Aircraft thermal control structure and its welding method

Abstract

This invention provides a thermal control structure for aircraft and its welding method. A layer of thermally conductive fiber-reinforced composite copper foil is applied to the outer surface of the aircraft's titanium alloy skin using a diffusion bonding method, following a designed heat flow direction. This serves as a diffusion channel, transferring heat from areas of high heat flux density to areas of low heat flux density, thereby achieving precise temperature control of critical components of the aircraft under service conditions. This invention integrates the aircraft's load-bearing structure with its thermal control function, while simultaneously solving the challenges of efficient and reliable welding and manufacturing. It can be widely applied to aircraft structural products with high requirements for heat resistance and reliability, and is particularly suitable for aircraft requiring high reusability.

Description

Technical Field

[0001] This invention belongs to the field of welding technology, specifically relating to an aircraft thermal control structure and its welding method. Background Technology

[0002] High-speed and long-endurance service environments require aircraft structures to be lightweight, have high-precision outlines, and withstand high temperatures for extended periods. Traditional structural designs with thermal jackets are no longer sufficient to meet the development requirements for lightweight, high-temperature resistant, and reusable components such as the fuselage, fuel tanks, and control wings. By combining lightweight, high-temperature resistant materials with an integrated heat dissipation structure, the weight of the fuselage, fuel tanks, and control wings can be reduced, and temperature resistance improved, thereby enhancing the performance of next-generation aircraft.

[0003] Currently, semi-active thermal protection typically involves placing heat pipes or other structures in areas requiring thermal protection through welding or direct embedding. If welding is used to connect the heat pipes to the structural surface, the weld joint is prone to cracking due to mismatch in thermal expansion coefficients, which can significantly impact the reliability of subsequent service. Furthermore, this method also has certain technological limitations for complex structural parts.

[0004] Patent CN107538768B provides a method for bonding and assembling a non-rigid thermal protection component to an aircraft. 1. Based on scanning data of the aircraft cabin's outer surface and the thermal protection component's inner surface, as well as the matching of pre-assembled surfaces, data is provided for the adhesive layer thickness during subsequent bonding. 2. The thermal protection component is positioned using a process simulation positioning method. 3. The adhesive layer is prepared using a double-sided air spraying method. 4. Full-area pressure is applied to the thermal protection component through vacuuming and mechanical loading to ensure the aerodynamic shape and bonding quality of the bonded component.

[0005] However, patent CN107538768B requires an additional bonding process. It ensures the assembly surface accuracy of the heat protection component by measuring the adhesive layer thickness in real time, and uses vacuuming and mechanical loading for loading and curing. The operation steps are complicated and cannot guarantee that it meets the usage requirements of its extreme service environment. Summary of the Invention

[0006] In view of the deficiencies in the prior art, the purpose of this invention is to provide a thermal control structure for an aircraft and its welding method.

[0007] A welding method for a thermal control structure of an aircraft according to the present invention includes:

[0008] S1, assembling the aircraft's titanium alloy skin, frame, and brazing filler metal;

[0009] S2, cut the thermally conductive fiber reinforced composite copper foil strip into strips of a certain shape, and spot weld the thermally conductive fiber reinforced composite copper foil strip to the titanium alloy skin of the outer layer of the aircraft structure according to the designed heat flow direction;

[0010] S3, the thermally conductive fiber reinforced composite copper foil strip is assembled with the titanium alloy skin of the outer layer of the aircraft structure and then placed in a vacuum furnace for heating.

[0011] S4, the titanium alloy skin and the thermally conductive fiber reinforced composite copper foil strip are connected by micro-area instantaneous liquid phase diffusion through tooling pressure;

[0012] S5 involves plasma spraying a zirconium oxide coating onto the entire outer surface of the aircraft structure to ensure a smooth final surface.

[0013] S6. After plasma spraying, the entire aircraft structure is placed in a vacuum furnace for annealing to obtain the aircraft thermal control structure.

[0014] Preferably, the titanium alloy of the titanium alloy skin is any one of Ti55, Ti60, Ti65, and Ti2AlNb.

[0015] Preferably, the reinforcing phase of the thermally conductive fiber-reinforced composite copper foil strip is carbon fiber or graphene.

[0016] Preferably, the thickness of the thermally conductive fiber reinforced composite copper foil strip is 0.5 mm to 1 mm.

[0017] Preferably, the instantaneous liquid-phase diffusion bonding process includes: evacuating the furnace to a vacuum level of 1×10⁻⁶. -2 After Pa, heating begins. Once the temperature reaches 890–920°C, a pressure of 0.2–0.5 MPa is applied to the surface of the thermally conductive fiber-reinforced composite copper foil strip. The heat is maintained for 10–20 minutes. After cooling to room temperature in the furnace, the aircraft structure is removed.

[0018] Preferably, the plasma spraying method involves first spraying a zirconium oxide layer of equal thickness to the thermally conductive fiber-reinforced composite copper foil strip onto the area where the thermally conductive fiber-reinforced composite copper foil strip is not added, and then spraying a zirconium oxide layer of 0.5 to 1 mm thickness onto the entire surface until the surface is smooth.

[0019] Preferably, the annealing process includes: after plasma spraying, placing the entire aircraft structure in a vacuum furnace, and annealing it at a temperature of 1×10⁻⁶ mm. -1 Annealing is performed by heating to 500-600℃ under vacuum conditions below Pa and holding for 2-4 hours.

[0020] Preferably, the welding method for the aircraft thermal control structure further includes: during pre-assembly, a layer of graphite paper of equal thickness is pre-placed in the area where the thermally conductive fiber reinforced composite copper foil strip has not been added to level it.

[0021] Preferably, the graphite paper is coated with a solder resist agent.

[0022] The present invention also provides an aircraft thermal control structure, which is manufactured by the aircraft thermal control structure welding method described in any one of the above.

[0023] Compared with the prior art, the present invention has the following beneficial effects:

[0024] 1. This invention addresses the long-term thermal protection requirements of high-heat-flux areas in aircraft components. It utilizes high thermal conductivity materials to rapidly transfer heat from high-heat-flux areas to low-heat-flux areas. Through large-area side heat dissipation on the cabin and deck, severe aerodynamic heat is transferred and released via heat conduction and radiation, thereby reducing the surface temperature of high-heat-flux areas and ensuring the long-term normal operation of metal materials in these areas. This invention proposes utilizing the high thermal conductivity of copper, combining it with a high-temperature resistant titanium alloy skin to form an integrated load-bearing and heat-protection structure with heat flow conduction capabilities.

[0025] 2. This invention leverages the low-melting-point eutectic property of copper and titanium. During brazing, pre-fabricated thermally conductive fiber-reinforced composite copper foil strips are directly diffused and welded onto the substrate using existing processes (temperature, pressure, and time). This achieves a high-strength bond by forming a thin, transient liquid phase, without adding extra welding steps, thus improving structural manufacturing efficiency. Finally, a smooth aerodynamic shape is obtained by covering the entire surface with zirconium oxide coatings of varying thicknesses and using a homogenization annealing method, preventing risks such as subsequent interface cracking. The thermally conductive fiber-reinforced composite copper foil strips used have high thermal conductivity and are lightweight, further reducing weight and improving structural thermal conductivity.

[0026] 3. This invention provides an integrated design method for aircraft load-bearing structure and thermal control function, and solves the problem of efficient and reliable welding manufacturing of the structure. It can be widely used in aircraft structural products with high requirements for heat resistance and reliability, and is especially suitable for aircraft with high requirements for reusability. Attached Figure Description

[0027] Other features, objects, and advantages of the present invention will become more apparent from the following detailed description of non-limiting embodiments with reference to the accompanying drawings:

[0028] Figure 1 A flowchart of the welding method for the thermal control structure of an aircraft;

[0029] Figure 2 This is a schematic diagram of the thermal control structure of an aircraft.

[0030] Figure 3 This is a schematic diagram showing the effect of welding on the thermal control structure of the aircraft cabin in Example 1;

[0031] Figure 4 This is a schematic diagram showing the effect of welding the aircraft rudder-type thermal control structure in Example 1.

[0032] The diagram shows: 1. Titanium alloy skin; 2. Skeleton; 3. Brazing filler metal; 4. Thermally conductive fiber reinforced composite copper foil strip; 5. Graphite paper; 6. Tooling. Detailed Implementation

[0033] The present invention will now be described in detail with reference to specific embodiments. These embodiments will help those skilled in the art to further understand the present invention, but do not limit the invention in any way. It should be noted that those skilled in the art can make several changes and improvements without departing from the concept of the present invention. These all fall within the protection scope of the present invention.

[0034] like Figure 1 As shown, this embodiment provides a welding method for an aircraft thermal control structure, the specific steps of which include:

[0035] S1, assemble the titanium alloy skin 1, frame 2 and brazing filler metal 3 of the aircraft;

[0036] S2, cut the thermally conductive fiber reinforced composite copper foil strip 4 into strips with a certain shape, and spot weld the thermally conductive fiber reinforced composite copper foil strip 4 to the titanium alloy skin 1 of the outer layer of the aircraft structure according to the designed heat flow direction.

[0037] S3, the thermally conductive fiber reinforced composite copper foil strip 4 is assembled with the titanium alloy skin 1 of the outer layer of the aircraft structure and then placed in a vacuum furnace for heating.

[0038] S4, the titanium alloy skin 1 and the thermally conductive fiber reinforced composite copper foil strip 4 are connected by micro-area instantaneous liquid phase diffusion through tooling pressure;

[0039] S5 involves plasma spraying a zirconium oxide coating onto the entire outer surface of the aircraft structure to ensure a smooth final surface.

[0040] S6. After plasma spraying, the entire aircraft structure is placed in a vacuum furnace for annealing to obtain the aircraft thermal control structure.

[0041] In one embodiment, the titanium alloy of the titanium alloy skin 1 is any one of Ti55, Ti60, Ti65, and Ti2AlNb.

[0042] In one embodiment, the reinforcing phase of the thermally conductive fiber-reinforced composite copper foil strip 4 is carbon fiber or graphene.

[0043] In one embodiment, the thickness of the thermally conductive fiber reinforced composite copper foil strip 4 is 0.5 mm to 1 mm.

[0044] In one embodiment, the transient liquid-phase diffusion bonding process includes: evacuating the furnace to a vacuum level of 1×10⁻⁶. -2 After Pa, heating begins. Once the temperature reaches 890–920°C, a pressure of 0.2–0.5 MPa is applied to the surface of the thermally conductive fiber-reinforced composite copper foil strip 4. The heat is maintained for 10–20 minutes. After cooling to room temperature in the furnace, the aircraft structure is removed.

[0045] In one embodiment, the plasma spraying method involves first spraying a zirconium oxide layer of equal thickness to the thermally conductive fiber-reinforced composite copper foil strip 4 onto the area where the thermally conductive fiber-reinforced composite copper foil strip 4 is not added, and then spraying a zirconium oxide layer of 0.5 to 1 mm thickness onto the entire surface until the surface is smooth.

[0046] In one embodiment, the annealing process includes: after plasma spraying, placing the entire aircraft structure in a vacuum furnace, and annealing it at a temperature of 1×10⁻⁶ mm. -1 Annealing is performed by heating to 500-600℃ under vacuum conditions below Pa and holding for 2-4 hours.

[0047] In one embodiment, the welding method for the aircraft thermal control structure further includes: during pre-assembly, a layer of graphite paper 5 of equal thickness is pre-placed in the area where the thermally conductive fiber reinforced composite copper foil strip 4 has not been added to level it.

[0048] In one embodiment, the graphite paper 5 is coated with a solder resist agent.

[0049] like Figure 2 As shown, the aircraft thermal control structure can be manufactured on tooling 6 using the method provided in this embodiment.

[0050] This method enables the application of a layer of thermally conductive fiber-reinforced composite copper foil strip 4 on the outer surface of the aircraft's titanium alloy skin 1, which is connected by a diffusion method according to the designed heat flow direction. This serves as a diffusion channel to transfer heat from local high heat flux density areas to low heat flux density areas, thereby achieving precise temperature control of critical parts of the aircraft under service conditions.

[0051] The welding method for the thermal control structure of aircraft provided in this embodiment integrates the aircraft's load-bearing structure with its thermal control function. It also solves the problem of efficient and reliable welding manufacturing of this structure. It can be widely used in aircraft structural products with high requirements for heat resistance and reliability, and is especially suitable for aircraft with high requirements for reusability.

[0052] Example 1

[0053] First, the titanium alloy skin 1, frame 2, and brazing filler metal 3 of the aircraft are assembled. Then, the thermally conductive fiber reinforced composite copper foil strip 4 is cut into strips of a certain shape and spot-welded to the outer titanium alloy skin 1 of the aircraft cabin or rudder structure according to the designed heat flow direction.

[0054] The titanium alloy skin 1 is made of Ti55 high-temperature resistant titanium alloy. The thickness of the thermally conductive fiber reinforced composite copper foil strip 4 is 0.5mm. The heat flow direction covers both the high-temperature and low-temperature areas of the product. The thickness, width, and distribution of the thermally conductive fiber reinforced composite copper foil strip 4 are designed according to the required heat flow. In areas where the thermally conductive fiber reinforced composite copper foil strip 4 is not added, a layer of graphite paper 5 of equal thickness is pre-placed to level the surface. Solder resist is sprayed onto the graphite paper 5 to prevent adhesion. After the entire assembly with the tooling 6 is completed, it is placed in a vacuum furnace, and the furnace is evacuated to a vacuum level of 1×10⁻⁶. -2 Heating begins after Pa, and once the temperature reaches 890℃, a pressure of 0.2MPa is applied to the surface of the thermally conductive fiber-reinforced composite copper foil strip 4. The surface is held at this temperature for 10 minutes, and then cooled to room temperature in the furnace before removal. Finally, a zirconium oxide coating is plasma-sprayed onto the entire outer surface. First, a zirconium oxide layer of equal thickness to the thermally conductive fiber-reinforced composite copper foil strip 4 is sprayed onto the areas without it, followed by a 0.5mm thick zirconium oxide layer to ensure a smooth final surface. After plasma spraying, the product is placed in a vacuum furnace at 1×10⁻⁶. -1 Annealing treatment is performed by heating to 500℃ and holding at that temperature for 2 hours below a vacuum level (Pa), resulting in the aforementioned aircraft structure with thermal control function. Figure 3 as well as Figure 4 As shown.

[0055] Example 2

[0056] First, the titanium alloy skin 1, frame 2, and brazing filler metal 3 of the aircraft are assembled. Thermally conductive fiber reinforced composite copper foil strip 4 is cut into strips of a specific shape. Following the designed heat flow direction, the thermally conductive fiber reinforced composite copper foil strip 4 is spot-welded to the outer titanium alloy skin 1 of the aircraft cabin or rudder structure. The titanium alloy skin 1 is Ti60 high-temperature resistant titanium alloy. The thickness of the thermally conductive fiber reinforced composite copper foil strip 4 is 0.75mm, and the heat flow direction covers both the high-temperature and low-temperature areas of the product. The thickness, width, and distribution of the thermally conductive fiber reinforced composite copper foil strip 4 are designed according to the required heat flow. A layer of graphite paper 5 of equal thickness is pre-placed in areas where the thermally conductive fiber reinforced composite copper foil strip 4 is not added to level the surface. Solder resist is sprayed onto the graphite paper 5 to prevent adhesion. After the entire assembly with the tooling 6 is completed, it is placed in a vacuum furnace, and the furnace is evacuated to a vacuum level of 1×10⁻⁶. -2Heating begins after Pa, and once the temperature reaches 900℃, a pressure of 0.3MPa is applied to the surface of the thermally conductive fiber-reinforced composite copper foil strip 4. The surface is held at this temperature for 15 minutes, and then cooled to room temperature in the furnace before removal. Finally, a zirconium oxide coating is plasma-sprayed onto the entire outer surface. First, a zirconium oxide layer of equal thickness to the thermally conductive fiber-reinforced composite copper foil strip 4 is sprayed onto the areas without it, followed by a 0.75mm thick zirconium oxide layer to ensure a smooth final surface. After plasma spraying, the product is placed in a vacuum furnace at 1×10⁻⁶. -1 The aircraft structure with thermal control function is obtained by annealing the aircraft by heating it to 550°C and holding it at a vacuum of less than Pa for 3 hours.

[0057] Example 3

[0058] First, the titanium alloy skin 1, frame 2, and brazing filler metal 3 of the aircraft are assembled. Thermally conductive fiber reinforced composite copper foil strip 4 is cut into strips of a specific shape. Following the designed heat flow direction, the thermally conductive fiber reinforced composite copper foil strip 4 is spot-welded to the outer titanium alloy skin 1 of the aircraft cabin or rudder structure. The titanium alloy is Ti65 high-temperature resistant titanium alloy. The thickness of the thermally conductive fiber reinforced composite copper foil strip 4 is 0.75mm, and the heat flow direction covers both the high-temperature and low-temperature areas of the product. The thickness, width, and distribution of the thermally conductive fiber reinforced composite copper foil strip 4 are designed according to the required heat flow. A layer of graphite paper 5 of equal thickness is pre-placed in areas where the thermally conductive fiber reinforced composite copper foil strip 4 is not added to level the surface. Solder resist is sprayed onto the graphite paper 5 to prevent adhesion. After the entire assembly with the tooling 6 is completed, it is placed in a vacuum furnace, and the furnace is evacuated to a vacuum level of 1×10⁻⁶. -2 Heating begins after Pa, and once the temperature reaches 910℃, a pressure of 0.4MPa is applied to the surface of the thermally conductive fiber-reinforced composite copper foil strip 4. The surface is held at this temperature for 15 minutes, and then cooled to room temperature in the furnace before removal. Finally, a zirconium oxide coating is plasma-sprayed onto the entire outer surface. First, a zirconium oxide layer of equal thickness to the thermally conductive fiber-reinforced composite copper foil strip 4 is sprayed onto the areas without it, followed by a 0.75mm thick zirconium oxide layer to ensure a smooth final surface. After plasma spraying, the product is placed in a vacuum furnace at 1×10⁻⁶. -1 The aircraft structure with thermal control function is obtained by annealing the aircraft by heating it to 550°C and holding it at a vacuum of less than Pa for 3 hours.

[0059] Example 4

[0060] First, the titanium alloy skin 1, frame 2, and brazing filler metal 3 of the aircraft are assembled. Thermally conductive fiber reinforced composite copper foil strip 4 is cut into strips of a specific shape. Following the designed heat flow direction, the thermally conductive fiber reinforced composite copper foil strip 4 is spot-welded to the outer titanium alloy skin 1 of the aircraft cabin or rudder structure. The titanium alloy is a Ti2AlNb high-temperature resistant titanium alloy. The thickness of the thermally conductive fiber reinforced composite copper foil strip 4 is 1mm, and the heat flow direction covers both the high-temperature and low-temperature areas of the product. The thickness, width, and distribution of the thermally conductive fiber reinforced composite copper foil strip 4 are designed according to the required heat flow. A layer of graphite paper 5 of equal thickness is pre-placed in areas where the thermally conductive fiber reinforced composite copper foil strip 4 is not added to level the surface. Solder resist is sprayed onto the graphite paper 5 to prevent adhesion. After the entire assembly with the tooling 6 is completed, it is placed in a vacuum furnace, and the furnace is evacuated to a vacuum level of 1×10⁻⁶. -2 Heating begins after Pa, and once the temperature reaches 920℃, a pressure of 0.5MPa is applied to the surface of the thermally conductive fiber-reinforced composite copper foil strip 4. The surface is held at this temperature for 20 minutes, and then cooled to room temperature in the furnace before removal. Finally, a zirconium oxide coating is plasma-sprayed onto the entire outer surface. First, a zirconium oxide layer of equal thickness to the thermally conductive fiber-reinforced composite copper foil strip 4 is sprayed onto the areas without it, followed by a 1mm thick zirconium oxide coating to ensure a smooth final surface. After plasma spraying, the product is placed in a vacuum furnace and heated at 1×10⁻⁶. -1 The aircraft structure with thermal control function is obtained by annealing at a vacuum level below Pa to 600°C and holding for 4 hours.

[0061] In the description of this application, it should be understood that the terms "upper", "lower", "front", "back", "left", "right", "vertical", "horizontal", "top", "bottom", "inner", "outer", etc., indicate the orientation or positional relationship based on the orientation or positional relationship shown in the accompanying drawings. They are only for the convenience of describing this application and simplifying the description, and do not indicate or imply that the device or element referred to must have a specific orientation, or be constructed and operated in a specific orientation. Therefore, they should not be construed as limitations on this application.

[0062] Specific embodiments of the present invention have been described above. It should be understood that the present invention is not limited to the specific embodiments described above, and those skilled in the art can make various changes or modifications within the scope of the claims, which do not affect the essence of the present invention. Unless otherwise specified, the embodiments and features described in this application can be arbitrarily combined with each other.

Claims

1. A welding method for a thermal control structure of an aircraft, characterized in that, include: S1, assemble the titanium alloy skin (1), frame (2) and brazing filler (3) of the aircraft; S2, cut the thermally conductive fiber reinforced composite copper foil strip (4) into strips with a certain shape, and spot weld the thermally conductive fiber reinforced composite copper foil strip (4) to the titanium alloy skin (1) of the outer layer of the aircraft structure according to the designed heat flow direction; S3, the thermally conductive fiber reinforced composite copper foil strip (4) is assembled with the titanium alloy skin (1) of the outer layer of the aircraft structure to form an integral whole, the whole is assembled with the tooling (6), and after the assembly is completed, it is placed in a vacuum furnace for heating. S4, the titanium alloy skin (1) and the thermally conductive fiber reinforced composite copper foil strip (4) are connected by micro-area instantaneous liquid phase diffusion through tooling pressure; S5 involves plasma spraying a zirconium oxide coating onto the entire outer surface of the aircraft structure to ensure a smooth final surface. S6. After plasma spraying, the entire aircraft structure is placed in a vacuum furnace for annealing to obtain the aircraft thermal control structure.

2. The welding method for the thermal control structure of an aircraft according to claim 1, characterized in that, The titanium alloy of the titanium alloy skin (1) is any one of Ti55, Ti60, Ti65, and Ti2AlNb.

3. The welding method for the thermal control structure of an aircraft according to claim 1, characterized in that, The reinforcing phase of the thermally conductive fiber-reinforced composite copper foil strip (4) is carbon fiber or graphene.

4. The welding method for the thermal control structure of an aircraft according to claim 1, characterized in that, The thickness of the thermally conductive fiber reinforced composite copper foil strip (4) is 0.5 mm to 1 mm.

5. The welding method for the thermal control structure of an aircraft according to claim 1, characterized in that, The micro-region instantaneous liquid-phase diffusion bonding process includes: evacuating the furnace to a vacuum level of 1×10⁻⁶. -2 After Pa, heating begins. When the temperature reaches 890-920℃, a pressure of 0.2-0.5MPa is applied to the surface of the thermally conductive fiber reinforced composite copper foil strip (4). The heat preservation time is 10-20min. After cooling to room temperature with the furnace, the aircraft structure is taken out.

6. The welding method for the thermal control structure of an aircraft according to claim 1, characterized in that, The plasma spraying method involves first spraying a zirconium oxide layer of equal thickness to the thermally conductive fiber-reinforced composite copper foil strip (4) onto the area where the thermally conductive fiber-reinforced composite copper foil strip (4) has not been added, and then spraying a zirconium oxide layer of 0.5 to 1 mm thickness onto the entire surface until the surface is smooth.

7. The welding method for the thermal control structure of an aircraft according to claim 1, characterized in that, The annealing process includes: after plasma spraying, placing the entire aircraft structure in a vacuum furnace, and annealing it at a temperature of 1×10⁻⁶. -1 Annealing is performed by heating to 500-600℃ under vacuum conditions below Pa and holding for 2-4 hours.

8. The welding method for the thermal control structure of an aircraft according to claim 1, characterized in that, The welding method for the thermal control structure of the aircraft also includes: during pre-assembly, a layer of graphite paper (5) of equal thickness is pre-placed in the area where the thermally conductive fiber reinforced composite copper foil strip (4) has not been added to level it.

9. The welding method for the thermal control structure of an aircraft according to claim 8, characterized in that, The graphite paper (5) is coated with a solder resist agent.

10. A thermal control structure for an aircraft, characterized in that, It is manufactured by the aircraft thermal control structure welding method described in any one of claims 1 to 9.

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