Two-phase open-loop pump-driven hypersonic vehicle thermal protection system

Abstract

The present application relates to the technical field of hypersonic vehicle thermal protection, and proposes a hypersonic vehicle thermal protection system based on open pump-driven two-phase, which comprises a wall thermal protection structure, a heat exchange unit and a coolant circulation supply unit. The hypersonic vehicle thermal protection system based on open pump-driven two-phase combines the design idea of flat plate heat pipe and the mode of pump-driven coolant phase change heat transfer to cool the cabin wall, which can effectively block the heat flow into the cabin and release the fully utilized coolant to the outside of the cabin. At the same time, the idea of cold cascade utilization is adopted, and the coolant is used to efficiently convectively cool part of high-heat-flux high-precision electronic equipment before cooling the cabin wall, further ensuring that the cabin environment temperature will not exceed the normal range.

Description

TECHNICAL FIELD

[0001] The present application relates to the technical field of hypersonic vehicle thermal protection, and in particular to a hypersonic vehicle thermal protection system based on open pump-driven two-phase. BACKGROUND

[0002] With the continuous breakthrough of technology and the continuous expansion of application fields, hypersonic vehicles will play an increasingly important role in civilian fields, representing an important direction in the future aerospace field. Hypersonic vehicles achieve transatmospheric and in-atmosphere maneuvering flight at a flight speed exceeding 5Ma, with characteristics of high speed, strong maneuverability, super-long range, and strong penetration. Hypersonic vehicles are developing towards longer endurance and higher speed. However, further increases in flight speed will result in increased aerodynamic heating, and future electronic equipment to be installed will be more precise and require higher cabin environmental temperature. In addition to the need for special cooling methods for thermal protection of the sharp leading edge stagnation point of the vehicle surface, the temperature control of the cabin wall and inside the cabin urgently requires a stable and efficient thermal protection system.

[0003] When the Mach number Ma of the hypersonic vehicle reaches 10 or more, the total temperature of the gas around the cabin approaches 4000K, and the thermal load on the vehicle is large. The vehicle flies for a longer time, in a more severe environment, and the amount of heat leakage through the thermal barrier layer into the cabin increases. If there is no external available cold source, the temperature inside the cabin will often exceed the allowable temperature range of electronic equipment. Currently, there are some technical means in the prior art to solve the cabin temperature rise caused by high speed, long time, and long distance flight of hypersonic vehicles.

[0004] Chinese patent application CN117326047A discloses a passive thermal insulation and active convection cooling combined thermal protection system applied to long-time high-speed vehicles and a corresponding control method. This cooling method can realize autonomous control of coolant flow, has good mechanical properties and good thermal insulation effect, but it is only suitable for cooling small-area high heat flux density areas and cannot play a role in thermal insulation of the fuselage of a hypersonic vehicle. Since the cooling structure is not convenient for large-area installation, the temperature inside the cabin of a hypersonic vehicle after long-time flight will still exceed the allowable operating temperature of the equipment. At the same time, the main components of the cooling structure need to be connected by multiple perforations and sutures, and the processing and installation process is also relatively complex.

[0005] A wall surface structure based on metal nano-modified self-suction sweat cooling and a hypersonic aircraft are disclosed in Chinese patent application CN117944865A. The wall surface structure can not only supply coolant without pump power consumption, but also efficiently and quickly cool the wall surface under a high-heat-flux external environment, thereby having a strong thermal protection effect. However, as the aerodynamic heat of the hypersonic aircraft increases, the self-suction porous structure is difficult to match the increase in heat flux density, and the phase change interface may be locally dry, thereby deteriorating the cabin environment. Moreover, to achieve a good cooling effect, the design thickness of the wall surface structure is relatively large, and the volume of the entire system is also large, which will increase the weight of the aircraft and is not conducive to the improvement of the maneuverability of the aircraft.

[0006] In summary, the existing hypersonic aircrafts in the prior art mostly use passive, semi-active heat protection measures or active cooling measures in local areas, have a large volume, are difficult to arrange in a large area, and have unstable ability to regulate the temperature in the cabin of the aircraft, and cannot simultaneously meet the temperature range requirements of the wall surface and high-precision electronic equipment in the thermal protection or system level. SUMMARY

[0007] Based on the needs and deficiencies of the prior art, the present application proposes a hypersonic aircraft thermal protection system based on open pump-driven two-phase, which uses an open pump to drive the refrigerant to control the phase change and efficient convection cooling of the wall surface and high-precision electronic equipment of the hypersonic aircraft. The technical solution of the present application is as follows:

[0008] A hypersonic aircraft thermal protection system based on open pump-driven two-phase, comprising a wall surface thermal protection structure, a heat exchange unit, and a coolant circulation supply unit;

[0009] The coolant circulation supply unit supplies coolant to the heat exchange unit and the wall surface thermal protection structure;

[0010] The wall surface thermal protection structure comprises, from inside to outside, a thermal insulation layer, an alloy inner shell, a coolant flow channel, a metal wire mesh capillary core, and a cabin wall surface, and the flow direction of the coolant in the wall surface thermal protection structure is opposite to the flight direction of the hypersonic aircraft;

[0011] The heat exchange unit is used for cooling high-precision electronic equipment.

[0012] Preferably, the wall surface thermal protection structure further comprises support and sealing elements arranged on the front side, the rear side, the left side, and the right side.

[0013] Preferably, the coolant flow channel is provided with support ribs for supporting the metal wire mesh capillary core and dividing the coolant flow channel into a plurality of sub-flow channels.

[0014] Preferably, the front and rear of the coolant flow channel are not provided with the support ribs, thereby forming an inlet liquid storage cavity and an outlet liquid storage cavity, respectively.

[0015] Preferably, the outer surface of the wire mesh capillary core or the inner surface of the cabin wall surface is provided with a plurality of vapor grooves.

[0016] Preferably, the support seal on the rear side of the wall thermal protection structure is provided with an exhaust port and a coolant outlet, and the support seal on the front side of the wall thermal protection structure is provided with a coolant inlet.

[0017] Preferably, the coolant circulation supply unit comprises a coolant, a liquid storage container, a circulation pipeline and a flow regulating pump, and the coolant flows through the flow regulating pump, the heat exchange unit, the wall thermal protection structure and the liquid storage container in sequence through the circulation pipeline.

[0018] Preferably, the heat exchange unit is a cold plate heat exchanger or a shell-and-tube heat exchanger.

[0019] Preferably, the support seal is made of an alloy material and is connected to the cabin wall surface and the aluminum alloy inner shell by welding or bolting.

[0020] Preferably, the support ribs are made of the same material as the aluminum alloy inner shell and are integrally formed with the aluminum alloy inner shell, and the outer surface and the inner surface of the wire mesh capillary core are tightly matched with the cabin wall surface and the support ribs, respectively.

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

[0022] 1. The open-pump-driven two-phase-based hypersonic vehicle thermal protection system provided by the present application can not only efficiently control the wall temperature of the hypersonic vehicle, but also actively and efficiently cool the core high-precision electronic equipment in the cabin, realize the step-by-step utilization of the cooling capacity of the coolant, and ensure that the temperature in the cabin is maintained within a suitable range.

[0023] 2. Compared with the traditional passive thermal insulation technology and the self-suction sweating cooling technology, the open-pump-driven two-phase-based hypersonic vehicle thermal protection system provided by the present application can avoid the situations of excessive heat flow and over-temperature failure, insufficient or mismatched coolant supply, and can reasonably adjust the carrying amount of the coolant according to the actual flight requirements, and cooperate with the precise control of the coolant flow in the coolant circulation supply unit, thereby greatly improving the flight duration of the hypersonic vehicle.

[0024] 3. The wall thermal protection structure provided by the present application occupies a small and thin space, is flexible to install, and the coolant is discharged outside the cabin after being fully utilized, thereby improving the space utilization rate in the cabin under limited conditions, and being conducive to realizing large-area arrangement and coolant coverage, and being able to effectively alleviate the uneven distribution of the thermal load of the cabin wall of the hypersonic vehicle.

[0025] 4. The wall thermal protection structure has an inlet liquid storage cavity and an outlet liquid storage cavity, and the flow direction of the coolant is opposite to the flight direction of the hypersonic aircraft, which fully utilizes the acceleration effect of the hypersonic aircraft and the inertia of the coolant, saves pump power consumption, and improves system operation adaptability. BRIEF DESCRIPTION OF DRAWINGS

[0026] In order to more clearly illustrate the technical solutions in the embodiments of the present application or the prior art, the drawings needed in the embodiments will be briefly introduced below. The features and advantages of the present application can be more clearly understood by referring to the drawings. The drawings are schematic and should not be construed as any limitation on the present application. For those skilled in the art, other drawings can be obtained without creative labor on the basis of these drawings.

[0027] Figure 1 is a structural schematic diagram of a hypersonic aircraft thermal protection system based on an open pump-driven two-phase in the embodiment.

[0028] Figure 2 is a schematic diagram of two setting modes of the wall thermal protection structure steam channel in the embodiment.

[0029] Figure 3 is a schematic diagram of the wall thermal protection structure circumferentially arranged in the cabin of the hypersonic aircraft in the embodiment.

[0030] Among them, 1-cooling water, 2-liquid storage container, 3-circulation pipeline, 4-flow regulating pump, 5-cold plate heat exchanger, 6-high-precision electronic equipment, 7-coolant inlet, 8-front side sealing support, 9-inlet liquid storage cavity, 10-metal wire mesh capillary core, 11-cabin wall, 12-exhaust port, 13-rear side sealing support, 14-support rib, 15-outlet liquid storage cavity, 16-coolant outlet, 17-aluminum alloy inner shell, 18-thermal insulation layer, 19-cabin wall steam channel, 20-coolant flow channel, 21-metal wire mesh capillary core steam channel, 22-left side sealing support, 23-right side sealing support. DETAILED DESCRIPTION

[0031] In order to more clearly illustrate the above-mentioned purposes, features and advantages of the present application, the present application will be further described in detail below in combination with the drawings and specific embodiments. It should be noted that the embodiments of the present application and the features in the embodiments can be combined with each other without conflict.

[0032] Many specific details are set forth in the following description in order to provide a full understanding of the invention. However, the invention may also be practiced in other ways different from those described herein, and therefore the scope of protection of the invention is not limited to the specific embodiments disclosed below.

[0033] like Figures 1-2 As shown, this embodiment proposes a hypersonic vehicle thermal management system based on an open-loop pump-driven two-phase system, including cooling water 1, a liquid storage container 2, a circulation pipeline 3, a flow regulating pump 4, a cold plate heat exchanger 5, high-precision electronic equipment 6, a coolant inlet 7, a front sealing support 8, an inlet liquid storage chamber 9, a metal wire mesh capillary core 10, a cabin wall 11, an exhaust port 12, a rear sealing support 13, a support rib 14, an outlet liquid storage chamber 15, a coolant outlet 16, an aluminum alloy inner shell 17, a heat insulation layer 18, a coolant flow channel 20, a left sealing support 22, and a right sealing support 23.

[0034] Figure 2 The diagram illustrates two configurations of the steam channels: one is a steam channel 19 located on the inner surface of the cabin wall, and the other is a steam channel 21 located on the outer surface of the wire mesh capillary core. In the region between the wire mesh capillary core 10 and the cabin wall 11, the cooling water 1 undergoes a phase change, transforming from a liquid to a gaseous state. Utilizing the sensible heat and latent heat of phase change of the liquid, the external aerodynamic heat entering the cabin through the cabin wall 11 is dissipated. The vaporized cooling water 1 can be collected through either the steam channel 19 or the wire mesh capillary core steam channel 21, and the high-temperature gas is discharged through the exhaust port 12 on the rear sealing support 13.

[0035] The different structures described above can be connected in the following ways:

[0036] The front sealing support 8, the rear sealing support 13, the left sealing support 22, and the right sealing support 23 can be made of alloy materials that can withstand certain temperatures and thermal strains, such as aluminum alloys. These sealing supports are connected to the cabin wall 11 by welding or bolting.

[0037] The support rib 14 and the aluminum alloy inner shell 17 are connected by welding or bolts. Preferably, both are made of the same aluminum alloy and are integrally formed during processing.

[0038] The front sealing support 8, rear sealing support 13, left sealing support 22, and right sealing support 23 are connected to the aluminum alloy inner shell 17 by welding or bolting. Preferably, these sealing supports and the aluminum alloy inner shell 17 can also be integrally formed by stamping.

[0039] The heat insulation layer 18 is connected with the aluminum alloy inner shell 17 by means of adhesion, and the heat insulation layer 18 can further reduce the penetration of heat flow to the cabin. The heat insulation layer 18 can be made of common heat insulation materials such as glass wool and foamed plastic.

[0040] The metal wire mesh wick 10 is tightly clamped between the cabin wall 11 and the support rib 14 through close cooperation.

[0041] As shown in Figure 3 , by arranging the wall thermal protection structure circumferentially on the cabin of the hypersonic vehicle, comprehensive coverage can be achieved, thereby realizing circumferential large-area thermal protection. Each wall thermal protection structure can be processed and formed in blocks, and then installed in cooperation. The shape of the hypersonic vehicle can not be completely circular, for example, it can be elliptical or round.

[0042] In the case of ensuring the heat dissipation effect, in order to reduce the mass of the entire wall thermal protection structure, the total thickness of the metal wire mesh wick 10, the support rib 14 and the aluminum alloy inner shell 17 is controlled to be between 3-5 mm, which embodies the characteristics of lightness and thinness.

[0043] The size of the cabin wall steam channel 19 and the metal wire mesh wick steam channel 21 is designed according to the actual situation to achieve better heat transfer effect. Preferably, the surface of these steam channels is treated to generate macro or micro-nano structures that are beneficial to heat exchange, for example, small solid carriers with different thermal conductivities are arranged on the surface of the steam channel, the surface of the steam channel is roughened, a thermal conductive coating (such as graphite, carbon nanotubes) and a film are added, etc.

[0044] The metal wire mesh wick 10 is sintered from several thin layers of metal wire mesh. The wire diameter and aperture of metal wire mesh with different mesh counts are different. The mesh count of the metal wire mesh is 200, and the heat dissipation heat flow density can reach 50 W / cm 2 . To improve the capillary force of the metal wire mesh, physical and chemical treatment methods can be used to generate small particles or linear and needle-shaped structures on the metal wire mesh to increase the bubble nucleation sites in the boiling heat transfer process; or metal wire meshes with different mesh counts can be mixed according to certain rules and then sintered to form a suction gradient, which produces a greater suction force on the cooling water 1 on the side close to the support rib 14, which is beneficial to the planar spreading of the cooling water 1 on the metal wire mesh and the flow to the cabin wall, and on the side close to the cabin wall 11, the bubbles generated by phase change need to be detached from the metal wire mesh and enter the cabin wall steam channel 19 or the metal wire mesh wick steam channel 21.

[0045] Cooling water 1 enters the inlet liquid storage cavity 9 through the coolant inlet 7, and when flowing through the coolant flow channel 20, most of the cooling water is sucked to the interface area with the cabin wall 11 by the metal wire mesh capillary core 10 to exchange heat through boiling phase change, and the remaining cooling water without phase change enters the outlet liquid storage cavity 15, and then returns to the liquid storage container 2 through the coolant outlet 16.

[0046] Compared with the coolant outlet 16, the coolant inlet 7 is located at a more forward position of the hypersonic aircraft, that is, the cooling water flowing trend in the coolant flow channel 20 is opposite to the flight direction of the hypersonic aircraft. The main function of the inlet liquid storage cavity 9 and the outlet liquid storage cavity 15 is to distribute and collect the cooling water, and secondarily, when the hypersonic aircraft is accelerated to overload, a small amount of cooling water stored in the inlet liquid storage cavity 9 can supplement the coolant flow channel 20, offset the flow channel resistance of the coolant flow channel 20 by using the acceleration effect and inertia, and facilitate the matching regulation and control of the flow by the flow regulating pump 4.

[0047] For the entire open-pump-driven two-phase-based hypersonic aircraft thermal management system, the cooling water 1 in the liquid storage container 2 is delivered to the cold plate heat exchanger 5 in the area where the high-precision electronic equipment 6 is located by the flow regulating pump 4, and first cools the high-precision electronic equipment 6 which has a relatively high temperature requirement, and the temperature of the cooling water will rise to 40-50 DEG C, and then enters the wall thermal protection structure through the circulating pipeline 3 from the coolant inlet 7 to efficiently cool the cabin wall 11 through phase change, most of the cooling water becomes water vapor and is discharged through the exhaust port 12, and a small part returns to the liquid storage container 2 through the circulating pipeline 3 to continue to participate in the subsequent circulating cooling process.

[0048] The open-pump-driven two-phase-based hypersonic aircraft thermal management system in the embodiment fully utilizes the refrigerant cooling capacity, and utilizes the sensible heat and latent heat of the cooling water in stages, controls the temperature of the high-precision electronic equipment, and efficiently cools the cabin wall, so as to accurately regulate and control the temperature in the cabin to prevent the temperature from exceeding the working range. The liquid storage container should be placed at a position more forward than the coolant inlet to utilize the acceleration effect of the aircraft and the inertia of the cooling water to promote the delivery of the cooling water and effectively reduce the system pump power consumption.

[0049] In the present application, unless otherwise clearly specified and limited, the terms such as "mounting", "connection", "connecting", "fixing" and the like should be understood in a broad sense, for example, can be fixed connection, can also be detachable connection, or can be integrated; can be mechanical connection, or can be electrical connection; can be directly connected, or can be indirectly connected through an intermediate medium; can be the communication inside two elements or the interaction relationship between two elements. For those skilled in the art, the specific meaning of the above terms in the present application can be understood according to the specific circumstances.

[0050] In the present application, unless otherwise explicitly specified and limited, "on" or "under" of a first feature to a second feature can include that the first and second features are in direct contact, or that the first and second features are not in direct contact but are in contact through another feature between them. Moreover, "on", "above" and "over" of a first feature to a second feature includes that the first feature is directly above and obliquely above the second feature, or only indicates that the first feature is higher in horizontal height than the second feature. "Under", "below" and "underneath" of a first feature to a second feature includes that the first feature is directly below and obliquely below the second feature, or only indicates that the first feature is lower in horizontal height than the second feature.

[0051] In the present application, the terms "first", "second", "third", "fourth" are only for descriptive purpose, and should not be understood as indicating or implying relative importance. The term "a plurality of" means two or more, unless otherwise explicitly limited.

[0052] The above only describes the preferred embodiments of the present application and is not intended to limit the present application. The present application can have various modifications and changes for those skilled in the art. Any modification, equivalent replacement, improvement, etc. made within the spirit and principle of the present application shall be included in the protection scope of the present application.

Claims

1. A thermal protection system for hypersonic vehicles based on open-pump driven two-phase systems, characterized in that, This includes a wall thermal protection structure, a heat exchange unit, and a coolant circulation supply unit; The coolant circulation supply unit supplies coolant to the heat exchange unit and the wall thermal protection structure; The wall thermal protection structure includes, from the inside out, a heat insulation layer, an alloy inner shell, a coolant channel, a metal wire mesh capillary core, and a cabin wall. The coolant flow direction in the wall thermal protection structure is opposite to the flight direction of the hypersonic vehicle. The heat exchange unit is used for cooling high-precision electronic equipment; The wall thermal protection structure also includes supporting seals disposed on the front, rear, left, and right sides; The coolant flow channel is provided with support ribs to support the metal wire mesh capillary core and to divide the coolant flow channel into several sub-channels; The supporting ribs are not provided in the front and rear parts of the coolant flow channel, thus forming an inlet liquid storage chamber and an outlet liquid storage chamber, respectively; The rear support seal of the wall thermal protection structure is provided with an exhaust port and a coolant outlet, and the front support seal of the wall thermal protection structure is provided with a coolant inlet.

2. The hypersonic vehicle thermal protection system based on open-pump driven two-phase as described in claim 1, characterized in that, The outer surface of the metal wire mesh capillary core or the inner surface of the cabin wall is provided with multiple steam channels.

3. The hypersonic vehicle thermal protection system based on open-pump driven two-phase as described in claim 1, characterized in that, The coolant circulation supply unit includes coolant, a storage container, a circulation pipeline, and a flow regulating pump. The coolant flows sequentially through the flow regulating pump, the heat exchange unit, the wall thermal protection structure, and the storage container via the circulation pipeline.

4. The hypersonic vehicle thermal protection system based on open-type pump-driven two-phase as described in claim 1, characterized in that, The heat exchange unit is a cold plate heat exchanger or a shell-and-tube heat exchanger.

5. The hypersonic vehicle thermal protection system based on open-pump driven two-phase as described in claim 1, characterized in that, The supporting seal is made of alloy material and is connected to the cabin wall and aluminum alloy inner shell by welding or bolting.

6. The hypersonic vehicle thermal protection system based on open-pump driven two-phase as described in claim 1, characterized in that, The material of the support rib is the same as that of the alloy inner shell, and it is integrally formed with the alloy inner shell. The outer and inner surfaces of the metal wire mesh capillary core are tightly fitted with the cabin wall and the support rib, respectively.

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