Aircraft fuselage heat diverting system

Abstract

The invention relates to an aircraft fuselage heat diverting system. Comprising heat exchange sections arranged on the aircraft, a first heat exchange section is arranged on the windward side of the aircraft, and a second heat exchange section is arranged on the leeward side of the aircraft; the first heat exchange section communicates with the second heat exchange section through a communicating assembly. The first heat exchange section is filled with a cooling working medium; and after heat absorption phase change, the cooling working medium flows to the second heat exchange section through the communicating assembly. Efficient heat conduction can be achieved, the destructive problems of deformation and the like of a fuselage structure caused by thermal stress are solved, and the effect of optimizing space is achieved.

Description

Technical Field

[0001] The present invention relates to the technical field of aircraft thermal protection, and in particular to a heat conduction system for an aircraft fuselage. Background Art

[0002] With the development of aerospace technology, thermal protection has become the key to the design of many aircraft. For spacecraft, its external environment faces heat input from solar radiation, space particle radiation and other sources. For high-speed aircraft, due to the aerodynamic heating effect, its surface temperature will rise sharply, which puts extremely high demands on the materials and structures of the aircraft.

[0003] The advantage of metal thermal protection systems is that the heat protection layer is integrated with the load-bearing structure, which is suitable for aircraft such as reusable carriers that need to withstand high temperatures and aerodynamic loads for a long time. However, under the long-term cumulative effect of extremely high heat loads, the thermal control performance of metal thermal protection systems is constrained by wall thickness, material properties, etc., and active or semi-passive thermal protection technologies need to be combined to achieve more efficient thermal management, which has many limitations in practical applications. Summary of the invention

[0004] Based on this, it is necessary to provide an aircraft fuselage heat conduction system that does not require an additional cooling medium drive device and achieves efficient thermal management in order to address the above technical problems.

[0005] An aircraft fuselage heat conduction system comprises: a heat exchange section arranged on the aircraft, wherein a first heat exchange section is arranged on the windward side of the aircraft, and a second heat exchange section is arranged on the leeward side of the aircraft; the first heat exchange section and the second heat exchange section are connected via a connecting component; The first heat exchange section is filled with a cooling medium; after the cooling medium undergoes a heat absorption phase change, it flows to the second heat exchange section through the connecting component.

[0006] In one embodiment, there are two sets of interconnecting components; the two sets of interconnecting components are respectively connected to both ends of the first heat exchange section and the second heat exchange section to form a circulating flow of the cooling medium.

[0007] In one of the embodiments, a support member is provided in the heat exchange section, the heat exchange section is supported by the support member, and a microchannel is formed in the heat exchange section to facilitate the flow of the cooling medium.

[0008] In one embodiment, there are more than two support members, and the more than two support members are arranged at intervals to form microchannels in the interval area; the microchannels are connected to each other to facilitate the flow of the cooling medium.

[0009] In one embodiment, the support member adopts a lattice, honeycomb or corrugated configuration.

[0010] In one embodiment, the interconnecting assembly includes more than one interconnecting pipe; and the more than one interconnecting pipe is arranged in parallel between the first heat exchange section and the second heat exchange section.

[0011] In one of the embodiments, a porous liquid absorbent core is further provided in the heat exchange section; the porous liquid absorbent core is provided in the microchannel, and the porous liquid absorbent core is coaxial with the connecting tube.

[0012] In one embodiment, the porous liquid absorbent core is made of a porous material with strong hydrophilicity and high temperature resistance.

[0013] In one embodiment, the cooling medium is liquid sodium, liquid potassium or liquid lithium.

[0014] In one of the embodiments, the heat exchange section, the support member, the connecting assembly and the fuselage of the aircraft are integrally formed.

[0015] Compared with the prior art, the aircraft fuselage heat conduction system provided by the present invention has the following effects: 1. By setting up heat exchange sections on the windward and leeward sides of the aircraft and using the phase change of the cooling medium to absorb heat, efficient heat conduction can be achieved, the temperature peak of the aircraft fuselage can be effectively reduced, the temperature gradient distribution can be reduced, and destructive problems such as deformation of the fuselage structure caused by thermal stress can be solved.

[0016] 2. Effectively utilize the temperature gradient characteristics of the aircraft at different postures to realize the free conversion of heat in the first heat exchange section and the second heat exchange section without the need to carry an additional cooling medium drive device. This can not only reduce the weight of the aircraft and achieve lightweight design, but also optimize the space. BRIEF DESCRIPTION OF THE DRAWINGS

[0017] In order to more clearly illustrate the embodiments of the present invention or the technical solutions in the prior art, the drawings required for use in the embodiments or the description of the prior art will be briefly introduced below. Obviously, the drawings described below are only some embodiments of the present invention. For ordinary technicians in this field, other drawings can be obtained based on the structures shown in these drawings without paying creative work.

[0018] Figure 1 It is a schematic diagram of the axial structure of arranging a heat conduction system on an aircraft fuselage in one embodiment; Figure 2 Schematic diagram of the structure of a heat conduction system in one embodiment.

[0019] Description of reference numerals: The first heat exchange section 1, the second heat exchange section 2, the connecting pipe 3, the porous liquid wick 4, the support 5, the cooling medium 6, the fuselage 7, the control rudder 8, and the wing surface 9.

[0020] The purpose, features and advantages of the present invention will be further described with reference to the accompanying drawings in conjunction with the embodiments. DETAILED DESCRIPTION

[0021] The following will be combined with the drawings in the embodiments of the present invention to clearly and completely describe the technical solutions in the embodiments of the present invention. Obviously, the described embodiments are only part of the embodiments of the present invention, not all of the embodiments. Based on the embodiments of the present invention, all other embodiments obtained by ordinary technicians in this field without creative work are within the scope of protection of the present invention.

[0022] It should be noted that all directional indications (such as up, down, left, right, front, back, etc.) in the embodiments of the present invention are only used to explain the relative position relationship, movement status, etc. between the components under a certain specific posture (as shown in the accompanying drawings). If the specific posture changes, the directional indication will also change accordingly.

[0023] In addition, in the present invention, descriptions such as "first", "second", etc. are only used for descriptive purposes and cannot be understood as indicating or implying their relative importance or implicitly indicating the number of the indicated technical features. Therefore, the features defined as "first" or "second" may explicitly or implicitly include at least one of the features. In the description of the present invention, the meaning of "plurality" is at least two, such as two, three, etc., unless otherwise clearly and specifically defined.

[0024] In the present invention, unless otherwise clearly specified and limited, the terms "connection", "fixation", etc. should be understood in a broad sense. For example, "fixation" can be a fixed connection, a detachable connection, or an integral connection; it can be a mechanical connection, an electrical connection, a physical connection, or a wireless communication connection; it can be a direct connection, or an indirect connection through an intermediate medium, or it can be the internal connection of two elements or the interaction relationship between two elements, unless otherwise clearly defined. For ordinary technicians in this field, the specific meanings of the above terms in the present invention can be understood according to specific circumstances.

[0025] It can be understood that the technical solutions between the various embodiments of the present invention can be combined with each other, but it must be based on the fact that ordinary technicians in the field can implement it. When the combination of technical solutions is contradictory or cannot be implemented, it should be deemed that such combination of technical solutions does not exist and is not within the scope of protection required by the present invention.

[0026] The following will describe the implementation of the present invention in detail with reference to the accompanying drawings in the embodiment diagram of the present invention.

[0027] It can be understood that during the ascent and descent of the aircraft, due to the change in attitude, there is a distinction between the windward side and the leeward side. The windward side is subjected to stronger aerodynamic heating, and the temperature is significantly higher than that of the leeward side, so a large temperature difference will be formed. Based on this, the present invention designs an aircraft fuselage heat conduction system, which effectively utilizes the large temperature gradient distribution characteristics of the aircraft under different attitudes, uses the windward side aerodynamic heat as a power source, and utilizes the cooling medium 6 in the first heat exchange section 1 to absorb the heat of the windward side. Under the action of the phase change pressure difference, the cooling medium 6 flows to the leeward side, so that the heat is conducted to the leeward side. At the same time, the lower temperature of the leeward side is utilized to realize the condensation of the cooling medium 6, and the capillary driving force provided by the porous liquid wick 4 is realized to realize the process of the condensed cooling medium 6 flowing back to the first heat exchange section 1 for cooling cycle. Through such a circulation process, the temperature of the windward and leeward areas can be uniformized, and the temperature peak can be reduced. In addition, the integrated design of the load-bearing and thermal control of the fuselage 7 can also be realized.

[0028] See also Figure 1 and Figure 2 , which is a heat conduction system for aircraft fuselage provided in this embodiment, is provided with a control rudder 8 and a wing surface 9 on the fuselage surface of the aircraft, wherein the control rudder 8 is located at the upper part of the tail cabin of the fuselage 7, and the wing surface 9 is located at both sides of the fuselage 7. Heat exchange sections and connecting components are arranged in the metal heat protection layer of the fuselage 7; wherein the first heat exchange section 1 is arranged on the windward side of the aircraft, and the second heat exchange section 2 is arranged on the leeward side of the aircraft; the first heat exchange section 1 is connected with the second heat exchange section 2 through the connecting component; the first heat exchange section 1 is filled with a cooling medium 6; when the cooling medium 6 undergoes a heat absorption phase change, it flows to the second heat exchange section 2 through the connecting component.

[0029] It is understandable that the windward side and the leeward side depend on the flight attitude of the aircraft and are not limited to Figure 1 The structural state shown. The first heat exchange section 1 arranged on the windward side of the fuselage 7 is mainly used to absorb heat; the second heat exchange section 2 arranged on the leeward side of the fuselage 7 is mainly used to release heat, and the heat exchange is carried out by the phase change of the cooling medium 6 therein. When the windward side and the leeward side change, the functions of the first heat exchange section 1 and the second heat exchange section 2 change accordingly, that is, the second heat exchange section 2 absorbs heat, and the first heat exchange section 1 releases heat. In addition, in this embodiment, the fuselage 7 adopts 3D printing technology, and it is integrally formed with the first heat exchange section 1, the second heat exchange section 2, the connecting component and the support member 5 to achieve efficient integration of load bearing and thermal control.

[0030] Specifically, the first heat exchange section 1 and the second heat exchange section 2 are laid on the windward side and the leeward side along the plane direction of the fuselage 7, and their shapes are adapted to the shape of the fuselage 7. For example, in the present embodiment, the first heat exchange section 1 arranged on the windward side is an arc-shaped section, and the second heat exchange section 2 arranged on the leeward side is a straight section. The length and width are determined according to the situation so as to be able to cover the windward side and the leeward side as a whole to meet the heat exchange requirements.

[0031] The first heat exchange section 1 and the second heat exchange section 2 are cavity structures, and a support member 5 is arranged in the cavity structure. The support of the support member 5 increases the structural strength of the first heat exchange section 1 and the second heat exchange section 2, so as to increase the strength of the fuselage 7; at the same time, a microchannel is formed in the first heat exchange section 1 and the second heat exchange section 2 through the support member 4. On the one hand, the microchannel facilitates the flow of the cooling medium 6. On the other hand, when the cooling medium 6 flows in the microchannel, the support member 5 changes the flow path and velocity distribution of the cooling medium 6, thereby playing a role in fluid disturbance, enhancing the heat transfer effect, and improving the heat exchange efficiency. It is worth noting that the cavity height of the first heat exchange section 1 and the second heat exchange section 2 is the same as the thickness of the fuselage 7, so as to be integrally formed with the fuselage 7.

[0032] The number of the support members 5 is more than two, and the two or more support members 5 are arranged at intervals, so as to form more than three microchannels in the interval area; the microchannels are connected to each other so as to flow the cooling medium 6. The distance between the support members 5 can be equidistant or non-equidistant to form microchannel sizes with the same or different widths. In terms of configuration, the support members 5 can adopt a lattice, honeycomb or corrugated structure. In this embodiment, a lattice configuration is adopted, such as Figure 2 As shown, the support member 5 is a small unit component, which is arranged in an array, so as to form a microchannel and achieve disturbance of the fluid. In addition, it can also reduce the overall weight of the aircraft and achieve a lightweight design. When a honeycomb or corrugated configuration is adopted, the webs of the shape of honeycomb, corrugation, etc. are arranged at intervals in the horizontal direction to form a microchannel; at the same time, a number of connecting holes are opened on the web, and the microchannels are connected through the connecting holes to achieve the disturbance of the local fluid, thereby enhancing the effect of heat exchange. At the same time, the design of the web connecting holes can also achieve a lightweight design of the aircraft. The support member 5 can be made of heat-resistant materials such as stainless steel and titanium alloy, preferably titanium alloy.

[0033] A porous wick 4 is also provided in the first heat exchange section 1 and the second heat exchange section 2; the porous wick 4 is filled in the microchannel along the path of the microchannel and is coaxial with the connecting tube 3. The porous wick 4 is strip-shaped, and its width is adapted to the width of the microchannel. It is preferably made of a porous material with strong hydrophilicity and high temperature resistance, such as a sintered porous metal material. Furthermore, it can be made of metal foam aluminum. By using the porous wick 4, the overall weight of the fuselage can be further reduced while improving the heat exchange performance, thereby realizing a lightweight design of the aircraft.

[0034] The cooling medium 6 is liquid sodium, liquid potassium or liquid lithium, which has the characteristics of low melting point, high boiling point, large specific heat capacity, good thermal conductivity, etc. It is filled in the first heat exchange section 1 at the initial stage of setting and adsorbed by the porous liquid wick 4.

[0035] There are two sets of interconnecting components, and the two sets of interconnecting components are respectively connected to the two ends of the first heat exchange section 1 and the second heat exchange section 2 to form a circulating flow of the cooling medium 6. Specifically, the interconnecting components include more than one interconnecting pipe 3; more than one interconnecting pipe 3 is arranged in parallel between the first heat exchange section 1 and the second heat exchange section 2 to improve the uniformity and flow stability of the cooling medium 6 during temperature transfer. Specifically, the parallel arrangement is as follows: in the plane direction, the connecting pipes 3 are arranged at intervals, and the spacing can be equal spacing or non-equal spacing; in the vertical direction, the connecting pipe 3 can be a single layer or multiple layers. When it is a multi-layer, the multi-layer connecting pipe 3 is stacked, and each layer of the connecting pipe 3 is tightly fitted when stacked. Preferably, the inner diameter of the stacked connecting pipe 3 is 2-5mm, so that the stacked tube bundle can play the role of thermal siphon, thereby promoting the flow of the cooling medium 6. It is worth noting that the overall thickness of the stacked connecting pipe 3 is the same as the thickness of the fuselage 7, so that it can be integrally formed with the fuselage 7.

[0036] During operation, when the aircraft is in flight, the heat flow on the windward side is greater than that on the leeward side. The cooling medium 6 in the first heat exchange section 1 on the windward side undergoes a phase change after absorbing heat, and its volume expands. The cavity pressure in the first heat exchange section 1 increases. Under the action of the phase change pressure difference, the cooling medium 6 mixed with liquid and vapor flows through the connecting pipe 3 to the second heat exchange section 2 on the leeward side, and in the process, takes away the heat of the first heat exchange section 1; the vaporized medium condenses in the second heat exchange section 2 with a lower temperature, releases heat and returns to a liquid state, and flows back to the first heat exchange section 1 under the action of the capillary driving force provided by the porous liquid wick 4; a cooling cycle is formed, thereby playing a role in efficient heat conduction.

[0037] When the aircraft is in different flight postures, the lateral aerodynamic force and thermal environment of the windward side and leeward side of the fuselage 7 change. The first heat exchange section 1 and the second heat exchange section 2 can be freely converted with the conversion of the windward side and leeward side, thereby meeting different flight posture requirements and achieving reusability.

[0038] The technical features of the above embodiments may be combined arbitrarily. To make the description concise, not all possible combinations of the technical features in the above embodiments are described. However, as long as there is no contradiction in the combination of these technical features, they should be considered to be within the scope of this specification.

[0039] The above-mentioned embodiments only express several implementation modes of the present invention, and the description thereof is relatively specific and detailed, but it cannot be understood as limiting the scope of the invention. It should be pointed out that, for a person of ordinary skill in the art, several modifications and improvements can be made without departing from the concept of the present invention, and these all belong to the protection scope of the present invention. Therefore, the protection scope of the present invention shall be subject to the attached claims.

Claims

1. An aircraft fuselage heat relief system, characterized in that: include: A heat exchange section arranged on the fuselage of the aircraft, wherein a first heat exchange section is arranged on the windward side of the aircraft, and a second heat exchange section is arranged on the leeward side of the aircraft; the first heat exchange section and the second heat exchange section are connected via a connecting component; The first heat exchange section is filled with a cooling medium; after the cooling medium undergoes a heat absorption phase change, it flows to the second heat exchange section through the connecting component.

2. The aircraft fuselage heat conduction system according to claim 1, characterized in that: There are two sets of interconnecting components; the two sets of interconnecting components are respectively connected to the two ends of the first heat exchange section and the second heat exchange section to form a circulating flow of the cooling medium.

3. The aircraft fuselage heat conduction system according to claim 1 or 2, characterized in that: A support member is provided in the heat exchange section, the heat exchange section is supported by the support member, and a microchannel is formed in the heat exchange section to facilitate the flow of the cooling medium.

4. The aircraft fuselage heat conduction system according to claim 3, characterized in that: There are more than two support members, and the more than two support members are arranged at intervals to form microchannels in the interval area; the microchannels are connected to each other to facilitate the flow of cooling medium.

5. The aircraft fuselage heat conduction system according to claim 4, characterized in that: The support member adopts a lattice, honeycomb or corrugated configuration.

6. The aircraft fuselage heat conduction system according to claim 4 or 5, characterized in that: The interconnecting assembly includes more than one interconnecting pipe; the more than one interconnecting pipe is arranged in parallel between the first heat exchange section and the second heat exchange section.

7. The aircraft fuselage heat conduction system according to claim 6, characterized in that: A porous liquid wick is also arranged in the heat exchange section; the porous liquid wick is arranged in the microchannel, and the porous liquid wick is coaxial with the connecting pipe.

8. The aircraft fuselage heat conduction system according to claim 7, characterized in that: The porous liquid-absorbing core is made of a porous material with strong hydrophilicity and high temperature resistance.

9. The aircraft fuselage heat conduction system according to claim 1 or 2, characterized in that: The cooling medium is liquid sodium, liquid potassium or liquid lithium.

10. The aircraft fuselage heat conduction system according to claim 4 or 5, characterized in that: The heat exchange section, the support member, the connecting component and the fuselage of the aircraft are integrally formed.

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