A method and device for actively regulating and controlling aerodynamic heat flow of a hypersonic carrier rocket

Abstract

The application discloses a hypersonic launch vehicle aerodynamic heat flow active regulation control method and a regulation and control device, wherein the hypersonic launch vehicle aerodynamic heat flow active regulation control method comprises the following steps: determining a heat flow concentration area; arranging arrayed micro-jet nozzles and heat flow sensors on the outer surface of the heat flow concentration area; judging whether the heat flow data collected by the heat flow sensors exceeds a safety threshold; if the safety threshold is exceeded, entering an active regulation state, opening low-temperature heat flow and adjusting nozzle parameters; judging whether the heat flow data returns to the safety threshold; if the safety threshold is returned to, reducing the jetting intensity or closing the nozzle, and exiting the active regulation state. The application can significantly reduce the local heat flow peak value of the key part of the rocket under the hypersonic flight condition, effectively alleviate the heat flow concentration phenomenon, and improve the overall heat flow distribution uniformity, so that the structural thermal load and the thermal damage risk are reduced.

Description

Technical Field

[0001] This application relates to the aerospace field, specifically to a method and device for active regulation and control of aerodynamic heat flux in hypersonic launch vehicles. Background Technology

[0002] In the design and engineering applications of existing hypersonic launch vehicles and related aircraft, passive thermal protection is the most common technical solution to address the intense aerodynamic heating caused by high Mach number flight conditions. This type of solution typically involves applying ablative heat-resistant materials, heat-insulating tile structures, or high-temperature resistant metal and composite material skins to the rocket's outer surface, relying on the materials' inherent heat resistance, ablation heat absorption, or insulation capabilities to resist aerodynamic heat flow. In areas prone to heat concentration, such as the rocket's nose, interstage sections, and transitional regions, structural safety is ensured under extreme aerodynamic thermal environments by increasing the thickness of the thermal protection structure, selecting higher-grade heat-resistant materials, or increasing structural design margins. Furthermore, some existing technologies optimize flight trajectory and attitude control strategies to minimize prolonged exposure to high aerodynamic heat loads, reducing overall thermal protection pressure from a mission planning perspective. In recent years, a few studies have attempted to introduce cooling or flow control methods, but these primarily rely on passive cooling structures or fixed-parameter flow field control, typically designed for specific operating conditions, and have not yet formed a complete technical system capable of dynamic adjustment according to changes in flight conditions. While the aforementioned solutions can meet the thermal protection requirements of traditional hypersonic flight missions to a certain extent, their shortcomings have gradually become apparent in the engineering applications of next-generation hypersonic launch vehicles. First, once the structure and materials of passive thermal protection schemes are determined, their protective capabilities are essentially fixed, making it impossible to adjust them according to real-time changes in the aerodynamic and thermal environment during flight. This makes them ineffective in addressing unsteady and sudden heat flux peaks caused by enhanced shock-boundary layer interactions, local flow separation, or flight state disturbances. Second, to ensure safety margins under extreme conditions, passive protection often requires significantly increasing material thickness or using high-density heat-resistant materials, leading to an increase in the mass of the thermal protection system and reducing the launch vehicle's payload capacity and overall performance. Furthermore, ablation-type thermal protection structures consume materials during use, hindering the reusability and lifespan management of launch vehicles, while also increasing maintenance costs and mission uncertainty. The few existing schemes involving flow field control or cooling often lack real-time sensing and feedback control capabilities, typically operating only under preset conditions and failing to adapt to the complex and variable aerodynamic and thermal environment during hypersonic flight.

[0003] Therefore, how to provide a method that can achieve real-time sensing and active control of local aerodynamic heat flow while ensuring lightweight design has become an urgent problem to be solved in this field. Summary of the Invention

[0004] To address the aforementioned problems, this application proposes an active aerodynamic heat flux control device for hypersonic launch vehicles, characterized by comprising: a heat flux concentration region determination module, an array of microjet nozzles, a high-temperature heat flux sensor, a cryogenic gas supply system, and a control unit; the heat flux concentration region determination module is used to determine the heat flux concentration region; the array of microjet nozzles is arranged on the outer surface of the heat flux concentration region and connected to the cryogenic gas supply system built into the rocket body, used to accurately inject the cryogenic medium into the rocket's heat flux concentration region; the high-temperature heat flux sensor is arranged in the heat flux concentration region, used to collect local heat flux density data in real time; the cryogenic gas supply system is used to provide the cryogenic injection medium in the active control state; the control unit is used to receive the high-temperature heat flux sensor signal and compare it with a safe heat flux threshold, and enter the active control state according to the heat flux exceeding the limit, automatically adjusting the nozzle opening and injection parameters.

[0005] The hypersonic launch vehicle aerodynamic heat flow active control device described above includes heat flow concentration areas in the rocket nose stagnation area, interstage section, and areas of geometric abrupt change.

[0006] The hypersonic launch vehicle aerodynamic heat flow active control device described above includes a cryogenic gas supply system comprising a cryogenic storage tank, a decompression and control component, and liquid nitrogen as the injection medium.

[0007] The hypersonic launch vehicle aerodynamic thermal flux active control device described above includes injection parameters such as injection pressure and injection flow rate.

[0008] The hypersonic launch vehicle aerodynamic heat flow active control device described above uses an array of microjet nozzles made of high-temperature resistant materials.

[0009] A method for active regulation and control of aerodynamic heat flux of a hypersonic launch vehicle based on any of the above-mentioned devices includes the following steps: determining a heat flux concentration region; arranging an array of microjet nozzles and a heat flux sensor on the outer surface of the heat flux concentration region; determining whether the heat flux data collected by the heat flux sensor exceeds a safety threshold; if it exceeds the safety threshold, entering an active regulation state, activating cryogenic heat flux and adjusting nozzle parameters; determining whether the heat flux data returns to within the safety threshold; if it returns to within the safety threshold, reducing the jet intensity or closing the nozzle, and exiting the active regulation state.

[0010] The above-mentioned active aerodynamic heat flow control method for hypersonic launch vehicles uses the stagnation region of the rocket head, the interstage section, and the parts where the geometry changes abruptly as heat flow concentration areas.

[0011] The hypersonic launch vehicle aerodynamic heat flux active control method described above includes determining whether the heat flux data collected by the heat flux sensor exceeds the safety threshold by comparing the collected heat flux density data with a pre-set safety heat flux threshold. If the heat flux density data exceeds the safety heat flux threshold, the active control state is entered.

[0012] The above-described active control method for aerodynamic thermal flux of hypersonic launch vehicles includes, in which entering the active control state, activating cryogenic thermal flux, and adjusting nozzle parameters, the following steps are taken: determining the degree of thermal flux exceeding the limit; and activating cryogenic thermal flux and adjusting nozzle parameters based on the degree of thermal flux exceeding the limit.

[0013] The hypersonic launch vehicle aerodynamic heat flux active regulation and control method described above involves exiting the active regulation state and returning to the monitoring state. When the heat flux density is detected to exceed the safety threshold again, a new round of active regulation is initiated.

[0014] This application has the following beneficial effects: (1) This application can significantly reduce the local heat flux peak value of key parts of the rocket under hypersonic flight conditions, effectively alleviate the heat flux concentration phenomenon, and improve the overall heat flux distribution uniformity, thereby reducing the structural heat load and the risk of thermal damage. This application achieves the thermal protection effect by actively adjusting the aerodynamic flow field, which can improve the thermal protection capability without significantly increasing the thickness and mass of the thermal protection structure, and is conducive to improving the effective payload capacity and overall performance of the launch vehicle.

[0015] (2) The closed-loop control method adopted in this application can adapt to the complex and variable aerodynamic thermal environment during hypersonic flight, and has good response capability to unsteady heat flow and sudden heat flow peaks, significantly improving the safety and reliability of the thermal protection system. At the same time, it also has good engineering integration and can be used in conjunction with existing passive thermal protection structures, providing a smart aerodynamic thermal protection solution with practical engineering application value for hypersonic launch vehicles and reusable launch vehicles. Attached Figure Description

[0016] To more clearly illustrate the technical solutions in the embodiments of this application or the prior art, the drawings used in the description of the embodiments or the prior art will be briefly introduced below. Obviously, the drawings described below are only some embodiments recorded in this application. For those skilled in the art, other drawings can be obtained based on these drawings.

[0017] Figure 1 This is a flowchart illustrating the active aerodynamic thermal flux control method for hypersonic launch vehicles provided in the embodiments of this application. Figure 2 This is a schematic diagram of the internal structure of the active aerodynamic heat flow control device for hypersonic launch vehicles provided in the embodiments of this application. Detailed Implementation

[0018] The technical solutions of the embodiments of this application will be clearly and completely described below with reference to the accompanying drawings. Obviously, the described embodiments are only some, not all, of the embodiments of this application. All other embodiments obtained by those skilled in the art based on the embodiments of this application without creative effort are within the scope of protection of this application.

[0019] This invention proposes an active aerodynamic heat flux control device and its control strategy for hypersonic launch vehicles, overcoming the limitations of traditional passive thermal protection technologies in terms of adjustment capability and adaptability. This invention introduces cryogenic active jets into the concentrated heat flux regions of the rocket, actively altering the local aerodynamic flow field structure and weakening shock wave-boundary layer interactions and wall heat transfer intensity from an aerodynamic mechanism perspective, thereby effectively reducing local heat flux peaks. Simultaneously, this invention deploys high-temperature heat flux sensors at key locations to acquire aerodynamic thermal environment information in real time and constructs a closed-loop control strategy based on heat flux thresholds. This allows the timing, intensity, and duration of the cryogenic jet to be adaptively adjusted according to the actual heat flux state, achieving precise and on-demand control of aerodynamic heat flux. Through these solutions, this invention improves the adaptability of hypersonic launch vehicles to complex and unsteady aerodynamic thermal environments without significantly increasing structural mass, enhancing the safety, reliability, and engineering applicability of the overall thermal protection system.

[0020] Example 1 like Figure 1 As shown in the figure, this embodiment provides an active aerodynamic thermal flux regulation and control method for hypersonic launch vehicles, which specifically includes the following steps: Step S1: Determine the area of ​​concentrated heat flow.

[0021] During the overall rocket design phase, aerodynamic and thermal numerical simulation and flight envelope analysis were used to identify and assess areas where heat flow may be concentrated during hypersonic flight, determining the stagnation point area at the rocket nose, the interstage section, and the locations of abrupt changes in geometric shape as areas of concentrated heat flow.

[0022] Step S2: Arrange an array of miniature jet nozzles and a heat flow sensor on the outer surface of the heat flow concentration area.

[0023] An array of micro-jet nozzles is arranged on the outer surface of the heat-concentrated area to ensure that the cryogenic jet can act uniformly and directionally on the key heated parts, achieving precise disturbance of the local flow field. The nozzles adopt a high-temperature resistant structural design and are connected to the cryogenic gas supply system inside the rocket body.

[0024] The cryogenic gas supply system is located inside the rocket body. This system includes a cryogenic storage tank, depressurization and control components, and preferably uses liquid nitrogen as the injection medium. The liquid nitrogen is stored in the storage tank and, under the command of the control system, is depressurized and vaporized before being ejected from microjet nozzles in the corresponding areas. During the injection process, the cryogenic gas jet interacts with the hypersonic incoming flow, actively disturbing the near-wall boundary layer structure, weakening the shock-boundary layer interaction intensity, thickening the local boundary layer, and reducing the wall temperature gradient, thereby reducing the local aerodynamic heat flux peak.

[0025] Furthermore, to achieve real-time and adaptive control of the process, high-temperature heat flux sensors are synchronously arranged in the same area as the nozzle array. The distribution of the high-temperature heat flux sensors matches the array layout of the nozzles, ensuring that the sensors can accurately collect real-time heat flux density data for each jet control area.

[0026] Understandably, areas with higher heat flux density and larger heat flux concentration areas require more nozzles and sensors. The number of sensors should be matched with the nozzle array as needed to meet the requirements of real-time, full-coverage acquisition of local heat flux density data; there is no need for a one-to-one correspondence between sensors and nozzles.

[0027] Step S3: Determine whether the heat flow data collected by the heat flow sensor exceeds the safety threshold.

[0028] The sensor signals are transmitted to the control unit in real time. The control unit compares the collected heat flux density data with a preset safe heat flux threshold. When the heat flux in a certain area exceeds the safe heat flux threshold, step S4 is executed.

[0029] If the safety threshold is exceeded, proceed to step S4. If the safety threshold is not exceeded, proceed to step S7.

[0030] Step S4: Enter active control mode, turn on low temperature heat flow and adjust nozzle parameters.

[0031] The control unit automatically determines that the area has entered an active control state, triggers the corresponding micro-jet nozzle to open, and adjusts the injection pressure and injection flow rate according to the degree of heat flow exceeding the limit.

[0032] Different safe heat flux thresholds can be set for each heat flux concentration area. The difference between the real-time heat flux density of each area and the corresponding safe heat flux threshold is calculated to obtain the heat flux exceedance value of the corresponding area. The larger the exceedance value, the higher the aerodynamic heat flux risk in the area and the stronger the jet intervention intensity required.

[0033] If the over-limit is low (over-limit value ≤ 20% of the safety threshold): turn on the low-speed jet, use low pressure + small flow rate injection, and only slightly disturb the near-wall flow field with weak jet to save liquid nitrogen working fluid to the maximum extent while meeting the heat flow control requirements.

[0034] If the over-limit is moderate (20% < over-limit value ≤ 50%): switch to medium-speed jet, use medium pressure + medium flow rate to moderately enhance the flow field disturbance effect and quickly suppress the upward trend of heat flow.

[0035] If the exceedance is high (exceedance value > 50%): directly activate the high-level jet, using high pressure + high flow rate injection, to significantly thicken the local boundary layer through strong jet, quickly weaken the shock wave-boundary layer interaction, rapidly reduce the peak heat flux, and cope with the high-risk heat flux state.

[0036] Step S5: Determine whether the heat flow data has recovered to within the safe threshold.

[0037] If the temperature returns to within the safe heat flux threshold, proceed to step S6. Otherwise, return to step S4.

[0038] Step S6: Reduce the spray intensity or close the nozzle to exit the active control state.

[0039] During the jet action, the control unit continuously receives feedback data from the heat flux sensor and dynamically corrects the injection parameters based on changes in heat flux. When the heat flux is detected to have decreased to within a safe threshold, the control unit gradually reduces the injection intensity or closes the nozzle, exiting the active control state, thus forming a complete closed-loop control process based on real-time heat flux monitoring.

[0040] Step S7: Return to monitoring status.

[0041] Upon returning to normal operation, the array-type micro-jet nozzles cease ejecting cryogenic gas, or maintain only a very low standby ejection level. The cryogenic gas supply system no longer supplies high-pressure, high-flow-rate vaporized cryogenic medium to the nozzles. High-temperature heat flux sensors located in areas of concentrated heat flux maintain their normal acquisition frequency, continuously and in real-time capturing local heat flux density data on the rocket's outer surface and transmitting the data uninterruptedly to the control unit without any interruption or frequency reduction. The control unit continues to compare and analyze the real-time heat flux data transmitted by the sensors with preset safe heat flux thresholds, operating in a standby response mode for real-time monitoring and immediate judgment. Once the heat flux density is detected to exceed the safe threshold again, a new round of active jet control procedures will be immediately triggered.

[0042] Example 2 like Figure 2 As shown in the figure, a hypersonic launch vehicle aerodynamic heat flow active control device provided in this application embodiment includes: a heat flow concentration area determination module 210, an array of micro jet nozzles 220, a high temperature heat flow sensor 230, a cryogenic gas supply system 240, and a control unit 250.

[0043] The heat flow concentration area determination module 210 is used to determine the heat flow concentration area.

[0044] During the overall rocket design phase, aerodynamic and thermal numerical simulation and flight envelope analysis were used to identify and assess areas where heat flow may be concentrated during hypersonic flight, determining the stagnation point area at the rocket nose, the interstage section, and the locations of abrupt changes in geometric shape as areas of concentrated heat flow.

[0045] An array of micro-jet nozzles 220 is arranged on the outer surface of the aforementioned heat concentration area and is connected to the cryogenic gas supply system built into the rocket body to precisely spray the cryogenic medium into the rocket's heat concentration area.

[0046] An array of micro-jet nozzles is arranged on the outer surface of the heat flow concentration area to ensure that the low temperature jet can act uniformly and directionally on the key heated parts, and achieve precise disturbance of the local flow field. The nozzle adopts a high temperature resistant structural design.

[0047] The high-temperature heat flux sensor 230 is placed in the area of ​​concentrated heat flux to collect local heat flux density data in real time.

[0048] The distribution of high-temperature heat flux sensors matches the array layout of the nozzles, ensuring that the sensors can accurately collect real-time heat flux density data for each jet control area.

[0049] The cryogenic gas supply system 240 is used to provide cryogenic injection medium in an actively controlled state.

[0050] The cryogenic gas supply system is located inside the rocket body. This system includes a cryogenic storage tank, depressurization and control components, and preferably uses liquid nitrogen as the injection medium. The liquid nitrogen is stored in the storage tank and, under the command of the control system, is depressurized and vaporized before being ejected from microjet nozzles in the corresponding areas. During the injection process, the cryogenic gas jet interacts with the hypersonic incoming flow, actively disturbing the near-wall boundary layer structure, weakening the shock wave-boundary layer interaction intensity, thickening the local boundary layer, and reducing the wall temperature gradient, thereby reducing the local aerodynamic heat flux peak.

[0051] The control unit 250 is used to receive sensor signals and compare them with a safe heat flow threshold, and automatically adjust the opening of the nozzle, the injection pressure and the flow rate according to the heat flow exceeding the limit.

[0052] The control unit 250 compares the collected heat flow data with a preset safe heat flow threshold. When the heat flow in a certain area exceeds the safe heat flow threshold, the low-temperature heat flow is activated and the nozzle parameters are adjusted.

[0053] The control system automatically determines that the area has entered an active control state, triggers the corresponding micro-jet nozzle to open, and adjusts the injection pressure and injection flow rate according to the degree of heat flow exceeding the limit.

[0054] Different safe heat flux thresholds can be set for each heat flux concentration area. The difference between the real-time heat flux density of each area and the corresponding safe heat flux threshold is calculated to obtain the heat flux exceedance value of the corresponding area. The larger the exceedance value, the higher the aerodynamic heat flux risk in the area and the stronger the jet intervention intensity required.

[0055] If the over-limit is low (over-limit value ≤ 20% of the safety threshold): turn on the low-speed jet, use low pressure + small flow rate injection, and only slightly disturb the near-wall flow field with weak jet to save liquid nitrogen working fluid to the maximum extent while meeting the heat flow control requirements.

[0056] If the over-limit is moderate (20% < over-limit value ≤ 50%): switch to medium-speed jet, use medium pressure + medium flow rate to moderately enhance the flow field disturbance effect and quickly suppress the upward trend of heat flow.

[0057] If the exceedance is high (exceedance value > 50%): directly activate the high-level jet, using high pressure and high flow rate to significantly thicken the local boundary layer through strong jet, quickly weaken the shock wave-boundary layer interaction, rapidly reduce the heat flux peak, and cope with the high-risk heat flux state.

[0058] The control unit 250 also needs to determine whether the heat flux data has returned to within the safe heat flux threshold. If it has returned to the safe threshold, the injection intensity is reduced or the nozzle is shut off. If it has not returned to the safe threshold, the low-temperature heat flux continues to operate and the nozzle parameters are adjusted.

[0059] This application also provides a computer storage medium storing computer instructions, which, when invoked, are used to execute the hypersonic launch vehicle aerodynamic thermal flux active regulation and control method.

[0060] The embodiments disclosed in this invention provide a computer-readable storage medium storing computer program instructions. When the computer program instructions are executed on a computer, the computer performs the above-described active aerodynamic thermal flux control method for hypersonic launch vehicles.

[0061] This invention provides a processor for processing the above-described active aerodynamic thermal flux regulation and control method for hypersonic launch vehicles.

[0062] In this embodiment of the invention, the processor can be an integrated circuit chip with signal processing capabilities. The processor can be a general-purpose processor, a digital signal processor (DSP), an application-specific integrated circuit (ASIC), a field-programmable gate array (FPGA), or other programmable logic devices, discrete gate or transistor logic devices, or discrete hardware components.

[0063] The various methods, steps, and logic diagrams disclosed in the embodiments of this invention can be implemented or executed. The general-purpose processor can be a microprocessor or any conventional processor. The steps of the methods disclosed in the embodiments of this invention can be directly implemented by a hardware decoding processor, or implemented by a combination of hardware and software modules in the decoding processor. The software modules can reside in random access memory, flash memory, read-only memory, programmable read-only memory, electrically erasable programmable memory, registers, or other mature storage media in the art. The processor reads information from the storage medium and, in conjunction with its hardware, completes the steps of the above methods.

[0064] The storage medium can be memory, such as volatile memory or non-volatile memory, or may include both volatile and non-volatile memory.

[0065] Non-volatile memory can be read-only memory (ROM), programmable read-only memory (PROM), erasable programmable read-only memory (EPROM), electrically erasable programmable read-only memory (EEPROM), or flash memory. Volatile memory can be random access memory (RAM), which is used as an external cache. By way of example, but not limitation, many forms of RAM are available, such as Static Random Access Memory (SRAM), Dynamic Random Access Memory (DRAM), Synchronous DRAM (SDRAM), Double Data Rate Synchronous DRAM (DDRSDRAM), Enhanced Synchronous DRAM (ESDRAM), Synchlink DRAM (SLDRAM), and Direct Rambus RAM (DRRAM).

[0066] This application has the following beneficial effects: (1) This application can significantly reduce the local heat flux peak value of key parts of the rocket under hypersonic flight conditions, effectively alleviate the heat flux concentration phenomenon, and improve the overall heat flux distribution uniformity, thereby reducing the structural heat load and the risk of thermal damage. This application achieves the thermal protection effect by actively adjusting the aerodynamic flow field, which can improve the thermal protection capability without significantly increasing the thickness and mass of the thermal protection structure, and is conducive to improving the effective payload capacity and overall performance of the launch vehicle.

[0067] (2) The closed-loop control method adopted in this application can adapt to the complex and variable aerodynamic thermal environment during hypersonic flight, and has good response capability to unsteady heat flow and sudden heat flow peaks, significantly improving the safety and reliability of the thermal protection system. At the same time, it also has good engineering integration and can be used in conjunction with existing passive thermal protection structures, providing a smart aerodynamic thermal protection solution with practical engineering application value for hypersonic launch vehicles and reusable launch vehicles.

[0068] Although the examples referenced in this application are described for illustrative purposes only and not for limiting the scope of this application, changes, additions and / or deletions to the implementation may be made without departing from the scope of this application.

[0069] The above description is merely a specific embodiment of this application, but the scope of protection of this application is not limited thereto. Any variations or substitutions that can be easily conceived by those skilled in the art within the scope of the technology disclosed in this application should be included within the scope of protection of this application. Therefore, the scope of protection of this application should be determined by the scope of the claims.

Claims

1. A hypersonic launch vehicle aerodynamic heat flux active control device, characterized in that, include: The heat flow concentration area determination module, array-type micro jet nozzles, high-temperature heat flow sensor, low-temperature gas supply system and control unit; The heat flux concentration area determination module is used to determine the heat flux concentration area; An array of micro-jet nozzles is arranged on the outer surface of the heat concentration area and connected to the cryogenic gas supply system built into the rocket body to precisely spray the cryogenic medium into the heat concentration area of ​​the rocket. High-temperature heat flux sensors are placed in areas of concentrated heat flux to collect local heat flux density data in real time. A cryogenic gas supply system is used to provide cryogenic injection media in an actively controlled state. The control unit receives signals from the high-temperature heat flux sensor and compares them with the safe heat flux threshold. If the heat flux exceeds the limit, it enters an active control state and automatically adjusts the opening of the nozzle and the injection parameters.

2. The active aerodynamic heat flux control device for hypersonic launch vehicles as described in claim 1, characterized in that, Areas of concentrated heat flow include the stagnation point area at the rocket nose, the interstage section, and areas where the geometry changes abruptly.

3. The active aerodynamic heat flux control device for hypersonic launch vehicles as described in claim 1, characterized in that, The cryogenic gas supply system includes a cryogenic storage tank, a pressure reduction and control component, and uses liquid nitrogen as the injection medium.

4. The active aerodynamic heat flux control device for hypersonic launch vehicles as described in claim 1, characterized in that, Injection parameters include injection pressure and injection flow rate.

5. The active aerodynamic heat flux control device for hypersonic launch vehicles as described in claim 6, characterized in that, The array-type micro jet nozzles are made of high-temperature resistant materials.

6. A method for active regulation and control of aerodynamic thermal flux of a hypersonic launch vehicle based on the device described in any one of claims 1 to 5, characterized in that, Includes the following steps: Identify areas of concentrated heat flow; An array of miniature jet nozzles and heat flow sensors are arranged on the outer surface of the heat flow concentration area; Determine whether the heat flux data collected by the heat flux sensor exceeds the safety threshold; If the safety threshold is exceeded, the system enters active control mode, activates low-temperature heat flow, and adjusts nozzle parameters. Determine whether the heat flux data has returned to within the safe threshold; If the spray intensity returns to within the safe threshold, reduce the spray intensity or close the nozzle to exit the active control state.

7. The active aerodynamic thermal flux control method for hypersonic launch vehicles as described in claim 6, characterized in that, The stagnation area at the rocket's nose, the interstage section, and the areas where the geometry changes abruptly are considered as areas of concentrated heat flow.

8. The active aerodynamic thermal flux control method for hypersonic launch vehicles as described in claim 7, characterized in that, Determining whether the heat flux data collected by the heat flux sensor exceeds the safety threshold involves comparing the collected heat flux density data with a pre-set safety heat flux threshold. If the heat flux density data exceeds the safety heat flux threshold, the system enters an active control state.

9. The active aerodynamic thermal flux control method for hypersonic launch vehicles as described in claim 8, characterized in that, Entering active control mode, activating low-temperature heat flow, and adjusting nozzle parameters include: Determine the extent of heat flux exceeding the limit; The low-temperature heat flow is activated and the nozzle parameters are adjusted according to the degree of heat flow exceeding the limit.

10. The active aerodynamic thermal flux control method for hypersonic launch vehicles as described in claim 6, characterized in that, After exiting the active control state, it returns to the monitoring state. When the heat flux density is detected to exceed the safety threshold again, it enters a new round of active control.

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