Thermal barrier between fuselage and wing

Abstract

The present application relates to a kind of heat insulation structures between fuselage and wing, and the heat insulation structure includes at least one heat exchange pipe, the heat exchange pipe includes evaporative end and condensing end, evaporative end is set at the fuselage near engine tail flame or nozzle, condensing end is set at the wing needing to prevent ice, deicing;The heat exchange pipe further includes liquid channel and steam channel, and liquid channel is wrapped steam channel;Heat exchange is carried out between fuselage and wing using the principle of capillary suction.The present application, liquid working medium in evaporative end absorbs the heat transmitted to the nearby fuselage by engine tail flame or nozzle, thereby controlling the temperature of structure, improving the material selection limitation caused by temperature rise, and further utilizing this part of heat, at the same time, by setting condensing end at the wing needing to prevent ice, deicing, heat is used for preventing ice, deicing, which can reduce engine compressor bleed air, avoid the increase of fuel consumption caused by a large amount of engine bleed air in prior art, thereby reducing cost.

Description

Technical Field

[0001] This invention relates to the field of aircraft skin heat exchange management structure technology, and in particular to a heat insulation structure between the fuselage and the wing. Background Technology

[0002] Normally, a safe distance is maintained between the engine nacelle and the fuselage structure to ensure that the fuselage structure is not affected by the high temperature of the exhaust plume. With the rapid development of aircraft speed and altitude, new requirements have been placed on engine nacelle mounting. To reduce fuselage height, for structures with engine nacelles mounted on the belly or back of the fuselage, the distance between the engine nacelle and the fuselage must be minimized. However, the exhaust plume temperature can reach over 800°C, causing the structural temperature of the fuselage structure near the engine exhaust plume area to exceed 150°C.

[0003] Without a heat-resistant structural design, traditional aluminum alloy structures are ill-suited to withstand the high temperatures in the exhaust plume's affected area. Replacing the structural materials would incur significant costs and weight.

[0004] Therefore, it is necessary to design a heat-insulating structure to transfer the heat load of the exhaust flame area to components such as the leading edge of the tail fin that require anti-icing and de-icing. Summary of the Invention

[0005] To address the technical problems existing in the prior art, the present invention aims to provide a heat insulation structure between the fuselage and the wing, which can realize heat exchange between the fuselage and the wing. On the one hand, it controls the temperature of the fuselage structure to meet the allowable requirements of aluminum alloy, and on the other hand, it can reduce the amount of hot air bleed from the engine used for anti-icing and de-icing.

[0006] To achieve the above-mentioned objectives, the present invention provides a heat insulation structure between the fuselage and the wing, which is installed on an aircraft.

[0007] The heat insulation structure includes at least one heat exchange tube, which includes an evaporation end and a condensation end. The evaporation end is located on the fuselage near the engine exhaust or nozzle, and the condensation end is located on the wing where anti-icing and de-icing are required.

[0008] The heat exchange tube further includes a liquid channel and a steam channel, wherein the liquid channel encloses the steam channel;

[0009] Heat exchange is carried out between the fuselage and the wing using the principle of capillary suction.

[0010] According to one technical solution of the present invention, the space between the evaporation end and the condensation end is an insulation section, and the insulation section is wrapped with insulation material.

[0011] According to one technical solution of the present invention, the heat insulation section is disposed in the compartment between the fuselage and the wing, and does not contact the fuselage or the wing.

[0012] According to one technical solution of the present invention, the evaporating end is provided with at least one evaporator, and any one of the evaporators is connected to the engine body near the engine exhaust or nozzle by a heat-conducting material; or

[0013] The evaporator is embedded inside the skin of the fuselage near the engine exhaust or nozzle.

[0014] According to one technical solution of the present invention, the evaporation end has a U-shaped structure with multiple bends.

[0015] According to one technical solution of the present invention, the condensing end is provided with at least one condenser, and any one of the condensers is connected to the wing that needs to be de-iced or de-iced via a thermally conductive material; or

[0016] The condenser is integrated with the wing, which requires anti-icing and de-icing.

[0017] According to one technical solution of the present invention, the liquid channel utilizes capillary suction to transport the working fluid, which has been cooled and liquefied by the condenser, to the evaporator.

[0018] According to one technical solution of the present invention, the condenser end is provided with an outer shell, and the outer shell is provided with a mounting interface for connecting with the tail fin.

[0019] According to one technical solution of the present invention, the cross-section of the outer shell is a V-shaped structure adapted to the tail fin;

[0020] The mounting interface is a countersunk hole, and the mounting interface is symmetrically arranged on two sides of the V-shaped structure.

[0021] Compared with the prior art, the present invention has the following beneficial effects:

[0022] According to one aspect of the present invention, the heat insulation structure includes an evaporation end and a condensation end. The liquid working fluid in the evaporation end absorbs the heat transferred from the engine exhaust or nozzle to the nearby fuselage, thereby controlling the structure temperature. This can improve the limitations of material selection caused by temperature rise in the prior art, and further utilize this heat. At the same time, by setting the condensation end at the wing that needs anti-icing and de-icing, the heat is used for anti-icing and de-icing. Compared with traditional anti-icing and de-icing solutions, it can reduce the amount of engine compressor bleed air, avoid the increased fuel consumption caused by the large amount of engine bleed air used in the prior art, and thus reduce costs.

[0023] Furthermore, compared to traditional technologies, this invention can significantly reduce the structural temperature in the area affected by the engine exhaust flame or nozzle, control the fuselage structural temperature to meet the allowable requirements for aluminum alloys, thereby ensuring the lightweight of the aircraft, expanding the range of optional structural materials, improving the load-bearing capacity of the structure, and at the same time, reducing the amount of hot air bleed from the engine used for anti-icing and de-icing, ensuring the normal operating power of the aircraft, and improving aircraft performance.

[0024] A heat exchange tube is a highly efficient two-phase heat transfer device suitable for long-distance heat transfer. It utilizes the capillary action of a liquid channel to drive the circulation of the working fluid. The working fluid absorbs heat at the evaporation end, evaporating into a gas. The gaseous working fluid then flows along the vapor channel through the adiabatic end to the condensation end, where it releases heat and condenses back into a liquid state. Driven by capillary suction, the liquid working fluid flows back to the evaporation end, achieving efficient heat transfer. Heat exchange tubes offer advantages such as flexibility, ease of installation, low flow resistance, and the ability to transfer heat over long distances. Attached Figure Description

[0025] To more clearly illustrate the technical solutions in the embodiments of the present invention or the prior art, the drawings used in the embodiments will be briefly described below. Obviously, the drawings described below are merely some embodiments of the present invention, and those skilled in the art can obtain other drawings based on these drawings without creative effort.

[0026] Figure 1 This diagram illustrates the structure of the heat exchange tube in an embodiment of the present invention.

[0027] Figure 2 This schematic diagram illustrates the structure of the heat insulation structure between the fuselage and the wing in an embodiment of the present invention.

[0028] Figure 3 The schematic diagram illustrates the cross-sectional structure of the condenser end in an embodiment of the present invention.

[0029] in, Figures 1 to 3 The correspondence between the reference numerals and component names in the attached drawings is as follows:

[0030] 10. Heat exchange tube; 20. Fuselage; 30. Wing; 101. Evaporator end; 102. Condenser end; 103. Liquid passage; 104. Steam passage; 105. Insulation section; 1021. Outer shell; 1022. Mounting interface. Detailed Implementation

[0031] The description of the embodiments in this specification should be taken in conjunction with the accompanying drawings, which should form part of the complete specification. In the drawings, the shape or thickness of the embodiments may be exaggerated and may be indicated in a simplified or convenient manner. Furthermore, parts of the various structures in the drawings will be described separately; it is worth noting that elements not shown in the figures or not described in words are in a form known to those skilled in the art.

[0032] The descriptions of the embodiments herein, including any references to directions and orientations, are for ease of description only and should not be construed as limiting the scope of the invention. The following description of preferred embodiments involves combinations of features, which may exist independently or in combination; the invention is not particularly limited to the preferred embodiments. The scope of the invention is defined by the claims.

[0033] like Figures 1 to 3 As shown, a heat insulation structure between the fuselage and wing of the present invention is installed on an aircraft and includes:

[0034] The heat insulation structure includes at least one heat exchange tube 10, which includes an evaporation end 101 and a condensation end 102. The evaporation end 101 is located on the fuselage 20 near the engine exhaust or nozzle, and the condensation end 102 is located on the wing 30 where anti-icing and de-icing are required.

[0035] The heat exchange tube 10 also includes a liquid channel 103 and a steam channel 104, with the liquid channel 103 enclosing the steam channel 104;

[0036] Heat exchange is carried out between the fuselage 20 and the wing 30 using the principle of capillary suction.

[0037] In this embodiment, the high temperature of the engine exhaust or nozzle will cause the structure of the affected area of ​​the fuselage 20 to heat up, while the wing 30 has a low temperature and is prone to icing. The heat insulation structure of the present invention includes an evaporation end 101 and a condensation end 102. The liquid working fluid in the evaporation end 101 absorbs the heat transferred from the engine exhaust or nozzle to the nearby fuselage 20, thereby controlling the structural temperature. This can improve the limitations of material selection caused by temperature rise in the prior art, and can further utilize this heat. At the same time, by setting the condensation end 102 at the wing 30 that needs to be de-iced, the heat is used for de-icing. Compared with the traditional de-icing scheme, the engine compressor bleed air can be reduced, avoiding the increased fuel consumption caused by the large use of engine bleed air in the prior art, thereby reducing costs.

[0038] Furthermore, compared to traditional technologies, this invention can significantly reduce the structural temperature in the area affected by the engine exhaust flame or nozzle, control the fuselage 20 structural temperature to meet the allowable requirements for aluminum alloys, thereby ensuring the lightweight of the aircraft, expanding the range of optional structural materials, improving the load-bearing capacity of the structure, and at the same time, reducing the amount of hot air bleed from the engine used for anti-icing and de-icing, ensuring the normal operating power of the aircraft, and improving aircraft performance.

[0039] The heat exchange tube 10 is a highly efficient two-phase heat transfer device suitable for long-distance heat transfer. It utilizes the capillary action of the liquid channel 103 to drive the circulating flow of the working fluid. The working fluid absorbs heat in the evaporation end 101, evaporating into a gas. The gaseous working fluid flows along the vapor channel through the adiabatic end to the condensation end 102, where it outputs heat and condenses back into a liquid state. Driven by capillary suction, the liquid working fluid flows back to the evaporation end 101, achieving efficient heat transfer. The heat exchange tube 10's pipeline has advantages such as easy bending, easy installation, low flow resistance, and the ability to transfer heat over long distances.

[0040] Additionally, the area 20 near the engine exhaust or nozzle refers to the region on the fuselage 20 whose temperature is affected by the engine exhaust or nozzle.

[0041] In one embodiment of the present invention, preferably, there is an insulation section 105 between the evaporation end 101 and the condensation end 102, and the insulation section 105 is wrapped with insulation material to prevent heat loss.

[0042] In one embodiment of the present invention, preferably, the heat insulation section 105 is disposed in the cabin between the fuselage 20 and the wing 30, and does not contact the fuselage 20 and the wing 30, which helps to further prevent heat loss.

[0043] In one embodiment of the present invention, preferably, the evaporator end 101 is provided with at least one evaporator, and any evaporator is connected to the fuselage 20 near the engine exhaust or nozzle by a heat-conducting material; or

[0044] The evaporator is embedded inside the skin of the fuselage 20, which is close to the engine exhaust or nozzle.

[0045] In this embodiment, one or more evaporators are configured to extract heat from the exhaust flame area; one or more condensers are configured to transfer the heat extracted by the one or more evaporators to one or more components requiring de-icing.

[0046] The absorbed heat is transferred to one or more condensers via a working fluid, and each evaporator in the evaporator is connected to at least one condenser via at least one heat pipe fluid passage in which the heat transfer working fluid circulates.

[0047] In one embodiment of the present invention, preferably, the evaporation end 101 has a U-shaped structure with multiple bends.

[0048] In one embodiment of the present invention, preferably, the condensing end 102 is provided with at least one condenser, and any condenser is connected to the wing 30 that needs to be de-iced or de-iced via a thermally conductive material; or

[0049] The condenser is integrated with the wing 30, which requires anti-icing and de-icing.

[0050] In one embodiment of the present invention, preferably, the liquid channel 103 uses capillary suction to transport the working fluid liquefied by the condenser to the evaporator.

[0051] In one embodiment of the present invention, preferably, the condenser end 102 is provided with an outer shell 1021, and the outer shell 1021 is provided with an installation interface 1022 for connection with the tail fin.

[0052] In one embodiment of the present invention, preferably, the cross-section of the outer shell 1021 is a V-shaped structure adapted to the tail fin;

[0053] The mounting interface 1022 is a countersunk hole. The mounting interface 1022 is symmetrically arranged on both sides of the V-shaped structure to improve connection stability and thus ensure a stable heat exchange process.

[0054] According to one aspect of the present invention, an aircraft is provided, which is equipped with a heat-insulating structure between the fuselage 20 and the wing 30 as described in any of the above embodiments.

[0055] Figure 3 As shown Figure 2 A cross-sectional view of the condenser end 102 of the mid-tail fin shows that the outer shell 1021 adopts a conformal design, with its shape consistent with the aerodynamic shape. The liquid channel 103 of the condenser end 102 is located inside the outer shell 1021, and a vapor channel for the condenser end 102 is located inside the liquid channel 103. The gaseous working fluid outputs heat at the condenser end 102 at the leading edge of the tail fin, condensing into a liquid state, thereby transferring the heat transferred from the engine exhaust to the fuselage structure 20 to the leading edge of the tail fin.

[0056] The present invention discloses a heat insulation structure between the fuselage and the wing. The heat insulation structure includes at least one heat exchange tube, which includes an evaporation end and a condensation end. The evaporation end is located on the fuselage near the engine exhaust or nozzle, and the condensation end is located on the wing where anti-icing and de-icing are required. The heat exchange tube also includes a liquid channel and a vapor channel, with the liquid channel enclosing the vapor channel. Heat exchange is carried out between the fuselage and the wing using the principle of capillary suction.

[0057] The heat insulation structure of this invention includes an evaporation end and a condensation end. The liquid working fluid in the evaporation end absorbs the heat transferred from the engine exhaust or nozzle to the nearby fuselage, thereby controlling the structure temperature. This can improve the limitations of material selection caused by temperature rise in the prior art, and can further utilize this heat. At the same time, by setting the condensation end at the wing that needs anti-icing and de-icing, the heat is used for anti-icing and de-icing. Compared with traditional anti-icing and de-icing solutions, it can reduce the amount of engine compressor bleed air, avoid the increased fuel consumption caused by the large amount of engine bleed air used in the prior art, and thus reduce costs.

[0058] Furthermore, compared to traditional technologies, this invention can significantly reduce the structural temperature in the area affected by the engine exhaust flame or nozzle, control the fuselage structural temperature to meet the allowable requirements for aluminum alloys, thereby ensuring the lightweight of the aircraft, expanding the range of optional structural materials, improving the load-bearing capacity of the structure, and at the same time, reducing the amount of hot air bleed from the engine used for anti-icing and de-icing, ensuring the normal operating power of the aircraft, and improving aircraft performance.

[0059] The foregoing has shown and described the basic principles, main features, and advantages of the present invention. Those skilled in the art should understand that the present invention is not limited to the above embodiments. The embodiments and descriptions in the specification are merely preferred examples and are not intended to limit the invention. Various changes and modifications can be made to the invention without departing from its spirit and scope, and all such changes and modifications fall within the scope of the present invention as claimed. The scope of protection of the present invention is defined by the appended claims and their equivalents.

Claims

1. A heat insulation structure between the fuselage and the wing, characterized in that, The heat insulation structure includes at least one heat exchange tube (10), the heat exchange tube (10) includes an evaporation end (101) and a condensation end (102), the evaporation end (101) is located on the fuselage (20) near the engine exhaust or nozzle, and the condensation end (102) is located on the wing (30) where anti-icing and de-icing are required. The heat exchange tube (10) also includes a liquid channel (103) and a steam channel (104), wherein the liquid channel (103) encloses the steam channel (104). Heat exchange is carried out between the fuselage (20) and the wing (30) using the principle of capillary suction; The evaporation end (101) has a U-shaped structure with multiple bends; The condenser end (102) is provided with an outer shell (1021), and the outer shell (1021) is provided with a mounting interface (1022) for connecting with the tail fin. The cross-section of the outer shell (1021) is a V-shaped structure adapted to the tail fin.

2. The heat insulation structure between the fuselage and wing according to claim 1, characterized in that, The space between the evaporation end (101) and the condensation end (102) is an insulation section (105), and the insulation section (105) is wrapped with insulation material.

3. The heat insulation structure between the fuselage and wing according to claim 2, characterized in that, The insulation section (105) is located in the compartment between the fuselage (20) and the wing (30) and does not contact the fuselage (20) and the wing (30).

4. The heat insulation structure between the fuselage and the wing according to claim 2, characterized in that, The evaporator end (101) is provided with at least one evaporator, and any one of the evaporators is connected to the fuselage (20) near the engine exhaust or nozzle by a heat-conducting material; or The evaporator is embedded inside the skin of the fuselage (20) near the engine exhaust or nozzle.

5. The heat insulation structure between the fuselage and wing according to claim 4, characterized in that, The condensing end (102) is provided with at least one condenser, and any one of the condensers is connected to the wing (30) that needs to be de-iced or de-iced via a thermally conductive material; or The condenser is integrated with the wing (30) which requires anti-icing and de-icing.

6. The heat insulation structure between the fuselage and wing according to claim 5, characterized in that, The liquid channel (103) uses capillary suction to transport the working fluid, which has been cooled and liquefied by the condenser, to the evaporator.

7. The heat insulation structure between the fuselage and wing according to claim 5, characterized in that, The mounting interface (1022) is a countersunk hole, and the mounting interface (1022) is symmetrically arranged on the two sides of the V-shaped structure.

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