Heterogeneous immersed dry ice jet two-stage cooling device for thermal protection of end of hypersonic aircraft and use method of heterogeneous immersed dry ice jet two-stage cooling device

Abstract

The invention relates to an active immersed jet cooling device for thermal protection of a hypersonic aircraft end and a use method thereof in the field of aircrafts, the front part of a machine body of the device is provided with a porous medium end, and the porous medium end is provided with a plurality of micropores sealed by a low-melting-point coating; a continuous liquid phase water cavity with a pressure bearing structure is arranged in the porous medium end, a dry ice spraying device is further arranged in the machine body, a nozzle of the dry ice spraying device is immersed in water in the continuous liquid phase water cavity, and the spraying front of the nozzle is aligned with the center of the porous medium end. During use, the interior of the hypersonic aircraft is cooled through the strong cooling capacity generated by the immersed dry ice jet cooling system, in the cooling process, dry ice is heated and sublimated to generate a large amount of gas, the gas and water vapor generated after immersion liquid water is heated permeate into the outer side of the end together, and therefore a gas layer is formed on the outer side of the end; and the air film cooling effect is achieved, and therefore it is guaranteed that the aircraft works more safely and reliably.

Description

Technical Field

[0001] The invention relates to a hypersonic aircraft, and in particular to a cooling device for heat protection of a tip of a hypersonic aircraft. Background Art

[0002] With the rapid development of modern aerospace technology, hypersonic vehicles have attracted widespread attention due to their important role in the military and civilian fields. Such vehicles can cross long distances in a very short time, provide rapid response capabilities, and significantly improve global arrival speeds. However, when an aircraft flies at speeds exceeding Mach 5, its front-end components will experience extreme aerodynamic heating, with temperatures reaching thousands of degrees Celsius. This high temperature not only threatens the integrity and performance of the aircraft's structural materials, but may also cause problems such as thermal stress damage, material degradation or failure, thereby affecting flight safety and mission success.

[0003] Traditional thermal protection systems (TPS) mainly rely on the application of passive insulation materials such as high-temperature resistant ceramics and carbon-carbon composites. Although these materials have certain anti-ablation properties, they are often heavy, expensive, and difficult to repair once damaged. In addition, traditional methods mostly use static designs, which are not flexible enough when facing dynamic changes under different flight conditions. Therefore, in order to meet the needs of future hypersonic vehicles for efficient, lightweight, and reliable thermal protection, exploring new active cooling strategies has become one of the research hotspots.

[0004] Active thermal protection technology plays a vital role in the field of aerospace. In the Chinese patent database, a thermal protection and drag reduction system for a hyperspeed aircraft is disclosed, with publication number: CN104608942A; publication date: 20150513. The device includes a cold source arranged inside a sealed cavity of the hyperspeed aircraft, and a cold source driving device for converting the cold source into high-pressure gas and ejecting it; at least part of the wall surface of the cavity wall of the hyperspeed aircraft is a sandwich structure, and the sandwich structure includes a transition layer through which the cold source gas can pass and an outer surface layer located on the surface of the transition layer, and the outer surface layer is provided A plurality of micropores are used to connect the transition layer with the outside of the cavity; the cold source driving device includes a cold source storage, an air pump and a buffer; the air pump is connected to the cold source storage; the buffer includes a buffer inlet and a buffer outlet, the buffer inlet is connected to the cold source storage, the buffer outlet is connected to the transition layer of the cavity wall, and a sealing valve is arranged at the connecting part between the buffer outlet and the transition layer; in the working state, the air pump provides compressed air to the cold source storage, the cold source enters the buffer and vaporizes under the action of air pressure, the sealing valve is opened, the gas is sprayed into the transition layer from the buffer outlet, and then sprayed out of the cavity from the micropores of the outer surface layer to form an air film. The micropores are non-circular holes; the diameter of the micropores is preferably 0.05 mm-2.0 mm. The cold source is liquid nitrogen, dry ice, compressed air, or other cooling substances produced by chemical reactions. Its disadvantage is that the air film forms a protective film on the outer surface, showing strong thermal insulation performance. However, although the jet cooling has efficient cooling capacity, it performs poorly when facing extremely high temperature and high heat flux density working conditions. This is mainly because jet cooling relies on a relatively low flow rate of jet fluid, which can easily cause the coolant on the hot surface to dry up, thus losing the cooling effect. In other words, the cooling capacity of existing means such as liquid nitrogen, dry ice, and compressed air is insufficient. Summary of the invention

[0005] In order to overcome the shortcomings of the prior art, the present invention proposes an immersion dry ice spraying two-stage cooling device for thermal protection of the end of a hypersonic aircraft, aiming to solve the problems in the prior art that passive thermal insulation materials are heavy, costly and have poor adaptability, and active cooling but insufficient cooling capacity. Through the immersion dry ice spraying two-stage cooling device, an internal cooling mechanism is adopted to quickly absorb the heat entering the aircraft structure, and the external air film effectively isolates the influence of the external high temperature environment on the aircraft. The cooling intensity can be adjusted according to actual needs in different flight stages, thereby ensuring that the aircraft works safer and more reliable.

[0006] To this end, the technical solution of the present invention is as follows: a heterogeneous immersion dry ice blasting two-stage cooling device for thermal protection of a hypersonic aircraft tip, comprising a fuselage, a porous medium tip being provided at the front of the fuselage, the inner and outer walls of the porous medium tip being connected through a plurality of micropores in the medium, and a low-melting-point coating being provided on the outer surface of the porous medium tip to seal the outer openings of the micropores; a continuous liquid phase water cavity having a pressure-bearing structure is provided in the porous medium tip, a dry ice blasting device is also provided in the fuselage, the nozzle of the dry ice blasting device being immersed in the water in the continuous liquid phase water cavity, and the front of the nozzle spraying is aligned with the center of the porous medium tip.

[0007] Furthermore, a carbon dioxide gas storage tank, a temperature sensor, a pressure sensor, a flow meter and a control valve are also provided in the body; the outlet of the carbon dioxide gas storage tank is connected to the inlet of the control valve, the outlet of the control valve is connected to the dry ice blasting nozzle via the flow meter, and the temperature sensor and the pressure sensor are arranged on the pipe connecting the outlet of the control valve and the nozzle.

[0008] Furthermore, the surface of the porous medium end is densely covered with micropores, and the average pore size of the micropores is 0.05 mm.

[0009] The present invention also discloses a method for using a heterogeneous immersion dry ice blasting two-stage cooling device for thermal protection of a hypersonic aircraft end. When the hypersonic aircraft performs hypersonic flight and generates aerodynamic heat by friction with the external atmosphere, after the low-melting-point coating melts, the heterogeneous immersion dry ice blasting two-stage cooling device for thermal protection of the aircraft end is started, and dry ice particles are sprayed out from a dry ice blasting nozzle immersed in liquid water. The dry ice particles are heated and sublimated / vaporized at the moment of spraying, thereby generating a dry ice blasting downstream airflow. At the same time, the continuous liquid phase water cavity used for immersion is also vaporized into water vapor by the action of aerodynamic heat, and a mixed gas of carbon dioxide and water vapor is generated inside the body of the porous medium end of the hypersonic aircraft. The inner and outer walls of the medium are connected in the porous medium end through multiple micropores in the medium. The mixed gas penetrates from the inside of the body to the outside of the body through the porous medium, and the water vapor and carbon dioxide mixed gas generates a mixed thermal protection gas layer of water vapor and carbon dioxide attached to the outside of the body, forming a cooling thermal protection gas film.

[0010] The mass ratio of the dry ice to water is 1:(0.6-2.2).

[0011] Furthermore, the melting temperature of the low melting point coating is 400-600° C. The low melting point coating can be made of metal or metal alloy material. When the temperature reaches 400-600° C., the low melting point coating melts and is blown away by the high-speed airflow.

[0012] Compared with the prior art, the beneficial effects of the present invention are as follows: by spraying dry ice particles into water, the present invention utilizes the combined effects of low-temperature gas generated by dry ice sublimation and water vapor formed by liquid evaporation to enhance the heat transfer effect and spray out from micropores to form an external air film barrier. The internal cooling mechanism quickly absorbs the heat entering the aircraft structure, while the external air film effectively isolates the aircraft from the impact of the external high-temperature environment. This two-stage cooling strategy not only improves the cooling efficiency, but also reduces the additional mass, and enhances the flexibility and reliability of the system. In addition, the system can adjust the cooling intensity in real time according to different flight conditions to ensure that the aircraft can obtain optimal protection in various high-speed environments. This innovative active cooling method provides a more efficient and lightweight thermal protection solution for hypersonic aircraft.

[0013] This two-stage cooling mechanism combines the advantages of strong internal heat transfer and external air film isolation. It can absorb a large amount of heat in a short time and effectively block the external heat from being transferred to the aircraft body. Compared with the traditional single passive protection method, the solution provided by the present invention is more flexible and adaptable, and can adjust the cooling intensity according to actual needs in different flight phases. More importantly, it also has a lower added mass, which helps to reduce the overall burden of the aircraft and improve fuel efficiency and range.

[0014] In summary, the submerged dry ice spray two-stage cooling system proposed in this invention provides an innovative thermal protection solution for hypersonic aircraft, solves the challenges faced by current technology, and is expected to play an important role in the design of future high-performance aircraft. With the development and improvement of related technologies, it is expected that this new cooling system will be applied to more types of high-speed flight platforms, promoting the advancement of aerospace technology. BRIEF DESCRIPTION OF THE DRAWINGS

[0015] Figure 1 This is a working principle diagram of the heterogeneous immersion dry ice spray two-stage cooling for the hypersonic vehicle tip thermal protection in the present invention.

[0016] Figure 2 It is a partial enlarged view of the two-stage cooling process of the aircraft end heterogeneous immersion dry ice spraying in the present invention.

[0017] Figure 3 The present invention shows the aircraft end heterogeneous immersion dry ice spraying two-stage cooling device and installation diagram.

[0018] Figure 4 It is a detailed connection diagram of the dry ice blasting system in the present invention.

[0019] In the figure: 1 porous medium end, 2 dry ice particles, 3 pipeline, 4 continuous liquid phase water chamber, 5 nozzle, 6 mixed gas, 7 dry ice jet downstream airflow, 8 air film, 9 pneumatic heat, 10 carbon dioxide gas storage tank, 11 body, 12 temperature sensor, 13 pressure sensor, 14 flow meter, 15 control valve. DETAILED DESCRIPTION

[0020] The present invention will be further described below in conjunction with the accompanying drawings and embodiments:

[0021] like Figure 1 As shown, a heterogeneous immersion dry ice blasting two-stage cooling device for thermal protection of a hypersonic aircraft tip comprises a body 11 of a hypersonic aircraft, a porous medium tip 1 is provided at the front of the body 11, the inner and outer walls of the medium are connected through a plurality of micropores in the porous medium tip 1, and a low-melting-point coating is provided on the outer surface of the porous medium tip 1 to seal the outer openings of the micropores; a continuous liquid phase water cavity 4 with a pressure-bearing structure is provided in the porous medium tip 1, the continuous liquid phase water cavity 4 can be an independent water cavity, or a cavity connected to a pressure water source through a pipeline 3, and pressurized water can be injected into the continuous liquid phase water cavity 4; a dry ice blasting device is also provided in the body 11, the dry ice blasting device comprises a connecting pipeline 3 and a nozzle 5, the nozzle 5 is immersed in the water in the continuous liquid phase water cavity 4, and the front of the nozzle 5 is aligned with the center of the porous medium tip 1.

[0022] The dry ice blasting device is composed of a pipeline 3, a nozzle 5, a carbon dioxide gas storage tank 10, a temperature sensor 12, a pressure sensor 13, a flow meter 14 and a control valve 15; the carbon dioxide gas storage tank 10 is arranged in a body 11, the outlet of the carbon dioxide gas storage tank 10 is connected to the inlet of the control valve 15, the outlet of the control valve 15 is connected to the dry ice blasting nozzle 5 via the flow meter 14, and the temperature sensor 12 and the pressure sensor 13 are arranged on the pipeline 3 connecting the outlet of the control valve 15 and the nozzle 5.

[0023] The surface of the porous medium end 1 is densely covered with micropores, and the average pore diameter of the micropores is 0.05 mm.

[0024] The above-mentioned heterogeneous immersion dry ice spraying two-stage cooling device for thermal protection of the end of a hypersonic aircraft is used as follows:

[0025] During operation, liquid water is filled into the porous medium end 1. The hypersonic vehicle generates aerodynamic heat 9 by friction with the external atmosphere during hypersonic flight. After the low-melting-point coating melts, the heterogeneous immersion dry ice jetting two-stage cooling device for thermal protection of the aircraft end is started. Dry ice particles 2 are sprayed out from a nozzle 5 immersed in liquid water. The dry ice particles 2 are heated and sublimated / vaporized at the moment of spraying, thereby generating a dry ice jetting downstream airflow 7. At the same time, the continuous liquid water cavity 4 used for immersion is also vaporized into water vapor by the action of aerodynamic heat 9. A mixed gas 6 of carbon dioxide and water vapor is generated inside the body 11 of the porous medium end 1 of the hypersonic vehicle. The inner and outer sides of the porous medium end 1 are connected to the inner and outer walls of the medium through the pores in the medium. The mixed gas 6 penetrates from the inside of the body 11 through the porous medium to the outside of the body 11 through the porous medium. The water vapor and carbon dioxide mixed gas generates a mixed thermal protection gas layer of water vapor and carbon dioxide attached to the outside of the body 11, forming a cooling thermal protection air film 8.

[0026] Furthermore, the mass ratio of dry ice to water is 1:(0.6-2.2); because the latent heat of vaporization of water is 137 kcal / kg for dry ice and 539 kcal / kg for water, within this ratio range, the latent heat of vaporization absorbed by the mixed gas is 2-3 times that of the total mass of dry ice alone, which can obviously take away more heat. This is not something that dry ice alone can do. The melting temperature of the low-melting-point coating is 400-600°C. The low-melting-point coating can be made of metal or metal alloy material. When the temperature reaches 400-600°C, the low-melting-point coating melts and is blown away by the high-speed airflow. A small amount of coating material in the infiltrated micropores will also flow away due to the extrusion of the mixed gas 6 of carbon dioxide and water vapor.

[0027] Figure 2 In the process, liquid water is filled in the porous medium end 1, which works together with the airflow downstream of the dry ice jet to internally cool the aerodynamic heat 9 transmitted into the body 11. Through the phase change effect of the internal cooling, a mixture of water vapor and carbon dioxide is generated that penetrates from the inside of the body 11 through the porous medium to the outside of the body 11, forming a mixed thermal protective gas layer of water vapor and carbon dioxide attached to the outside of the body 11, that is, an air film 8. The air film 8 forms a heat-insulating effect, which can reduce the heat transfer from the extremely high aerodynamic heat 9 caused by hypersonic flight to the inside of the body 11.

[0028] The relevant quantitative parameter information is as follows:

[0029] The average pore size of the porous medium material of the aircraft end body 11 is generally 50 microns.

[0030] In film cooling, the thickness of the film attached to the end of the aircraft is generally in the millimeter range. If the film thickness H is 1 mm, the thermal conductivity of water vapor and carbon dioxide λ is 0.023-0.024 W / (m·K). If it is 0.0235 W / (m·K), then only 1 square meter (S=1m 2 )The thermal resistance of the thermal protection layer θ=H / λS, and its value is only 0.04K / W. It is difficult to provide effective thermal protection for today's hypersonic aircraft by relying solely on traditional air film 8 cooling.

[0031] The specific calculation is as follows: the melting point of common aviation high temperature materials such as titanium is 1668℃, and the failure temperature should be below 1000℃. Therefore, in order to ensure the long-term stability of the fuselage 11, it is assumed that the temperature inside the air film 8 is 1000℃, and the temperature outside the air film 8 can reach 4000℃ when the aircraft is flying at hypersonic speed. Therefore, under the temperature difference △T of 3000℃, the heat penetrating into 1 square meter of air film 8 and reaching the fuselage 11 is Q=△T / θ, which is as high as 75000W, and its heat flux density q is as high as 75000W / m 2 That is to say, the interior of the body 11 needs to be protected from the infiltrated 75000W / m 2 The heat flow must be effectively dissipated, otherwise the temperature of the body 11 will further increase, causing the material to fail and the body to disintegrate.

[0032] In the face of such a huge heat flux density, gas cooling alone is ineffective. Internal cooling conforms to the convection cooling formula q = h (T h -T l ), where h is the convective heat transfer coefficient, which is only ~60 W / (K·m for gas cooling without phase change. 2 ), T h is the heat source temperature (here the body temperature is 1000°C), T l is the temperature of the cooling source (here is the temperature of the gas used for cooling), in order to achieve 75000W / m 2 The heat dissipation capacity requires T h With T l The temperature difference reaches 1250°C (requiring the temperature of gas cooling to reach -250°C), which is impossible, and the above calculation is only an ideal state. The actual heat penetrating into the body 11 will be greater, so it is invalid to rely solely on gas cooling for thermal protection. In other words, it is impossible to cool down by relying solely on the phase change of dry ice.

[0033] This project introduces phase change cooling of liquid water and dry ice sublimation cooling. The convective heat transfer coefficient h of phase change jet cooling is at least 10 3 -10 4 W / (K·m 2), which is hundreds or even thousands of times that of non-phase change gas cooling. This is due to the latent heat of phase change of the substance (the latent heat of vaporization of dry ice is 137kcal / kg, and that of water is 539kcal / kg). The two substances work together to face 75000W / m 2 The heat dissipation requirement is only T h With T l The temperature difference of a few or tens of degrees Celsius can be satisfied. For a high-temperature heat source of 1000℃, this can be achieved very easily, and there is a large design margin (usually the temperature difference can reach more than 900 degrees because the operating temperature of dry ice and the water temperature will not exceed 100℃), which can calmly face more severe situations.

Claims

1. A heterogeneous immersion dry ice spraying two-stage cooling device for thermal protection of the end of a hypersonic vehicle, characterized in that: The machine body comprises a porous medium end head, the front part of which is provided with a porous medium end head, the inner and outer walls of the medium are connected through a plurality of micropores in the medium, and the outer surface of the porous medium end head is provided with a low melting point coating to seal the outer openings of the micropores; a continuous liquid phase water cavity with a pressure-bearing structure is provided in the porous medium end head, and a dry ice blasting device is also provided in the machine body, the nozzle of the dry ice blasting device is immersed in the water in the continuous liquid phase water cavity, and the front of the nozzle spraying is aligned with the center of the porous medium end head.

2. A heterogeneous immersion dry ice spraying two-stage cooling device for thermal protection of a hypersonic vehicle tip according to claim 1, characterized in that: A carbon dioxide gas storage tank, a temperature sensor, a pressure sensor, a flow meter and a control valve are also provided in the machine body; the outlet of the carbon dioxide gas storage tank is connected to the inlet of the control valve, the outlet of the control valve is connected to the dry ice blasting nozzle via the flow meter, and the temperature sensor and the pressure sensor are arranged on the pipeline connecting the outlet of the control valve and the nozzle.

3. The heterogeneous immersion dry ice blasting two-stage cooling method for hypersonic vehicle tip thermal protection according to claim 1, characterized in that: The surface of the porous medium end is densely covered with micropores, and the average pore diameter of the micropores is 0.05 mm.

4. A method for using a heterogeneous immersion dry ice spraying two-stage cooling device for thermal protection of a hypersonic vehicle tip according to any one of claims 1 to 3, characterized in that: When a hypersonic vehicle performs hypersonic flight, friction with the external atmosphere generates aerodynamic heat. After the low-melting-point coating melts, the heterogeneous immersion dry ice jet two-stage cooling device for thermal protection of the aircraft end is started. Dry ice particles are sprayed out from a dry ice jet nozzle immersed in liquid water. The dry ice particles are heated and sublimated / vaporized at the moment of spraying to generate a dry ice jet downstream airflow. At the same time, the continuous liquid water cavity used for immersion is also vaporized into water vapor by the action of aerodynamic heat. A mixed gas of carbon dioxide and water vapor is generated on the inside of the body of the porous medium end of the hypersonic aircraft. The inner and outer walls of the medium are connected through multiple micropores in the porous medium end. The mixed gas penetrates from the inside of the body through the porous medium to the outside of the body. The water vapor and carbon dioxide mixture generates a mixed thermal protection gas layer of water vapor and carbon dioxide attached to the outside of the body, forming a cooling thermal protection air film.

5. A two-stage cooling method of heterogeneous immersion dry ice spraying for thermal protection of a hypersonic vehicle tip according to claim 4, characterized in that: The mass ratio of dry ice to water is 1:(0.6-2.2).

6. The heterogeneous immersion dry ice blasting two-stage cooling method for hypersonic vehicle tip thermal protection according to claim 4, characterized in that: The melting temperature of the low melting point coating is 400-600℃.

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