Aircraft thermal control structure and welding method thereof

Abstract

According to the aircraft thermal control structure and the welding method thereof, the outer surface of an aircraft titanium alloy skin is covered with a layer of heat conduction fiber reinforced composite copper foil belt through a diffusion bonding method according to the designed heat flow trend, the heat conduction fiber reinforced composite copper foil belt serves as a diffusion channel, and heat in a local high-heat-flow-density area is conducted to a low-heat-flow-density area; therefore, accurate temperature control of the key part of the aircraft in a service environment is realized. The aircraft bearing structure and the thermal control function are integrated, meanwhile, the problem of efficient and high-reliability welding and manufacturing of the structure is solved, and the structure can be widely applied to aircraft structure products with high requirements for heat resistance and reliability and is particularly suitable for aircrafts with high requirements for repeated use.

Description

TECHNICAL FIELD

[0001] The present application belongs to the technical field of welding, and particularly relates to an aircraft thermal control structure and a welding method thereof. BACKGROUND

[0002] High-speed and long-endurance service environments require aircraft structures to be lightweight, have high contour accuracy, and be resistant to high temperatures for a long time. The traditional structure plus heat shield scheme has been unable to meet the development requirements of lightweight, high-temperature resistance, and reusability of cabin bodies, fuel tanks, rudders, and wings. By combining lightweight high-temperature-resistant materials with heat dissipation integrated structures, the weight of cabin bodies, fuel tanks, and rudder structures can be reduced, the temperature resistance performance can be improved, and the efficiency of the new generation of aircraft can be improved.

[0003] Currently, semi-active heat protection structures are usually placed in parts requiring heat protection by welding or directly embedding heat pipes. If heat pipes are connected to the surface of the structure by welding, cracks in the welding joint are likely to occur due to the mismatch of the thermal expansion coefficients, which greatly affects the subsequent service reliability. In addition, for complex structural parts, the use of this method also has certain process limitations.

[0004] Patent CN107538768B provides a method for bonding and assembling a non-rigid heat protection member and an aircraft. 1. According to the scanning data of the aircraft cabin body contour surface and the heat protection member inner surface and the pre-assembly surface matching, data basis is provided for the thickness of the adhesive layer during subsequent bonding; 2. The heat protection member is positioned by using a process simulation part; 3. The adhesive layer is prepared by using the air spraying method of double-sided glue spraying; 4. The full-area pressurization of the heat protection member is realized by vacuumizing and mechanically loading, and the aerodynamic contour and bonding quality of the bonded heat protection member are ensured.

[0005] However, patent CN107538768B needs to add an additional bonding process to ensure the assembly surface accuracy of the heat protection member by measuring the thickness of the adhesive layer in real time, and to load and solidify by vacuumizing and mechanically loading, which is complex in operation steps and cannot guarantee that it meets the use requirements of extreme service environments. SUMMARY

[0006] In view of the defects in the prior art, the purpose of the present application is to provide an aircraft thermal control structure and a welding method thereof.

[0007] According to the aircraft thermal control structure welding method provided by the present application, the following steps are included: S1, assembling the titanium alloy skin, framework and filler material of the aircraft; S2, cutting the heat-conducting fiber-reinforced composite copper foil tape into a tape with a certain shape, and point-welding and assembling the heat-conducting fiber-reinforced composite copper foil tape and the titanium alloy skin of the outer layer of the aircraft structure according to the designed heat flow direction. S3, the heat-conducting fiber-reinforced composite copper foil tape is placed in a vacuum furnace after being assembled with the titanium alloy skin of the outer layer of the aircraft structure; S4, the titanium alloy skin and the heat-conducting fiber-reinforced composite copper foil tape are micro-regionally and instantaneously connected by liquid-phase diffusion through pressure of a tool; S5, the outer surface of the aircraft structure is entirely plasma sprayed with a zirconia coating layer, so as to ensure that the final outer shape is flat; S6, after the plasma spraying, the aircraft structure is placed in a vacuum furnace, and annealing treatment is performed on the product, so as to obtain the aircraft thermal control structure.

[0008] Preferably, the titanium alloy of the titanium alloy skin is any one of Ti55, Ti60, Ti65 and Ti2AlNb.

[0009] Preferably, the reinforcing phase of the heat-conducting fiber-reinforced composite copper foil tape is carbon fiber or graphene.

[0010] Preferably, the thickness of the heat-conducting fiber-reinforced composite copper foil tape is 0.5 mm to 1 mm.

[0011] Preferably, the instant liquid-phase diffusion connection process comprises the following steps: after the vacuum degree in the furnace reaches 1x10 -2 Pa, heating is started, 0.2 to 0.5 MPa pressure is applied to the surface of the heat-conducting fiber-reinforced composite copper foil tape when the temperature reaches 890 to 920 ℃, the heat-conducting fiber-reinforced composite copper foil tape is kept for 10 to 20 min, and the aircraft structure is taken out after the furnace is cooled to room temperature.

[0012] Preferably, the plasma spraying method comprises the following steps: a zirconia layer with a thickness equal to that of the heat-conducting fiber-reinforced composite copper foil tape is first sprayed in a region without the heat-conducting fiber-reinforced composite copper foil tape, and then a 0.5 to 1 mm thick zirconia layer is entirely sprayed, so that the surface is entirely smooth.

[0013] Preferably, the annealing treatment process comprises the following steps: after the plasma spraying, the aircraft structure is placed in a vacuum furnace, and heating is performed to 500 to 600 ℃ under a vacuum degree of 1x10 -1 Pa for 2 to 4 h for annealing treatment.

[0014] Preferably, the aircraft thermal control structure welding method further comprises the following step: a layer of graphite paper pad with the same thickness is pre-arranged in a region without the heat-conducting fiber-reinforced composite copper foil tape during assembly before welding.

[0015] Preferably, the graphite paper is sprayed with a welding inhibitor.

[0016] The application further provides an aircraft thermal control structure manufactured by the aircraft thermal control structure welding method.

[0017] Compared with the prior art, the present application has the following beneficial effects: 1. The present application is aimed at the long-time thermal protection requirement of local high heat flow area of aircraft components, uses high thermal conductivity material to quickly transfer the heat of the high heat flow area to the low heat flow area, and releases the severe aerodynamic heat in the form of heat conduction and radiation by means of the large-surface side heat dissipation of the cabin body and cabin surface, so as to reduce the surface temperature of the high heat flow area and ensure that the metal material of the high temperature area can be normally used for a long time. The present application proposes to use the high thermal conductivity of copper to form a load-bearing and heat-proof integrated structure with the diffusion bonding of the high-temperature-resistant titanium alloy skin.

[0018] 2. Based on the characteristics of copper and titanium forming a low melting point eutectic, the present application uses the original process (temperature, pressure, time) to directly diffusion weld the prefabricated heat-conducting fiber reinforced composite copper foil tape in the brazing process, achieves the effect of high-strength connection through the formation of a thin layer of transient liquid phase, and at the same time does not increase the additional welding process, thereby improving the structure manufacturing efficiency. Finally, a smooth aerodynamic shape is obtained by covering the whole with different thicknesses of zirconium oxide coating and homogenizing annealing method, and the risk of cracking of the subsequent bonding interface is prevented. The heat-conducting fiber reinforced composite copper foil tape used has the characteristics of high heat conduction efficiency and light weight, which can further reduce the weight and improve the structure heat conduction efficiency.

[0019] 3. The present application provides an integrated design method of aircraft load-bearing structure and thermal control function, which simultaneously solves the problems of high-efficiency and high-reliability welding manufacturing of the structure, and can be widely applied to aircraft structure products with high requirements for heat resistance and reliability, especially to aircraft with high requirements for reusability. BRIEF DESCRIPTION OF DRAWINGS

[0020] Other features, objects and advantages of the present application will become more apparent from the following detailed description of non-limiting embodiments with reference to the following drawings: Figure 1 The flow chart of the welding method for the aircraft thermal control structure; Figure 2 The structural schematic diagram of the aircraft thermal control structure; Figure 3 The effect schematic diagram of the aircraft cabin body type thermal control structure after welding in embodiment 1; Figure 4 The effect schematic diagram of the aircraft rudder wing type thermal control structure after welding in embodiment 1.

[0021] The figure shows: titanium alloy skin 1, framework 2, brazing filler metal 3, heat-conducting fiber reinforced composite copper foil tape 4, graphite paper 5, tooling 6. DETAILED DESCRIPTION

[0022] 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.

[0023] like Figure 1 As shown, this embodiment provides a welding method for an aircraft thermal control structure, the specific steps of which include: S1, assemble the titanium alloy skin 1, frame 2 and brazing filler metal 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 and then 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.

[0024] In one embodiment, the titanium alloy of the titanium alloy skin 1 is any one of Ti55, Ti60, Ti65, and Ti2AlNb.

[0025] In one embodiment, the reinforcing phase of the thermally conductive fiber-reinforced composite copper foil strip 4 is carbon fiber or graphene.

[0026] In one embodiment, the thickness of the thermally conductive fiber reinforced composite copper foil strip 4 is 0.5 mm to 1 mm.

[0027] 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.

[0028] In an embodiment, the plasma spraying method is to spray a layer of zirconia with a thickness equal to that of the thermally conductive fiber-reinforced composite copper foil tape 4 in the area where the thermally conductive fiber-reinforced composite copper foil tape 4 is not added, and then spray a layer of zirconia with a thickness of 0.5-1 mm, so that the surface is smooth.

[0029] In an embodiment, the annealing process includes: after plasma spraying, the aircraft structure is placed in a vacuum furnace as a whole, heated to 500-600℃ under a vacuum degree of 1x10 -1 Pa, and annealed by being heated to 500-600℃ under a vacuum degree of 1x10

[0030] In an embodiment, the aircraft thermal control structure welding method further includes: during assembly before welding, a layer of graphite paper 5 with the same thickness is pre-positioned as a pad in the area where the thermally conductive fiber-reinforced composite copper foil tape 4 is not added.

[0031] In an embodiment, the graphite paper 5 is sprayed with a welding inhibitor.

[0032] As shown in Figure 2 , the aircraft thermal control structure can be obtained by manufacturing the aircraft thermal control structure on the tooling 6 through the method provided in the embodiment.

[0033] The method realizes that a layer of thermally conductive fiber-reinforced composite copper foil tape 4 is covered on the outer surface of the aircraft titanium alloy skin 1 according to the designed heat flow direction through the diffusion bonding method, which serves as a diffusion channel to conduct heat from the local high heat flux density area to the low heat flux density area, so as to realize precise temperature control of the aircraft key parts under the service environment.

[0034] The aircraft thermal control structure welding method provided in the embodiment realizes the integration of the aircraft load-bearing structure and the thermal control function, and at the same time solves the problem of efficient and reliable welding manufacturing of the structure, and can be widely applied to aircraft structure products with high heat resistance and reliability requirements, and is especially suitable for aircraft with high reusability requirements.

[0035] Embodiment 1 First, the titanium alloy skin 1, the framework 2, and the filler material 3 of the aircraft are assembled, the thermally conductive fiber-reinforced composite copper foil tape 4 is cut into a tape with a certain shape, the thermally conductive fiber-reinforced composite copper foil tape 4 is point-welded and assembled with the outer titanium alloy skin 1 of the aircraft cabin body or the rudder wing structure according to the designed heat flow direction, The material of the titanium alloy skin 1 is Ti55 high-temperature-resistant titanium alloy, the thickness of the heat-conducting fiber-reinforced composite copper foil strip 4 is 0.5 mm, the heat flow direction covers the high-temperature area and the low-temperature area of the product, and the thickness, width and distribution position of the heat-conducting fiber-reinforced composite copper foil strip 4 are designed according to the heat flow to be guided, a layer of graphite paper 5 with the same thickness is prearranged in the area without the heat-conducting fiber-reinforced composite copper foil strip 4, and a soldering inhibitor is sprayed on the graphite paper 5 to prevent adhesion. After the whole assembly is completed, the assembly is placed in a vacuum furnace, the furnace is vacuumized to 1x10 -2 Pa, heating is started, after the temperature reaches 890℃, 0.2 MPa pressure is applied to the surface of the heat-conducting fiber-reinforced composite copper foil strip 4, the holding time is 10 min, the product is taken out after the furnace is cooled to room temperature. Finally, a zirconia coating is plasma sprayed on the outer surface, a layer of zirconia with the same thickness as the heat-conducting fiber-reinforced composite copper foil strip 4 is first sprayed in the area without the heat-conducting fiber-reinforced composite copper foil strip 4, and then a 0.5 mm thick zirconia layer is sprayed on the whole, to ensure the final shape surface is flat. After plasma spraying, the product is placed in a vacuum furnace, heated to 500℃ under a vacuum of 1x10 -1 Pa, and annealed for 2 h to obtain the aircraft structure with the heat control function. As shown in Figure 3 and Figure 4 .

[0036] Example 2 First, the titanium alloy skin 1, the framework 2 and the filler metal 3 of the aircraft are assembled, the heat-conducting fiber-reinforced composite copper foil strip 4 is cut into a strip with a certain shape, the heat-conducting fiber-reinforced composite copper foil strip 4 is spot-welded and assembled with the outer titanium alloy skin 1 of the aircraft cabin or rudder wing structure according to the designed heat flow direction, the titanium alloy skin 1 is Ti60 high-temperature-resistant titanium alloy, the thickness of the heat-conducting fiber-reinforced composite copper foil strip 4 is 0.75 mm, the heat flow direction covers the high-temperature area and the low-temperature area of the product, and the thickness, width and distribution position of the heat-conducting fiber-reinforced composite copper foil strip 4 are designed according to the heat flow to be guided, a layer of graphite paper 5 with the same thickness is prearranged in the area without the heat-conducting fiber-reinforced composite copper foil strip 4, and a soldering inhibitor is sprayed on the graphite paper 5 to prevent adhesion. After the whole assembly is completed, the assembly is placed in a vacuum furnace, the furnace is vacuumized to 1x10 -2 Pa, heating is started, after the temperature reaches 900℃, 0.3 MPa pressure is applied to the surface of the heat-conducting fiber-reinforced composite copper foil strip 4, the holding time is 15 min, the product is taken out after the furnace is cooled to room temperature. Finally, a zirconia coating is plasma sprayed on the outer surface, a layer of zirconia with the same thickness as the heat-conducting fiber-reinforced composite copper foil strip 4 is first sprayed in the area without the heat-conducting fiber-reinforced composite copper foil strip 4, and then a 0.75 mm thick zirconia layer is sprayed on the whole, to ensure the final shape surface is flat. After plasma spraying, the product is placed in a vacuum furnace, heated to 500℃ under a vacuum of 1x10 -1Pa below the vacuum degree to 550℃ heating 3h annealing treatment, get the aircraft structure with heat control function.

[0037] Example 3 First, the aircraft titanium alloy skin 1, framework 2 and filler metal 3 are assembled, the heat-conducting fiber reinforced composite copper foil strip 4 is cut into a certain shape, and the heat-conducting fiber reinforced composite copper foil strip 4 is spot-welded and assembled with the outer titanium alloy skin 1 of the aircraft cabin body or rudder wing structure according to the designed heat flow direction. The titanium alloy is Ti65 high-temperature titanium alloy, the thickness of the heat-conducting fiber reinforced composite copper foil strip 4 is 0.75mm, the heat flow direction covers the high-temperature area and the low-temperature area of the product, and the thickness, width and distribution position of the heat-conducting fiber reinforced composite copper foil strip 4 are designed according to the required heat flow. A layer of graphite paper 5 with the same thickness is pre-installed in the area without the heat-conducting fiber reinforced composite copper foil strip 4 to pad flat, and a soldering inhibitor is sprayed on the graphite paper 5 to prevent adhesion. After the whole assembly is completed, it is placed in a vacuum furnace, the furnace is evacuated to 1×10 -2 Pa, and then heating starts. After the temperature reaches 910℃, 0.4MPa pressure is applied to the surface of the heat-conducting fiber reinforced composite copper foil strip 4, the holding time is 15min, and the product is taken out after the furnace cools to room temperature. Finally, a layer of zirconia coating is plasma sprayed on the outer surface, a layer of zirconia with the same thickness as the heat-conducting fiber reinforced composite copper foil strip 4 is first sprayed in the area without the heat-conducting fiber reinforced composite copper foil strip 4, and then a layer of 0.75mm thick zirconia is sprayed on the whole to ensure the final outer surface is flat. After plasma spraying, the product is placed in a vacuum furnace, heated to 550℃ under a vacuum degree of 1×10 -1 Pa, and annealed for 3h to obtain the aircraft structure with heat control function.

[0038] Example 4 First, the aircraft titanium alloy skin 1, framework 2 and filler metal 3 are assembled, the heat-conducting fiber reinforced composite copper foil strip 4 is cut into a certain shape, and the heat-conducting fiber reinforced composite copper foil strip 4 is spot-welded and assembled with the outer titanium alloy skin 1 of the aircraft cabin body or rudder wing structure according to the designed heat flow direction. The titanium alloy is Ti2AlNb high-temperature titanium alloy, the thickness of the heat-conducting fiber reinforced composite copper foil strip 4 is 1mm, the heat flow direction covers the high-temperature area and the low-temperature area of the product, and the thickness, width and distribution position of the heat-conducting fiber reinforced composite copper foil strip 4 are designed according to the required heat flow. A layer of graphite paper 5 with the same thickness is pre-installed in the area without the heat-conducting fiber reinforced composite copper foil strip 4 to pad flat, and a soldering inhibitor is sprayed on the graphite paper 5 to prevent adhesion. After the whole assembly is completed, it is placed in a vacuum furnace, the furnace is evacuated to 1×10 -2After the heating starts, when the temperature reaches 920℃, 0.5MPa pressure is applied to the surface of the heat-conductive fiber reinforced composite copper foil tape 4, and the holding time is 20min. The product is taken out after cooling to room temperature in the furnace. Finally, the outer surface is plasma sprayed with a zirconia coating. A layer of zirconia with a thickness equal to that of the heat-conductive fiber reinforced composite copper foil tape 4 is first sprayed in the area without the heat-conductive fiber reinforced composite copper foil tape 4, and then a 1mm-thick zirconia layer is sprayed overall to ensure the final outer surface is flat. After plasma spraying, the product is placed in a vacuum furnace and heated to 600℃ under a vacuum of 1x10 -1 Pa to perform annealing treatment for 4h, obtaining the aircraft structure with heat control function.

[0039] In the description of the present application, it should be understood that the terms "upper", "lower", "front", "rear", "left", "right", "vertical", "horizontal", "top", "bottom", "inner", "outer" and the like indicate the orientation or positional relationship shown in the drawings, and are only for the convenience of describing the present application and simplifying the description, and do not indicate or imply that the device or element referred to must have a particular orientation, be constructed and operated in a particular orientation, and therefore cannot be understood as limiting the present application.

[0040] The specific embodiments of the present application are described above. It should be understood that the present application 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 does not affect the essential content of the present application. The embodiments of the present application and the features in the embodiments can be arbitrarily combined with each other without conflict.

Claims

1. A method of welding a thermal control structure for an aircraft, characterized in that, The method comprises the following steps: S1, assembling the titanium alloy skin (1), the framework (2) and the filler (3) of the aircraft; S2, cutting the heat-conducting fiber-reinforced composite copper foil tape (4) into a tape with a certain shape, and point-welding and assembling the heat-conducting fiber-reinforced composite copper foil tape (4) and the titanium alloy skin (1) of the outer layer of the aircraft structure according to the designed heat flow direction; S3, placing the heat-conducting fiber-reinforced composite copper foil tape (4) and the titanium alloy skin (1) of the outer layer of the aircraft structure in a vacuum furnace after assembly; S4, performing micro-zone transient liquid-phase diffusion connection on the titanium alloy skin (1) and the heat-conducting fiber-reinforced composite copper foil tape (4) through tooling pressure; S5, integrally plasma spraying a layer of zirconia coating on the outer surface of the aircraft structure to ensure the smoothness of the final outer shape; S6, placing the aircraft structure in a vacuum furnace after plasma spraying to perform annealing treatment on the aircraft structure, thereby obtaining the aircraft thermal control structure.

2. The aircraft thermal control structure welding method of claim 1, wherein, The titanium alloy of the titanium alloy skin (1) is any one of Ti55, Ti60, Ti65 and Ti2AlNb.

3. The aircraft thermal control structure welding method of claim 1, wherein, The reinforcing phase of the heat-conducting fiber-reinforced composite copper foil tape (4) is carbon fiber or graphene.

4. The aircraft thermal control structure welding method of claim 1, wherein The thickness of the heat-conducting fiber-reinforced composite copper foil tape (4) is 0.5mm-1mm.

5. The aircraft thermal control structure welding method of claim 1, wherein, The micro-region transient liquid phase diffusion bonding process includes: vacuumizing in the furnace to 1x10 -2 After the heating starts, when the temperature reaches 890-920℃, a pressure of 0.2-0.5MPa is applied to the surface of the heat-conducting fiber reinforced composite copper foil tape (4), and the holding time is 10-20min. After cooling to room temperature in the furnace, the aircraft structure is taken out.

6. The aircraft thermal control structure welding method of claim 1, wherein, The plasma spraying method is to first spray a layer of zirconia with a thickness equal to that of the heat-conducting fiber-reinforced composite copper foil tape (4) in the area without adding the heat-conducting fiber-reinforced composite copper foil tape (4), and then integrally spray a layer of 0.5-1mm-thick zirconia until the surface is smooth.

7. The aircraft thermal control structure welding method of claim 1, wherein The annealing process includes: after plasma spraying, the whole aircraft structure is placed in a vacuum furnace, heated to 500-600℃ under 1x10 -1 Pa vacuum degree for 2-4h.

8. The aircraft thermal control structure welding method of claim 1, wherein, The aircraft thermal control structure welding method further comprises the following step: prepositioning a layer of graphite paper (5) with the same thickness in the area without adding the heat-conducting fiber-reinforced composite copper foil tape (4) during assembly before welding.

9. The aircraft thermal control structure welding method of claim 8, wherein, The graphite paper (5) is sprayed with a welding inhibitor.

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

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