Blade system of an aircraft engine compressor and aircraft engine

Abstract

The present disclosure relates to the technical field of aero-engines, and particularly provides a blade system of an aero-engine compressor and an aero-engine. The blade system of the aero-engine compressor comprises a controller and a plurality of stationary vanes. The stationary vane comprises an inner casing, an outer casing and a plurality of blades arranged between the inner casing and the outer casing. The blade comprises a support framework, a flexible blade skin and a liquid metal capable of changing from a solid phase to a liquid phase with temperature change. The flexible blade skin is wrapped outside the support framework to form a closed space, and the liquid metal is filled in the closed space. The support framework is capable of deformation. A heating element is arranged in the closed space. The controller is used to control at least the working state of the heating element, so that the liquid metal changes phase, and the flexible blade skin deforms under the action of the liquid metal and the support framework, thereby realizing dynamic adjustment of the blade profile and effectively adjusting the airflow direction, and improving the efficiency and stability of the engine.

Description

Technical Field

[0001] This disclosure relates to the field of aero-engine technology, and in particular to a blade system for an aero-engine compressor and an aero-engine. Background Technology

[0002] The compressor is one of the key components of an aero-engine. It consists of moving blades and stationary blades. The compressor compresses air, which then mixes and burns with fuel in the combustion chamber, producing high-temperature, high-pressure gas. This gas drives the turbine, which then ejects the gas through the exhaust nozzle, providing thrust for the aircraft. The performance of the compressor directly affects the overall efficiency and thrust of the engine.

[0003] However, the compressor blades of the related technologies cannot meet the airflow requirements under different operating conditions, resulting in low compressor efficiency, which in turn affects the overall efficiency and stability of the engine. Summary of the Invention

[0004] In order to solve the above-mentioned technical problems, or at least partially solve the above-mentioned technical problems, this disclosure provides a blade system for an aero-engine compressor and an aero-engine, which can effectively adjust the airflow direction and improve the overall efficiency and stability of the engine.

[0005] In a first aspect, this disclosure provides a blade system for an aero-engine compressor, including a controller and a multi-stage stator blade, wherein the stator blade includes an inner casing, an outer casing surrounding the inner casing, and a plurality of blades disposed between the inner casing and the outer casing, the plurality of blades being arranged at circumferential intervals along the inner casing.

[0006] The blade includes a support frame, a flexible blade skin, and a liquid metal that can undergo a solid-liquid phase change with temperature. The flexible blade skin covers the outside of the support frame to form a closed space, and the liquid metal fills the closed space.

[0007] The supporting frame is deformable, and a heating element electrically connected to the controller is provided in the enclosed space. The heating element is used to heat the liquid metal, and the controller is at least used to control the working state of the heating element to cause the liquid metal to undergo a phase change, so that the flexible blade skin deforms under the action of the liquid metal and the supporting frame.

[0008] Optionally, the support frame includes an intermediate shaft and multiple support frames;

[0009] The plurality of support frames are arranged circumferentially around the intermediate axis, and each support frame is connected to the intermediate axis and can move around the intermediate axis.

[0010] The flexible blade skin covers the outside of the multiple support frames.

[0011] Optionally, one end of the support frame has a collar, which is sleeved on the intermediate shaft;

[0012] The intermediate shaft is rotatable, and a first meshing tooth is provided on a portion of the outer peripheral wall of the intermediate shaft. A second meshing tooth is provided on a portion of the inner ring wall of the collar. The second meshing tooth can mesh with the first meshing tooth, so that the support frame can move around the intermediate shaft, and the flexible blade skin deforms under the action of the liquid metal and the support frame.

[0013] Optionally, each of the aforementioned support frames includes a support rod and a connecting rod;

[0014] The support rod is arranged radially within the flexible blade skin along the inner casing. The end face of the flexible blade skin corresponding to the outer casing is in contact with the inner wall of the outer casing under the action of the multiple support rods, and the end face of the flexible blade skin corresponding to the inner casing is in contact with the outer wall of the inner casing under the action of the multiple support rods.

[0015] One end of the connecting rod is connected to and fixed relative to the support rod, and the other end of the connecting rod is connected to the intermediate shaft to drive the support rod to move around the intermediate shaft.

[0016] Optionally, the flexible blade skin includes a deformable flexible inner support layer and an elastic protective layer disposed on the outside of the flexible inner support layer.

[0017] Optionally, the flexible inner support layer is a mesh-like shape memory alloy layer;

[0018] The elastic protective layer includes a ceramic fiber braided layer disposed on the outside of the flexible inner support layer and an elastic coating disposed on the outside of the ceramic fiber braided layer.

[0019] Optionally, a temperature sensor is installed inside the enclosed space;

[0020] The temperature sensor is electrically connected to the controller. The temperature sensor is used to detect the temperature of the liquid metal, and the controller is used to control the working state of the heating element according to the temperature detected by the temperature sensor.

[0021] Optionally, the blade system of the engine compressor further includes a pressure control module;

[0022] The pressure control module includes a pressure regulator and a pressure sensor disposed within the enclosed space; the pressure sensor and the pressure regulator are electrically connected to the controller, respectively.

[0023] The pressure sensor is used to detect the pressure of the liquid metal, and the controller is also used to control the pressure regulator to adjust the pressure of the liquid metal according to the pressure value detected by the pressure sensor.

[0024] Optionally, in the multi-stage stator blades, the melting point of the liquid metal in the blades of the last stage stator blade of the compressor is higher than that of the liquid metal in the blades of the inlet stage stator blade of the compressor.

[0025] Secondly, this disclosure provides an aircraft engine including the blade system of the aircraft engine compressor as described above.

[0026] The blade system of the aero-engine compressor and the aero-engine disclosed herein, by configuring the stator blade as including a support frame, a flexible blade skin, and a liquid metal that can undergo a solid-liquid phase change with temperature, the flexible blade skin covers the outside of the support frame to form a closed space, and the liquid metal is filled in the closed space. By setting a controller and a heating element electrically connected to the controller in the closed space, the liquid metal is heated by the heating element, causing the liquid metal to undergo a phase change. That is to say, according to different flight conditions, the heating element is controlled to heat the liquid metal. Heating or stopping the liquid metal enables a solid-liquid phase transition. Because the supporting frame can deform, the flexible blade skin deforms under the synergistic effect of the liquid metal and the supporting frame during the solid-liquid phase transition. This allows for dynamic adjustment of the stator blade profile, effectively regulating the airflow direction, improving the compressor's adaptability to different flow fields, increasing its efficiency and stability, and ultimately enhancing the engine's overall efficiency, thrust, and operational stability. This enables the aero-engine to maintain optimal performance under various operating conditions.

[0027] It should be understood that both the foregoing general description and the following detailed description are exemplary and intended to provide further illustration of the claimed technology. Attached Figure Description

[0028] The above and other objects, features, and advantages of this disclosure will become more apparent from the more detailed description of the embodiments thereof in conjunction with the accompanying drawings. The drawings are provided to further illustrate the embodiments of this disclosure and form part of the specification. They are used together with the embodiments of this disclosure to explain the disclosure and do not constitute a limitation thereof. In the drawings, the same reference numerals generally represent the same components or steps.

[0029] Figure 1 This is a structural schematic diagram of one state of the blades of an aero-engine compressor blade system provided in an embodiment of the present disclosure;

[0030] Figure 2This is a schematic diagram of another state of the blades of the blade system of an aero-engine compressor provided in an embodiment of the present disclosure;

[0031] Figure 3 A schematic diagram of the support frame in the blade system of an aero-engine compressor according to an embodiment of this disclosure;

[0032] Figure 4 A top view of the connection between the intermediate shaft and the support frame in the blade system of an aero-engine compressor according to an embodiment of this disclosure;

[0033] Figure 5 A partial cross-sectional view of the flexible blade skin in the blade system of an aero-engine compressor according to an embodiment of this disclosure;

[0034] Figure 6 This is a schematic diagram of the blades and pressure regulating components in the blade system of an aero-engine compressor according to an embodiment of the present disclosure.

[0035] Among them, 100, stationary blade; 1, inner casing; 2, outer casing; 3, blade; 31, support frame; 311, intermediate shaft; 312, support frame; 313, support rod; 314, connecting rod; 315, collar; 316, first meshing tooth; 317, second meshing tooth; 32, flexible blade skin; 321, flexible inner support layer; 322, elastic protective layer; 323, ceramic fiber braided layer; 324, elastic coating; 325, clearance hole; 33, enclosed space; 4, heating element; 5, temperature sensor; 6, pressure sensor; 7, pressure regulating component. Detailed Implementation

[0036] To make the objectives, technical solutions, and advantages of this disclosure more apparent, exemplary embodiments according to this disclosure will now be described in detail with reference to the accompanying drawings. Obviously, the described embodiments are merely some embodiments of this disclosure, and not all embodiments of this disclosure. It should be understood that this disclosure is not limited to the exemplary embodiments described herein.

[0037] The term "comprising" and its variations as used herein are open-ended, meaning "including but not limited to". The term "based on" means "at least partially based on". The term "one embodiment" means "at least one embodiment"; the term "another embodiment" means "at least one additional embodiment"; the term "some embodiments" means "at least some embodiments". Definitions of other terms will be given in the description below. It should be noted that the concepts of "first", "second", etc., used in this disclosure are only used to distinguish different devices, modules, or units, and are not intended to limit the order of functions performed by these devices, modules, or units or their interdependencies.

[0038] It should be noted that the terms "a" and "a plurality of" used in this disclosure are illustrative rather than restrictive, and those skilled in the art should understand that, unless otherwise expressly indicated in the context, they should be understood as "one or more".

[0039] The compressor is one of the key components of an aero-engine, and its performance directly affects the overall efficiency and thrust of the engine. The compressor consists of moving blades and stationary blades. The moving blades are high-speed rotating components that do work on the airflow through high-speed rotation. The stationary blades mainly function to adjust the direction of the airflow. The compressor compresses the air, providing high-pressure gas to the combustion chamber. That is, the compressed air mixes with fuel and burns in the combustion chamber, producing high-temperature, high-pressure gas. This high-temperature, high-pressure gas drives the turbine to rotate, causing the gas to be ejected through the exhaust nozzle, providing thrust for the aircraft's flight.

[0040] However, the compressor blades of the related technologies cannot meet the airflow requirements under different operating conditions, resulting in low compressor efficiency, which in turn affects the overall efficiency and stability of the engine.

[0041] Based on this, the present disclosure provides a blade system for an aero-engine compressor and an aero-engine. By including a support frame, a flexible blade skin, and liquid metal in the blade, the dynamic adjustment of the blade profile of the stationary blade is achieved through the synergistic effect of the liquid metal phase change, the flexible blade skin, and the deformable support frame. This effectively adjusts the airflow direction, enabling the compressor to meet the airflow requirements under different operating conditions, improving the efficiency and stability of the compressor, and thus improving the overall efficiency, thrust, and stability of the engine.

[0042] The blade system of the aero-engine compressor and the aero-engine provided in this disclosure will be specifically described below with reference to the accompanying drawings and through specific embodiments:

[0043] Reference Figures 1 to 6 As shown, this disclosure provides a blade system for an aero-engine compressor, including a controller (not shown) and multi-stage stator blades 100.

[0044] The stationary blade 100 includes an inner casing 1, an outer casing 2 surrounding the outer side of the inner casing 1, and a plurality of blades 3 disposed between the inner casing 1 and the outer casing 2. The plurality of blades 3 are arranged at intervals along the circumference of the inner casing 1.

[0045] In practice, the outer casing 2 of the multi-stage stator vanes 100 is an integral structure, which can be understood as the multi-stage stator vanes 100 sharing a single outer casing 2. Each stage stator vane 100 has its own inner casing 1.

[0046] The blade 3 includes a supporting frame 31, a flexible blade skin 32, and a liquid metal that undergoes a solid-liquid phase change with temperature. It is understood that when the temperature rises to a preset temperature, the liquid metal changes to a liquid state; when the temperature falls below the preset temperature, the liquid metal solidifies into a solid state.

[0047] The flexible blade skin 32 covers the outside of the supporting frame 31 to form a closed space 33, and liquid metal is filled in the closed space 33.

[0048] The support frame 31 can deform. A heating element 4 electrically connected to the controller is provided in the enclosed space 33. The heating element 4 is used to heat the liquid metal. The controller is used to control the working state of the heating element 4 to cause the liquid metal to undergo a phase change, so that the flexible blade skin 32 deforms under the action of the liquid metal and the support frame 31.

[0049] For example, the heating element 4 may be a Bi2Te3 / PbTe composite element. The heating element 4 may be disposed on the support frame 31.

[0050] For example, the support frame 31 can be made of materials that are resistant to high temperature, high strength and long fatigue life, such as titanium alloy, carbon fiber composite material, etc.

[0051] In practice, the enclosed space 33 of the blade 3 can be divided into multiple regions by the support frame 31. Each region is filled with liquid metal and heating element 4. The liquid metal in the corresponding region is heated by the heating element 4, thereby adjusting the profile of the flexible blade skin 32 in each region. This allows the blade profile to be adjusted, improves the adjustment accuracy of the blade profile, and achieves effective adjustment of the airflow direction.

[0052] The blade system of the aero-engine compressor provided in this embodiment configures the blade 3 of the stator blade 100 as including a support frame 31, a flexible blade skin 32, and a liquid metal that can undergo a solid-liquid phase change with temperature. The flexible blade skin 32 covers the outside of the support frame 31 to form a closed space 33, and the liquid metal is filled in the closed space 33. By setting a controller and a heating element 4 electrically connected to the controller in the closed space 33, the liquid metal is heated by the heating element 4, causing the liquid metal to undergo a phase change. That is, according to different flight conditions, the control... The heating element heats or stops heating the liquid metal, achieving a solid-liquid phase change. Furthermore, since the support frame 31 can deform, when the liquid metal undergoes a solid-liquid phase change, the flexible blade skin 32 deforms under the synergistic effect of the liquid metal and the support frame 31, achieving dynamic adjustment of the stator blade profile 3. This allows for effective adjustment of the airflow direction, improving the compressor's adaptability to different flow fields, increasing the compressor's efficiency and stability, and ultimately enhancing the engine's overall efficiency, thrust, and operational stability. This enables the aero-engine to maintain optimal performance under various operating conditions.

[0053] Reference Figure 2 and Figure 3 As shown, in some embodiments, the support frame 31 includes an intermediate shaft 311 and a plurality of support frames 312. The plurality of support frames 312 are arranged circumferentially around the intermediate shaft 311, and each support frame 312 is connected to the intermediate shaft 311 and can move around the intermediate shaft 311. The flexible blade skin 32 covers the outside of the plurality of support frames 312.

[0054] Understandably, each support frame 312 can move around the central axis 311, thereby allowing the entire support frame 31 to deform. This, in turn, works in conjunction with the phase-change liquid metal to deform the flexible blade skin 32, achieving dynamic adjustment of the entire blade surface 3.

[0055] Furthermore, combined Figures 2 to 4 As shown, one end of the support frame 312 has a collar 315, which is sleeved on the intermediate shaft 311. The intermediate shaft 311 is rotatable, and a first meshing tooth 316 is provided on a portion of the outer peripheral wall of the intermediate shaft 311. A second meshing tooth 317 is provided on a portion of the inner ring wall of the collar 315. The second meshing tooth 317 can mesh with the first meshing tooth 316, so that the support frame 312 can move around the intermediate shaft 311, causing the flexible blade skin 32 to deform under the action of the liquid metal and the support frame 31.

[0056] With this configuration, when the intermediate shaft 311 rotates, the interaction between the first meshing tooth 316 and the second meshing tooth 317 causes the support frame 312 to move stably relative to the intermediate shaft 311.

[0057] In a specific implementation, through holes can be provided on the flexible blade skin 32 and the outer casing 2. One end of the intermediate shaft 311 passes through the through holes of the flexible blade skin 32 and the outer casing 2 sequentially to the outside of the outer casing 2. The intermediate shaft 311 is sealed to the wall of the through hole of the flexible blade skin 32. A drive motor or other drive component is provided on the outside of the outer casing 2, and the intermediate shaft 311 is connected to the drive component. The drive component drives the intermediate shaft 311 to rotate, thereby realizing the displacement of the support frame 31.

[0058] Continue to refer to Figure 3 As shown, in some embodiments, each support frame 312 includes a support rod 313 and a connecting rod 314. The support rod 313 is radially disposed within the flexible blade skin 32 along the inner casing 1. The end face of the flexible blade skin 32 corresponding to the outer casing 2 is in contact with the inner wall of the outer casing 2 under the action of the multiple support rods 313, and the end face of the flexible blade skin 32 corresponding to the inner casing 1 is in contact with the outer wall of the inner casing 1 under the action of the multiple support rods 313. One end of the connecting rod 314 is connected to and relatively fixed to the support rod 313, and the other end of the connecting rod 314 is connected to the intermediate shaft 311 to drive the support rod 313 to move around the intermediate shaft 311.

[0059] This design not only improves the overall support effect of the support frame 31 on the flexible blade skin 32, but also ensures that the blade 3 has a certain deformation space, thereby further ensuring the smooth adjustment of the blade 3 profile.

[0060] Compared to a scheme where the blades are rotatably connected to the inner and outer casings to adjust the blade angle and thus regulate the airflow direction, this method inevitably results in gaps between the rotating blades and the walls of the inner and outer casings. It also leads to compressor thrust loss and increased local flow separation and turbulence. However, the embodiment of this disclosure, through the aforementioned arrangement, ensures that the flexible blade skin 32 fits snugly against the corresponding end faces of the inner and outer casings. This reduces gaps to a certain extent, improves the sealing effect, effectively blocks gas leakage paths, and significantly reduces local airflow separation and increased turbulence. Therefore, it not only reduces compressor thrust loss but also improves the overall performance and reliability of the engine.

[0061] For example, the connecting rod 314 and the support rod 313 can be integrally formed.

[0062] In addition, the support rod 313 can be made into a hollow structure, and reinforcing ribs can be set in the inner cavity of the support rod 313. The reinforcing ribs can be set into a mesh structure, which not only reduces weight and flight drag, but also improves the structural strength of the support rod 313 by setting reinforcing ribs, thereby ensuring effective support for the flexible blade skin 32 and ensuring the structural strength of the blade 3.

[0063] The connecting rod 314 can also be configured as a hollow structure, and reinforcing ribs can be provided in the inner cavity of the connecting rod 314. The reinforcing ribs can be configured as a mesh structure, which not only reduces weight and flight drag, but also improves the structural strength of the connecting rod 314 by providing reinforcing ribs. This ensures the connection stability between the intermediate shaft 311 and the support rod 313, and thus ensures the effective support of the support frame 31 for the flexible blade skin 32, and ensures the structural strength of the blade 3.

[0064] For example, the support frame 31 can be a Ti-6Al-4V honeycomb structure formed by selective laser melting (SLM) (wall thickness 0.8mm, pore diameter Φ2mm, porosity 62%), wherein the support rod 313 can be a W-25Re alloy wedge support (e.g., front edge thickness 3.2mm, rear edge tapering to 1.5mm).

[0065] In practice, the support rod 313 can be placed in key areas of the pressure / suction surface of the blade 3 (such as the leading edge and throat), thereby ensuring effective support of the support frame 31 for the flexible blade skin 32, while retaining liquid adjustability in non-critical areas. This design ensures the structural strength of key parts while providing deformation freedom for non-critical areas.

[0066] Reference Figure 5 As shown, in some embodiments, the flexible blade skin 32 includes a deformable flexible inner support layer 321 and an elastic protective layer 322 disposed on the outside of the flexible inner support layer 321.

[0067] This design further balances the flexible deformation capability and support capability of the flexible blade skin 32.

[0068] Furthermore, the flexible inner support layer 321 is a mesh-like shape memory alloy layer. The elastic protective layer 322 includes a ceramic fiber braided layer 323 disposed on the outside of the flexible inner support layer 321 and an elastic coating 324 disposed on the outside of the ceramic fiber braided layer 323.

[0069] By setting the outer layer of the flexible blade skin 32 as a high-temperature resistant ceramic fiber braided layer 323 and coating the surface with an elastic coating 324 to maintain an aerodynamically smooth surface, and setting the inner layer as a shape memory alloy wire mesh support, the flexible blade skin 32 can absorb thermal stress by utilizing its super elasticity to prevent wrinkles. This not only meets the temperature resistance requirements of the compressor in high-temperature environments, but also achieves good aerodynamic performance and structural stability through the above structure.

[0070] For example, the elastic coating 324 can be an ultra-thin polyimide layer, or of course, other high-temperature resistant elastic materials.

[0071] For example, the shape memory alloy wire mesh can be made of Ni50.2-Ti49.3-Cu0.5 (at%), and the deformation temperature can be, for example, 120°C.

[0072] In some embodiments, a temperature sensor 5 is provided in the enclosed space 33. The temperature sensor 5 is electrically connected to the controller. The temperature sensor 5 is used to detect the temperature of the liquid metal. The controller is also used to control the working state of the heating element 4 according to the temperature detected by the temperature sensor 5.

[0073] For example, temperature sensor 5 can be a Pt1000 thin film sensor.

[0074] The temperature sensor 5 detects the temperature change of the liquid metal in real time, enabling the controller to control the heating element 4 to heat the liquid metal, continue heating, or stop heating based on the detected temperature, thereby improving control accuracy and response speed.

[0075] Combination Figure 2 , Figure 3 and Figure 6 As shown, in some embodiments, the blade system of the aero-engine compressor further includes a pressure control module. The pressure control module includes a pressure regulator 7 and a pressure sensor 6 disposed within an enclosed space 33. The pressure sensor 6 and the pressure regulator 7 are electrically connected to a controller.

[0076] Pressure sensor 6 is used to detect the pressure of liquid metal, and the controller is also used to control pressure regulator 7 to adjust the pressure of liquid metal based on the pressure value detected by pressure sensor 6.

[0077] For example, the pressure sensor 6 can be an aluminum nitride thin film sensor that can withstand temperatures up to 600°C.

[0078] For example, the pressure sensor 6 is located in the enclosed space 33, such as at the root of the support rod 313 of the support frame 31 or in the stress concentration area of ​​the flexible blade skin 32.

[0079] By incorporating a pressure control module, precise pressure control is achieved, ensuring that the liquid metal pressure and aerodynamic load are balanced during the deformation of blade 3. Pressure sensor 6 monitors real-time pressure changes in the liquid metal, further ensuring that the liquid metal can counteract aerodynamic forces while in a liquid state.

[0080] For example, refer to Figure 6 As shown, the pressure regulating component 7 includes a retractable hydraulic rod. Specifically, the outer casing 2 has a through hole, and the flexible blade skin 32 has a clearance hole 325 at a position corresponding to the through hole. The hydraulic rod passes through the through hole and the clearance hole 325, and is sealed to the wall of the clearance hole 325. The controller controls the extension and retraction of the hydraulic rod based on the pressure value detected by the pressure sensor 6.

[0081] When the pressure of the liquid metal is detected to be lower than the preset pressure value, the hydraulic rod extends into the enclosed space 33 to press against the liquid metal, thereby pressurizing the liquid metal so that it can resist the pneumatic force.

[0082] Specifically, the hydraulic rod can be connected to a hydraulic system, which then drives the rod to extend or retract. For example, the hydraulic rod can provide a working pressure of 3–5 MPa.

[0083] During the blade profile switching process, for example, when the liquid metal is in a liquid state, the pressure regulator 7 provides a pressure of about 3 MPa to ensure that the liquid metal can withstand aerodynamic loads; when the liquid metal solidifies, the pressure regulator 7 automatically releases excess pressure to avoid structural damage. This dynamic balancing mechanism enables the compressor to adapt to changes in aerodynamic loads under different operating conditions while maintaining good aerodynamic performance.

[0084] For example, the blade 3 can be divided into multiple independent phase change control regions (such as leading edge, middle section, trailing edge, etc.), each region is equipped with an independent temperature sensor 5, pressure sensor 6 and pressure regulating component 7, thereby achieving high-precision, fast response and multi-region coordinated control of liquid metal temperature and pressure, and thus accurately adjusting the blade 3 profile.

[0085] In some embodiments, in the multi-stage stationary vanes 100, the melting point of the liquid metal in the blade 3 of the compressor's last stage stationary vane 100 is higher than the melting point of the liquid metal in the blade 3 of the compressor's inlet stage stationary vane 100.

[0086] This allows for the selection of different materials for the liquid metal based on the varying thermal environments of each stage of the compressor blades. For example, the inlet stage blades operate at lower temperatures, so a low-melting-point gallium-based alloy (such as a Ga-In-Sn alloy) can be used for their liquid metal. The intermediate stages can utilize bismuth-based alloys (such as Bi-Pb alloys). The final stage, with its higher temperatures, can employ a high-melting-point Al-Cu-Si-Mg-Li multi-element alloy.

[0087] For example, the liquid metal of the inlet stage blade 3 can be Ga62In21Sn17, the liquid metal of the intermediate stage blade 3 can be Bi53Pb32In15, and the liquid metal of the final stage blade 3 can be Al86Cu7Si5Mg1Li1. The phase transition temperature of the liquid metal is, for example, 360℃.

[0088] This regional material selection strategy further ensures that the phase change temperature of the liquid metal is slightly higher than the operating temperature of the blade 3. When there is no heating, the liquid metal solidifies due to cooling by the external airflow. When the profile changes, the heating element 4 can quickly melt the liquid metal to achieve structural deformation of the blade 3, so as to effectively regulate the airflow and further improve the stability of the blade 3 under different operating conditions and adapt to different flow fields.

[0089] This embodiment utilizes the solid-liquid phase change characteristics of liquid metal and the synergistic effect of the internal movable frame to achieve continuous change and precise matching of the blade profile 3. The specific working process is as follows: During profile switching, the controller activates the heating element 4 to melt the liquid metal, for example, ensuring that the liquid metal in the target area completes a phase change within a preset time (e.g., 1 second). Simultaneously, the pressure control system applies appropriate pressure to the liquid metal to ensure it can withstand aerodynamic loads (e.g., the pressure of a high-pressure compressor, approximately 3 MPa). The internal movable support frame 31 displaces, adjusting its position and guiding the liquid metal to the area containing the target shape. Finally, the controller stops heating the heating element 4, and the liquid metal rapidly solidifies under the cooling effect of the external airflow and its own properties, thereby fixing the new blade profile 3 to adapt to the current airflow conditions and ensure the efficient operation of the compressor.

[0090] Liquid metal possesses extremely high mechanical strength in its solid state, capable of withstanding the aerodynamic loads of a compressor during operation. In its liquid state, it exhibits excellent fluidity and plasticity, allowing it to accommodate the shape variations of the internal framework and simultaneously meet the requirements for aerodynamic performance and structural strength. Furthermore, liquid metal has excellent thermal conductivity, aiding in heat dissipation and temperature control. During shape switching, temperature control and pressure control work together to form a complete control closed loop.

[0091] The embodiments disclosed herein employ a collaborative approach of "internal movable support frame + flexible blade skin + liquid metal filling + precise temperature control + precise pressure control" to achieve a dynamic balance system of "material phase change - structural deformation - aerodynamic load". This enables continuous and precise matching of the three blade profiles, meeting aerodynamic requirements under different operating conditions, thereby improving the efficiency and stability of the compressor, and consequently improving the overall efficiency, thrust and stability of the engine.

[0092] This disclosure also provides an aero-engine, including a blade system, which can be specifically applied to an aircraft.

[0093] The blade system in this embodiment has the same specific structure and implementation principle as the blade system of the aero-engine compressor provided in the above embodiments, and can bring the same or similar technical effects. It will not be described in detail here, but can be referred to the description of the above embodiments.

[0094] The above description is merely an embodiment of this disclosure and an explanation of the technical principles employed. Those skilled in the art should understand that the scope of this disclosure is not limited to technical solutions formed by specific combinations of the above-described technical features, but should also cover other technical solutions formed by arbitrary combinations of the above-described technical features or their equivalents without departing from the above-described concept. For example, technical solutions formed by substituting the above features with (but not limited to) technical features disclosed in this disclosure that have similar functions.

[0095] While specific embodiments of this disclosure have been described in detail by way of example, those skilled in the art should understand that the examples are for illustrative purposes only and not intended to limit the scope of this disclosure. Those skilled in the art should understand that modifications can be made to the above embodiments without departing from the scope and spirit of this disclosure. The scope of this disclosure is defined by the appended claims.

Claims

1. A blade system for an aero-engine compressor, characterized in that, It includes a controller and multi-stage stationary blades. The stationary blades include an inner casing, an outer casing surrounding the inner casing, and multiple blades disposed between the inner casing and the outer casing. The multiple blades are arranged at intervals along the circumference of the inner casing. The blade includes a support frame, a flexible blade skin, and a liquid metal that can undergo a solid-liquid phase change with temperature. The flexible blade skin covers the outside of the support frame to form a closed space, and the liquid metal fills the closed space. The supporting frame is deformable, and a heating element electrically connected to the controller is provided in the enclosed space. The heating element is used to heat the liquid metal, and the controller is at least used to control the working state of the heating element to cause the liquid metal to undergo a phase change, so that the flexible blade skin deforms under the action of the liquid metal and the supporting frame.

2. The blade system of the aero-engine compressor according to claim 1, characterized in that, The support frame includes an intermediate shaft and multiple support frames; The plurality of support frames are arranged circumferentially around the intermediate axis, and each support frame is connected to the intermediate axis and can move around the intermediate axis. The flexible blade skin covers the outside of the multiple support frames.

3. The blade system of the aero-engine compressor according to claim 2, characterized in that, One end of the support frame has a collar, which is sleeved on the intermediate shaft; The intermediate shaft is rotatable, and a first meshing tooth is provided on a portion of the outer peripheral wall of the intermediate shaft. A second meshing tooth is provided on a portion of the inner ring wall of the collar. The second meshing tooth can mesh with the first meshing tooth, so that the support frame can move around the intermediate shaft, and the flexible blade skin deforms under the action of the liquid metal and the support frame.

4. The blade system of the aero-engine compressor according to claim 2, characterized in that, Each of the aforementioned support frames includes a support rod and a connecting rod; The support rod is arranged radially within the flexible blade skin along the inner casing. The end face of the flexible blade skin corresponding to the outer casing is in contact with the inner wall of the outer casing under the action of the multiple support rods, and the end face of the flexible blade skin corresponding to the inner casing is in contact with the outer wall of the inner casing under the action of the multiple support rods. One end of the connecting rod is connected to and fixed relative to the support rod, and the other end of the connecting rod is connected to the intermediate shaft to drive the support rod to move around the intermediate shaft.

5. The blade system of the aero-engine compressor according to claim 1, characterized in that, The flexible blade skin includes a deformable flexible inner support layer and an elastic protective layer disposed on the outside of the flexible inner support layer.

6. The blade system of the aero-engine compressor according to claim 5, characterized in that, The flexible inner support layer is a mesh-like shape memory alloy layer; The elastic protective layer includes a ceramic fiber braided layer disposed on the outside of the flexible inner support layer and an elastic coating disposed on the outside of the ceramic fiber braided layer.

7. The blade system of an aero-engine compressor according to any one of claims 1 to 6, characterized in that, A temperature sensor is installed inside the enclosed space; The temperature sensor is electrically connected to the controller. The temperature sensor is used to detect the temperature of the liquid metal, and the controller is used to control the working state of the heating element according to the temperature detected by the temperature sensor.

8. The blade system of an aero-engine compressor according to any one of claims 1 to 6, characterized in that, The blade system of the aero-engine compressor also includes a pressure control module; The pressure control module includes a pressure regulator and a pressure sensor disposed within the enclosed space; the pressure sensor and the pressure regulator are electrically connected to the controller, respectively. The pressure sensor is used to detect the pressure of the liquid metal, and the controller is also used to control the pressure regulator to adjust the pressure of the liquid metal according to the pressure value detected by the pressure sensor.

9. The blade system of an aero-engine compressor according to any one of claims 1 to 6, characterized in that, In the multi-stage stator blades, the melting point of the liquid metal in the blades of the last stage stator blade of the compressor is higher than that of the liquid metal in the blades of the inlet stage stator blade of the compressor.

10. An aircraft engine, characterized in that, Includes the blade system of an aero-engine compressor as described in any one of claims 1 to 9.

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