drawings described below are only some embodiments of the present invention. For those skilled in the art, other drawings can be obtained based on these drawings without creative effort. The above and other objects, features, and advantages of the present invention will become clearer through the accompanying drawings. The same reference numerals indicate the same parts in all the drawings. The drawings are not intentionally drawn to scale to actual dimensions; the focus is on illustrating the main points of the invention.

[0038] Figure 1 This is a three-dimensional structural diagram of a turbine rotor blade with a ceramic matrix composite skeleton structure for an aero-engine.

[0039] Figure 2 This is a partial cross-sectional view of a turbine rotor blade with a ceramic matrix composite skeleton structure for an aero-engine.

[0040] Figure 3 This is a three-dimensional structural diagram of the force transmission skeleton.

[0041] Figure 4 This is a schematic diagram of the internal cross-sectional structure of the force transmission frame.

[0042] Figure 5 This is a schematic diagram showing the connection between the force transmission frame and the external flow channel.

[0043] Figure 6 This is a schematic diagram of the fiber orientation in the force transmission skeleton.

[0044] Figure 7 This is a schematic diagram of the cold air flow on the turbine rotor blades.

[0045] Figure 8 This is a three-dimensional structural diagram of the flow channel.

[0046] Figure 9 A schematic diagram of the blade tip sealing block installation.

[0047] Among them, 1. Turbine rotor blades; 2. Force transmission frame; 3. Flow channel shape; 4. Blade tip sealing block; 5. Pin; 6. Dovetail-shaped single tooth tenon; 7. Transition section; 8. Branched force transmission beam; 9. Pin hole; 10. Cooling air passage; 11. Air film vent; 12. Blade body; 13. Edge plate; 14. Lug. Detailed Implementation

[0048] To make the objectives, technical solutions, and advantages of the embodiments of the present invention clearer, the technical solutions of the embodiments of the present invention will be clearly and completely described below with reference to the accompanying drawings. Obviously, the described embodiments are only some embodiments of the present invention, and not all embodiments. The components of the embodiments of the present invention described and shown in the accompanying drawings can generally be arranged and designed in various different configurations.

[0049] Therefore, the following detailed description of the embodiments of the invention provided in the accompanying drawings is not intended to limit the scope of the claimed invention, but merely to illustrate selected embodiments of the invention. All other embodiments obtained by those skilled in the art based on the embodiments of the invention without inventive effort are within the scope of protection of the invention.

[0050] It should be noted that similar labels and letters in the following figures indicate similar items. Therefore, once an item is defined in one figure, it does not need to be further defined and explained in subsequent figures.

[0051] The structures, proportions, sizes, etc. illustrated in this specification are only for the purpose of assisting those skilled in the art in understanding and reading the content disclosed herein, and are not intended to limit the conditions under which the present invention can be implemented. Therefore, they have no substantial technical significance. Any modifications to the structure, changes in the proportions, or adjustments to the size, without affecting the effects and objectives that the present invention can produce, should still fall within the scope of the technical content disclosed in the present invention.

[0052] like Figures 1-2 As shown, this invention provides a turbine rotor blade with a ceramic matrix composite skeleton structure for aero-engines, comprising a force-transmitting skeleton 2 made of ceramic matrix composite material, a flow channel shape 3, a blade tip sealing block 4, and multiple pins 5. All components of the turbine rotor blade 1 are made of SiC fiber reinforced SiC matrix composite material, which enables the turbine rotor blade 1 to maintain excellent high-temperature strength, oxidation resistance, and creep resistance even at temperatures above 1400℃, and its density is only about 1 / 3 that of high-temperature alloys.

[0053] Specifically, such as Figures 2-6As shown, in one specific embodiment of the force transmission skeleton 2, continuous SiC fiber bundles are used. Through automatic layup technology, the fibers are laid up in a direction perpendicular to the engine axis (i.e., the circumferential direction of the wheel disk) to form an integrated fiber preform containing a dovetail-shaped single-tooth tenon 6, a transition section 7, and five branched force transmission beams 8. This design allows the force transmission skeleton 2 to most effectively withstand the huge centrifugal force generated by the high-speed rotation of the blades, ensuring that the fiber direction is consistent with the direction of the principal tensile stress, and fully leveraging the advantages of CMC fiber reinforcement. Each branched force transmission beam 8 is inclined upwards, and the angle between each branched force transmission beam 8 and the main trunk of the transition section 7 is 20°~40°. This angle range ensures a smooth force transmission path, effectively transferring the load to the main trunk, and conforms to the composite material layup design principles, facilitating continuous fiber laying and avoiding fiber bending and performance loss caused by sharp turns.

[0054] By optimizing contact geometry parameters, such as the contact width and contact angle of the dovetail-shaped single-tooth tenon 6, the assembly stress and thermal stress caused by the difference in thermal expansion coefficients between the CMC blade and the metal turbine disk can be better adjusted and managed, thus improving the matching between the blade and the metal disk.

[0055] The thickness of the branched force transmission beam 8 is 1.5mm to 3mm. This range is the result of careful consideration: if the thickness is less than 1.5mm, the beam will not be stiff enough to effectively fix the flow channel shape 3 and transmit the load; if the thickness is greater than 3mm, the force transmission frame 2 will be too thick, increasing the weight, and in subsequent matrix densification processes such as chemical vapor infiltration, the excessively thick cross section is prone to defects due to uneven gas infiltration.

[0056] As one specific arrangement of the five branched force transmission beams 8, the multiple branched force transmission beams 8 are divided into three groups from top to bottom; the first group of branched force transmission beams 8 includes a branched force transmission beam 8 located at the tip of the blade 12, with its end attached to the inner wall of the blade's back side; the remaining groups of branched force transmission beams 8 each include two branched force transmission beams 8 located on the two side walls of the transition section 7, with the ends of the two branched force transmission beams attached to the inner wall of the blade's base side and the inner wall of the blade's back side, respectively. The above design enables multiple branched force transmission beams 8 to support the flow channel shape 3 at multiple points, ensuring the stability of the flow channel shape 3 and providing it with multi-path load transfer.

[0057] The thickness of the entire force transmission frame 2 gradually decreases from the root of the dovetail-shaped single-tooth tenon 6 to the tip, achieving equal strength distribution of the material and minimizing the weight of the frame while ensuring safety.

[0058] like Figure 2 , Figure 7 and Figure 8As shown, as a specific configuration of the flow channel shape 3, the flow channel shape 3 includes a blade 12, a rim plate 13, and two lugs 14; the blade 12 is located on the upper end face of the rim plate 13, and the two lugs 14 are located on the lower end face of the rim plate 13; the leading edge of the blade 12 is provided with multiple film-forming holes 11 communicating with its interior, and the blade 12 is provided with multiple connecting pin holes; as a specific manufacturing method of the flow channel shape 3, the flow channel shape 3 is manufactured by using a three-dimensional braiding machine to weave SiC fibers into an integral braided body with a specific airfoil cross-section and including the rim plate 13 and lugs 14 structures. The wall thickness of the blade 12 gradually changes from the blade root to the blade tip according to the three-dimensional digital model design.

[0059] Preferably, but not limited to, the leaf blade wall thickness gradually increases from the leaf tip to the leaf root, with a minimum wall thickness of not less than 1.5 mm and a maximum wall thickness of not more than 3 mm. The root bears the greatest centrifugal force, so a thicker wall is required; the tip experiences less force, so the wall can be thinned to reduce weight. The wall thickness range is controlled to a minimum of not less than 1.5 mm and a maximum of not more than 3 mm to ensure good process feasibility while withstanding aerodynamic and thermal loads.

[0060] like Figure 9 As shown, the blade tip sealing block 4 is assembled at the tip of the blade body 12 and matches the inner shape of the blade body 12 to form a closed cavity; specifically, it is made by laying SiC fiber cloth and then sewing it, or by simple weaving and forming it, with the shape matching the inner cavity of the blade tip.

[0061] The specific manufacturing method of pin 5 is as follows: after impregnating SiC fiber bundles with adhesive, they are pultruded into thin rods of a specified diameter and then cut into the required length.

[0062] The transition section 7 of the force transmission frame 2 and the branched force transmission beam 8 are located inside the blade 12. The dovetail-shaped single-tooth tenon 6 is located below the edge plate 13. The inner sidewalls of the two lugs 14 are fixedly connected to the two sidewalls of the dovetail-shaped single-tooth tenon 6 respectively. Multiple branched force transmission beams 8 are fixedly connected to the inner wall of the blade 12 by means of pins 5 passing through pin holes on the blade 12 and pin holes 9 on the branched force transmission beam 8.

[0063] Preferably, but not limited to, the diameter of pin 5 is 0.1 mm larger than the diameter of the pin hole and pin hole 9 to which it is assembled. This creates a slight interference fit. A certain preload is generated during assembly, which helps eliminate microscopic gaps between components, improves connection stiffness and force transmission efficiency, and maintains tightness of the connection during thermal cycling.

[0064] After the prefabricated components of the force transmission frame 2, flow channel shape 3, blade tip sealing block 4, and multiple pins 5 are assembled to form turbine rotor blade 1, they are cured by an integrated matrix to form a continuous matrix. Finally, the continuous matrix is ​​subjected to post-processing steps such as dimensional precision machining and surface grinding and polishing to obtain the finished turbine rotor blade 1 of the aero-engine ceramic matrix composite skeleton structure of the present invention.

[0065] This invention discloses a turbine rotor blade with a ceramic matrix composite skeleton structure for an aero-engine. The force transmission skeleton 2 is the main load-bearing structure of the blade, functioning similarly to the skeleton of the human body, responsible for transmitting the centrifugal force, aerodynamic loads, and other loads borne by the blade to the turbine disk. The flow channel shape 3 constitutes the aerodynamic surface of the blade, its function being to form an efficient gas flow channel, converting gas energy into mechanical work. It is similar to the "skin" covering the skeleton. The blade tip sealing block 4 is assembled inside the tip of the blade body 12, its function being to seal the tip of the internal cavity formed by the flow channel shape 3 and the force transmission skeleton 2, preventing cooling gas from leaking from the blade tip and ensuring cooling effect. The mechanical connection between the components mainly relies on pins 5.

[0066] The turbine rotor blade 1 is mounted in the corresponding slot of the metal turbine disk via its dovetail-shaped single-tooth tenon 6. When the engine is running, the turbine disk rotates at high speed, and the turbine rotor blade 1 is subjected to huge centrifugal force. This force is transmitted to the branched force transmission beam 8 through the flow channel shape 3, then converges to the main trunk of the force transmission frame 2, and finally is safely transmitted to the turbine disk through the contact surface of the dovetail-shaped single-tooth tenon 6. The high-temperature combustion gas flows through the flow channel of the turbine rotor blade 1, driving the turbine rotor blade 1 to do work. At the same time, the cooling air drawn from the compressor enters the air inlet at the bottom of the dovetail-shaped single-tooth tenon 6 from the disk, flows through the cold air channel 10 inside the force transmission frame 2, enters the hollow cavity inside the blade body 12, cools the hot inner wall of the flow channel shape 3, and then exits from the film gas hole 11 at the leading edge to form a protective film gas. Due to the inherent high-temperature resistance of ceramic matrix composites, combined with effective internal convection and film cooling, the blades can operate reliably for extended periods at combustion gas temperatures of 1200℃~1400℃, an improvement of approximately 100℃~300℃ compared to advanced high-temperature alloy blades. This increased temperature resistance directly translates to a reduction in the required cooling air volume. The cooling air requirement of the blades in this invention can be reduced by 60%~80% compared to traditional blades, allowing more air to be used for combustion and significantly improving engine thermal efficiency and power output.

[0067] The above are merely preferred embodiments of the present invention and are not intended to limit the present invention. For those skilled in the art, the present invention can have various modifications and variations. Any modifications, equivalent substitutions, improvements, etc., made within the spirit and principles of the present invention should be included within the scope of protection of the present invention.

Claims

1. A turbine rotor blade with a ceramic matrix composite skeleton structure for an aero-engine, characterized in that, It includes a force transmission frame (2) made of ceramic matrix composite material, a flow channel shape (3), a blade tip sealing block (4), and multiple pins (5); The force transmission frame (2) includes an integrally formed dovetail-shaped single-tooth tenon (6) and a transition section (7); multiple branched force transmission beams (8) are arranged sequentially from top to bottom on both sides of the transition section (7), each branched force transmission beam (8) extends outward from the side wall of the transition section (7) and each branched force transmission beam (8) is provided with a pin hole (9); a cold air channel (10) is provided inside the force transmission frame (2), the bottom end of the cold air channel (10) is connected to the bottom end of the dovetail-shaped single-tooth tenon (6), and the top end of the cold air channel (10) extends to the transition section (7) and is connected to one side wall of the transition section (7); The flow channel shape (3) includes a blade (12), a rim plate (13), and two lugs (14); the blade (12) is disposed on the upper end face of the rim plate (13), and the two lugs (14) are disposed on the lower end face of the rim plate (13); the leading edge of the blade (12) is provided with a plurality of air film holes (11) communicating with its interior, and the blade (12) is provided with a plurality of connecting pin holes; The leaf tip sealing block (4) is assembled at the tip of the leaf body (12) and matches the inner shape of the leaf body (12) to form a closed cavity; The transition section (7) of the force transmission frame (2) and the branched force transmission beam (8) are located inside the blade (12). The dovetail-shaped single-tooth tenon (6) is located below the edge plate (13). The inner sidewalls of the two lugs (14) are fixedly connected to the two sidewalls of the dovetail-shaped single-tooth tenon (6). The multiple branched force transmission beams (8) are fixedly connected to the inner wall of the blade (12) by passing the pin through the pin hole on the blade (12) and the pin hole (9) on the branched force transmission beam. After the prefabricated components of the force transmission frame (2), flow channel shape (3), blade tip sealing block (4) and multiple pins (5) are assembled to form turbine rotor blades (1), they are solidified by an integrated matrix to form a continuous matrix.

2. The turbine rotor blade with a ceramic matrix composite skeleton structure for aero-engines according to claim 1, characterized in that, The matrix material of the force transmission skeleton (2), the flow channel shape (3), the blade tip sealing block (4) and the pin (5) are all SiC, and the reinforcing fibers are all SiC fibers.

3. The turbine rotor blade with a ceramic matrix composite skeleton structure for aero-engines according to claim 1 or 2, characterized in that, In the fiber preform of the force transmission skeleton (2), the fibers are oriented in a direction perpendicular to the engine axis.

4. The turbine rotor blade with a ceramic matrix composite skeleton structure for aero-engines according to claim 1, characterized in that, Each of the branched force transmission beams (8) is inclined upwards, and the angle between each branched force transmission beam (8) and the main trunk of the transition section (7) is 20°~40°.

5. The turbine rotor blade with a ceramic matrix composite skeleton structure for aero-engines according to claim 1, characterized in that, The thickness of the branched force transmission beam (8) is 1.5mm to 3mm.

6. The turbine rotor blade with a ceramic matrix composite skeleton structure for aero-engines according to claim 1, characterized in that, The multiple branched force transmission beams (8) are divided into three groups of branched force transmission beams (8) from top to bottom; the first group of branched force transmission beams (8) includes a branched force transmission beam (8) located at the tip of the leaf body (12) and its end is attached to the inner wall of the back side of the leaf body (12); The remaining groups of branched force transmission beams (8) each include two branched force transmission beams (8) located on the two side walls of the transition section (7) respectively, and the ends of the two branched force transmission beams (8) are respectively attached to the inner wall of the leaf basin side (20) and the inner wall of the leaf back side of the leaf body (12).

7. The turbine rotor blade with a ceramic matrix composite skeleton structure for aero-engines according to claim 1, characterized in that, Each of the said branched force transmission beams (8) is connected to the blade (12) by at least two of the said pins (5).

8. The turbine rotor blade with a ceramic matrix composite skeleton structure for aero-engines according to claim 1, characterized in that, The thickness of the force transmission frame (2) gradually decreases from the root to the tip of the dovetail-shaped single-tooth tenon (6).

9. The turbine rotor blade with a ceramic matrix composite skeleton structure for aero-engines according to claim 1, characterized in that, The thickness of the leaf blade (12) gradually increases from the leaf tip to the leaf root, with a minimum wall thickness of not less than 1.5 mm and a maximum wall thickness of not more than 3 mm.

10. The turbine rotor blade with a ceramic matrix composite skeleton structure for aero-engines according to claim 1, characterized in that, The diameter of the pin (5) is 0.1 mm larger than the diameter of the pin hole to which it is assembled and the diameter of the pin hole (9).

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