Integrated heat exchanger, parameter determination method, and start-up launch device

Abstract

The present invention discloses an integrated heat exchanger, which is applied to a launch device. The launch device includes a launch tube, and the heat exchanger includes a plurality of heat exchange tubes, which are circumferentially distributed in the tube wall of the launch tube; the heat exchange tube includes a condensation section, an insulation section and an evaporation section which are sequentially connected to form a closed tube cavity, and the tube diameter of the closed tube cavity gradually decreases from the condensation section to the evaporation section; a liquid wick is arranged on the inner wall of the closed tube cavity, and a working medium capable of phase change between gaseous and liquid states is arranged in the closed tube cavity, and the liquid wick can absorb the liquid working medium and guide the liquid working medium to flow from the condensation section to the evaporation section. The present invention can collect the heat generated by the unmanned aerial vehicle at the front of the launch tube due to friction and use it to heat the tail cavity of the launch tube, so as to prevent carbon dioxide from condensing into a solid, and also balance the temperature of the launch tube to ensure its structural integrity. The present invention also provides a parameter determination method, a launch starter, an electronic device and a computer-readable storage medium.

Description

Technical Field

[0001] The present invention relates to the technical field of high-pressure gas starting and launching, and in particular to an integrated heat exchanger and a parameter determination method, and a starting and launching device. Background Art

[0002] As a novel power system with simple structure and convenient maintenance, high-pressure gas-started launch device has been gradually developed and applied in the aerospace field. It avoids the irreversible damage of high-temperature and high-speed tail flame to the system during the self-starting propulsion of unmanned aerial vehicles. The working medium used by the high-pressure gas-started launch device is a key factor affecting its working performance, which determines the aerodynamic ejection capability and function realization of the launch system.

[0003] Compared with traditional high-pressure media such as compressed air, supercritical carbon dioxide has higher specific internal energy and density under the same working fluid state, and a large amount of carbon dioxide working fluid can be loaded in a smaller high-pressure storage tank. When the carbon dioxide working fluid enters the launch tube from the storage tank, it can release more propulsion work, providing a new way to miniaturize and compact the launch device.

[0004] However, ordinary heat exchangers find it difficult to adapt to the drastic changes in the physical properties of carbon dioxide during the launch process, which has an adverse effect on the startup of the launch device. Summary of the invention

[0005] The present invention aims to solve one of the technical problems in the related art to a certain extent. To this end, the present invention provides an integrated heat exchanger and a parameter determination method and a start-up launch device.

[0006] As a first aspect of the present invention, an integrated heat exchanger is provided for use on a launching device, wherein the launching device comprises a launching tube, and the heat exchanger comprises a plurality of heat exchange tubes, wherein the plurality of heat exchange tubes are circumferentially distributed in the tube wall of the launching tube; the heat exchange tube comprises a condensation section, an insulation section and an evaporation section which are sequentially connected to form a closed tube cavity, and the tube diameter of the closed tube cavity gradually decreases from the condensation section to the evaporation section; a liquid wick is arranged on the inner wall of the closed tube cavity, and a working medium capable of phase change between gas and liquid is arranged in the closed tube cavity, and the liquid wick can absorb the liquid working medium and guide the liquid working medium to flow from the condensation section to the evaporation section.

[0007] The application of this technical solution has the following beneficial effects: when the unmanned aerial vehicle is started and launched, the integrated heat exchanger in the high-pressure gas start-up launch device quickly transfers the large amount of friction heat generated by the aircraft in the launch tube to the low-temperature and low-pressure chamber of the launch tube through the flow and phase change of the working fluid in the tube, thereby optimizing the temperature at the front end of the launch tube and increasing the temperature of the carbon dioxide in the high-pressure and low-pressure chamber of the launch tube, ensuring the temperature level of the launch device valve inlet to prevent the formation of dry ice, improving the energy conversion efficiency and realizing the reliable operation of the supercritical carbon dioxide start-up launch tube.

[0008] Optionally, the extension direction of the heat exchange tube is consistent with the launch direction of the launch tube, the condensation section of the heat exchange tube is located at the air inlet end of the launch tube, and the evaporation section of the heat exchange tube is located at the launch end of the launch tube; a heat exchange groove group is arranged between the evaporation section of the heat exchange tube and the inner wall of the launch tube, and the heat exchange groove group includes a plurality of annular grooves arranged at intervals along the axial direction of the heat exchange tube, and the openings of the annular grooves pass through the inner wall of the launch tube.

[0009] Optionally, the spacing between the plurality of annular grooves gradually decreases in a direction from the condensation section to the evaporation section.

[0010] Optionally, the heat exchange tube, the heat exchange tank group and the launch tube are configured as an integrated structure.

[0011] As a second aspect of the present invention, a method for determining parameters of an integrated heat exchanger is provided, which is applied to the heat exchanger provided in the first aspect of the present invention, and the method comprises the following steps:

[0012] Determine the side area of ​​the heat exchange tube according to the taper of the heat exchange tube, the side area is the heat exchange area of ​​the heat exchange tube, and establish a mirror image gradient function of the heat exchange area along the emission direction;

[0013] According to the motion speed control equation and friction heat control equation of the unmanned aerial vehicle, the friction heat gradient function along the launching direction is established;

[0014] The heat exchange area mirror gradient function and the friction heat gradient function are matched, and the taper of the heat exchange tube is output through iterative calculation.

[0015] Optionally, determining the side area of ​​the heat exchange tube according to the taper of the heat exchange tube, the side area being the heat exchange area of ​​the heat exchange tube, and establishing a mirror image gradient function of the heat exchange area along the emission direction includes:

[0016] The heat exchange area of ​​the heat exchange tube is determined according to the side area of ​​the heat exchange tube. The calculation formula of the heat exchange area is as follows:

[0017] (1-1)

[0018] Where A is the heat exchange area, h is the height of the heat exchange tube, C is the taper of the heat exchange tube, r 0 is the bottom radius of the heat exchange tube, r 1 is the top radius of the heat exchange tube; the taper of the heat exchange tube is expressed as:

[0019] (1-2)

[0020] Among them, h ’ is the height of the heat exchange tube from the position corresponding to the pneumatic piston of the launch tube to the top of the heat exchange tube, r 2 is the radius of the heat exchange tube at the position of the pneumatic piston in the launch tube;

[0021] The axial height of the launch tube is defined as the Z axis, and the direction of the launch tube toward the launch end is defined as the positive direction of the Z axis. The heat exchange area gradient of the heat exchange tube along the positive direction of the Z axis is calculated as follows:

[0022] (1-3)

[0023] Let x = hz, and calculate the axial mirror gradient of the heat exchange area of ​​the heat exchange tube:

[0024] (1-4)

[0025] The axial image gradient of the heat exchange area of ​​the heat exchange tube is processed by finite difference and normalization to obtain the heat exchange area image gradient function:

[0026] (1-5)

[0027] Optionally, the friction heat gradient function is established according to the motion speed control equation and the friction heat control equation of the unmanned aerial vehicle, including:

[0028] According to the parameters of the launch device, the motion speed control equation of the unmanned aerial vehicle is calculated as:

[0029] (2-1)

[0030] Where P 1 is the pressure of the working fluid in the low-pressure chamber of the launch tube, and P 2 is the pressure of the working fluid in the high-pressure storage tank of the launch tube, m is the total mass of the unmanned aerial vehicle and the pneumatic piston, A v is the maximum radial cross-sectional area of ​​the UAV and the pneumatic piston, g is the gravity constant; the air friction resistance F is expressed as:

[0031] (2-2)

[0032] Where ρ is the air density, C v is the friction coefficient; then the friction heat control equation is established as:

[0033] (2-3)

[0034] Based on formula (2-1)-formula (2-3), the function of friction heat with respect to the emission distance is obtained. The function of friction heat with respect to the emission distance is processed by finite difference and normalization to obtain the friction heat gradient function. The friction heat gradient function characterizes the rate of change of friction heat along the emission direction:

[0035] (2-4)

[0036] Optionally, f a (hz) and f e (z) is matched, dA / dz is determined by iterative calculation with the minimum absolute error as the goal, and the taper C of the heat exchange tube is finally output; wherein the goal of the minimum absolute error is expressed as:

[0037] (3-1)

[0038] In the formula, n is the number of axial nodes evenly divided from bottom to top, that is:

[0039] (3-2)

[0040] As a third aspect of the present invention, a starting and launching device is provided, wherein the starting and launching device comprises a launching tube, and the launching tube comprises the integrated heat exchanger provided in the first aspect of the present invention.

[0041] As a fourth aspect of the present invention, an electronic device is provided, comprising: at least one processor; and a memory communicatively connected to the at least one processor; wherein the memory stores a computer program executable by the at least one processor, and the computer program is executed by the at least one processor so that the at least one processor can execute the integrated heat exchanger parameter determination method provided in the second aspect of the present invention.

[0042] As a fifth aspect of the present invention, a computer-readable medium is provided, wherein the computer-readable storage medium stores computer instructions, and the computer instructions are used to enable a processor to implement the integrated heat exchanger parameter determination method provided in the second aspect of the present invention when executed.

[0043] These features and advantages of the present invention will be disclosed in detail in the following specific embodiments and drawings. The best embodiments or means of the present invention will be described in detail in conjunction with the drawings, but they are not intended to limit the technical solutions of the present invention. In addition, there are multiple features, elements, and components that appear in each of the following texts and drawings, and different symbols or numbers are marked for convenience, but they all represent components with the same or similar structures or functions. BRIEF DESCRIPTION OF THE DRAWINGS

[0044] The present invention will be further described below in conjunction with the accompanying drawings:

[0045] Figure 1 It is a schematic structural diagram of the integrated heat exchanger of the present invention.

[0046] Figure 2 It is a schematic diagram of the structure of the launching device of the present invention.

[0047] Figure 3 This is a top view of the launch device of the present invention.

[0048] Figure 4 A bottom view of the launch device of the present invention.

[0049] Figure 5 It is the launch speed diagram of the unmanned aerial vehicle of the present invention.

[0050] Figure 6 This is a graph showing the change of friction heat with time according to the present invention.

[0051] Figure 7 This is a graph showing the variation of friction heat with respect to emission distance according to the present invention.

[0052] Figure 8 This is the gradient function result diagram of the present invention.

[0053] Among them, 1. launching tube; 11. pneumatic piston; 12. low-pressure chamber; 2. heat exchange tube; 21. liquid absorption core; 22. closed tube cavity; 23. heat exchange tank group; 3. starting valve; 4. high-pressure storage tank; 5. filling port. DETAILED DESCRIPTION

[0054] Embodiments of the present invention are described in detail below, examples of which are shown in the accompanying drawings, wherein the same or similar reference numerals throughout represent the same or similar elements or elements with the same or similar functions. The embodiments in the implementation manner are intended to be used to explain the present invention and should not be construed as limiting the present invention.

[0055] References in this specification to "one embodiment" or "an example" or "an example" mean that a particular feature, structure, or characteristic described in conjunction with the embodiment itself may be included in at least one embodiment of the present patent disclosure. The appearance of the phrase "in one embodiment" in various places in the specification does not necessarily refer to the same embodiment.

[0056] In the related technology, when carbon dioxide is used as a working fluid in the launch process of an unmanned aerial vehicle, the temperature of the working fluid drops rapidly due to expansion, approaching its triple point, causing supercritical carbon dioxide to liquefy or even form solid dry ice, which may lead to blockage of related pipelines and valves, and structural damage due to excessively low temperature at the tail of the launch tube. On the other hand, when the unmanned aerial vehicle moves at high speed in the launch tube, friction generates a lot of heat, and the temperature at the front of the launch tube is too high, which also has an adverse effect on the structure of the tube body.

[0057] In view of this, please refer to the attached Figure 1-4 The first aspect of the present invention provides an integrated heat exchanger for use in a launch device, wherein the launch device includes a launch tube 1, and the heat exchanger includes a plurality of heat exchange tubes, which are circumferentially distributed in the tube wall of the launch tube 1; the heat exchange tube includes a condensation section, an insulation section and an evaporation section which are sequentially connected to form a closed tube cavity 22, and the tube diameter of the closed tube cavity 22 gradually decreases from the condensation section to the evaporation section; a liquid absorbent core 21 is arranged on the inner wall of the closed tube cavity, and a working medium capable of phase change between gas and liquid is placed in the closed tube cavity, and the liquid absorbent core 21 can absorb the liquid working medium and guide the liquid working medium to flow from the condensation section to the evaporation section.

[0058] The structure of the closed tube cavity 22 gradually thickens from top to bottom, so that when the working medium moves up from the bottom to the top under the action of the liquid wick, the tube diameter gradually becomes thinner and the capillary force gradually increases, and the liquid phase working medium is effectively pushed. This structure compensates for the adverse effect of gravity on the reflux of the working medium. On the other hand, the friction heat generated when the unmanned aerial vehicle is started and launched and the low temperature of the low-pressure chamber cause a large axial temperature gradient in the launch tube. The gradually tapering truncated cone pipeline structure optimizes the heat distribution inside the pipeline. In addition, the upper end of the thinner pipeline has a larger heat flux density and a smaller thermal resistance, while the lower end of the thicker pipeline provides a larger heat exchange area, ensuring the full condensation of the gas phase working medium, improving the heat exchange efficiency, and preventing overheating; the heat generated by the unmanned aerial vehicle at the front of the launch tube due to friction is collected and used to heat the cavity at the tail of the launch tube to prevent carbon dioxide from condensing into a solid, and the temperature of the launch tube is also balanced to ensure its structural integrity.

[0059] The extension direction of the heat exchange tube is consistent with the emission direction of the emission tube. The condensation section of the heat exchange tube is located at the air inlet end of the emission tube 1, and the evaporation section of the heat exchange tube is located at the emission end of the emission tube 1. According to the axial height of the heat exchange tube, the structural integrated heat exchanger is divided into a condensation section, an insulation section and an evaporation section, wherein the condensation section is a pipe section outside the low-pressure cavity, and the gas phase working medium in this section of the pipeline is condensed and releases heat to the low-pressure cavity, thereby increasing the temperature in the cavity and preventing valve clogging. Then the liquid phase working medium rises to the insulation section through capillary action; the insulation section is located between the upper end of the pneumatic piston and the starter. The middle and lower part of the starting device is where the aircraft is in the initial starting state, with a low movement speed, little frictional heat, and a relatively gentle temperature gradient. Therefore, this section minimizes the outward loss of heat of the working fluid in the pipe while ensuring the flow of the working fluid in the pipe; the evaporation section is located from the middle to the top of the starting device. The ambient temperature outside the heat exchanger in this section rises extremely quickly due to the high-speed movement of the unmanned aerial vehicle. The liquid working fluid in the pipeline quickly absorbs heat and evaporates in this area to achieve phase change, thereby reducing the temperature of the front part of the launch tube. Then the gas phase working fluid flows along the axial temperature gradient and pressure difference to the low temperature at the bottom to form a cycle.

[0060] A heat exchange groove group 23 is arranged between the evaporation section of the heat exchange tube and the inner wall of the launch tube. The heat exchange groove group 23 includes a plurality of annular grooves arranged at intervals along the axial direction of the heat exchange tube, and the openings of the annular grooves penetrate the inner wall of the launch tube. In the evaporation section, due to the presence of a pneumatic piston, it is impossible to build a heat exchange structure such as a fin into the launch tube. Therefore, a multi-layer annular groove is opened on the inner side of the structural member near the structural integrated heat exchanger to increase the heat exchange area and the turbulence intensity of the high-temperature gas generated by the friction of the aircraft, thereby improving the heat exchange efficiency and compensating for the smaller heat transfer area of ​​the thinner upper end of the structural integrated heat exchanger. At the same time, in some embodiments, the axial density of the multi-layer annular grooves matches the extremely large temperature gradient at the front end of the launch tube, that is, the higher the axial direction, the denser the grooves.

[0061] The heat exchange tube, the heat exchange groove group 23 and the launch tube 1 are configured as an integrated structure.

[0062] As a second aspect of the present invention, a method for determining parameters of an integrated heat exchanger is provided, which is applied to the heat exchanger provided in the first aspect of the present invention, and the method comprises the following steps:

[0063] S1, determining the side area of ​​the heat exchange tube according to the taper of the heat exchange tube, the side area being the heat exchange area of ​​the heat exchange tube, and establishing a mirror image gradient function of the heat exchange area along the emission direction; specifically comprising:

[0064] The side area of ​​the heat exchange tube is determined according to the taper of the heat exchange tube. The side area is the heat exchange area of ​​the heat exchange tube. The calculation formula of the heat exchange area is as follows:

[0065] (1-1)

[0066] Where A is the heat exchange area, h is the height of the heat exchange tube, C is the taper of the heat exchange tube, r 0 is the bottom radius of the heat exchange tube, r 1 is the top radius of the heat exchange tube; the taper of the heat exchange tube is expressed as:

[0067] (1-2)

[0068] Among them, h ’ is the height of the heat exchange tube from the position corresponding to the pneumatic piston of the launch tube to the top of the heat exchange tube, r 2 h is the radius of the heat exchange tube at the position of the pneumatic piston in the launch tube; ’ Based on the height of the launch tube, r 2 The value is based on the diameter of the launch tube.

[0069] The axial height of the launch tube is defined as the Z axis, and the direction of the launch tube toward the launch end is defined as the positive direction of the Z axis. The heat exchange area gradient of the heat exchange tube along the positive direction of the Z axis is calculated as follows:

[0070] (1-3)

[0071] Let x = hz, and calculate the axial mirror gradient of the heat exchange area of ​​the heat exchange tube:

[0072] (1-4)

[0073] The axial image gradient of the heat exchange area of ​​the heat exchange tube is processed by finite difference and normalization to obtain the heat exchange area image gradient function:

[0074] (1-5).

[0075] S2, according to the motion speed control equation and friction heat control equation of the unmanned aerial vehicle, establish the friction heat gradient function, which specifically includes:

[0076] According to the parameters of the launch device, the motion speed control equation of the unmanned aerial vehicle is calculated as:

[0077] (2-1)

[0078] Where P 1 is the pressure of the working fluid in the low-pressure chamber of the launch tube, and P 2 is the pressure of the working fluid in the high-pressure storage tank of the launch tube, m is the total mass of the unmanned aerial vehicle and the pneumatic piston, A v is the maximum radial cross-sectional area of ​​the UAV and the pneumatic piston, g is the gravity constant; the air friction resistance F is expressed as:

[0079] (2-2)

[0080] Where ρ is the air density, C v is the friction coefficient; then the friction heat control equation is established as:

[0081] (2-3)

[0082] Based on formula (2-1)-formula (2-3), the function of friction heat with respect to the emission distance is obtained. The function of friction heat with respect to the emission distance is processed by finite difference and normalization to obtain the friction heat gradient function. The friction heat gradient function characterizes the rate of change of friction heat along the emission direction:

[0083] (2-4).

[0084] S3, matching the heat exchange area mirror gradient function and the friction heat gradient function, and outputting the taper of the heat exchange tube by iterative calculation, specifically comprising: a (hz) and f e (z) is matched, dA / dz is determined by iterative calculation with the minimum absolute error as the goal, and the taper C of the heat exchange tube is finally output; wherein the goal of the minimum absolute error is expressed as:

[0085] (3-1)

[0086] In the formula, n is the number of axial nodes evenly divided from bottom to top, that is:

[0087] (3-2).

[0088] According to the friction heat gradient in the launch tube, the appropriate heat transfer area of ​​the heat transfer tube is designed, that is, the friction heat gradient function in the launch tube is matched with the mirror image gradient of the heat transfer area to obtain the appropriate taper; then the top radius r of the heat transfer tube can be obtained by reverse calculation. 1 With bottom radius r 0 .

[0089] Specifically, in one embodiment of the present invention, referring to Figure 2 For example, a UAV with a total mass of 1000kg and a pneumatic piston is used for launching. The diameter of the UAV and the pneumatic piston are both 0.5m, the length of the launch tube is 7m, and the volume is 0.1m 3 The temperature and pressure of the supercritical carbon dioxide to be filled in the high-pressure storage tank are 323.15K and 9MPa respectively, and the volume is also 0.1m 3The initial carbon dioxide temperature and pressure in the low-pressure chamber are 303.15K and 0.1MPa respectively. The start valve is a linear valve with an opening time of 0.5 seconds and a diameter of 0.1m. The high-pressure gas start-up launch device is designed and operated according to the following method:

[0090] like Figure 1-Figure 4 As shown, on the structural member serving as the wall of the launch tube, near the inner wall of the launch tube, several adjacent axial truncated cone channels (i.e., heat exchange tubes) are processed along the circumference and evenly arranged, with the lower end of the channel extending to the low-temperature area at the bottom of the low-pressure chamber, and the upper end extending to the high-temperature area at the top of the launch device, and the outside of the channel is closed;

[0091] For each heat exchange tube, the radial cross-sectional area of ​​the low-pressure cavity from the top to the bottom of the axial launch device is continuously increased, that is, the channel is gradually thickened from top to bottom. Then, a liquid wick is added along the axial direction of the truncated cone channel to provide a capillary pump effect, and acetone with a low boiling point (56.5°C) and a low freezing point (-94.6°C) is filled as a channel working fluid, which is easy to achieve gas-liquid phase change and prevent condensation. The channel as a whole acts as a heat pipe and is located inside the structural member, such as Figure 4 As shown, it is called a structural integrated heat exchanger;

[0092] The heat exchange tubes gradually become thicker from top to bottom, and their diameter parameters are designed according to the friction heat gradient in the launch tube:

[0093] S1, calculation of heat exchange area of ​​a single heat exchanger based on the side area of ​​the heat exchange tube:

[0094] (1-1)

[0095] Where A is the heat exchange area, h is the height of the heat exchange tube, C is the taper of the heat exchange tube, r 0 is the bottom radius of the heat exchange tube, r 1 is the top radius of the heat exchange tube; the taper of the heat exchange tube is expressed as:

[0096] (1-2)

[0097] Among them, h ’ The height of the heat exchange tube from the position corresponding to the pneumatic piston of the launch tube to the top of the heat exchange tube is 7m, r 2 is the radius of the heat exchange tube corresponding to the position of the pneumatic piston in the launch tube, which is one twelfth of the launch tube radius, i.e. r 2 is 0.025m; in each small axial interval, the micro-element slope height calculated by the taper C is used as the unit height in the heat exchange area gradient, the axial height of the launch tube is defined as the Z axis, the direction of the launch tube toward the launch end is the positive direction of the Z axis, and the heat exchange area gradient of the heat exchange tube along the positive direction of the Z axis is calculated, and the calculation formula is as follows:

[0098] (1-3)

[0099] Let x = hz, and calculate the axial mirror gradient of the heat exchange area of ​​the heat exchange tube:

[0100] (1-4)

[0101] The axial image gradient of the heat exchange area of ​​the heat exchange tube is processed by finite difference and normalization to obtain the heat exchange area image gradient function:

[0102] (1-5)

[0103] S2, according to the motion speed control equation and friction heat control equation of the unmanned aerial vehicle, establish the friction heat gradient function, which specifically includes:

[0104] According to the parameters of the launch device, the motion speed control equation of the unmanned aerial vehicle is calculated as:

[0105] (2-1)

[0106] Where P 1 is the pressure of the working fluid in the low-pressure chamber of the launch tube 0.1MPa, and P 2 is the pressure of the working fluid in the high-pressure storage tank of the launch tube 9MPa, m is the total mass of the unmanned aerial vehicle and the pneumatic piston 1000kg, A v is the maximum radial cross-sectional area of ​​the unmanned aerial vehicle and the pneumatic piston 0.25 п, g is the gravity constant 9.81; the air friction resistance F is expressed as:

[0107] (2-2)

[0108] Where ρ is the air density, which is 1.225 kg / m 3 , C v The friction coefficient is taken as 0.5; the calculated speed of the unmanned aerial vehicle from the start to the departure from the tube is as follows Figure 5 As shown in the figure, it is a concave curve with an opening upward, which proves that the launch speed and acceleration of the unmanned aerial vehicle are constantly increasing. At the same time, its friction force should continue to increase, and the friction heat gradually increases. Then the friction heat control equation is established as follows:

[0109] (2-3)

[0110] Based on formula (2-1)-formula (2-3), the function of friction heat with respect to the emission distance is obtained. The function of friction heat with respect to the emission distance is processed by finite difference and normalization to obtain the friction heat gradient function. The friction heat gradient function characterizes the rate of change of friction heat along the emission direction:

[0111] (2-4).

[0112] Figure 6 and Figure 7 The following are the calculation results of friction heat normalization with respect to time and launch distance of the aircraft. They are all concave curves, that is, as the aircraft continues to launch and accelerate, the friction heat increases faster and the temperature at the front end of the launch tube continues to increase. The second-order polynomial function of friction heat with respect to the launch distance x is obtained by fitting:

[0113] (2-5)

[0114] The root mean square error and determination coefficient of the curve fitting result are 0.0057 and 0.9996 respectively, which proves the accuracy of the fitting curve. Then, based on the fitting curve, we can get f e (x):

[0115] (2-6)

[0116] Finally, the heat exchange area mirror gradient function and the friction heat gradient function are matched, and the taper of the heat exchange tube is output through iterative calculation, which specifically includes:

[0117] f a (hz) and f e (z) is matched, dA / dz is determined by iterative calculation with the minimum absolute error as the goal, and the taper C of the heat exchange tube is finally output; wherein the goal of the minimum absolute error is expressed as:

[0118] (3-1)

[0119] In the formula, n is the number of axial nodes evenly divided from bottom to top, that is:

[0120] (3-2).

[0121] The final determined taper C is 0.00286, r1 is 0.003m, the channel bottom diameter r0 is 0.0266m, and the heat exchange area of ​​a single heat exchanger is 0.651m 2 . Figure 8 f a (hz) and f e(z) Result diagram: the two gradient curves are very close and almost indistinguishable, which proves the reliability of the taper C calculation results.

[0122] When the UAV is started and launched, supercritical carbon dioxide is filled into the high-pressure storage tank through the filling port, and the initial state inside the structural integrated heat exchanger is vacuum to facilitate the reflux and heat exchange of the working fluid in the tube;

[0123] The start valve opens and fills the carbon dioxide in the high-pressure storage tank into the low-pressure chamber, starting to push the UAV and generating increasing friction heat which is absorbed by the structural integrated heat exchanger. The liquid acetone flows upward from the condensation section at the bottom through the capillary action of the wick, and the working medium temperature in the adiabatic section and evaporation section of the pipeline gradually rises.

[0124] As the opening of the start valve increases and the UAV gradually accelerates, the temperature of the carbon dioxide in the low-pressure chamber rapidly decreases due to expansion. At this time, the friction heat gradually increasing in the evaporation section of the launch tube is quickly transferred to the structural integrated heat exchanger. The liquid acetone in the tube collects heat by evaporation, reducing the temperature of the front end of the launch tube. Then, the gaseous acetone flows back to the low-temperature condensation section outside the low-pressure chamber, liquefies and releases heat into the low-pressure chamber, buffering the temperature drop of the low-pressure chamber and its inlet valve, preventing the carbon dioxide from reaching the triple point and forming dry ice to block the valve;

[0125] When the starting valve is fully opened, a large amount of friction heat generated by the high-speed movement of the UAV in the evaporation section is continuously transferred to the low-pressure chamber through the structural integrated heat exchanger. The acetone in the pipeline increases the flow rate and heat exchange efficiency through rapid phase change in the condensation section and the evaporation section, thereby ensuring the temperature level of the low-pressure chamber and realizing the reliable operation of the high-pressure gas starting launch device.

[0126] According to the designed supercritical carbon dioxide starting launch tube, after the unmanned aerial vehicle is started, the acetone in the pipeline enters the adiabatic section and evaporation section from the liquid phase of the low-pressure chamber condensation section to vaporize and then enters the low-pressure chamber again to condense to form a closed cycle. While efficient heat exchange reduces the temperature at the front end of the launch tube and increases the temperature of the low-pressure chamber, the energy of frictional heat is converted into the internal energy of carbon dioxide in the low-pressure chamber, thereby improving the energy utilization efficiency of the system and further improving the pneumatic ejection capability of the starting launch system.

[0127] As a third aspect of the present invention, a starting and launching device is provided, wherein the starting and launching device comprises a launching tube, and the launching tube comprises the integrated heat exchanger provided in the first aspect of the present invention.

[0128] As a fourth aspect of the present invention, an electronic device is provided, comprising: at least one processor; and a memory communicatively connected to the at least one processor; wherein the memory stores a computer program executable by the at least one processor, and the computer program is executed by the at least one processor so that the at least one processor can execute the integrated heat exchanger parameter determination method provided in the second aspect of the present invention.

[0129] As a fifth aspect of the present invention, a computer-readable medium is provided, wherein the computer-readable storage medium stores computer instructions, and the computer instructions are used to enable a processor to implement the integrated heat exchanger parameter determination method provided in the second aspect of the present invention when executed.

[0130] The above are only specific embodiments of the present invention, but the protection scope of the present invention is not limited thereto. Those skilled in the art should understand that the present invention includes but is not limited to the contents described in the drawings and the above specific embodiments. Any modification that does not deviate from the functional and structural principles of the present invention will be included in the scope of the claims.

Claims

1. An integrated heat exchanger, applied to a launching device, the launching device comprising a launching tube (1), characterized in that: The heat exchanger comprises a plurality of heat exchange tubes, which are circumferentially distributed in the wall of the emitter tube (1); the heat exchange tube comprises a condensation section, an insulation section and an evaporation section which are sequentially connected to form a closed tube cavity (22), the diameter of the closed tube cavity (22) gradually decreasing from the condensation section to the evaporation section; a liquid wick (21) is arranged on the inner wall of the closed tube cavity, and a working medium capable of phase change between gaseous and liquid states is placed in the closed tube cavity, the liquid wick (21) being capable of absorbing liquid working medium and guiding the liquid working medium to flow from the condensation section to the evaporation section.

2. The integrated heat exchanger according to claim 1, characterized in that: The extension direction of the heat exchange tube is consistent with the emission direction of the emission tube, the condensation section of the heat exchange tube is located at the air inlet end of the emission tube (1), and the evaporation section of the heat exchange tube is located at the emission end of the emission tube (1); a heat exchange groove group (23) is arranged between the evaporation section of the heat exchange tube and the inner wall of the emission tube, and the heat exchange groove group (23) includes a plurality of annular grooves arranged at intervals along the axial direction of the heat exchange tube, and the openings of the annular grooves pass through the inner wall of the emission tube.

3. The integrated heat exchanger according to claim 2, characterized in that: In a direction from the condensation section to the evaporation section, the spacing between the plurality of annular grooves gradually decreases.

4. The integrated heat exchanger according to claim 3, characterized in that: The heat exchange tube, the heat exchange groove group (23) and the launch tube (1) are provided as an integrated structure.

5. A method for determining parameters of an integrated heat exchanger, characterized in that: Applied to the heat exchanger according to any one of claims 1 to 4, the method comprises the following steps: Determine the side area of ​​the heat exchange tube according to the taper of the heat exchange tube, the side area is the heat exchange area of ​​the heat exchange tube, and establish a mirror image gradient function of the heat exchange area along the emission direction; According to the motion speed control equation and friction heat control equation of the unmanned aerial vehicle, the friction heat gradient function along the launching direction is obtained; The heat exchange area mirror gradient function and the friction heat gradient function are matched, and the taper of the heat exchange tube is output through iterative calculation.

6. The parameter determination method according to claim 5, characterized in that: The side area of ​​the heat exchange tube is determined according to the taper of the heat exchange tube, the side area is the heat exchange area of ​​the heat exchange tube, and a mirror image gradient function of the heat exchange area along the emission direction is established, including: The heat exchange area of ​​the heat exchange tube is determined according to the side area of ​​the heat exchange tube. The calculation formula of the heat exchange area is as follows: （1-1） In the formula, A is the heat exchange area, h is the height of the heat exchange tube, C is the taper of the heat exchange tube, r0 is the bottom radius of the heat exchange tube, and r1 is the top radius of the heat exchange tube; the taper of the heat exchange tube is expressed as: （1-2） Among them, h ’ is the height of the heat exchange tube from the position corresponding to the pneumatic piston of the launch tube to the top of the heat exchange tube, and r2 is the radius of the tube cavity of the heat exchange tube corresponding to the position of the pneumatic piston in the launch tube; The axial height of the launch tube is defined as the Z axis, and the direction of the launch tube toward the launch end is defined as the positive direction of the Z axis. The heat exchange area gradient of the heat exchange tube along the positive direction of the Z axis is calculated as follows: （1-3） Let x = hz, and calculate the axial mirror gradient of the heat exchange area of ​​the heat exchange tube: （1-4） The axial image gradient of the heat exchange area of ​​the heat exchange tube is processed by finite difference and normalization to obtain the heat exchange area image gradient function: （1-5） 7. The parameter determination method according to claim 6, characterized in that: The friction heat gradient function along the launch direction is obtained according to the motion speed control equation and the friction heat control equation of the unmanned aerial vehicle, including: According to the parameters of the launch device, the motion speed control equation of the unmanned aerial vehicle is calculated as: （2-1） Where, P1 is the pressure of the working fluid in the low-pressure chamber of the launch tube, P2 is the pressure of the working fluid in the high-pressure storage tank of the launch tube, m is the total mass of the unmanned aerial vehicle and the pneumatic piston, A v is the maximum radial cross-sectional area of ​​the UAV and the pneumatic piston, g is the gravity constant; the air friction resistance F is expressed as: （2-2） Where ρ is the air density, C v is the friction coefficient; then the friction heat control equation is established as: （2-3） Based on formula (2-1)-formula (2-3), the function of friction heat with respect to the emission distance is obtained. The function of friction heat with respect to the emission distance is processed by finite difference and normalization to obtain the friction heat gradient function. The friction heat gradient function characterizes the rate of change of friction heat along the emission direction: （2-4） 8. The parameter determination method according to claim 7, characterized in that: include: f a (hz) and f e (z) is matched, dA / dz is determined by iterative calculation with the minimum absolute error as the goal, and the taper C of the heat exchange tube is finally output; wherein the goal of the minimum absolute error is expressed as: （3-1） In the formula, n is the number of axial nodes evenly divided from bottom to top, that is: （3-2） 9. A launching device, characterized in that: The starting and launching device comprises the integrated heat exchanger according to any one of claims 1 to 8.

10. An electronic device, characterized in that: The electronic device includes: at least one processor; and a memory communicatively connected to the at least one processor; wherein the memory stores a computer program executable by the at least one processor, and the computer program is executed by the at least one processor so that the at least one processor can execute the integrated heat exchanger parameter determination method described in any one of claims 5-6.

11. A computer readable medium, characterized in that The computer-readable storage medium stores computer instructions, and the computer instructions are used to enable a processor to implement the integrated heat exchanger parameter determination method according to any one of claims 5 to 6 when executed.

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