阅读说明：本技术 一种用于飞行器的热管理系统 (Thermal management system for aircraft ) 是由 苗辉 李亚忠 周琨 魏宽 于 2020-12-22 设计创作，主要内容包括：发明公开了一种用于飞行器的热管理系统,包括吸热单元、导热单元、换热单元和燃料预热单元,吸热单元设置在飞行器需散热位置处,换热单元经导热单元与所述吸热单元连接,燃料预热单元与所述换热单元相连接。本申请通过吸热单元吸收飞行器上需要进行散热区域的热量,并且通过导热单元将热量传递给换热单元,燃料预热单元吸收换热单元的热量加热燃料,燃料进行预热以后进入飞行器的动力单元燃烧；本申请中热管理系统的构成简单,可靠性高,能量利用率高,且飞行器中燃料燃烧的热效率高。(The invention discloses a thermal management system for an aircraft, which comprises a heat absorption unit, a heat conduction unit, a heat exchange unit and a fuel preheating unit, wherein the heat absorption unit is arranged at the position of the aircraft where heat is required to be dissipated, the heat exchange unit is connected with the heat absorption unit through the heat conduction unit, and the fuel preheating unit is connected with the heat exchange unit. The heat absorption unit absorbs heat of a region needing heat dissipation on the aircraft, the heat conduction unit transfers the heat to the heat exchange unit, the fuel preheating unit absorbs the heat of the heat exchange unit to heat fuel, and the fuel enters the power unit of the aircraft to be combusted after being preheated; the heat management system in this application's constitution is simple, and the reliability is high, and energy utilization is high, and the thermal efficiency of fuel burning in the aircraft is high.)

1. A thermal management system for an aircraft, comprising:

the heat absorption unit is arranged at a position, needing heat dissipation, of the aircraft and can absorb heat at the position, needing heat dissipation, of the aircraft;

the heat exchange unit is connected with the heat absorption unit through a heat conduction unit, and the heat conduction unit can transfer the heat absorbed by the heat absorption unit to the heat exchange unit;

and the fuel preheating unit is connected with the heat exchange unit and can absorb the heat of the heat exchange unit to heat the fuel.

2. The thermal management system for an aircraft according to claim 1, wherein: the heat conduction unit comprises at least one set of circulation pipeline, one end of the circulation pipeline is communicated with the heat absorption unit, the other end of the circulation pipeline is communicated with the heat exchange unit, a heat exchange medium is arranged in the circulation pipeline, a medium driving element is further installed on the circulation pipeline, and the medium driving element can drive the heat exchange medium to circularly flow in the circulation pipeline.

3. The thermal management system for an aircraft according to claim 2, characterized in that: the medium driving element is connected with a gas power unit, and the gas power unit can convert the kinetic energy of the airflow outside the aircraft into the power of the medium driving element.

4. The thermal management system for an aircraft according to claim 3, wherein: the gas power unit is connected with an air suction engine of the aircraft.

5. The thermal management system for an aircraft according to claim 4, wherein: the gas power unit is connected with the air suction engine through the gas drainage unit.

6. The thermal management system for an aircraft according to claim 5, wherein: the inlet of the gas guide unit is arranged at the nozzle position of the air suction engine.

7. The thermal management system for an aircraft according to claim 2, characterized in that: the heat exchange medium of the medium circulation heat exchange element is liquid metal.

8. The thermal management system for an aircraft according to claim 7, wherein: the liquid metal is gallium-indium alloy and/or potassium-sodium alloy.

9. The thermal management system for an aircraft according to claim 2, characterized in that: the medium drive element is connected with a power unit of the aircraft.

10. The thermal management system for an aircraft according to claim 1, wherein: the front edge of the aircraft is provided with a skin, and the heat absorption unit is in contact with the skin.

Technical Field

The disclosure belongs to the technical field of aircrafts, and particularly relates to a thermal management system for an aircraft.

Background

The speeding up of aircraft has extremely important military and civilian value. However, due to the existence of the aerodynamic heating effect, the temperature of the air sensed by the surface of the aircraft is greatly increased along with the increase of the flying speed, for example, when the stratosphere flies in a Ma5 state, the leading edge (including the nose leading edge and the wing leading edge) of the aircraft is subjected to the aerodynamic heating temperature of 1000 ℃, which is far higher than the safety temperature of the aircraft skin material, and thermal protection measures are required.

Thermal protection measures are divided into two categories, passive and active. The passive thermal protection mainly adopts a heat insulation layer or an ablation material, and is suitable for occasions of one-time and short-time flight such as missiles and the like. The surface thermal protection of the hypersonic aircraft which can be repeatedly used during long-endurance navigation inevitably selects an active thermal protection measure.

In the heat recovery device in the prior art, the thermoelectric conversion material is used for recovering and reusing the aerodynamic heat generated by high-speed incoming flow in the working process of the aircraft and the heat transfer of an engine combustion chamber, so that the effective heat management of the aircraft and the comprehensive utilization of energy are realized, and the total energy utilization efficiency is improved. The disadvantage is that the cooling is limited by the thermoelectric conversion material, the efficiency is extremely low, and the high-temperature components cannot be reliably cooled.

Disclosure of Invention

To address at least one of the above technical problems, the present disclosure provides a thermal management system for an aircraft.

The technical scheme of the disclosure is as follows:

a thermal management system for an aircraft, comprising:

the heat absorption unit is arranged at a position, needing heat dissipation, of the aircraft and can absorb heat at the position, needing heat dissipation, of the aircraft;

the heat exchange unit is connected with the heat absorption unit through a heat conduction unit, and the heat conduction unit can transfer the heat absorbed by the heat absorption unit to the heat exchange unit;

and the fuel preheating unit is connected with the heat exchange unit and can absorb the heat of the heat exchange unit to heat the fuel.

Optionally, the heat conducting unit includes at least one set of circulation pipeline, one end of the circulation pipeline is communicated with the heat absorbing unit, the other end of the circulation pipeline is communicated with the heat exchanging unit, a heat exchanging medium is disposed in the circulation pipeline, and a medium driving element is further installed on the circulation pipeline and can drive the heat exchanging medium to circularly flow in the circulation pipeline.

Optionally, the medium driving element is connected to a gas power unit, and the gas power unit can convert kinetic energy of airflow outside the aircraft into power of the medium driving element.

Optionally, the aerodynamic unit is connected to an air-breathing engine of the aircraft.

Optionally, the gas power unit is connected to the air-breathing engine through a gas diversion unit.

Optionally, the inlet of the gas guiding unit is arranged at the nozzle position of the air suction engine.

Optionally, the heat exchange medium of the medium circulation heat exchange element is liquid metal.

Optionally, the liquid metal is a gallium indium alloy and/or a potassium sodium alloy.

Optionally, the medium drive element is connected to a power unit of the aircraft.

Optionally, a skin is disposed on the leading edge of the aircraft, and the heat absorbing unit is in contact with the skin.

According to the thermal management system of the aircraft, the heat absorption unit absorbs the heat of a region needing heat dissipation on the aircraft, the heat conduction unit transfers the heat to the heat exchange unit, the fuel preheating unit absorbs the heat of the heat exchange unit to heat the fuel, and the fuel enters the power unit of the aircraft to be combusted after being preheated; the heat management system in this application's constitution is simple, and the reliability is high, and energy utilization is high, and the thermal efficiency of fuel burning in the aircraft is high.

Drawings

The accompanying drawings, which are included to provide a further understanding of the disclosure and are incorporated in and constitute a part of this specification, illustrate exemplary embodiments of the disclosure and together with the description serve to explain the principles of the disclosure.

FIG. 1 is a block diagram of a thermal management system for an aircraft according to the present disclosure;

FIG. 2 is a schematic structural view of a thermal management system for an aircraft with heat sinking of skin locations in accordance with the present disclosure;

FIG. 3 is a schematic structural diagram of a thermal management system for an aircraft with heat dissipation to an engine location according to the present disclosure;

FIG. 4 is a schematic structural diagram of a thermal management system for an aircraft with heat sinking of electronic devices in the present disclosure;

FIG. 5 is a schematic structural diagram of a thermal management system for an aircraft according to a second embodiment of the disclosure;

fig. 6 is a schematic structural diagram of a thermal management system for an aircraft according to a third embodiment of the disclosure.

Detailed Description

The present disclosure will be described in further detail with reference to the drawings and embodiments. It is to be understood that the specific embodiments described herein are for purposes of illustration only and are not to be construed as limitations of the present disclosure. It should be further noted that, for the convenience of description, only the portions relevant to the present disclosure are shown in the drawings.

It should be noted that the embodiments and features of the embodiments in the present disclosure may be combined with each other without conflict. The present disclosure will be described in detail below with reference to the accompanying drawings in conjunction with embodiments.

Example 1

As shown in fig. 1, the invention provides a thermal management system for an aircraft, which includes a heat absorption unit 1, a heat conduction unit 2, a heat exchange unit 3 and a fuel preheating unit 4, wherein the heat absorption unit 1 is arranged at a position of the aircraft where heat is to be dissipated, the heat exchange unit 3 is connected with the heat absorption unit 1 through the heat conduction unit 2, and the fuel preheating unit 4 is connected with the heat exchange unit 3; the heat unit can absorb the heat of aircraft position department that needs the heat dissipation, and the heat conduction unit can transmit the heat that the heat absorption unit absorbed to the heat transfer unit, and the fuel preheats the heat heating fuel that the unit can absorb the heat of heat transfer unit.

In one embodiment, as shown in fig. 2, the leading edge of the aircraft is provided with a skin 7, and the heat absorption unit 1 can be in contact with the skin 7 to absorb heat and cool the skin 7 of the aircraft; as shown in fig. 3, the heat absorption unit may also be installed at an engine position of the aircraft to absorb heat and cool the engine; as shown in fig. 4, the heat absorption unit 1 may be installed at the position of the electronic device 8 of the aircraft to absorb heat and cool down the electronic device. The heat absorption unit 1 can be a heat absorption plate, has a simple heat absorption plate structure, is attached and contacted with the shape of a heat dissipation position, has a large contact area and high heat dissipation efficiency, is not influenced by the size of a device needing heat dissipation, and can be freely customized according to the shape and the size of a field space. The heat absorption unit 1 may also be a heat sink, which absorbs heat and cools a position to be cooled through a heat dissipation medium flowing in a circulating manner, and a heat dissipation pipeline of the heat sink may be embedded into the device to be cooled, so that heat dissipation may be performed from the inside of the device, and thus, the position to be cooled may be accurately cooled. For example, a flow passage is formed in a housing of an engine, and heat is dissipated through a heat dissipating medium; for another example, a flow channel is formed in the aircraft skin, and heat is dissipated through the aircraft skin by a heat dissipation medium.

In one embodiment, the heat conducting unit 2 may be a heat conducting sheet, and the material thereof may be tungsten-copper alloy, which is capable of resisting high temperature. One end of the heat conduction unit 2 is connected with the heat absorption unit 1, and the other end of the heat conduction unit 2 is connected with the heat exchange unit 3 for transferring heat. In another embodiment, the heat conducting unit 2 may also be a heat conducting glue, a heat conducting silicone grease, or a heat conducting insulating elastic rubber, which can conduct heat to the electronic component, and can conduct heat and insulate. In another embodiment, the heat conducting unit 2 may also absorb the heat of the heat absorbing unit 1 by using a heat conducting medium that circulates, and transfer the heat to the heat exchanging unit.

In one embodiment, the heat exchange unit 3 is a dividing wall type heat exchanger, has a mature structure and high heat exchange efficiency, and can transfer the heat absorbed by the heat absorption unit 1 to the heat exchange unit. For another example, the heat exchange unit 3 is a metal-based heat dissipation plate, and the heat exchange plate is made of a heat exchange material to perform direct heat conduction and heat exchange, so that the structure is simple. For another example, the heat exchange unit 3 is a shell-and-tube heat exchanger, and a spiral tube bundle design, so that the turbulence effect can be increased to the maximum extent, and the heat exchange efficiency is increased.

In one embodiment, the fuel preheating unit 4 may be installed on a fuel tank of an aircraft, and preheat and keep warm the fuel in the whole fuel tank, so as to maintain the fuel in the fuel tank at a stable temperature; alternatively, the fuel preheating unit 4 may be installed between the fuel tank 5 and the power unit 6 of the aircraft to preheat the fuel to be introduced into the power unit 6 of the aircraft, and the fuel in the fuel tank 5 is introduced into the power unit 6 of the aircraft for combustion after passing through the fuel preheating unit 4. The fuel preheating unit 4 and the heat exchange unit 3 can be integrally connected to form a dividing wall type heat exchanger or a shell-and-tube type heat exchanger, and heat transferred by the heat conduction unit 2 exchanges heat with fuel through the dividing wall type heat exchanger or the shell-and-tube type heat exchanger, so that the fuel is preheated.

So, absorb the aircraft through heat absorption unit 1 and go on the heat that the heat dissipation is regional to give heat transfer unit 3 with the heat through heat conduction unit 2, fuel preheating unit 4 absorbs heat heating fuel of heat transfer unit 3, and fuel gets into the power pack 6 burning of aircraft after preheating, has avoided power pack 6 because fuel temperature operation is unstable and exert oneself and phenomenon such as reduce. The heat management system is simple in structure, high in reliability and energy utilization rate, and high in thermal efficiency of fuel combustion in the aircraft.

Example 2

In this embodiment, as shown in fig. 5, the heat conducting unit 2 includes one set of circulating pipes 21 (a plurality of sets of circulating pipes 21 may also be provided), the circulation pipeline 21 can be made of copper, nickel or stainless steel, the material cost is low, the circulation pipeline 21 can also be made of ceramic materials, one end of the circulation pipeline 21 is communicated with the heat absorption unit 1, the heat absorbing unit 1 may be a radiator, the other end of the circulation line 21 is communicated with the heat exchanging unit 3, the heat exchanging unit 3 may be a heat exchanging fin, wherein, a heat exchange medium is arranged in the circulating pipeline 21, a medium driving element 22 is also arranged on the circulating pipeline 21, the medium driving member 22 may be a vane pump, an injection pump, a centrifugal pump, a mechanical pump, etc., and the medium driving member 22 drives the heat exchange medium to circulate between the heat conduction unit 2 and the heat exchange unit through the circulation line 21.

If the aircraft needs to dissipate heat, the electronic equipment is arranged at the position, the heat exchange medium can be water, the cost is low, and if the heat exchange medium leaks, the heat exchange medium is easy to supplement; or the heat exchange medium is a fluorocarbon cooling liquid, so that the heat stability is high and the chemical stability is good;

if the aircraft needs to radiate heat at the engine, the heat exchange medium can be oil, and even if the oil leaks, the electronic equipment on the engine cannot be influenced;

if the aircraft skin is the position needing heat dissipation of the aircraft, the heat exchange medium can be fused salt, the fused salt is heated up quickly, the heat dissipation efficiency is high, and fused salt decomposition is not easy to cause; or the heat exchange medium is liquid metal, so that the heat exchange capacity is high, the liquid metal cannot undergo phase change, and the system safety is good; the liquid metal can also be one of indium tin alloy, InSn, InSnGa and InSnGaZn alloy; or an alloy of gallium, indium, tin, lead, etc.

In another embodiment of the present application, the liquid metal is a gallium-indium alloy and/or a potassium-sodium alloy, which has a very high thermal conductivity and relatively high heat absorption, a thermal conductivity much higher than that of conventional metals,

in a preferred embodiment of the present application, as shown in fig. 5, the medium driving element 22 is connected to a gas power unit 23, the gas power unit 23 being capable of converting kinetic energy of the aircraft exterior gas flow into power for the medium driving element 22. The gas power unit 23 can be an air turbine, high-speed air passes through the air turbine and then drives the air turbine to rotate, the air turbine transmits power to the medium driving element 22, and the medium driving element 22 drives the heat exchange medium to circularly flow between the heat absorption unit 1 and the heat exchange unit 3. The gas power unit 23 can also be a fan, the structure is simple, and the process requirement is low; the gas power unit 23 may also be a wind power motor, and after being converted into electric energy, the electric energy drives the medium driving element 22 to do work. The gas power unit 23 may be mounted in an aircraft and, in use, extend from the aircraft to the exterior of the aircraft, the high velocity air on the exterior of the aircraft being able to drive the gas power unit 23 to perform work. The aircraft can also be specially provided with a diversion air duct capable of being opened and closed, the gas power unit 23 is fixedly arranged in diversion, when the diversion air duct is opened, high-speed air outside the aircraft enters the diversion air duct to drive the gas power unit to do work, and when the gas power unit 23 is not needed, only the diversion air duct needs to be closed.

In another embodiment of the present application, the gas power unit 23 is connected to an air intake engine of an aircraft, the gas power unit 23 preferably employs an air turbine, the air turbine is driven to rotate by high-pressure and high-temperature air ejected from the air intake engine of the aircraft, and the air turbine is technically mature in connection with systems in the aircraft and stable in performance; the air turbine can be directly and mechanically connected with the medium driving element 22, for example, the air turbine and a rotating shaft in the medium driving element are coaxially arranged, so that the power transmission efficiency is high, and the structure is simple; the power output by the air turbine can also be transmitted to the medium driving element 22 through an electric system or a hydraulic system, and the power output distance is long.

Example 3

As shown in fig. 6, the present embodiment is different from embodiment 2 in that the aerodynamic unit 23 is connected to the intake engine through the air guiding unit 24, the high-temperature and high-pressure air ejected from the intake engine can be guided to the aerodynamic unit 23 through the air guiding unit 24, the air flow is concentrated, the power is strong, and the aerodynamic force ejected from the intake engine can be effectively utilized. The inlet of the gas guide unit 24 can be arranged at the nozzle position of the air suction engine, the other end of the gas guide unit is arranged at the air inlet position of the gas power unit 23, high-temperature and high-pressure gas can be guided to a safe area to drive the gas power unit 23, and the gas power unit 23 can be effectively prevented from being burnt out due to the fact that the temperature of the just-sprayed high-temperature and high-pressure gas is too high; the cross section of the gas drainage unit can be a round pipe, a rectangular pipe or a special-shaped structure, and the inner wall of the gas drainage unit can be sprayed with high-temperature-resistant materials.

In the preferred embodiment of the present application, the inlet of the gas guiding unit 23 is arranged at the nozzle position of the air-breathing engine, because the temperature of the gas after combustion of the air-breathing engine is high, the temperature of the gas after being sprayed out of the air-breathing engine is preferably properly reduced, and the high-temperature and high-pressure gas after being properly reduced in temperature is guided to the position of the gas power unit 23 by collecting part of the gas by the gas guiding unit 24. The inlet position of the gas guiding unit 23 may also be installed according to the working condition or actual condition of the engine, for example, at the position of the tail section of the engine.

Example 4

In this example, the present application differs from example 2 in that; the medium drive element 22 is connected with the power unit 6 of the aircraft; the power unit 6 can directly drive the medium driving element 22 to do work through the speed reducing mechanism, and the structure is simple; the power unit 6 can also drive the medium driving element 22 to do work through a hydraulic system or an electric system, the power transmission distance is long, the hydraulic system and the electric system can be accessed everywhere in the aircraft, and the installation position selectivity of the medium driving element 22 is also high.

In the description herein, reference to the description of the terms "one embodiment/mode," "some embodiments/modes," "example," "specific example," or "some examples," etc., means that a particular feature, structure, material, or characteristic described in connection with the embodiment/mode or example is included in at least one embodiment/mode or example of the application. In this specification, the schematic representations of the terms used above are not necessarily intended to be the same embodiment/mode or example. Furthermore, the particular features, structures, materials, or characteristics described may be combined in any suitable manner in any one or more embodiments/modes or examples. Furthermore, the various embodiments/aspects or examples and features of the various embodiments/aspects or examples described in this specification can be combined and combined by one skilled in the art without conflicting therewith.

Furthermore, the terms "first", "second" and "first" are used for descriptive purposes only and are not to be construed as indicating or implying relative importance or implicitly indicating the number of technical features indicated. Thus, a feature defined as "first" or "second" may explicitly or implicitly include at least one such feature. In the description of the present application, "plurality" means at least two, e.g., two, three, etc., unless specifically limited otherwise.

It will be understood by those skilled in the art that the foregoing embodiments are merely for clarity of illustration of the disclosure and are not intended to limit the scope of the disclosure. Other variations or modifications may occur to those skilled in the art, based on the foregoing disclosure, and are still within the scope of the present disclosure.