A cooling system and aircraft

Abstract

This application relates to the technical field of spacecraft, and in particular to a cooling system and a spacecraft. The system is applied to a spacecraft whose fuselage consists of an inner main body and an outer skin connected together. A flow channel extending from the head to the tail of the fuselage is provided between the main body and the skin. The fuselage has film vents penetrating the main body and the skin. The cooling system is located within the fuselage cavity and includes a storage component, a cooling component, and a circulation component. The storage component stores the cooling medium, and the cooling component cools the cooling medium. The circulation component includes a power pump and a turbine. The power pump pressurizes the cooling medium and pumps the pressurized cooling medium from the head of the fuselage into the flow channel. The turbine's inlet communicates with the flow channel between the tail and neck of the fuselage, and the turbine's exhaust outlet communicates with the film vents at one end of the fuselage cavity. The system of this application can rapidly reduce the temperature of the spacecraft's head.

Description

Technical Field

[0001] This application relates to the technical field of spacecraft, and in particular to a cooling system and a spacecraft. Background Technology

[0002] Near space refers to the airspace 20 to 100 kilometers above the Earth's surface. This airspace includes the stratosphere (18 to 55 kilometers above the ground), the mesosphere (55 to 85 kilometers above the ground), and a small portion of the thermosphere (85 to 800 kilometers above the ground), and is predominantly composed of homogeneous atmosphere. High-speed near-space vehicles are aircraft that can or will only operate in near space for extended periods, or suborbital vehicles, or hypersonic cruise vehicles that operate in near space; they have extremely important applications in the defense field.

[0003] High-speed near-space vehicles generate a large amount of heat through friction with the atmosphere during high-speed flight, causing a rapid increase in the temperature of the vehicle's nose section. The nose section materials may fail due to excessive heat, reducing the vehicle's safety. To improve safety, high-temperature resistant alloys or flame-retardant materials are commonly used to protect the nose section. However, high-temperature resistant alloys have a high density, significantly increasing the overall mass of the vehicle, while flame-retardant materials are disposable and difficult to manufacture.

[0004] Therefore, how to quickly cool down the nose of an aircraft is a technical challenge. Summary of the Invention

[0005] Therefore, it is necessary to provide a cooling system and aircraft capable of rapidly cooling the nose of an aircraft in response to the aforementioned technical problems.

[0006] In a first aspect, this application provides a cooling system applied to an aircraft. The aircraft's fuselage is composed of an inner main body and an outer skin connected together. A flow channel extending from the head to the tail of the fuselage is provided between the main body and the skin. The fuselage has film vents penetrating the main body and the skin. The cooling system is located within the interior cavity of the fuselage and includes a storage component, a cooling component, and a circulation component, wherein:

[0007] The storage component is used to store the cooling medium, and the cooling component is used to cool the cooling medium.

[0008] The circulation assembly includes a power pump and a turbine, wherein the power pump is used to pressurize the cooling medium and pump the pressurized cooling medium from the head of the machine body into the flow channel;

[0009] The turbine's air inlet is connected to the flow channel between the tail and neck of the machine body, and the turbine's exhaust port is connected to the air film hole at one end of the machine body's inner cavity.

[0010] In one embodiment, the output shaft of the turbine is connected to the power shaft of the power pump, and the turbine's air inlet is equipped with a starter.

[0011] In one embodiment, the system further includes an expansion valve, the inlet of which is connected to the exhaust port of the turbine, and the outlet of which is connected to one end of the film gas hole located in the inner cavity of the machine body.

[0012] In one embodiment, a collector is provided on one side of the main body located within the inner cavity of the machine body, the collector being a hollow cavity;

[0013] The collector is connected to the outlet of the expansion valve and to one end of the air film orifice located in the inner cavity of the machine body.

[0014] In one embodiment, the storage assembly includes a storage tank, a filling line, and a pressure relief line, wherein:

[0015] The storage tank is used to store the cooling medium. The storage tank is connected to a filling pipeline, which is used to fill the storage tank with the cooling medium. The storage tank is also connected to a pressure relief pipeline, which is used to relieve pressure in the storage tank.

[0016] The storage tank is connected to the inlet of the power pump via a flow control valve.

[0017] In one embodiment, the storage box comprises an inner box and an outer box, wherein:

[0018] The inner box is located inside the outer box, and the inner box and the outer box are connected by a connector;

[0019] The inner casing is used to store the cooling medium, and the space between the inner casing and the outer casing is a vacuum.

[0020] In one embodiment, the cooling assembly includes a refrigerator, a circulating pump, a heat shield, and an air-cooled panel, wherein:

[0021] The circulating pump and the refrigeration unit are both located inside the heat insulation box, and the air-cooled screen is located inside the storage box;

[0022] The input end of the circulating pump is connected to the return end of the air-cooled screen, the output end of the circulating pump is connected to the input end of the refrigerator, and the output end of the refrigerator is connected to the inlet end of the air-cooled screen.

[0023] In one embodiment, the cooling medium is liquid ammonia.

[0024] Secondly, this application also provides an aircraft, the fuselage of which is composed of an inner main body and an outer skin connected together, and a flow channel extending from the head to the tail of the fuselage is provided between the main body and the skin;

[0025] The body has air film pores that connect the main body and the skin;

[0026] The aircraft is also equipped with a cooling system as described in any one of the first aspects above.

[0027] In one embodiment, the air film pores are not connected to the flow channel.

[0028] The cooling assembly cools the cooling medium stored within the storage assembly, enabling the storage assembly to store and supply a cooler medium to the flow channel. A power pump pumps the cooler medium from the nose of the aircraft into the flow channel. Since the flow channel is located between the inner body and outer skin of the aircraft fuselage, and extends from the nose to the tail, the cooler medium absorbs heat from the nose of the aircraft fuselage, transforming into a hot gas with higher temperature and pressure. This hot gas then flows along the flow channel towards the tail of the aircraft. Furthermore, since the turbine's air inlet is connected to the flow channel between the tail and neck of the fuselage, hot gas can enter the turbine's air inlet and drive the turbine blades to rotate and perform work. After driving the turbine, the hot gas's temperature and pressure are further reduced, becoming neutral gas. The neutral gas is discharged from the turbine to the film cooling hole and then from the film cooling hole to the outside of the fuselage. Since the film cooling hole penetrates the main body and skin of the aircraft, the neutral gas will further absorb heat from the fuselage when it is discharged from the film cooling hole, thereby further cooling the fuselage. In other words, during the process of the cooling medium from the nose of the fuselage to the turbine, the cooling medium can absorb heat to cool the nose and neck of the fuselage for the first time; and during the process of the neutral gas being cooled and depressurized by the turbine and then discharged from the film cooling hole, the neutral gas can also cool the fuselage for the second time. Therefore, the cooling system provided in this application can quickly cool the nose of the fuselage, thereby improving the stability and safety of the nose of the aircraft. Attached Figure Description

[0029] To more clearly illustrate the technical solutions in the embodiments or related technologies of this application, the accompanying drawings used in the description of the embodiments or related technologies will be briefly introduced below. Obviously, the accompanying drawings described below are only some embodiments of this application. For those skilled in the art, other drawings can be obtained based on these drawings without creative effort.

[0030] Figure 1 This is a cross-sectional schematic diagram of the head of the machine body in one embodiment;

[0031] Figure 2 This is a schematic diagram illustrating the skin and flow channel structure in one embodiment;

[0032] Figure 3 This is a schematic cross-sectional view of the body illustrating the structure of the collector and the air film pores in one embodiment;

[0033] Figure 4 This is a schematic diagram illustrating the loop relationships of various components in a cooling system in one embodiment;

[0034] Figure 5 This is a schematic diagram of the storage component in one embodiment;

[0035] Figure 6 This is a schematic diagram of the cooling component in one embodiment;

[0036] Figure 7 This is a schematic diagram of the connection relationship of the loop components in one embodiment;

[0037] Figure 8 This is a schematic diagram of the state of the circulating medium at each node in the cooling system in another embodiment.

[0038] Figure label:

[0039] 100. Body; 110. Main body; 120. Skin; 130. Flow channel; 140. Collector; 150. Film gas vent; 200. Storage assembly; 210. Storage tank; 211. Inner casing; 212. Outer casing; 213. Connecting parts; 220. Filling pipeline; 221. Level gauge; 222. Check valve; 223. Filling port; 230. Pressure relief pipeline; 231. Back pressure valve; 232. Pressure relief valve; 233. Mass flow meter; 234, pressure relief port; 240, flow control valve; 300, cooling assembly; 310, refrigeration unit; 320, circulating pump; 330, quick-connect switch valve; 340, heat insulation box; 350, air-cooled screen; 400, circulation assembly; 410, power pump; 411, liquid inlet; 412, liquid outlet; 420, turbine; 421, air inlet; 422, exhaust port; 430, expansion valve; 440, starter. Detailed Implementation

[0040] To make the above-mentioned objectives, features, and advantages of this application more apparent and understandable, the specific embodiments of this application are described in detail below with reference to the accompanying drawings. Many specific details are set forth in the following description to provide a thorough understanding of this application. However, this application can be implemented in many other ways different from those described herein, and those skilled in the art can make similar modifications without departing from the spirit of this application; it should be understood that the specific embodiments described herein are merely illustrative of this application and are not intended to limit this application.

[0041] In the description of this application, it should be understood that if terms such as "center", "longitudinal", "lateral", "length", "width", "thickness", "upper", "lower", "front", "rear", "left", "right", "vertical", "horizontal", "top", "bottom", "inner", "outer", "clockwise", "counterclockwise", "axial", "radial", "circumferential" appear, these terms indicate the orientation or positional relationship based on the orientation or positional relationship shown in the accompanying drawings, and are only for the convenience of describing this application and simplifying the description, and do not indicate or imply that the device or element referred to must have a specific orientation, or be constructed and operated in a specific orientation, and therefore should not be construed as a limitation of this application.

[0042] Furthermore, where the terms "first" and "second" appear, these terms are for descriptive purposes only and should not be construed as indicating or implying relative importance or implicitly specifying the number of technical features indicated. Thus, a feature defined with "first" or "second" may explicitly or implicitly include at least one of that feature. In the description of this application, where the term "multiple" appears, "multiple" means at least two, such as two, three, etc., unless otherwise explicitly specified.

[0043] In this application, unless otherwise expressly specified and limited, the terms "installation," "connection," "joining," and "fixing," etc., should be interpreted broadly. For example, they can refer to a fixed connection, a detachable connection, or an integral part; they can refer to a mechanical connection or an electrical connection; they can refer to a direct connection or an indirect connection through an intermediate medium; they can refer to the internal communication of two components or the interaction between two components, unless otherwise expressly limited. Those skilled in the art can understand the specific meaning of the above terms in this application based on the specific circumstances.

[0044] In this application, unless otherwise expressly specified and limited, the use of descriptions such as "above" or "below" the second feature indicates that the first and second features are in direct contact or indirect contact via an intermediate medium. Furthermore, "above," "on top of," and "over" the second feature can mean that the first feature is directly above or diagonally above the second feature, or simply that the first feature is at a higher horizontal level than the second feature. Similarly, "below," "below," and "under" the second feature can mean that the first feature is directly below or diagonally below the second feature, or simply that the first feature is at a lower horizontal level than the second feature.

[0045] It should be noted that if an element is referred to as being "fixed to" or "set on" another element, it can be directly on the other element or there may be an intervening element. If an element is considered to be "connected to" another element, it can be directly connected to the other element or there may be an intervening element. If so, the terms "vertical," "horizontal," "upper," "lower," "left," "right," and similar expressions used in this application are for illustrative purposes only and do not represent the only possible implementation.

[0046] The cooling system provided in this application can be applied to near-space vehicles; here, the necessary structural description of the vehicle to which the cooling system is applicable will be given first. In an exemplary embodiment, refer to... Figure 1 and Figure 2 The aircraft fuselage 100 is composed of an inner main body 110 and an outer skin 120. A flow channel 130 extending from the head to the tail of the fuselage 100 is provided between the main body 110 and the skin 120. The fuselage 100 has an air film hole 150 that runs through the main body 110 and the skin 120.

[0047] In one embodiment, the nose of the aircraft can be conical. The material of the nose can be, but is not limited to, TA15 titanium alloy, and its processing technology includes the following steps: ① Grooving is performed on a flat plate by machining to form a flat plate containing a channel structure (flow channel); ② The skin 120 is welded to the channel structure of the main body 110 to form a plate of the main body 110 containing internal channels. The welding method includes, but is not limited to, brazing and diffusion welding; ③ The plate of the main body 110 containing internal channels is stamped to form the fuselage 100 of the aircraft; ④ The joints are welded firmly to meet the first-level weld quality requirements. Furthermore, the air film holes 150 opened on the fuselage 100 are not connected to the flow channel 130 to form their own independent passages.

[0048] The flow channel 130 may be located only at the head of the body 100, or it may extend from the head to the neck of the body 100, or further to the tail of the body 100. The path of each flow channel 130 may be along the central axis of the body 100; however, the path of the flow channel 130 may also be a tortuous curve, for example, the flow channel 130 may be a curved path extending from the head of the body 100 along a spiral path to the neck or tail of the body 100. The specific location and path of the flow channel 130 are not specifically limited in this embodiment.

[0049] Furthermore, referring to Figure 1 and Figure 3 A collector 140 is provided on one side of the main body 110 located in the inner cavity of the body 100. The collector 140 is a hollow cavity and is fixedly arranged around the side of the main body 110 located in the inner cavity of the body 100. The collector 140 is connected to one end of the air film hole 150 located in the inner cavity of the body 100.

[0050] In one exemplary embodiment, refer to Figure 4 The cooling system provided in this application includes: a storage component 200, a cooling component 300, and a circulation component 400, wherein:

[0051] Storage component 200 is used to store cooling medium, and cooling component 300 is used to cool the cooling medium. Circulation component 400 includes a power pump 410 and a turbine 420. The power pump 410 pressurizes the cooling medium and pumps the pressurized cooling medium from the head of the fuselage 100 into the flow channel 130. The air inlet 421 of the turbine 420 communicates with the flow channel 130 between the tail and neck of the fuselage 100, and the exhaust port 422 of the turbine 420 communicates with one end of the film cooling hole 150 located within the inner cavity of the fuselage 100. It should be further noted that all components of the cooling system provided in this application are located within the inner cavity of the aircraft fuselage 100.

[0052] The cooling assembly 300 can cool the cooling medium stored in the storage assembly 200, thereby enabling the storage assembly 200 to store and supply a lower-temperature cooling medium into the flow channel 130. The power pump 410 can pump the lower-temperature cooling medium stored in the storage assembly 200 from the nose of the aircraft into the flow channel 130. Since the flow channel 130 is located between the inner body 110 and the outer skin 120 of the aircraft fuselage 100, and the flow channel 130 extends from the nose to the tail of the aircraft, the lower-temperature cooling medium can absorb the heat from the nose of the aircraft fuselage 100 and become a hot gas with higher temperature and pressure. The hot gas can flow along the flow channel 130 towards the tail of the fuselage 100. Furthermore, since the air intake 421 of the turbine 420 is connected to the flow channel 130 between the tail and neck of the fuselage 100, hot gas can enter the air intake 421 of the turbine 420 and drive the blades of the turbine 420 to rotate and do work. After driving the turbine 420, the temperature and pressure of the hot gas are further reduced to become neutral gas. The neutral gas is discharged from the turbine 420 to the film gas hole 150 and discharged from the film gas hole 150 to the outside of the fuselage 100. Since the film gas hole 150 penetrates the main body 110 and skin 120 of the aircraft fuselage 100, the neutral gas will further absorb the heat of the fuselage 100 when it is discharged from the film gas hole 150 to the outside of the fuselage 100, thereby further cooling the fuselage 100.

[0053] In other words, during the process of the cooling medium flowing from the head of the fuselage 100 to the turbine 420, the cooling medium can absorb heat to cool the head and neck of the fuselage 100 for the first time; and during the process of the neutral gas being cooled and depressurized by the turbine 420 and discharged from the fuselage 100 through the gas film vent 150, the neutral gas can also cool the fuselage 100 for the second time. Therefore, the cooling system provided by this application can quickly cool the head of the fuselage 100, thereby improving the stability and safety of the head of the aircraft fuselage 100.

[0054] Furthermore, the following sections will elaborate on each component of the cooling system.

[0055] In one embodiment, reference Figure 4 and Figure 5 The storage assembly 200 includes a storage tank 210, a filling pipeline 220, and a pressure relief pipeline 230; wherein, the storage tank 210 is used to store cooling medium, the storage tank 210 is connected to the filling pipeline 220, the filling pipeline 220 is used to fill the storage tank 210 with cooling medium, and the storage tank 210 is connected to the pressure relief pipeline 230, the pressure relief pipeline 230 is used to relieve pressure on the storage tank 210.

[0056] Specifically, the storage box 210 consists of an inner box 211 and an outer box 212. The inner box 211 is located inside the outer box 212, and the inner box 211 and the outer box 212 are connected by a connector 213. The inner box 211 is used to store the cooling medium, and the space between the inner box 211 and the outer box 212 is a vacuum. The inner box 211 of the storage box 210 can be made of aluminum alloy, and furthermore, carbon fiber can be wound around the outside of the inner box 211, so that the inner box 211 is lightweight while also having greater structural rigidity. The outer box 212 of the storage box 210 can also be made of aluminum alloy, and other materials can also be used; no specific limitation is made in this embodiment.

[0057] The inner casing 211 and outer casing 212 of the storage box 210 are fixedly connected by connectors 213. Connectors 213 can be made of epoxy fiberglass with low thermal conductivity. Connectors 213 penetrate the wall of the inner casing 211 and connect to the outer casing 212. Multiple connectors 213 are provided between the inner casing 211 and the outer casing 212, thus ensuring relative stability between them. Since the space between the inner casing 211 and the outer casing 212 is a vacuum or near-vacuum, the heat exchange efficiency between the inner casing 211 and the outside is low. Furthermore, the connectors 213 are also made of a material with low thermal conductivity, allowing the cooling medium stored within the inner casing 211 to maintain a low temperature under the action of the cooling system.

[0058] Furthermore, referring to Figure 4 and Figure 5 The filling pipeline 220 includes a check valve 222 and a level gauge 221. A filling port 223 is provided on the storage tank 210, connected to the outlet of the check valve 222. The inlet of the check valve 222 is equipped with a level gauge 221. External cooling medium can enter the storage tank 210 through the check valve 222 and the filling port 223. The check valve 222 prevents backflow of the cooling medium, thereby improving the efficiency of filling the storage tank 210 with cooling medium. The level gauge 221 can detect the level of the cooling medium in the storage tank 210. It should be noted that the cooling medium can be either liquid or gas at room temperature. In one example of this application embodiment, liquid ammonia is used as the cooling medium. Ammonia is gaseous at room temperature and pressure, but its boiling point is -33.5℃ and its melting point is -77.73℃. When the temperature drops to a certain level, ammonia can liquefy into liquid ammonia. Of course, other media can also be used as the cooling medium, but no specific limitations are made in the embodiments of this application.

[0059] Furthermore, referring to Figure 4 and Figure 5The pressure relief pipeline 230 includes a back pressure valve 231, a pressure gauge, a pressure relief valve 232, and a mass flow meter 233. The storage tank 210 has a pressure relief port 234. The inlet end of the mechanical pressure relief valve 232 is connected to the pressure relief port 234 of the storage tank 210, the outlet end of the pressure relief valve 232 is connected to the inlet end of the back pressure valve 231, and the outlet end of the pressure relief valve 232 is connected to the mass flow meter 233. The back pressure valve 231 is also connected to a pressure gauge. When the pressure inside the storage tank 210 is high (e.g., exceeding a set pressure threshold), the pressure relief valve 232 and the back pressure valve 231 can open, thereby relieving pressure in the storage tank 210 and ensuring its safety.

[0060] In one embodiment, reference Figure 5 and Figure 6 The cooling assembly 300 includes a refrigerator 310, a circulating pump 320, a heat insulation box 340, a quick-connect valve 330, and an air-cooled screen 350. The circulating pump 320 and the refrigerator 310 are both located inside the heat insulation box 340, and the air-cooled screen 350 is located inside the storage box 210. The input end of the circulating pump 320 is connected to the return end of the air-cooled screen 350, the output end of the circulating pump 320 is connected to the input end of the refrigerator 310, and the output end of the refrigerator 310 is connected to the inlet end of the air-cooled screen 350.

[0061] Specifically, refer to Figure 5 and Figure 6 The air-cooled screen 350 is constructed from a series of bent connecting pipes made of a material with high thermal conductivity, thereby improving its heat exchange efficiency. One end of the air-cooled screen 350 is connected to the inlet of the refrigerator 310 via a quick-connect valve 330, and the outlet of the refrigerator 310 is connected to the inlet of the circulating pump 320. The outlet of the circulating pump 320 is connected to the other end of the air-cooled screen 350 via another quick-connect valve 330. A mass flow meter 233 is also connected between the outlet of the circulating pump 320 and the air-cooled screen 350. Furthermore, the heat insulation box 340 is fixedly connected to the inner cavity of the aircraft fuselage 100, and the refrigerator 310, circulating pump 320, quick-connect valve 330, and mass flow meter 233 are all housed within the heat insulation box 340 to reduce heat exchange with the outside environment.

[0062] The circulating pump 320 stores a cooling fluid, such as a hydrocarbon refrigerant. Under the action of the circulating pump 320, the fluid can circulate between the air-cooled screen 350 and the refrigerator 310. The fluid removes the heat of the cooling medium in the storage tank 210, and the fluid cooled by the refrigerant enters the air-cooled screen 350 again. The continuous circulation of the fluid cools the circulating medium. And because the heat exchange efficiency between the storage tank 210 and the outside is low, the cooling medium in the storage tank 210 can be maintained at a low temperature for a long period of time.

[0063] In one embodiment, the storage tank 210 is connected to the inlet 411 of the power pump 410 via a flow control valve 240; the inlet end of the flow control valve 240 is connected to the inner casing 211 of the storage tank 210, the outlet end of the flow control valve 240 is connected to the inlet 411 of the power pump 410, and the outlet 412 of the power pump 410 is connected to the flow channel 130 at the head position of the aircraft fuselage 100.

[0064] Reference Figure 7 The output shaft of the turbine 420 is connected to the power shaft of the turbine 420 pump, and the air inlet 421 of the turbine 420 is equipped with a starter 440. The air inlet 421 of the turbine 420 is connected to the flow channel 130 between the tail and neck of the machine body 100, and is used to introduce heated and pressurized hot gas (cooling medium) in the flow channel 130 to drive the blades of the turbine 420 to rotate.

[0065] Reference Figure 4 The circulation assembly 400 also includes an expansion valve 430, and the exhaust port 422 of the turbine 420 is connected to the inlet end of the expansion valve 430. The outlet end of the expansion valve 430 is connected to the film gas hole 150, which is used to discharge the neutral gas after being cooled and depressurized by the turbine 420 to the outside of the machine body 100 through the film gas hole 150.

[0066] Furthermore, the expansion valve 430 can adjust the flow rate of the gas output to the air film through the expansion valve 430 based on the external heat flow, wherein the external heat flow is the high-temperature heat flow generated by the friction between the aircraft body 100 and the atmosphere during flight. The expansion valve 430 can further reduce the temperature and pressure of the neutral gas discharged from the exhaust port 422 of the turbine 420, thereby improving the efficiency of the neutral gas in providing a second cooling effect on the aircraft body 100.

[0067] Furthermore, referring to Figure 4 and Figure 7 A starter 440 is provided at the air inlet 421 of the turbine 420. The starter 440 is used to start the turbine 420. The starter 440 can be an electric starter 440 or a gunpowder starter 440 to generate high-temperature and high-pressure gas to provide initial drive for the turbine 420. Since the power pump 410 is driven by the drive shaft of the turbine 420, after the turbine 420 is initially started by the starter 440, the power pump 410 starts to work, thereby pumping the cooling medium in the storage tank 210 into the flow channel 130. The cooling medium absorbs heat and rises in temperature in the flow channel 130 to become hot gas. The hot gas then enters the air inlet 421 of the turbine 420, thereby realizing circulation to maintain the continuous operation of the turbine 420 and the power pump 410.

[0068] Furthermore, referring to Figure 2 and Figure 4A collector 140 is provided on one side of the main body 110 located in the inner cavity of the body 100. The collector 140 is a hollow cavity. The collector 140 is connected to the outlet of the expansion valve 430 and to one end of the air film hole 150 located in the inner cavity of the body 100.

[0069] Specifically, the collector 140 is a hollow cavity and is fixedly installed around the main body 110 on one side of the inner cavity of the machine body 100. The collector 140 is connected to the outlet end of the expansion valve 430, and also to one end of the gas film orifice 150 located in the inner cavity of the machine body 100. The neutral gas discharged through the expansion valve 430 is buffered in the collector 140, which facilitates the uniform distribution of the neutral gas. The neutral gas is then discharged from the collector 140 to the gas film orifice 150 to achieve a second cooling of the machine body 100. Furthermore, since the gas film orifice 150 on the machine body 100 is not connected to the flow channel 130, the cooling medium in the flow channel 130 is prevented from leaking directly from the gas film orifice 150.

[0070] Furthermore, referring to Figure 4 A temperature and pressure assembly is installed on the machine body 100 to detect the temperature and pressure data of the machine body 100. The temperature and pressure assembly includes a temperature sensor and a pressure sensor. Multiple temperature and pressure assemblies can be installed inside the machine body 100; for example, one set of temperature and pressure assemblies can be installed at the head, neck, and tail positions of the inner cavity of the machine body 100. Furthermore, a temperature and pressure assembly is also installed on the connecting pipe between the power pump 410 and the flow channel 130 at the head position of the machine body 100 to detect the pressure and temperature data of the cooling medium entering the flow channel 130. A temperature and pressure assembly is also installed at the air inlet 421 of the turbine 420 to detect the pressure and temperature data of the gas entering the turbine 420 from the flow channel 130. Similarly, a temperature and pressure assembly is installed at the inlet of the expansion valve 430 to detect the pressure and temperature data of the neutral gas output from the turbine 420. A set of temperature and pressure components is installed inside the inner box 211 of the storage box 210 to detect the pressure and temperature data of the cooling medium stored in the storage box 210.

[0071] In one embodiment, the overall parameters of the aircraft including the aforementioned cooling system are derived and calculated. The flight time of the aircraft within the atmosphere is set to... The aircraft generates heat through friction with the atmosphere during flight, and its average heat flux density is... The area of ​​the high-temperature region at the nose of the aircraft is The heat absorbed by the aircraft while flying within the atmosphere It can be expressed by formula (1), where:

[0072] , formula (1).

[0073] Furthermore, the heat absorbed by the aircraft while flying within the atmosphere Part of the heat is absorbed by the aircraft's own heat capacity, another part is dissipated through radiation, and the remaining heat is absorbed twice by the cooling medium in the cooling system and discharged from the aircraft. The heat absorbed by the cooling medium (here, we assume the cooling medium is ammonia) is... It can be expressed by formula (2):

[0074] Formula (2). Wherein, The specific heat capacity of the aircraft's nose; The mass of the aircraft's nose; These are the stable temperature of the aircraft's nose and the temperature before flight, respectively. The surface emissivity of the aircraft material; is the blackbody radiation constant.

[0075] Furthermore, the pressure of ammonia inside storage tank 210 is set to... The temperature is After the flow control valve 240 is opened, liquid ammonia enters the power pump 410. The pressure of the liquid ammonia at the outlet 412 of the power pump 410 is... The temperature is enthalpy value Then, liquid ammonia enters flow channel 130 from the nose of the aircraft; after the liquid ammonia completes its first heat absorption, the pressure of the resulting ammonia gas (thermal gas) is... The temperature is enthalpy value Then, the high-temperature, high-pressure ammonia gas (a hot gas) enters the turbine 420 to do work. The pressure of the ammonia gas (a neutral gas) at the exhaust port 422 of the turbine 420 is... The temperature is The neutral gas discharged from turbine 420 passes through expansion valve 430, and the pressure of ammonia gas is... The temperature is enthalpy value The neutral gas discharged from expansion valve 430 re-enters collector 140 and then enters the cooling channel of gas film orifice 150. After absorbing heat again, the neutral gas is discharged outside the machine body 100, and the final pressure of the discharged ammonia gas is... The temperature is enthalpy value .

[0076] The enthalpy difference between the two heat absorptions of ammonia It can be expressed by formula (3):

[0077] , formula (3).

[0078] According to the law of conservation of energy, the total mass of ammonia required is... It can be expressed by formula (4), where:

[0079] , formula (4).

[0080] The required ammonia flow rate can be expressed by formula (5), where: , formula (5).

[0081] Furthermore, the length of the flow channel 130 at the nose position of the aircraft is set. Equivalent diameter is The resistance coefficient of ammonia in flow channel 130 is... It can be expressed by formula (6), where:

[0082] , formula (6); is the Reynolds number of ammonia at the corresponding temperature and pressure.

[0083] The flow of ammonia within the flow channel 130 at the nose of the aircraft is turbulent. The Nusselt numbers for ammonia under different conditions are... As shown in formulas (7) and (8) respectively, where:

[0084] Formula (7) - gaseous state;

[0085] Formula (8) - Liquid;

[0086] in, The temperature of ammonia in the flow channel at 130°C in the nose of the aircraft; The wall temperature of the flow channel 130 at the nose of the aircraft; This represents the Prandtl number of ammonia at the corresponding temperature and pressure.

[0087] Furthermore, the overall heat transfer coefficient of ammonia flowing within the flow channel 130 at the nose of the aircraft is... This can be expressed by formula (9):

[0088] , formula (9); is the thermal conductivity of ammonia at the corresponding temperature and pressure.

[0089] The total heat exchange area of ​​the ammonia flow channel at the nose of the aircraft is 130. It can be expressed by formula (10):

[0090] , formula (10).

[0091] The total pressure drop in the flow channel 130 at the nose of the aircraft is expressed by formula (11): , formula (11); This represents the flow rate of ammonia at the corresponding temperature and pressure.

[0092] The detailed structure of the flow channel 130 at the nose of the aircraft can be obtained based on the total heat exchange area, total pressure drop, equivalent diameter, and length of the flow channel 130.

[0093] Furthermore, such as Figure 4 or Figure 7 As shown, temperature sensors (T) and pressure sensors (P) are installed at the inlets and outlets of key parts of the aircraft and various components in the cooling system to continuously monitor temperature and pressure data. The flow rate of ammonia in the cooling system can be adjusted by the flow control valve 240 and the expansion valve 430 to ensure that the temperature of the aircraft body 100 is maintained at a normal level.

[0094] In one embodiment, the flight time of the aircraft within the atmosphere is set. The average heat flux density is 1200s. During flight, the aircraft generates heat through friction with the atmosphere. The power is 400kW, and the area of ​​the high-temperature region at the nose of the aircraft is... for Then, the amount of heat absorbed by the aircraft during flight within the atmosphere can be calculated using formula (1). ,in: .

[0095] In this embodiment, the heat absorbed by the aircraft while flying within the atmosphere Part of the heat is due to the aircraft's own heat capacity, part is dissipated through radiation, and the remaining heat is absorbed by ammonia. The amount of heat absorbed by the cooling medium ammonia can be calculated using formula (2). :

[0096] .

[0097] In this embodiment, the pressure of ammonia inside the inner chamber 211 of the storage tank 210 is set. for The temperature is With a pressure of 240K, after the flow control valve 240 opens, liquid ammonia enters the power pump 410. The pressure of the liquid ammonia at the outlet 412 of the power pump 410 is... 6MPa, temperature It is 242K, enthalpy value It can be found in Table (1); then the liquid ammonia enters the flow channel 130 in the head of the aircraft for the first time to absorb heat, and the liquid ammonia becomes a hot gas with a higher temperature and higher pressure. At this time, the pressure of the ammonia gas is... The pressure is 5.7 MPa, and the temperature is... It is 780K, and its enthalpy is The results can be found in Table (1); then, the high-temperature and high-pressure ammonia gas enters the turbine 420 to do work, and the pressure of the ammonia gas at the exhaust port 422 of the turbine 420 is... 1 MPa, temperature The pressure of ammonia gas after passing through expansion valve 430 is 585K. 0.3 MPa, temperature It is 465K, enthalpy value The neutral gas discharged from expansion valve 430 re-enters collector 140 and enters the cooling channel of gas film orifice 150; after the neutral gas undergoes a second heat absorption in gas film orifice 150, it is discharged to the outside of the machine body 100, and the final pressure of the discharged ammonia gas is... 0.2 MPa, temperature 800K, enthalpy value It can be found in Table (1).

[0098] The enthalpy difference of ammonia during the two heat absorptions can be calculated using formula (3): According to the law of conservation of energy, the total mass of ammonia required is... It can be calculated using formula (4): The required ammonia flow rate can be calculated using formula (5): .

[0099] In this embodiment, the length of the flow channel 130 at the nose of the aircraft is... It is 500mm, equivalent diameter If the diameter is 1.33 mm, then the resistance coefficient of ammonia in flow channel 130 is... It can be calculated using formula (6): .

[0100] Ammonia flows turbulently in channel 130 at the nose of the aircraft. The Nusselt number of ammonia varies under different conditions. It can be calculated using formulas (7) and (8), where:

[0101] ;

[0102] .

[0103] Furthermore, the overall heat transfer coefficient of ammonia flowing within the flow channel 130 at the nose of the aircraft is... It can be calculated using formula (9): .

[0104] The total heat exchange area of ​​the aircraft's nose flow channel 130 It can be calculated using formula (10):

[0105] .

[0106] Then the total heat transfer area of ​​the flow channel 130 at the nose of the aircraft can be obtained. .

[0107] The total pressure drop in the flow channel 130 at the nose of the aircraft can be calculated using formula (11), where:

[0108] .

[0109] The calculation results of this embodiment are summarized in Table (2). The relationship between the total temperature and pressure data of the cooling system calculated in this embodiment is illustrated as follows: Figure 8 As shown.

[0110] Table (1)

[0111]

[0112] Table (2)

[0113]

[0114] Unless otherwise specifically stated, the relative arrangement, numerical expressions, and values ​​of the components and steps described in these embodiments do not limit the scope of the invention. It should also be understood that, for ease of description, the dimensions of the various parts shown in the accompanying drawings are not drawn to actual scale.

[0115] The technical features of the above embodiments can be combined in any way. For the sake of brevity, 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.

[0116] The embodiments described above are merely illustrative of several implementation methods of this application, and while the descriptions are specific and detailed, they should not be construed as limiting the scope of this patent application. It should be noted that those skilled in the art can make various modifications and improvements without departing from the concept of this application, and these all fall within the protection scope of this application. Therefore, the protection scope of this application should be determined by the appended claims.

Claims

1. A cooling system, characterized in that, Applied to aircraft, the fuselage of the aircraft is composed of an inner main body and an outer skin connected together, and a flow channel extending from the head to the tail of the fuselage is provided between the main body and the skin; the fuselage is provided with air film vents that penetrate the main body and the skin; The cooling system is located within the internal cavity of the machine body, and the system includes a storage component, a cooling component, and a circulation component, wherein: The storage component is used to store the cooling medium, and the cooling component is used to cool the cooling medium. The circulation assembly includes a power pump and a turbine, wherein the power pump is used to pressurize the cooling medium and pump the pressurized cooling medium from the head of the machine body into the flow channel; The turbine's air inlet is connected to the flow channel between the tail and neck of the machine body. The turbine's exhaust port is connected to the film gas hole at one end of the inner cavity of the machine body. The hot gas discharged from the flow channel enters the turbine's air inlet and drives the turbine blades to rotate. The film gas hole is not connected to the flow channel. The output shaft of the turbine is connected to the power shaft of the power pump, and the turbine's air inlet is equipped with a starter for starting the turbine. The system also includes an expansion valve, the inlet of which is connected to the exhaust port of the turbine, and the outlet of which is connected to one end of the film gas hole located in the inner cavity of the machine body; A collector is provided on one side of the main body located in the inner cavity of the machine body. The collector is a hollow cavity. The collector is connected to the outlet of the expansion valve and to one end of the air film hole located in the inner cavity of the machine body.

2. The system according to claim 1, characterized in that, The starter is an electric starter or a gunpowder starter.

3. The system according to claim 1, characterized in that, The storage assembly includes a storage tank, a filling pipeline, and a pressure relief pipeline, wherein: The storage tank is used to store the cooling medium. The storage tank is connected to a filling pipeline, which is used to fill the storage tank with the cooling medium. The storage tank is also connected to a pressure relief pipeline, which is used to relieve pressure in the storage tank. The storage tank is connected to the inlet of the power pump via a flow control valve.

4. The system according to claim 3, characterized in that, The storage box consists of an inner box and an outer box, wherein: The inner box is located inside the outer box, and the inner box and the outer box are connected by a connector; The inner casing is used to store the cooling medium, and the space between the inner casing and the outer casing is a vacuum.

5. The system according to claim 4, characterized in that, The cooling assembly includes a refrigerator, a circulating pump, a heat insulation box, and an air-cooled shield, wherein: The circulating pump and the refrigeration unit are both located inside the heat insulation box, and the air-cooled screen is located inside the storage box; The input end of the circulating pump is connected to the return end of the air-cooled screen, the output end of the circulating pump is connected to the input end of the refrigerator, and the output end of the refrigerator is connected to the inlet end of the air-cooled screen.

6. The system according to any one of claims 1-5, characterized in that, The cooling medium is liquid ammonia.

7. An aircraft, characterized in that, The aircraft's fuselage is composed of an inner main body and an outer skin, with a flow channel extending from the head to the tail between the main body and the skin. The body has air film pores that connect the main body and the skin; The aircraft is also equipped with a cooling system as described in any one of claims 1-6.

8. The aircraft according to claim 7, characterized in that, The air film pores are not connected to the flow channel.

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