Composite thermal protection air rudder of aircraft and its manufacturing method

Abstract

This invention discloses a composite thermal protection aerodynamic rudder for aircraft and its manufacturing method, aiming to solve the problems of high manufacturing cost, poor thermal insulation, and poor overall mechanical performance in existing structures. The composite thermal protection aerodynamic rudder of this invention includes a metal frame, lightweight filler, and a composite material thermal protection structure arranged sequentially from the inside out. The metal frame includes a rudder shaft and a rudder surface, which are integrally formed. The rudder surface has an internal frame rib structure forming an internal cavity, which is filled with lightweight filler. The composite material thermal protection structure covers the outside of the rudder surface and the lightweight filler, forming the aerodynamic shape of the aerodynamic rudder. Through the frame rib structure design of the metal frame, the amount of material used is significantly reduced while ensuring load-bearing capacity, achieving lightweighting, which is beneficial to improving the thrust-to-weight ratio and maneuverability of the aircraft. Optimizing the rib layout according to the load path achieves efficient material utilization.

Description

Technical Field

[0001] This invention relates to the field of aircraft structural design and manufacturing technology, and in particular to a composite structure thermal protection air rudder for an aircraft and its manufacturing method. Background Technology

[0002] In the field of high-speed aircraft, aerodynamic control systems are key actuators in their attitude control systems, requiring them to withstand enormous aerodynamic loads while resisting extreme aerodynamic heating environments. Therefore, an ideal aerodynamic control system must possess high strength, high rigidity, excellent thermal protection capabilities, and lightweight characteristics.

[0003] Traditional technical solutions have two main drawbacks: First, the control surface made of all-metal integral processing is bulky, and the high thermal conductivity of metal makes it easy for heat to be quickly transferred to the control shaft connection, threatening the safety of the actuator; Second, the integral control surface made of carbon / silicon carbide and other ceramic matrix composite materials, although resistant to high temperature and lightweight, has problems such as high brittleness, poor impact resistance and extremely high manufacturing cost, which makes it difficult to meet the economic requirements of mass application.

[0004] Chinese patent CN120024056A discloses a lattice composite structure thermal protection air rudder for high-speed aircraft. It uses a metal lattice structure integrally formed by laser selective melting additive manufacturing process as the main load-bearing skeleton, and uses resin transfer molding and other processes to coat and fill the lattice structure with composite materials to form the final rudder body. The lattice design achieves the lightweight of the structure and provides thermal protection by using composite materials, showing a certain degree of progress.

[0005] However, this solution still has the following significant shortcomings, which limit its application in engineering and low-cost mass production: 1. The core metal lattice framework relies entirely on metal additive manufacturing technologies such as laser selective melting. This process has inherent drawbacks for large-size, complex spatial structures, including huge equipment investment, high raw material costs, slow forming speed, and cumbersome post-processing. This results in high unit manufacturing costs, making it difficult to meet the demands of modern weaponry for large-scale, cost-effective mass production. 2. It relies on external composite material layers to block heat. Its metal lattice skeleton is usually filled with composite material or is a cavity. The lattice structure itself is composed of a large number of interconnected thin metal rods. These metal rods provide a low-resistance path for heat conduction to the inside of the rudder body at high temperatures, resulting in poor heat insulation performance. 3. The mechanical properties, especially fatigue properties, of additively manufactured lattice structures are extremely sensitive to process parameters. They are prone to defects such as incomplete fusion and porosity, and stress concentration is prominent at the nodes. Summary of the Invention

[0006] (a) Purpose of the invention The purpose of this invention is to provide a composite structure thermal protection air rudder for aircraft and its manufacturing method, aiming to solve the problems of high manufacturing cost, poor heat insulation and poor overall mechanical performance in existing structures.

[0007] (II) Technical Solution To address the aforementioned problems, a first aspect of the present invention provides a composite structure thermal protection air rudder for an aircraft, the composite structure thermal protection air rudder comprising, from the inside out, a metal frame, a lightweight filler, and a composite material thermal protection structure; The metal frame includes a rudder shaft and a rudder surface, which are integrally formed. The rudder surface has a frame rib structure inside, forming an internal cavity, and the lightweight filler is filled into the internal cavity. The composite material heat-resistant structure covers the outside of the rudder surface and the lightweight filler, forming the aerodynamic shape of the air rudder.

[0008] Preferably, the rib distribution density of the frame rib structure in the root region where the rudder surface connects to the rudder shaft is greater than that in the tip region of the rudder surface.

[0009] Preferably, the outer surface of the lightweight filler is flush with the outer edge contour of the rudder surface of the metal frame.

[0010] Preferably, the material of the metal skeleton is aluminum alloy, titanium alloy or high-temperature alloy.

[0011] Preferably, the lightweight filler material is selected from at least one of fiber preform-reinforced modified phenolic resin-based composite materials, porous ceramics, or aerogel insulation materials.

[0012] Another aspect of the present invention provides a method for manufacturing a composite structure thermal protection air rudder, the method comprising the following steps: S1: Based on the thermal environment conditions, load conditions and application scenario requirements of the aircraft's air rudder, determine the structure of the metal frame and the composite material heat protection structure. The rudder shaft of the metal frame adopts a solid structure, and the rudder surface of the metal frame adopts a frame rib structure. The metal frame with rudder surface and frame rib structure is integrally formed by machining or investment casting process. S2: Based on the thermal environment conditions of the aircraft and the molding process of the composite material heat-resistant structure, determine the lightweight filler and prepare a lightweight filler preform with a matching shape according to the shape of the internal cavity of the metal frame control surface. Place the lightweight filler preform into the internal cavity of the control surface of the metal frame. S3: A composite material heat-resistant structure is formed by coating the metal skeleton filled with the lightweight filler through a composite material molding process.

[0013] Preferably, in step S1, the distribution and size of the frame rib structure are designed based on finite element analysis, as well as the composite material heat-resistant structure is designed to meet the predetermined stiffness, strength and weight indicators.

[0014] Preferably, step S2 further includes coating the surface of the cavity inside the rudder surface of the metal frame with a high-temperature resistant adhesive.

[0015] Preferably, in step S3, the composite material molding process is a resin transfer molding process or an integrated molding process.

[0016] Preferably, the manufacturing method further includes step S3a: performing thermal simulation analysis on the combination of the metal skeleton and the lightweight filler based on thermal analysis software, and designing the thickness distribution of the composite material heat-resistant structure in different regions of the rudder surface based on the temperature field and heat flow distribution results obtained from the analysis.

[0017] (III) Beneficial Effects The above-described technical solution of the present invention has the following beneficial technical effects: 1. By using a metal frame rib structure design, the amount of material used can be significantly reduced while ensuring load-bearing capacity, thus achieving lightweighting. This is beneficial for improving the thrust-to-weight ratio and maneuverability of the aircraft. The rib layout is optimized according to the load path to achieve efficient use of materials.

[0018] 2. The composite material heat-resistant structure directly faces the high-temperature environment, and its low thermal conductivity effectively blocks heat from being transferred to the interior; the lightweight filler further slows down heat conduction, protecting the metal skeleton from overheating failure. The multi-layer heat-resistant mechanism significantly improves the structural integrity and service life of the rudder at high temperatures.

[0019] 3. The metal skeleton is integrally molded, with mature technology and controllable manufacturing costs, avoiding the high costs and process uncertainties of additive manufacturing. The composite material heat-resistant structure can be molded, tightly integrated with the metal skeleton and lightweight filler, with a simple overall structure, reducing assembly links and the use of connectors, further reducing manufacturing costs, and improving production efficiency. Attached Figure Description

[0020] Figure 1 This is a schematic diagram of the composite thermal protection air rudder of an aircraft according to the present invention; Figure 2 This is a cross-sectional schematic diagram of a composite thermal protection air rudder for an aircraft according to the present invention; Figure 3 This is a flowchart of a method for manufacturing a composite structure thermal protection air rudder according to the present invention.

[0021] Figure label: 1. Metal frame; 11. Rudder shaft; 12. Rudder surface; 13. Frame rib structure; 2. Lightweight filler; 3. Composite material heat-resistant structure. Detailed Implementation

[0022] To make the objectives, technical solutions, and advantages of this invention clearer, the invention will be further described in detail below with reference to specific embodiments and the accompanying drawings. It should be understood that these descriptions are merely exemplary and not intended to limit the scope of the invention. Furthermore, descriptions of well-known structures and techniques are omitted in the following description to avoid unnecessarily obscuring the concept of the invention.

[0023] The accompanying drawings illustrate layer structure diagrams according to embodiments of the present invention. These drawings are not to scale, and some details have been enlarged for clarity, and some details may have been omitted. The shapes of the various regions and layers shown in the drawings, as well as their relative sizes and positional relationships, are merely exemplary and may deviate from reality due to manufacturing tolerances or technical limitations. Furthermore, those skilled in the art can design regions / layers with different shapes, sizes, and relative positions as needed.

[0024] Obviously, the described embodiments are only some, not all, of the embodiments of the present invention. All other embodiments obtained by those skilled in the art based on the embodiments of the present invention without inventive effort are within the scope of protection of the present invention.

[0025] Combination Figure 1 and Figure 2 The first aspect of the present invention provides a composite structure thermal protection air rudder for an aircraft. The composite structure thermal protection air rudder includes a metal frame 1, a lightweight filler 2, and a composite material thermal protection structure 3 arranged sequentially from the inside to the outside. The metal frame 1 includes a rudder shaft 11 and a rudder surface 12, which are integrally formed. The rudder surface 12 has a frame rib structure 13 inside, forming an internal cavity, and the lightweight filler 2 is filled in the internal cavity. The composite material thermal protection structure 3 covers the outside of the rudder surface 12 and the lightweight filler 2, and forms the aerodynamic shape of the air rudder.

[0026] Specifically, the metal frame 1 serves as the main load-bearing core of the air rudder, while the rudder shaft 11 undertakes the function of connecting and fixing with the aircraft, providing the installation foundation and torque transmission path for the air rudder. The frame rib structure 13 of the rudder surface 12 directly bears the aerodynamic load during high-speed flight, ensuring the structural stability of the rudder body under complex stress, while providing installation space and support for the lightweight filler 2. The lightweight filler 2 fills the internal cavity of the metal frame 1, playing a role in auxiliary support and optimizing weight distribution, while blocking the heat conduction path inside the metal frame 1, enhancing the overall heat insulation effect, and avoiding local overheating that could lead to structural failure. The composite material heat-resistant structure 3 covers the outside, directly resisting extreme aerodynamic heating during high-speed flight, forming the aerodynamic shape of the air rudder, ensuring thermal protection performance, meeting the aerodynamic layout requirements of the aircraft, and protecting the internal metal frame 1 and lightweight filler 2. The metal frame 1 is connected to the aircraft actuator via the rudder shaft 11. The actuator drives the rudder shaft 11 to deflect the control surface 12, thereby achieving attitude control of the aircraft. The frame rib structure 13 evenly transmits aerodynamic loads to the entire frame. It utilizes the integral molding structure to disperse stress and avoid localized stress concentration. After the lightweight filler 2 fills the internal cavity, it forms a synergistic force-bearing system with the metal frame 1. At the same time, it utilizes its own low thermal conductivity to reduce the conduction of external heat to the interior through the frame cavity. The composite material heat-resistant structure 3 is in direct contact with the high-temperature airflow. It blocks heat intrusion through its own high-temperature resistance. Its encapsulated structure ensures that heat cannot penetrate from the edge or gap of the control surface 12. From the inside out, a synergistic system of load-bearing, heat insulation, and heat protection is formed, which not only ensures the structural load-bearing capacity but also achieves efficient thermal protection.

[0027] Through this design, the frame rib structure 13 of the metal skeleton 1 significantly reduces material usage while ensuring load-bearing capacity, achieving lightweighting. This is beneficial for improving the thrust-to-weight ratio and maneuverability of the aircraft. Optimizing the rib layout according to the load path achieves efficient material utilization. The composite material heat-resistant structure 3 directly faces the high-temperature environment, and its low thermal conductivity effectively blocks heat transfer to the interior. The lightweight filler 2 further delays heat conduction, protecting the metal skeleton 1 from overheating failure. This multi-layered heat-resistant mechanism significantly improves the structural integrity and service life of the control body at high temperatures. The metal skeleton 1 is integrally molded, with mature technology and controllable manufacturing costs, avoiding the high costs and process uncertainties of additive manufacturing. The composite material heat-resistant structure 3 can be molded, tightly integrated with the metal skeleton 1 and lightweight filler 2, resulting in a simple overall structure, reducing assembly steps and the use of connectors, further reducing manufacturing costs while improving production efficiency.

[0028] In a preferred embodiment, the rib distribution density of the frame rib structure 13 is greater in the root region where the control surface 12 connects to the control shaft 11 than in the tip region of the control surface 12. Specifically, the high-density rib distribution in the root region of the control surface 12, i.e., the part connected to the control shaft 11, mainly bears the huge aerodynamic loads and moments concentrated on the root during high-speed flight, thus strengthening the load-bearing capacity of the root; the lower rib distribution density in the tip region reduces unnecessary material usage while meeting basic structural support requirements, thereby optimizing the overall weight distribution of the control surface 12. During high-speed flight, the aerodynamic load of the air rudder is mainly concentrated in the root region where the rudder surface 12 connects to the rudder shaft 11. This region needs to withstand greater bending and torsional stress. By increasing the rib distribution density in the root region, the structural cross-sectional moment of inertia and bearing area in this region can be increased, thereby improving the bending and torsional resistance and effectively dispersing the concentrated stress. Meanwhile, the load on the tip region of the rudder surface 12 is relatively small. Reducing the rib distribution density will not affect the structural strength, while reducing the weight of the tip region, optimizing the center of mass position of the rudder body, reducing the torque required for the actuator to drive the rudder surface 12 to deflect, and improving the control response speed.

[0029] With this configuration, the root bears the maximum load, while the load gradually decreases at the tip. The differentiated design of the rib distribution density achieves a precise match between high strength at high loads and light weight at low loads. The high-density ribs at the root can effectively resist concentrated stress and prevent structural failure at the root due to excessive load. The low-density ribs at the tip avoid material waste, improve overall load-bearing efficiency, and extend the service life of the aerodynamic control system. More importantly, reducing the rib distribution density in the tip region can reduce the weight of the control surface 12, bringing the center of gravity of the control body closer to the control axis 11. This reduces the moment of inertia when the control surface 12 deflects, lowers the drive power requirement of the actuator, and makes the deflection response of the aerodynamic control system faster and more sensitive. This improves the timeliness and accuracy of the aircraft's attitude control, while also enhancing flight safety and maneuverability.

[0030] In a preferred embodiment, the outer surface of the lightweight filler 2 is flush with the outer edge contour of the control surface 12 of the metal frame 1. Specifically, the lightweight filler 2 completely fills the metal fastener and can fit into the composite material heat-resistant structure 3 at the hollowed-out position of the metal frame 1. This provides a flat and continuous bonding surface for the composite material heat-resistant structure 3, ensuring the stability of the overlay molding. On the other hand, it avoids structural gaps caused by the protrusion or depression of the lightweight filler 2, ensuring the tightness of the connection between the three and maintaining the integrity of the internal structure of the control surface 12. After the lightweight filler 2 fills the internal cavity of the metal frame 1, its outer surface is flush with the outer edge of the frame, so that the composite material heat-resistant structure 3 can form a seamless interface with the metal frame 1 and the lightweight filler 2 during overlay, without gaps or stress concentrations caused by unevenness of the lightweight filler 2. The flat interface ensures that heat will not accumulate locally due to gaps during the heat transfer process, and at the same time improves the synergistic stress-bearing effect of the three. When subjected to external aerodynamic loads or thermal stress, the force can be evenly transferred to each component, avoiding excessive local stress.

[0031] This design, with its flush outer surface, eliminates gaps, allowing the composite heat-resistant structure 3, lightweight filler 2, and metal frame 1 to form a continuous thermal insulation system. Heat cannot penetrate through gaps, improving the overall effectiveness of thermal protection and ensuring that the internal metal frame 1 and actuators are within a safe temperature range. At the same time, the flat bonding surface maximizes the contact area between the three components. When the air rudder is subjected to aerodynamic loads or thermal stress, the load can be evenly transferred to the metal frame 1, lightweight filler 2, and composite heat-resistant structure 3 through the bonding surface. This avoids localized stress concentration caused by uneven interfaces, reduces the stress burden on individual components, lowers the risk of structural cracking and detachment, and improves the overall structural load-bearing reliability and durability. Furthermore, the flush design of the outer surface of the lightweight filler 2 provides a regular molding benchmark for the plastic molding or integral molding process of the composite material heat protection structure 3, avoiding problems such as uneven thickness and shape deviation of the heat protection structure caused by the irregular surface of the lightweight filler 2. This not only simplifies the operation difficulty of the molding process and reduces the scrap rate, but also ensures the consistency of the dimensions and structural performance of mass-produced air rudder products, meets the standardization requirements of aircraft components, and provides convenience for subsequent assembly and use.

[0032] In a preferred embodiment, the material of the metal frame 1 is aluminum alloy, titanium alloy, or high-temperature alloy. Aluminum alloy, titanium alloy, and high-temperature alloy, as materials for the metal frame 1, provide stable structural load-bearing capacity while adapting to different operating temperature environments, meeting the balance requirements of the aerodynamic control system in terms of strength, stiffness, temperature resistance, and weight. Material selection matching the operating environment ensures that the metal frame 1 maintains sufficient strength and stiffness at the corresponding temperature, allowing it to perform optimal load-bearing and temperature resistance under the corresponding environment. Aluminum alloy, with its low density, good machinability, and moderate strength, is suitable for high-speed flight scenarios with lower temperatures and moderate loads, and can be formed through machining, meeting lightweight requirements. Titanium alloy combines high strength, corrosion resistance, and good high-temperature resistance, maintaining stable mechanical properties even in medium- and high-temperature environments, and can withstand greater aerodynamic loads. High-temperature alloy possesses excellent high-temperature resistance and high-temperature strength, suitable for extreme thermal environments, effectively resisting high-temperature erosion caused by aerodynamic heating, and preventing structural softening or failure at high temperatures.

[0033] This design addresses the varying flight speeds and altitudes of different high-speed aircraft, resulting in differences in the temperature and load conditions of the aerodynamic control surfaces. Aluminum alloys are suitable for scenarios with lower Mach numbers and less intense aerodynamic heating, such as some short-range high-speed aircraft, where their low-density characteristics maximize weight reduction. Titanium alloys are suitable for medium- and long-range high-speed aircraft, maintaining structural stability under higher temperatures, such as 300-600℃, and heavier loads. High-temperature alloys are suitable for long-range or hypersonic aircraft, capable of withstanding extreme temperatures above 600℃. This flexibility in material selection allows aerodynamic control surfaces to adapt to the needs of different types of aircraft, expanding the product's application range. Aluminum alloys have relatively low raw material and processing costs. For mass production, cost-sensitive applications, and mild operating environments, choosing aluminum alloys can significantly reduce overall manufacturing costs while meeting basic performance requirements. Although titanium alloys and high-temperature alloys are more expensive, they can prevent structural failures and increased maintenance costs due to insufficient material performance under harsh environments, achieving long-term economic efficiency. For example, for conventional high-speed aircraft that are mass-produced, using an aluminum alloy frame can control costs; for advanced aircraft with high performance and long endurance, using titanium alloys or high-temperature alloys can ensure reliability, reduce maintenance frequency, and improve the overall economic efficiency throughout the entire life cycle.

[0034] In a preferred embodiment, the lightweight filler 2 is selected from at least one of fiber preform-reinforced modified phenolic resin-based composite materials, porous ceramics, or aerogel insulation materials. The lightweight filler 2 material includes fiber preform-reinforced modified phenolic resin-based composite materials, porous ceramics, or aerogel insulation materials. These materials have low density, low thermal conductivity, and certain temperature resistance, maintaining structural stability at high temperatures and effectively isolating heat. The fiber preform-reinforced modified phenolic resin-based composite material possesses both structural strength and thermal insulation performance, and can work in conjunction with the metal skeleton 1 to bear force while filling the cavity. The high-temperature resistance of its resin matrix can resist some heat transfer. The porous ceramic contains a large number of closed pores, and the still air within these pores can effectively block heat conduction. Simultaneously, the high-temperature resistance of the ceramic material ensures that it does not soften or decompose at high temperatures. The aerogel insulation material has extremely low thermal conductivity, making it a highly efficient insulation medium that can minimize heat transfer from the outside to the metal skeleton 1. By filling with these materials, a highly efficient insulation layer is formed inside the metal skeleton 1, while its lightweight properties reduce the overall weight.

[0035] With this design, the thermal conductivity of aerogel insulation material is much lower than that of ordinary insulation materials, allowing for significant heat attenuation during transfer and preventing the mechanical properties of the metal skeleton 1 from deteriorating due to high temperatures. The closed-pore structure of porous ceramics prevents heat convection and conduction, maintaining the internal temperature within a safe range even in extreme high-temperature environments, ensuring the normal operation of critical components such as actuators, and solving the problem of poor insulation performance in traditional structures. The density of fiber preform-reinforced modified phenolic resin-based composite materials is only 30%-50% of that of metal materials, while the density of porous ceramics and aerogels is even lower. After filling, they can effectively reduce the weight of the control body, reduce the fuel consumption of the aircraft, and improve range and maneuverability. At the same time, the lightweight design can also reduce the driving load of the actuators, improve the deflection response speed of the aerodynamic rudder, and enhance the flexibility of the aircraft's attitude control. Fiber preform-reinforced modified phenolic resin-based composite materials can be manufactured through machining, molding, or integrated compression molding processes, precisely matching the complex cavity shape of the metal skeleton 1. Porous ceramics can be made into customized shapes through sintering processes, while aerogels can fill irregular cavities through infusion, molding, and other methods. The molding flexibility of different materials allows the lightweight filler 2 to fit tightly with the metal frame 1, avoiding problems such as insufficient filling or loose fit, improving the integrity and compatibility of the structure, while also helping to improve the rigidity of the overall structure, reducing the risk of deformation of the metal frame 1, and further enhancing the structural stability of the air rudder.

[0036] Combination Figure 3 Another aspect of the present invention provides a method for manufacturing a composite structure thermal protection air rudder, the method comprising the following steps: S1: Based on the aircraft's thermal environment conditions, load conditions, and the application scenario requirements of the air rudder, the structure of the metal frame 1 and the composite material heat-resistant structure 3 are determined. The rudder shaft 11 of the metal frame 1 is a solid structure, and the rudder surface 12 of the metal frame 1 is a frame rib structure 13. The metal frame 1, with its rudder surface 12 and frame rib structure 13, is integrally formed using machining or investment casting. Specifically, the design parameters of the metal frame 1 are first determined through aircraft thermal environment simulation tests and load calculation analysis. The rudder shaft 11 is a solid structure to ensure connection strength and torque transmission. The rudder surface 12 is designed as a frame rib structure 13, and the rib distribution and dimensions are determined. Then, the material and initial thickness of the composite material heat-resistant structure 3 are selected based on thermal protection requirements. During manufacturing, if the metal frame 1 structure is relatively simple, it is integrally formed using machining, with the rudder shaft 11, rudder surface 12, and frame rib structure 13 directly machined using CNC machine tools or other equipment. If the structure is complex, investment casting is used to create a mold with the same shape as the frame. Molten metal is injected into the mold, and after cooling, the integral metal frame 1 is formed.

[0037] S2: Based on the aircraft's thermal environment conditions and the molding process of the composite material heat-resistant structure 3, the lightweight filler 2 is determined. A preform of the lightweight filler 2 with a matching shape is prepared according to the shape of the internal cavity of the control surface 12 of the metal frame 1. This preform of the lightweight filler 2 is then placed into the internal cavity of the control surface 12 of the metal frame 1. Specifically, considering parameters such as the temperature range and heat flux density of the aircraft's thermal environment, as well as the molding temperature and curing shrinkage rate of the composite material heat-resistant structure 3, a suitable lightweight filler 2 material is selected. The precise dimensions of the internal cavity of the control surface 12 of the metal frame 1 are obtained through 3D scanning. A preform of the lightweight filler 2 with a perfectly matching shape is prepared using machining, molding, and other processes. The preform is then aligned with the cavity position and placed in, ensuring complete filling without protruding from the outer surface of the metal frame 1.

[0038] S3: A composite material heat-resistant structure 3 is formed on the outside of the metal frame 1 filled with lightweight filler 2 using a composite material molding process. That is, a mold is set on the outside of the metal frame 1 into which the lightweight filler 2 is placed, and the shape of the mold is consistent with the aerodynamic shape of the air rudder; the composite material is injected into the mold through a resin transfer molding process, or the composite material is pressed into shape using an integrated molding process, so that the composite material completely covers the rudder surface 12 and the lightweight filler 2, and after curing, the composite material heat-resistant structure 3 is formed, completing the overall rudder manufacturing.

[0039] This manufacturing method requires low investment in machining or investment casting equipment and offers fast molding speeds, making it particularly suitable for mass production of large-size, complex structures. For example, machining and investment casting can form complex frame rib structures 13 in a single process, eliminating the need for cumbersome post-processing. Simultaneously, mature CNC equipment can be used to achieve efficient production, reducing unit manufacturing costs and production cycles, thus meeting the demands of large-scale, highly economical mass production of modern weaponry. The one-piece molding process of the metal skeleton 1 avoids gaps in the connection of separate structures, reducing heat conduction channels and stress concentration points. The lightweight filler 2 prefabricated body precisely matches the shape of the cavity, and the composite material heat-resistant structure 3 tightly covers the exterior. These three elements form a seamless whole, eliminating the risk of loosening or excess material, improving structural integrity, and making the performance of the air rudder more stable under high-temperature and high-load environments. This reduces the probability of failure due to structural gaps or loose connections, extending service life. Machining and investment casting processes enable diverse frame rib structure designs, such as tree-like, grid-like, and star-like distributions, which can be flexibly adjusted according to the load and center of gravity requirements of different aircraft. Composite material molding processes can precisely control the aerodynamic shape of the heat-resistant structure to meet the aerodynamic layout requirements of different aircraft. This makes the manufacturing method adaptable to the production of air rudders for various types of high-speed aircraft, expanding its application scope and providing more room for subsequent structural optimization.

[0040] In a preferred embodiment, in step S1, the distribution and dimensions of the frame rib structure 13 and the composite material heat-resistant structure 3 are designed based on finite element analysis to meet predetermined stiffness, strength, and weight specifications. Specifically, finite element analysis software, such as ANSYS or ABAQUS, is used to establish three-dimensional models of the metal frame 1 and the composite material heat-resistant structure 3. Thermal environment parameters and load parameters of the aircraft, such as temperature, heat flux density, aerodynamic loads, and inertial loads, are input as boundary conditions. The distribution and dimensions of the frame rib structure 13 are simulated to analyze the stress distribution, deformation, and weight data under different design schemes. Simultaneously, the material selection and thickness distribution of the composite material heat-resistant structure 3 are simulated to analyze its insulation effect and load-bearing capacity. The design parameters are iteratively optimized based on the simulation results until the predetermined stiffness, strength, and weight specifications are met. Finite element analysis discretizes the structure into multiple elements and accurately calculates the mechanical response of each element under complex loads and temperature fields. It can predict weak points in the structural design in advance, such as stress concentration areas in the frame rib structure 13 and the risk of thermal insulation failure due to insufficient thickness of the heat-resistant structure. By optimizing the distribution and size of the ribs, the weight of the metal frame 1 is minimized while meeting the strength and stiffness requirements. The optimization of the heat-resistant structure ensures that it achieves effective thermal insulation while avoiding material waste and weight redundancy, thus achieving a balance of various performance indicators.

[0041] This manufacturing method allows finite element analysis to quantify the stress and temperature response of a structure under different operating conditions. For example, simulation can determine the optimal spacing and dimensions of the frame rib structure 13, reducing the amount of ribs while ensuring strength and avoiding weight redundancy caused by over-design. Simultaneously, it can accurately identify key insulation areas of the heat-resistant structure, rationally allocate thickness, and avoid insulation failure or material waste due to uneven thickness, making the design more scientific and precise. Furthermore, finite element simulation allows for the verification and optimization of multiple design schemes before physical prototype manufacturing, eliminating the need for repeated prototype testing and reducing prototype manufacturing costs and testing expenses. Based on precise mathematical models and physical laws, finite element analysis ensures the rationality and stability of design parameters. Optimized design schemes can be directly used during mass production, avoiding product performance fluctuations due to design defects. Simultaneously, simulation analysis can identify potential structural risks in advance, such as stress concentration at frame rib nodes and thermal expansion mismatch between the heat-resistant structure and the metal skeleton 1, which can be mitigated through optimized design, improving product reliability under complex operating conditions and reducing the probability of failure during use.

[0042] In a preferred embodiment, step S2 further includes coating the surface of the cavity inside the rudder surface 12 of the metal frame 1 with a high-temperature resistant adhesive.

[0043] Specifically, a layer of high-temperature resistant adhesive, such as high-temperature epoxy adhesive or phenolic resin adhesive, is uniformly coated onto the surface of the cavity inside the rudder surface 12 of the metal frame 1. The coating thickness is determined according to the bonding requirements and the cavity size, typically 0.1-0.5 mm. The coating method can be brushing, spraying, or dipping, ensuring that all surfaces inside the cavity, including the rib surfaces, are covered with the adhesive without omissions or localized accumulation. After coating, within the effective bonding time of the adhesive, the lightweight filler 2 preform is placed into the cavity, and a certain pressure is applied to ensure a tight fit between the preform and the cavity surface. After the adhesive cures, a firm connection is achieved between the lightweight filler 2 and the metal frame 1. The high-temperature resistant adhesive forms a high-strength bonding interface after curing, tightly binding the lightweight filler 2 and the metal frame 1 together, preventing gaps or relative displacement between them. Under high-temperature conditions, the adhesive maintains stable bonding performance, without softening or failing, ensuring the continuity of the insulation layer. Simultaneously, the adhesive fills the tiny gaps between the lightweight filler 2 and the skeleton, further blocking heat conduction paths and enhancing the insulation effect. When the air rudder is subjected to aerodynamic loads or vibrations, the adhesive interface can transmit the force, allowing both to work together and preventing localized stress concentration.

[0044] This design, through the strong bonding interface formed by the high-temperature resistant adhesive, enhances the integrity of the metal skeleton 1, the composite material heat-insulating structure 3, and the lightweight filler 2. Together, they can resist vibration and impact, maintaining relative positional stability even under extreme conditions. This prevents structural failure due to displacement of the lightweight filler 2, thus improving the structural reliability of the aerodynamic rudder. The high-temperature resistant adhesive fills the tiny gaps between the lightweight filler 2 and the metal skeleton 1, forming a continuous heat-insulating interface that blocks convective heat transfer paths. Simultaneously, the adhesive itself possesses certain heat-insulating properties, working synergistically with the lightweight filler 2 to further enhance the overall heat insulation effect.

[0045] In a preferred embodiment, in step S3, the composite material molding process is a resin transfer molding process or an integrated compression molding process.

[0046] Specifically, the resin transfer molding process includes: fixing a metal skeleton 1 containing lightweight filler 2 inside a sealed mold, with the mold having pre-reserved resin injection and venting channels; injecting liquid composite resin, such as chopped fiber reinforced high-temperature resistant resin, into the mold using a pressure pump; under pressure, the resin fills the mold cavity, covering the metal skeleton 1 and lightweight filler 2, while simultaneously venting air from the cavity; after the resin cures, opening the mold, removing the molded air rudder, and completing the manufacturing of the composite heat-resistant structure 3. Pressure injection ensures that the resin can fully penetrate every corner of the mold, forming a tightly bonded interface with the metal skeleton 1 and lightweight filler 2, without gaps or air bubbles, ensuring the integrity of the structure; the cured resin forms a continuous heat-resistant structure, utilizing its high-temperature resistance to resist external high temperatures while transferring aerodynamic loads. Pressure injection allows the thickness tolerance of the heat-resistant structure to be controlled within ±0.1mm, while the integrated molding process ensures a smooth and flat surface without obvious defects, improving the molding accuracy and appearance quality of the product.

[0047] The integrated molding process includes: cutting a composite prepreg, such as chopped fiber reinforced high-temperature resistant resin prepreg, into the shape and size of the heat-resistant structure, wrapping it around a metal skeleton 1 containing lightweight filler 2, and placing them together into a mold of a molding machine; the molding machine applies predetermined pressure and temperature, causing the prepreg to solidify and form within the mold, tightly bonding with the metal skeleton 1 and lightweight filler 2 to form an integrated composite heat-resistant structure 3. Through the synergistic effect of pressure and temperature, the composite prepreg tightly covers the exterior of the internal structure, eliminating air from the material and increasing structural density and strength; molding ensures the shape accuracy of the heat-resistant structure meets aerodynamic requirements, while simultaneously forming a strong bond with internal components, achieving synergistic stress distribution and efficient heat insulation. Pressure ensures a tight fit between the composite material and internal components, enhancing interfacial bonding strength, avoiding the risk of structural delamination and detachment due to weak interfacial bonding, and improving overall load-bearing capacity and structural reliability.

[0048] In a preferred embodiment, the manufacturing method further includes step S3a: performing thermal simulation analysis on the combination of the metal skeleton 1 and the lightweight filler 2 using thermal analysis software, and designing the thickness distribution of the composite heat-resistant structure 3 in different regions of the control surface 12 based on the temperature field and heat flow distribution results obtained from the analysis. Specifically, a three-dimensional thermal simulation model of the combination of the metal skeleton 1 and the lightweight filler 2 is established using thermal analysis software such as ANSYS Icepak or Fluent, and key parameters of the aircraft's thermal environment, such as external airflow temperature, heat flux density, and flight time, are input; the temperature field distribution, heat flow transfer path, and thermal stress distribution of the combination under this thermal environment are simulated and calculated; based on the simulation results, the heat intensity of different regions of the control surface 12 is identified, including the strongest heat in the leading edge region and the weakest heat in the trailing edge region, and the thickness of the composite heat-resistant structure 3 in each region is designed accordingly. The thickness is increased in areas with strong heat to enhance the heat insulation effect, and the thickness is appropriately reduced in areas with weak heat to optimize weight, forming a differentiated thickness distribution scheme to guide the composite material molding process in step S3. Based on the fundamental principles of heat conduction, convection, and radiation, the thermal analysis software accurately simulates the thermal response of the internal assembly under high-temperature external environments, quantifying the temperature gradient and heat flux density in different regions. Simulations clearly define the heating priority of different parts of the control surface 12. For example, the leading edge of the control surface 12 directly faces the high-speed airflow, experiencing the greatest heat intensity and requiring a thicker thermal protection structure; the heating intensity gradually decreases in the middle and trailing edges of the control surface 12, allowing for a more appropriate reduction in thickness. This differentiated thickness design enables the thermal protection structure to achieve weight optimization while meeting insulation requirements, and simultaneously avoiding localized overheating or material waste.

[0049] This manufacturing method allows for precise identification of the heating conditions in each region through thermal simulation analysis. For example, simulations revealed that the heat flux density at the leading edge of the control surface 12 is 3-5 times that at the trailing edge. Based on this, the thickness of the leading edge thermal protection structure is designed to be 1.5-2 times that of the trailing edge, ensuring that each region can effectively withstand the corresponding intensity of high temperature, avoiding structural failure caused by localized overheating, while also avoiding material waste and achieving precise matching of thermal insulation performance. Optimizing the thickness distribution of the thermal protection structure based on the thermal simulation results, and reducing the thickness in areas with weaker heating, can reduce the overall weight of the thermal protection structure, thereby reducing the weight of the air rudder and even the entire aircraft. For example, for the air rudders of large, high-speed aircraft, differentiated thickness design can reduce the weight of the thermal protection structure by 15%-25%, effectively reducing fuel consumption and increasing range. At the same time, lightweight design reduces the driving load on the actuators, making the deflection response of the air rudder faster and enhancing the aircraft's maneuverability and attitude control accuracy.

[0050] It should be understood that the specific embodiments described above are merely illustrative or explanatory of the principles of the invention and do not constitute a limitation thereof. Therefore, any modifications, equivalent substitutions, improvements, etc., made without departing from the spirit and scope of the invention should be included within the protection scope of the invention. Furthermore, the appended claims are intended to cover all variations and modifications falling within the scope and boundaries of the appended claims, or equivalent forms of such scope and boundaries.

Claims

1. A composite thermal protection aerodynamic control system for an aircraft, characterized in that, The composite structure thermal protection air rudder includes a metal frame (1), a lightweight filler (2), and a composite material thermal protection structure (3) arranged sequentially from the inside to the outside. The metal frame (1) includes a rudder shaft (11) and a rudder surface (12). The rudder shaft (11) and the rudder surface (12) are integrally formed. The rudder surface (12) has a frame rib structure (13) inside to form an internal cavity. The lightweight filler (2) is filled in the internal cavity. The composite material heat-resistant structure (3) covers the outside of the rudder surface (12) and the lightweight filler (2), and forms the aerodynamic shape of the air rudder.

2. The composite structure thermal protection air rudder according to claim 1, characterized in that, The rib distribution density of the frame rib structure (13) in the root region where the rudder surface (12) and rudder shaft (11) are connected is greater than that in the tip region of the rudder surface (12).

3. The composite structure thermal protection air rudder according to claim 1, characterized in that, The outer surface of the lightweight filler (2) is flush with the outer edge contour of the rudder surface (12) of the metal frame (1).

4. The composite structure thermal protection air rudder according to claim 1, characterized in that, The metal skeleton (1) is made of aluminum alloy, titanium alloy or high temperature alloy.

5. The composite structure thermal protection air rudder according to claim 1, characterized in that, The material of the lightweight filler (2) is selected from at least one of fiber preform reinforced modified phenolic resin matrix composites, porous ceramics or aerogel insulation materials.

6. A method for manufacturing a composite thermal protection air rudder according to any one of claims 1-5, characterized in that, The manufacturing method includes the following steps: S1: Based on the thermal environment conditions, load conditions and application scenario requirements of the aircraft, determine the structure of the metal frame (1) and the composite material heat protection structure (3). The rudder shaft (11) of the metal frame (1) adopts a solid structure, and the rudder surface (12) of the metal frame (1) adopts a frame rib structure (13). The metal frame (1) with rudder surface (12) and frame rib structure (13) is integrally formed by machining or investment casting process. S2: Based on the thermal environment conditions of the aircraft and the molding process of the composite material heat protection structure (3), determine the lightweight filler (2), and prepare a lightweight filler (2) preform with matching shape according to the shape of the cavity inside the control surface (12) of the metal frame (1), and place the lightweight filler (2) preform into the cavity inside the control surface (12) of the metal frame (1). S3: A composite heat-resistant structure (3) is formed by coating the metal skeleton (1) filled with the lightweight filler (2) using a composite material molding process.

7. The manufacturing method according to claim 6, characterized in that, In step S1, the distribution and size of the frame rib structure (13) are designed based on finite element analysis, and the composite material heat-resistant structure (3) is designed to meet the predetermined stiffness, strength and weight indicators.

8. The manufacturing method according to claim 6, characterized in that, Step S2 further includes coating the surface of the cavity inside the rudder surface (12) of the metal frame (1) with a high-temperature resistant adhesive.

9. The manufacturing method according to claim 6, characterized in that, In step S3, the composite material molding process is either resin transfer molding or integrated compression molding.

10. The manufacturing method according to claim 6, characterized in that, The manufacturing method further includes step S3a: performing thermal simulation analysis on the combination of the metal skeleton (1) and the lightweight filler (2) based on thermal analysis software, and designing the thickness distribution of the composite material heat-resistant structure (3) in different regions of the rudder surface (12) based on the temperature field and heat flow distribution results obtained from the analysis.

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