Recyclable secondary rocket energy comprehensive management system and method

Abstract

This invention relates to the field of spacecraft reusability technology, specifically disclosing a comprehensive energy management system and method for a reusable second-stage rocket. By placing a heat capture and energy conversion module in the reentry thermal protection zone of the second-stage rocket, this invention utilizes cryogenic propellant to absorb aerodynamic heat, generating a high-temperature, high-pressure working fluid which is then converted into electrical energy. This is combined with a power distribution and management module to convert, store, and distribute the electrical energy, and a power execution and application module to drive the cryogenic propellant cycle and power subsequent missions. This solves the problems of single energy form, strong control coupling, serious energy waste, and lack of energy reuse functionality in existing technologies, thereby improving the energy utilization efficiency and mission adaptability of reusable second-stage rockets.

Description

Technical Field

[0001] This invention relates to the field of spacecraft reusability technology, and in particular to a reusable second-stage rocket energy integrated management system and method. Background Technology

[0002] As space transportation systems develop towards lower cost and higher frequency, reusable rocket technology has become a core pathway to improve launch economy and resource utilization efficiency. In existing reusable rockets, the second stage is responsible for delivering the payload into the designated orbit. After orbital deployment, it needs to re-enter the Earth's atmosphere at high speed. During re-entry, the rocket body experiences intense friction with the air, generating a large amount of aerodynamic heat and creating a harsh aerodynamic heating environment. If thermal protection measures are inadequate, the high temperature will lead to the degradation or even damage of the rocket's structural materials, directly affecting the rocket's reusability and flight safety. Therefore, developing efficient and reliable active cooling technology is of great significance for addressing the thermal protection problem of the second-stage rocket during re-entry.

[0003] To address the aforementioned issues, an active cooling technology solution has emerged that uses the rocket's own cryogenic propellant as the cooling medium. This technology uses liquid methane or liquid hydrogen as the cooling medium, introducing it into a heat exchanger located on the rocket's windward side. Through heat exchange, it absorbs aerodynamic heat, causing the cryogenic propellant to evaporate and form a high-temperature, high-pressure gas. This gas drives a turbine to rotate, and the turbine is rigidly connected to a cooling pump via a mechanical shaft, directly driving the cooling pump and forming a self-sustaining cooling closed loop to achieve thermal protection of the rocket body.

[0004] However, the aforementioned existing technical solutions have many limitations in practical engineering applications and cannot meet the requirements of efficient, reliable, and flexible energy management for reusable second-stage rockets. The specific limitations are as follows: First, the energy form is singular: the mechanical energy output by the turbine can only be directly transferred to the cooling pump through the mechanical shaft to drive the pump body. The energy form cannot be converted, so it cannot be flexibly distributed to other electrical equipment on the rocket, nor can energy be stored, which greatly limits the scope of energy application.

[0005] Second, the control coupling is strong: the cooling pump flow rate regulation depends entirely on the turbine speed, and the two are rigidly coupled through the mechanical shaft; the turbine speed changes with the fluctuation of aerodynamic heat flow, which makes it impossible to adjust the pump flow rate independently, making it difficult to accurately adapt to real-time heat flow changes, and the response capability to sudden heat flow impact is insufficient, which may affect the thermal protection effect.

[0006] Third, there is serious energy waste: during rocket reentry, there is a significant peak in aerodynamic heat flow, at which point the turbine output power is far higher than the actual demand of the cooling pump. Since the existing scheme lacks an effective energy recovery mechanism, the excess energy can only be lost through depressurization, dissipation and other means, which significantly reduces energy utilization efficiency.

[0007] Fourth, there is no energy reuse function: after the reentry phase, the working intensity of the active cooling circuit decreases, and the turbine output power decreases accordingly. The previously captured energy cannot be stored and reused, and cannot provide energy support for subsequent rocket missions (such as pressurization of the propellant tank before secondary ignition, standby power supply for onboard equipment, etc.), thus failing to maximize the value of energy across stages.

[0008] In summary, existing active cooling technologies have significant shortcomings in energy form conversion, independent flow control, excess energy recovery, and cross-stage energy reuse. There is an urgent need for a technical solution that can achieve comprehensive energy management to improve the energy utilization efficiency and mission adaptability of reusable second-stage rockets. Summary of the Invention

[0009] To address the aforementioned technical problems, this invention provides a reusable second-stage rocket energy integrated management system and method.

[0010] Firstly, this invention provides a reusable second-stage rocket energy integrated management system. During operation, in the initial stage of reentry, the energy storage unit powers an electric pump to pump cryogenic propellant into the active cooling circuit to absorb aerodynamic heat. The heated, high-temperature, high-pressure working propellant drives a turbine power generation unit to generate electricity, which is then rectified and fed into the DC bus. The integrated controller dynamically determines the allocation of power generation, energy storage charging / discharging power, and electric pump power based on parameters such as real-time heat flow, remaining energy in the energy storage unit, and the electric pump's demand. During peak heat flow, excess energy is stored in the energy storage unit; during low heat flow, the energy storage unit discharges to supplement the power gap; after reentry, remaining energy is used for subsequent tasks such as tank pressurization. Specifically, the technical solution of the reusable second-stage rocket energy integrated management system is as follows: The heat capture and energy conversion module is installed in the reentry heat protection zone of the second-stage rocket. It uses the cryogenic propellant carried by the second-stage rocket to absorb aerodynamic heat to generate a high-temperature and high-pressure working fluid, and converts the thermal energy of the high-temperature and high-pressure working fluid into electrical energy output. The power distribution and management module is used to convert the electrical energy output by the heat capture and energy conversion module into stable DC power, store the stable DC power, and distribute it according to control commands. The power execution and application module is used to drive the cryogenic propellant to circulate in the heat capture and energy conversion module according to the electrical energy allocated by the power distribution and management module, and to provide electrical energy for the subsequent mission of the second-stage rocket.

[0011] The beneficial effects of the reusable second-stage rocket energy integrated management system of the present invention are as follows: The system of this invention sets up a heat capture and energy conversion module in the reentry thermal protection zone of the second-stage rocket. It utilizes cryogenic propellant to absorb aerodynamic heat to generate a high-temperature, high-pressure working fluid and convert it into electrical energy. The power distribution and management module converts, stores, and distributes the electrical energy, and the power execution and application module drives the cryogenic propellant cycle and supplies power for subsequent missions. This solves the problems of single energy form, strong control coupling, serious energy waste, and lack of energy reuse function in the prior art, and improves the energy utilization efficiency and mission adaptability of reusable second-stage rockets.

[0012] Based on the above scheme, the reusable second-stage rocket energy integrated management system of the present invention can be further improved as follows.

[0013] In one alternative embodiment, the heat capture and energy conversion module includes: an active cooling loop and a gas turbine power generation unit; The active cooling circuit is located in the reentry thermal protection zone of the second-stage rocket and is used to absorb the aerodynamic heat and generate the high-temperature and high-pressure working fluid using the cryogenic propellant carried by the second-stage rocket. The gas turbine power generation unit is connected to the active cooling circuit and is used to receive the high-temperature and high-pressure working fluid and convert the thermal energy of the high-temperature and high-pressure working fluid into electrical energy output.

[0014] The beneficial effects of adopting the above-mentioned optional method are as follows: the heat capture and energy conversion module is further divided into an active cooling circuit and a gas turbine power generation unit, so that the cryogenic propellant absorbs aerodynamic heat in the re-entry heat protection zone to form a high-temperature and high-pressure working fluid. The heat energy is converted into electrical energy output through the gas turbine power generation unit, realizing aerodynamic heat capture and energy form conversion, solving the problem of the single energy form in the existing technology, and expanding the energy utilization methods.

[0015] In one alternative embodiment, the power distribution and management module includes a rectifier / voltage regulator unit, an energy storage unit, an integrated controller, and a power distribution unit; The rectifier / voltage regulator unit is connected to the gas turbine power generation unit and is used to convert the electrical energy output by the gas turbine power generation unit into the stable DC power and feed it into the DC bus. The energy storage unit is connected to the DC bus via a bidirectional DC / DC converter for storing the stable DC power. The integrated controller is connected to the sensor layer, which is deployed at the corresponding measurement points of the second-stage rocket. The integrated controller is used to acquire real-time operating parameters and generate control commands based on the real-time operating parameters. The power distribution unit is connected to the DC bus and the integrated controller, and is used to distribute the stable DC power on the DC bus according to the control command.

[0016] The advantages of adopting the above-mentioned optional method are as follows: the electrical energy output by the gas turbine generator unit is further converted into stable DC power and fed into the DC bus through the rectification / voltage regulation unit; the energy storage unit is connected to the DC bus through a bidirectional DC / DC converter to realize the storage of electrical energy; the integrated controller generates control commands according to real-time operating parameters; and the power distribution unit distributes electrical energy according to the control commands, thereby realizing the integrated management of electrical energy conversion, storage and intelligent distribution.

[0017] In one alternative embodiment, the power execution and application module includes an electric pump unit and an auxiliary function unit; The electric pump unit is connected to the power distribution unit and is used to drive the cryogenic propellant to circulate in the active cooling loop according to the stable DC power distributed by the power distribution unit. The auxiliary function unit is connected to the power distribution unit and is used to provide power for the subsequent missions of the second-stage rocket according to the stable DC power distributed by the power distribution unit.

[0018] The advantages of adopting the above-mentioned optional approach are as follows: the electric pump unit is further set up to drive the cryogenic propellant to circulate in the active cooling loop according to the electrical energy allocated by the power distribution unit, and the auxiliary functional unit provides electrical energy for the subsequent mission of the second stage rocket, realizing the dual functions of cooling cycle drive and mission power supply, solving the problem of no energy reuse function in the existing technology, and improving the level of comprehensive energy utilization.

[0019] In one alternative embodiment, the energy storage unit includes a battery pack and a supercapacitor, the battery pack and the supercapacitor being connected to the DC bus via the bidirectional DC / DC converter.

[0020] The advantages of adopting the above-mentioned optional method are as follows: further using battery packs and supercapacitors to form an energy storage unit, the two are connected to the DC bus through a bidirectional DC / DC converter, and the continuous energy characteristics of the battery pack and the instantaneous power characteristics of the supercapacitor complement each other, thereby optimizing the power response capability and energy storage performance of the energy storage unit and improving the flexibility and reliability of energy storage.

[0021] In one alternative approach, the integrated controller incorporates a predictive model, and the integrated controller is specifically used for: The turbine power generation prediction curve for the future time domain is output based on the real-time operating parameters and the prediction model, and the control command is generated based on the turbine power generation prediction curve and the real-time operating parameters.

[0022] The advantages of adopting the above-mentioned optional approach are as follows: by further embedding a prediction model in the integrated controller, the prediction curve of turbine power generation in the future time domain is output based on the real-time operating parameters, and control commands are generated based on the prediction curve and the real-time operating parameters, thereby realizing prediction-based feedforward control and improving the foresight and control accuracy of power allocation.

[0023] In one optional embodiment, the real-time operating parameters include reentry heat flux density, rocket body structure temperature, pressure and temperature of the cryogenic propellant, rotational speed of the electric pump unit, voltage and current of the DC bus, and remaining charge of the energy storage unit; the integrated controller is specifically used for: The predicted turbine power generation curve, the reentry heat flux density, the rocket body structure temperature, the pressure and temperature of the cryogenic propellant, the rotational speed of the electric pump unit, the voltage and current of the DC bus, and the remaining power of the energy storage unit are input into a multi-objective optimization model. The control command is generated by solving the multi-objective optimization model.

[0024] The advantages of adopting the above-mentioned optional method are as follows: the predicted curve of turbine power generation, reentry heat flux density, rocket body structure temperature, pressure and temperature of cryogenic propellant, rotational speed of electric pump unit, voltage and current of DC bus, and remaining power of energy storage unit are further input into multi-objective optimization model. By solving the model to generate control commands, optimization decision-making under multi-parameter coupling conditions is realized, and the scientificity and adaptability of control commands are improved.

[0025] In one alternative embodiment, the electric pump unit includes a motor and a centrifugal pump, the motor being connected to the power distribution unit and the centrifugal pump being connected to the motor, the motor being used to drive the centrifugal pump according to the stable DC power distributed by the power distribution unit, so that the centrifugal pump drives the cryogenic propellant to circulate in the active cooling loop.

[0026] The advantages of adopting the above-mentioned optional method are as follows: by further using an electric motor and a centrifugal pump to form an electric pump unit, the motor drives the centrifugal pump to operate according to the electrical energy allocated by the power distribution unit, thereby driving the cryogenic propellant to circulate in the active cooling circuit, realizing the conversion of electrical energy to mechanical energy and the cyclic drive of cryogenic propellant, eliminating the rigid mechanical coupling between the turbine and the cooling pump, and improving the independence and flexibility of flow regulation.

[0027] In one alternative embodiment, the auxiliary functional unit includes a tank booster, a navigation device, a communication device, and an attitude control system, wherein the tank booster, the navigation device, the communication device, and the attitude control system are respectively connected to the power distribution unit; The tank pressurizer is used to provide electrical energy to pressurize the tank of the second-stage rocket according to the stable DC power allocated by the power distribution unit. The navigation equipment, the communication equipment, and the attitude control system are used to provide operating power for the subsequent missions of the second-stage rocket according to the stable DC power allocated by the power distribution unit.

[0028] The advantages of adopting the above-mentioned optional approach are as follows: further setting up an auxiliary functional unit consisting of a tank pressurizer, navigation equipment, communication equipment and attitude control system, the tank pressurizer provides power for pressurizing the tank, and the navigation equipment, communication equipment and attitude control system provide operating power for subsequent missions, realizing the reuse of the captured energy in the reentry phase in diversified missions after the rocket lands, and improving the cross-stage utilization value of energy.

[0029] Secondly, this invention provides a comprehensive energy management method for a reusable second-stage rocket, employing the reusable second-stage rocket energy management system provided by this invention. The technical solution of this method is as follows: The heat capture and energy conversion module, located in the reentry heat protection zone of the second-stage rocket, utilizes the cryogenic propellant of the second-stage rocket to absorb aerodynamic heat and generate a high-temperature, high-pressure working fluid, which then converts the thermal energy of the high-temperature, high-pressure working fluid into electrical energy output. The power distribution and management module converts the electrical energy output by the heat capture and energy conversion module into stable DC power, stores the stable DC power, and distributes the stable DC power according to control commands. The power execution and application module drives the cryogenic propellant to circulate in the heat capture and energy conversion module according to the electrical energy allocated by the power distribution and management module, and provides electrical energy for the subsequent missions of the second-stage rocket.

[0030] The beneficial effects of the reusable second-stage rocket energy integrated management method of the present invention are as follows: The method of this invention sets up a heat capture and energy conversion module in the reentry thermal protection zone of the second-stage rocket. It utilizes cryogenic propellant to absorb aerodynamic heat to generate a high-temperature and high-pressure working fluid, which is then converted into electrical energy. In conjunction with a power distribution and management module, the electrical energy is converted, stored, and distributed. A power execution and application module drives the cryogenic propellant cycle and provides power for subsequent missions. This solves the problems of single energy form, strong control coupling, serious energy waste, and lack of energy reuse function in the prior art, and improves the energy utilization efficiency and mission adaptability of reusable second-stage rockets.

[0031] Thirdly, the technical solution of an electronic device according to the present invention is as follows: It includes a memory, a processor, and a program stored in the memory and running on the processor, wherein the processor executes the program to implement the steps of the reusable second-stage rocket energy integrated management method of the present invention.

[0032] Fourthly, the technical solution of a computer-readable storage medium provided by the present invention is as follows: The computer-readable storage medium stores instructions that, when read, cause the computer-readable storage medium to perform the steps of the reusable second-stage rocket energy integrated management method of the present invention.

[0033] The above description is merely an overview of the technical solution of the present invention. In order to better understand the technical means of the present invention and to implement it in accordance with the contents of the specification, and in order to make the above and other objects, features and advantages of the present invention more apparent and understandable, specific embodiments of the present invention are described below. Attached Figure Description

[0034] The accompanying drawings are for illustrative purposes only and are not intended to limit the invention. Furthermore, the same reference numerals denote the same parts throughout the drawings. In the drawings: Figure 1 This is a schematic diagram of an embodiment of a reusable second-stage rocket energy integrated management system according to the present invention; Figure 2 This is a schematic diagram of the system hardware architecture; Figure 3 A schematic diagram of the layered architecture for controlling hardware and software; Figure 4 This is a schematic diagram of the core control algorithm flow. Figure 5 This is a schematic diagram of the mode state transition; Figure 6 This is a flowchart illustrating an embodiment of the reusable second-stage rocket energy integrated management method of the present invention. Detailed Implementation

[0035] Exemplary embodiments of the invention will now be described in more detail with reference to the accompanying drawings. While exemplary embodiments of the invention are shown in the drawings, it should be understood that the invention can be implemented in various forms and should not be limited to the embodiments set forth herein.

[0036] Figure 1This diagram illustrates a structural schematic of an embodiment of a reusable second-stage rocket energy integrated management system 100 provided by the present invention. It should be noted that during operation, in the initial stage of reentry of the second-stage rocket, the energy storage unit powers an electric pump to pump cryogenic propellant into the active cooling circuit to absorb aerodynamic heat. The heated, high-temperature, high-pressure working propellant drives the turbine power generation unit to generate electricity, which is then rectified and fed into the DC bus. The integrated controller dynamically decides the allocation between power generation, energy storage charging / discharging power, and electric pump power based on parameters such as real-time heat flow, remaining energy in the energy storage unit, and electric pump demand. During peak heat flow phases, excess energy is stored in the energy storage unit; during low heat flow phases, the energy storage unit discharges to supplement the power gap; after reentry, the remaining energy is used for subsequent tasks such as tank pressurization. Figure 1 As shown, the reusable second-stage rocket energy integrated management system 100 includes: The heat capture and energy conversion module 101 is installed in the reentry heat protection zone of the second-stage rocket. It uses the cryogenic propellant of the second-stage rocket to absorb aerodynamic heat to generate a high-temperature and high-pressure working fluid, and converts the thermal energy of the high-temperature and high-pressure working fluid into electrical energy output.

[0037] In this context, a second-stage rocket refers to the second stage of a two- or multi-stage launch vehicle, responsible for delivering the payload into a predetermined orbit and re-entering the Earth's atmosphere after orbital deployment. For example, the second stage of a certain type of reusable rocket, after delivering a satellite into a 500km sun-synchronous orbit, separates from the satellite and re-enters the atmosphere at approximately Mach 8. Key components such as the nose cone, the windward surface of the fuselage, and the grid fins must withstand the harsh aerodynamic heating environment. The re-entry thermal protection zone refers to the areas of the second-stage rocket that require special protection during atmospheric re-entry due to the intense friction with the atmosphere. For instance, the re-entry thermal protection zone of the second stage of a certain type of reusable rocket includes the nose cone stagnation area, the windward surface of the fuselage, the surfaces of the four grid fins, and the outer wall of the propulsion module. These areas can reach surface temperatures exceeding 1200°C during re-entry.

[0038] Cryogenic propellants refer to rocket fuel or oxidizers stored in liquid form in rocket tanks with a boiling point below -160°C. For example, liquid methane carried by the second stage of a certain type of reusable rocket is used as a cryogenic propellant, stored in two independent tanks inside the rocket body, each with a volume of 40 m³ and a storage temperature of approximately -162°C. It serves as both a cooling medium in the active cooling loop and provides fuel for subsequent secondary ignition. High-temperature, high-pressure working fluids refer to cryogenic propellants that evaporate from a liquid state to a gaseous state after absorbing aerodynamic heat, forming a gaseous medium with significantly increased temperature and pressure. For example, liquid methane absorbs aerodynamic heat in the active cooling loop, raising its temperature from -162°C to 850°C and its pressure from 0.3 MPa to 12 MPa, forming high-temperature, high-pressure methane gas with the expansion capacity to drive the turbine.

[0039] The power distribution and management module 102 is used to convert the electrical energy output by the heat capture and energy conversion module 101 into stable DC power, store the stable DC power, and distribute it according to control commands.

[0040] Stable DC power refers to DC power whose voltage amplitude remains constant within a preset range and whose fluctuation range does not exceed a specified value. For example, the DC voltage output after rectification / voltage regulation is stable at 270V, with fluctuations controlled within ±5V, which can meet the power supply quality requirements of various electrical equipment on the rocket. Control commands refer to digital or analog signals generated by the integrated controller to guide the power distribution unit in distributing power and to each actuator in performing corresponding actions. For example, the integrated controller generates a set of control commands every second, which includes specific parameters such as "electric drive pump speed set to 12000r / min", "energy storage unit charging power set to 15kW", and "tank booster power supply voltage set to 270V".

[0041] The power execution and application module 103 is used to drive the cryogenic propellant to circulate in the heat capture and energy conversion module 101 according to the electrical energy allocated by the power distribution and management module 102, and to provide electrical energy for the subsequent mission of the second-stage rocket.

[0042] The subsequent tasks refer to the various operations that the second-stage rocket needs to perform after the reentry phase, before or after landing. For example, after the reentry phase of a certain type of reusable rocket is completed and before the parachute opens, the second-stage rocket body needs to perform a pressurization operation of the propellant tank before the second ignition, and at the same time provide standby power for the onboard navigation equipment, communication equipment and attitude control system. These operations are collectively referred to as the subsequent tasks.

[0043] The technical solution of this embodiment solves the problems of single energy form, strong control coupling, serious energy waste and lack of energy reuse function in the prior art by setting the heat capture and energy conversion module 101 in the reentry thermal protection zone of the second stage rocket. The cryogenic propellant absorbs aerodynamic heat to generate high temperature and high pressure working fluid and converts it into electrical energy. The power distribution and management module 102 converts, stores and distributes electrical energy, and the power execution and application module 103 drives the cryogenic propellant to circulate and provide power for subsequent missions. This improves the energy utilization efficiency and mission adaptability of the reusable second stage rocket.

[0044] In one alternative embodiment, the heat capture and energy conversion module 101 includes an active cooling circuit and a gas turbine power generation unit.

[0045] The active cooling circuit is located in the reentry thermal protection zone of the second-stage rocket and is used to absorb the aerodynamic heat using the cryogenic propellant carried by the second-stage rocket and generate the high-temperature and high-pressure working fluid.

[0046] The active cooling loop refers to a closed-loop circuit consisting of pipes, heat exchangers, and related connectors, used to introduce cryogenic propellant into the reentry thermal protection zone to absorb aerodynamic heat. For example, the active cooling loop uses titanium alloy pipes installed on the head and grid fin surfaces of the second stage of a certain type of reusable rocket. The pipes have an inner diameter of 12 mm and a wall thickness of 1.5 mm. Liquid methane flows in the pipes and absorbs heat. The inlet pressure of the loop is 0.5 MPa and the outlet pressure is 0.3 MPa.

[0047] The gas turbine power generation unit is connected to the active cooling circuit and is used to receive the high-temperature and high-pressure working fluid and convert the thermal energy of the high-temperature and high-pressure working fluid into electrical energy output.

[0048] Among them, the gas turbine power generation unit refers to a power generation device consisting of a turbine and a generator coaxially connected by a coupling, used to convert the thermal energy of a high-temperature and high-pressure working fluid into electrical energy output; for example, in the gas turbine power generation unit, the turbine inlet receives methane gas at 850℃ and 12MPa, the turbine speed reaches 50000r / min, driving the coaxially connected permanent magnet synchronous generator to rotate, and the generator outputs alternating current with a frequency variation range of 0 to 2000Hz.

[0049] In the above-mentioned optional approach, the heat capture and energy conversion module 101 is further divided into an active cooling circuit and a gas turbine power generation unit, so that the cryogenic propellant absorbs aerodynamic heat in the re-entry heat protection zone to form a high-temperature and high-pressure working fluid. The heat energy is converted into electrical energy output through the gas turbine power generation unit, realizing aerodynamic heat capture and energy form conversion, solving the problem of the single energy form in the prior art, and expanding the energy utilization methods.

[0050] In one alternative embodiment, the power distribution and management module 102 includes a rectification / voltage regulation unit, an energy storage unit, an integrated controller, and a power distribution unit.

[0051] The rectifier / voltage regulator unit is connected to the gas turbine generator unit and is used to convert the electrical energy output by the gas turbine generator unit into the stable DC power and feed it into the DC bus.

[0052] The rectifier / voltage regulator unit refers to a power electronic device composed of a rectifier circuit and a voltage regulator circuit, used to convert the variable-frequency AC power output from the generator into DC power with a constant voltage. For example, the rectifier / voltage regulator unit converts the AC power output from the gas turbine generator unit, which varies in frequency from 0 to 2000Hz, into DC power through a three-phase bridge rectifier circuit, and then stabilizes the voltage at 270V through a step-down voltage regulator circuit before feeding it into the DC bus. The DC bus refers to the common power transmission node connecting the output terminal of the rectifier / voltage regulator unit, the interface of the energy storage unit, and the input terminal of the power distribution unit. For example, the DC bus adopts a copper busbar structure with a rated voltage of 270V and a rated current of 200A. The electrical energy output from the rectifier / voltage regulator unit, the electrical energy charged and discharged by the energy storage unit, and the electrical energy taken by the power distribution unit all converge on the DC bus.

[0053] The energy storage unit is connected to the DC bus via a bidirectional DC / DC converter for storing the stable DC power.

[0054] The energy storage unit refers to an energy storage device composed of energy storage elements and their management circuits, used to absorb, store, or release energy from the DC bus. For example, an energy storage unit includes a set of lithium titanate batteries and a set of supercapacitors, with a total energy storage capacity of 10kWh. During the peak reentry heat flow phase, it absorbs excess energy from the DC bus at a power of 20kW, and during the power deficit phase, it releases energy back to the DC bus at a power of 15kW. The battery pack and supercapacitors are connected in parallel to the bidirectional DC-DC converter. The integrated controller dynamically allocates charging and discharging power according to the real-time power change rate: when the power change rate exceeds a preset threshold, the supercapacitor prioritizes responding to the instantaneous power demand; when the power change rate is lower than the preset threshold, the battery pack undertakes continuous power output. The two work together to optimize the power response speed and energy storage efficiency of the energy storage system.

[0055] Among them, a bidirectional DC / DC converter refers to a DC-DC conversion device that can realize bidirectional power transfer between the DC bus and the energy storage unit. For example, the bidirectional DC converter adopts a bidirectional buck-boost topology and has a rated power of 30kW. When the DC bus voltage is higher than the energy storage unit voltage, the converter operates in buck mode to charge the energy storage unit. When the DC bus voltage is lower than the energy storage unit voltage, the converter operates in boost mode to release the energy from the energy storage unit to the DC bus.

[0056] The integrated controller is connected to the sensor layer, which is deployed at the corresponding measurement points of the second-stage rocket. The integrated controller is used to acquire real-time operating parameters and generate control commands based on the real-time operating parameters.

[0057] The integrated controller refers to a high-performance embedded computer with a built-in real-time operating system and energy management algorithms, used to collect system operating parameters and generate control commands. For example, the integrated controller adopts a heterogeneous multi-core processor architecture with a main frequency of no less than 1.5 GHz, a built-in real-time operating system, executes a control cycle every second, collects no less than 25 real-time operating parameters sent by the sensor layer, and generates 12 control commands after running a model-based predictive dynamic energy management algorithm. The sensor layer refers to a sensing network composed of sensors and battery management units deployed in various key parts of the second-stage rocket, used to collect various physical parameters during system operation. For example, the sensor layer includes 8 K-type thermocouples, 5 piezoresistive pressure sensors, 3 electromagnetic flowmeters, 2 Hall current sensors, 2 voltage sensors, and a battery management unit. All sensors are connected to the integrated controller through the controller's local area network bus.

[0058] The corresponding measurement point location refers to the specific physical location of the sensor on the second stage rocket, which can accurately reflect the changes in the monitored parameters. For example, the reentry heat flux density sensor is located in the stagnation area of ​​the second stage rocket head, the rocket body structure temperature sensor is located on the inner surface of the windward skin of the rocket body, the cryogenic propellant pressure sensor is located at the inlet and outlet of the active cooling circuit, the electric drive pump unit speed sensor is located at the motor shaft end, and the DC bus voltage sensor is located at the input end of the power distribution unit.

[0059] Among them, real-time operating parameters refer to the physical quantity data collected by the sensor layer and transmitted to the integrated controller at the current moment during system operation. For example, the real-time operating parameters collected by the integrated controller at the beginning of each control cycle include: head heat flux density 380kW / m², rocket body surface temperature 650℃, active cooling circuit inlet pressure 0.5MPa, active cooling circuit outlet temperature 320℃, electric drive pump speed 11800r / min, DC bus voltage 271V, DC bus current 85A, and energy storage unit remaining power 62%.

[0060] The power distribution unit is connected to the DC bus and the integrated controller, and is used to distribute the stable DC power on the DC bus according to the control command.

[0061] The power distribution unit refers to an energy distribution device composed of multiple solid-state power controllers, used to distribute the power on the DC bus according to the control instructions of the integrated controller. For example, the power distribution unit contains 8 solid-state power controllers, each with a rated current of 50A. The integrated controller sends instructions to the power distribution unit through the controller local area network bus, and the power distribution unit distributes the 270V DC bus power to the electric drive pump, tank booster, navigation equipment, communication equipment and attitude control system respectively.

[0062] In the above-mentioned optional methods, the electrical energy output by the gas turbine generator unit is further converted into stable DC power and fed into the DC bus through a rectifier / voltage regulation unit. The energy storage unit is connected to the DC bus through a bidirectional DC / DC converter to realize energy storage. The integrated controller generates control commands based on real-time operating parameters, and the power distribution unit distributes electrical energy according to the control commands, thereby realizing integrated management of energy conversion, storage and intelligent distribution.

[0063] In one alternative embodiment, the power execution and application module 103 includes an electric pump unit and an auxiliary function unit.

[0064] The electric pump unit is connected to the power distribution unit and is used to drive the cryogenic propellant to circulate in the active cooling loop according to the stable DC power distributed by the power distribution unit.

[0065] The electric pump unit refers to an electromechanical integrated device consisting of an electric motor and a centrifugal pump, used to drive the cryogenic propellant to circulate in the active cooling circuit according to the allocated electrical energy. For example, the electric pump unit uses a brushless DC motor to drive the centrifugal pump. The motor has a rated power of 25kW and a rated speed of 15000r / min. The centrifugal pump has a rated flow rate of 120L / min. The motor drives the centrifugal pump to rotate according to the 20kW DC power allocated by the power distribution unit. The centrifugal pump draws liquid methane from the storage tank and sends it into the active cooling circuit.

[0066] The auxiliary function unit is connected to the power distribution unit and is used to provide power for the subsequent missions of the second-stage rocket according to the stable DC power distributed by the power distribution unit.

[0067] Auxiliary functional units refer to the collection of onboard equipment, excluding the electric pump unit, that needs to draw power from the DC bus to complete subsequent tasks. For example, auxiliary functional units include an electrically heated vaporizer for pressurizing the tank before secondary ignition, a combined navigation device of the Global Navigation Satellite System and the Inertial Navigation System for positioning and navigation, an S-band communication device for data transmission, and an attitude control nozzle solenoid valve for attitude adjustment.

[0068] In the above-mentioned optional approach, an electric pump unit is further configured to drive the cryogenic propellant to circulate in the active cooling loop according to the electrical energy allocated by the power distribution unit, and the auxiliary functional unit provides electrical energy for the subsequent mission of the second stage rocket. This realizes the dual functions of cooling cycle drive and mission power supply, solves the problem of no energy reuse function in the existing technology, and improves the level of comprehensive energy utilization.

[0069] In one alternative embodiment, the energy storage unit includes a battery pack and a supercapacitor, the battery pack and the supercapacitor being connected to the DC bus via the bidirectional DC / DC converter.

[0070] A battery pack refers to a secondary battery assembly composed of multiple cells connected in series or parallel, used to provide relatively continuous and stable energy storage and release. For example, a battery pack may consist of 24 lithium titanate cells connected in series, with a nominal voltage of 270V, a rated capacity of 30Ah, and a maximum continuous discharge current of 100A. It is mainly used to provide continuous power to the DC bus during periods of low heat flow. A supercapacitor refers to an energy storage device that utilizes the double-layer principle to store charge, used to provide instantaneous high-power energy absorption and release. For example, a supercapacitor module may consist of 60 cells connected in series, with a nominal voltage of 270V, a capacitance of 20F, and a maximum instantaneous power of 50kW. It is mainly used to rapidly absorb excess energy from the DC bus and suppress bus voltage fluctuations during peak reentry heat flow periods.

[0071] In the above-mentioned optional methods, a battery pack and a supercapacitor are further used to form an energy storage unit. The two are connected to the DC bus through a bidirectional DC / DC converter. The continuous energy characteristics of the battery pack and the instantaneous power characteristics of the supercapacitor complement each other, which optimizes the power response capability and energy storage performance of the energy storage unit and improves the flexibility and reliability of energy storage.

[0072] In one alternative approach, the integrated controller incorporates a predictive model, and the integrated controller is specifically used for: The turbine power generation prediction curve for the future time domain is output based on the real-time operating parameters and the prediction model, and the control command is generated based on the turbine power generation prediction curve and the real-time operating parameters.

[0073] The prediction model refers to a mathematical model built into the integrated controller, used to predict the future trend of turbine power generation in the time domain based on current real-time operating parameters. For example, the prediction model is constructed using the lumped parameter method, integrating the rocket reentry flight mechanics equations, aerodynamic heat conduction equations, and turbine power generation characteristic curves, with the current heat flux density input as 380 kW / m³. 2 With a flight Mach number of 7.2 and a coolant flow rate of 110 L / min, the predicted turbine power output sequence for the next 5 seconds is 38 kW, 37.5 kW, 36.8 kW, 35.9 kW, and 34.7 kW. The prediction model improves accuracy through a combination of offline training and online correction. In the offline phase, historical flight data is used to identify model parameters. In the online phase, recursive least squares is used to dynamically correct model parameters based on the deviation between real-time measurements and predicted values, keeping the prediction error within 5%.

[0074] The future time domain refers to a preset time interval extending into the future from the current moment, used to predict the output range of the model. For example, if the integrated controller sets the future time domain length to 5 seconds, the prediction model outputs a discrete sequence of turbine power generation values ​​within each control cycle, from the current moment to the end of the 5-second interval. The turbine power generation prediction curve refers to the functional relationship output by the prediction model that describes the trend of turbine power generation over time in the future time domain. For example, the turbine power generation prediction curve outputs 50 discrete points at 0.1-second intervals, forming a smooth curve that gradually decreases from the current power of 39kW to 35kW after 5 seconds. Based on the curve, the integrated controller determines that the system will have a 4kW power surplus within the next 5 seconds.

[0075] In the above-mentioned optional approach, a prediction model is further built into the integrated controller. Based on the real-time operating parameters, the prediction curve of the turbine power generation in the future time domain is output, and control commands are generated based on the prediction curve and the real-time operating parameters. This realizes prediction-based feedforward control and improves the foresight and control accuracy of power allocation.

[0076] In one optional embodiment, the real-time operating parameters include reentry heat flux density, rocket body structure temperature, pressure and temperature of the cryogenic propellant, rotational speed of the electric pump unit, voltage and current of the DC bus, and remaining charge of the energy storage unit; the integrated controller is specifically used for: The predicted turbine power generation curve, the reentry heat flux density, the rocket body structure temperature, the pressure and temperature of the cryogenic propellant, the rotational speed of the electric pump unit, the voltage and current of the DC bus, and the remaining power of the energy storage unit are input into a multi-objective optimization model. The control command is generated by solving the multi-objective optimization model.

[0077] Reentry heat flux density refers to the aerodynamic heating heat received per unit area of ​​the rocket body surface per unit time; for example, a reentry heat flux density sensor measures the current heat flux density at the stagnation point of the second stage head of a certain type of reusable rocket to be 380 kW / m³. 2 The numerical values ​​serve as an important basis for the integrated controller to adjust the speed of the electric drive pump and the charging and discharging power of the energy storage unit. The rocket body structure temperature refers to the measured temperature value of the structural materials of the second-stage rocket body during reentry; for example, if a K-type thermocouple installed on the inner surface of the windward skin of the rocket body measures the current rocket body structure temperature to be 650℃, the integrated controller will compare the temperature value with the preset safety threshold of 800℃ to determine whether the current thermal protection status is safe.

[0078] The pressure and temperature of the cryogenic propellant refer to the pressure and temperature values ​​of the cryogenic propellant at the inlet and outlet of the active cooling loop. For example, at the inlet of the active cooling loop, the pressure sensor measures a liquid methane pressure of 0.5 MPa and a temperature sensor measures a liquid methane temperature of -162°C. At the outlet of the active cooling loop, the pressure sensor measures a methane gas pressure of 0.3 MPa and a temperature sensor measures a methane gas temperature of 320°C. The integrated controller uses the difference between the inlet and outlet pressures to determine whether the cooling flow rate is sufficient. The rotational speed of the electric pump unit refers to the actual rotational speed of the motor rotor in the electric pump unit. For example, if the Hall sensor built into the electric pump unit measures the current motor speed as 11800 r / min, the integrated controller compares the rotational speed with the preset target speed of 12000 r / min and adjusts the output power of the power distribution unit to bring the rotational speed closer to the target value.

[0079] The DC bus voltage and current refer to the real-time voltage amplitude and real-time current flowing through the DC bus. For example, if the DC bus voltage sensor measures a current voltage of 271V and the current sensor measures a current of 85A, the integrated controller determines whether the bus power supply is stable based on the voltage value and calculates the total power consumption of the electric drive pump unit and auxiliary function units based on the current value. Remaining capacity refers to the percentage of electrical energy currently stored in the energy storage unit relative to its total capacity. For example, if the battery management unit calculates that the remaining capacity of the energy storage unit is 62% using the coulomb integral method, the integrated controller ensures that the charging and discharging operations do not cause the remaining capacity to fall below the safety lower limit of 20% or exceed the safety upper limit of 80%.

[0080] The multi-objective optimization model refers to a mathematical optimization model that comprehensively considers power balance constraints, battery safety constraints, and pump power limiting constraints, with the goal of minimizing a preset cost function. For example, the multi-objective optimization model aims to minimize the square integral of the difference between turbine power generation and average power generation. Constraints include that the sum of power generation and battery charging / discharging power equals the pump's required power, remaining battery capacity is between 20% and 80%, and pump power is between 15kW and 30kW. The optimal battery charging / discharging power sequence for the next 5 seconds is obtained through dynamic programming. The multi-objective optimization model introduces an adaptive weighting factor. The integrated controller dynamically adjusts the weights between peak shaving and valley filling objectives and energy recovery objectives based on the remaining battery capacity of the energy storage unit: when the remaining battery capacity is below 40%, the weight of the energy recovery objective is increased; when the remaining battery capacity is above 60%, the weight of the peak shaving and valley filling objective is increased to balance the system's performance requirements under different operating conditions.

[0081] In the above-mentioned optional methods, the predicted turbine power generation curve, reentry heat flux density, rocket body structure temperature, cryogenic propellant pressure and temperature, electric pump unit speed, DC bus voltage and current, and remaining energy of the energy storage unit are further input into a multi-objective optimization model. By solving the model, control commands are generated, realizing optimal decision-making under multi-parameter coupling conditions and improving the scientific nature and adaptability of the control commands.

[0082] In one alternative embodiment, the electric pump unit includes a motor and a centrifugal pump, the motor being connected to the power distribution unit and the centrifugal pump being connected to the motor, the motor being used to drive the centrifugal pump according to the stable DC power distributed by the power distribution unit, so that the centrifugal pump drives the cryogenic propellant to circulate in the active cooling loop.

[0083] The term "motor" refers to an electromagnetic device that converts electrical energy into mechanical energy to drive a centrifugal pump. For example, the brushless DC motor in an electric pump unit has a rated power of 25kW and a rated speed of 15000r / min. The motor stator windings receive 270V DC power through a power distribution unit, and the rotor drives the centrifugal pump shaft to rotate. A centrifugal pump is a fluid machine that uses the centrifugal force generated by the rotation of an impeller to draw cryogenic propellant from a storage tank, pressurize it, and then send it into an active cooling circuit. For example, a centrifugal pump with an impeller diameter of 120mm, a rated flow rate of 120L / min at a rated speed of 15000r / min, and a head of 200m, has its impeller driven to rotate by a motor shaft. Liquid methane is drawn in from the center of the impeller, accelerated by the blades, and discharged from the outlet.

[0084] In the above-mentioned optional methods, an electric pump unit is further constructed by using a motor and a centrifugal pump. The motor drives the centrifugal pump to operate according to the electrical energy allocated by the power distribution unit, thereby driving the cryogenic propellant to circulate in the active cooling loop. This realizes the conversion of electrical energy into mechanical energy and the cyclic drive of cryogenic propellant, decouples the rigid mechanical coupling between the turbine and the cooling pump, and improves the independence and flexibility of flow regulation.

[0085] In one alternative embodiment, the auxiliary functional unit includes a tank booster, a navigation device, a communication device, and an attitude control system, wherein the tank booster, the navigation device, the communication device, and the attitude control system are respectively connected to the power distribution unit.

[0086] The tank pressurizer is used to provide electrical energy to pressurize the tank of the second-stage rocket according to the stable DC power allocated by the power distribution unit. The navigation equipment, the communication equipment, and the attitude control system are used to provide operating power for the subsequent missions of the second-stage rocket according to the stable DC power allocated by the power distribution unit.

[0087] The tank pressurizer refers to a device used to pressurize the cryogenic propellant stored in the tank and deliver it to the engine or related equipment. For example, the tank pressurizer uses an electrically heated vaporizer, with the power distribution unit supplying 5kW of DC power to the vaporizer. The electric heating element vaporizes a small amount of liquid methane and fills the tank, increasing the tank pressure from 0.2MPa to 0.5MPa. The navigation equipment refers to electronic equipment used to provide navigation information such as the second-stage rocket's position, velocity, and attitude. For example, the navigation equipment uses a combined navigation system of Global Positioning System (GPS) and Inertial Navigation System (INS). The power distribution unit supplies 100W of DC power to the navigation equipment, and the system outputs the second-stage rocket's longitude, latitude, altitude, velocity vector, and attitude angle data in real time.

[0088] The communication equipment refers to the radio transceiver used for data exchange between the second-stage rocket and the ground control station. For example, the communication equipment uses an S-band transceiver, and the power distribution unit supplies 80W of DC power to the equipment. The equipment transmits real-time telemetry data of the second-stage rocket to the ground control station and simultaneously receives remote control commands from the ground. The attitude control system refers to the control system used to control the attitude and pointing of the second-stage rocket during reentry, ensuring that the rocket body flies in a predetermined attitude. For example, the attitude control system includes an attitude control computer and an attitude control nozzle solenoid valve. The power distribution unit supplies 200W of DC power to the attitude control system, and the attitude control computer calculates the control law based on the attitude angle data provided by the navigation equipment, driving the attitude control nozzle solenoid valve to open or close to adjust the rocket body attitude.

[0089] In the above-mentioned optional methods, an auxiliary functional unit is further set up, consisting of a tank pressurizer, navigation equipment, communication equipment and attitude control system. The tank pressurizer provides power for pressurizing the tank, and the navigation equipment, communication equipment and attitude control system provide operating power for subsequent missions. This realizes the reuse of the energy captured during the reentry phase in diversified missions after the rocket lands, and enhances the cross-phase utilization value of energy.

[0090] The core design concept of the technical solution in this embodiment is to break the rigid coupling mode of turbine mechanical energy directly driving the cooling pump, convert the mechanical energy output by the turbine in the reentry active cooling circuit into electrical energy, and introduce a high power density energy storage unit and integrated controller to build an integrated energy network that is dynamically adjustable, storable, and reusable. Through intelligent energy management algorithms, it achieves efficient energy capture, precise allocation, peak shaving and valley filling, and cross-stage value-added applications, which not only ensures the thermal protection requirements of the rocket body during the reentry stage, but also improves energy utilization efficiency and enhances the stability and reliability of system operation.

[0091] This embodiment adopts a modular design, mainly composed of three core modules: a heat capture and energy conversion module 101, a power distribution and management module 102, and a power execution and application module 103. These modules are organically connected through pipes, circuits, and signal lines to form a complete closed loop of energy capture, conversion, storage, distribution, and application. The specific connection relationships are as follows: Figure 2 As shown.

[0092] like Figure 2 As shown, the heat capture and energy conversion module 101 includes an active cooling loop and a gas turbine power generation unit. The active cooling loop, as the core component of aerodynamic heat capture, is located on the main windward surface during the second-stage rocket's reentry process, encompassing critical thermal protection areas such as the rocket's nose, body surface, grid fins, tail fins, and propulsion compartment. The active cooling loop employs a high-temperature resistant, corrosion-resistant, and lightweight piping design. The internal cooling medium uses the rocket's own cryogenic propellant, specifically liquid methane or liquid hydrogen, eliminating the need for additional cooling media. It also preheats the propellant, improving subsequent propulsion efficiency. The gas turbine power generation unit consists of a turbine and a generator coaxially connected via a rigid coupling, forming an integrated power generation structure. The turbine inlet is connected to the active cooling loop outlet via a high-temperature resistant pipe, receiving the high-temperature, high-pressure gaseous working fluid formed after heat exchange. The expansion force of the working fluid drives the turbine to rotate at high speed. The turbine outlet is connected to a dedicated exhaust device to safely discharge the cryogenic, low-pressure gas after work, preventing secondary thermal impact on the rocket structure. The generator output is connected to the rectifier / voltage regulator unit in the power distribution and management module 102 via a cable, which converts the mechanical energy transmitted by the turbine into electrical energy output. At this time, the output electrical energy is unstable electrical energy.

[0093] The power distribution and management module 102 includes a rectification / voltage regulation unit, an energy storage unit, an integrated controller, and a power distribution unit. The rectification / voltage regulation unit is connected to the gas turbine generator unit and processes the frequency-converted AC power output from the generator. It converts the AC power to DC power through a rectifier circuit, and then stabilizes the DC voltage within a preset range, which is adapted to the power supply requirements of the onboard equipment. The stabilized DC power processed by the rectification / voltage regulation unit is fed into the DC bus, providing a stable power source for subsequent power distribution. The energy storage unit adopts a hybrid energy storage scheme combining a high-power-density battery pack and a supercapacitor. It is connected to the DC bus through a bidirectional DC-DC converter, enabling bidirectional charging and discharging management of electrical energy. The supercapacitor is responsible for rapidly absorbing excess energy during peak reentry periods to cope with power surges; the battery pack is responsible for long-term energy storage, providing continuous power during low-heat periods and subsequent tasks. Together, they enhance the response speed and capacity of the energy storage system. The integrated controller, as the core control unit of the system, uses a high-performance embedded computer with a built-in real-time operating system and core energy management algorithms. The integrated controller communicates with all sensors within the system via signal lines to collect various operating parameters in real time. Connected to the power distribution unit and each actuator driver via control lines, the integrated controller outputs precise control commands, enabling comprehensive control and dynamic optimization of the entire system. The power distribution unit consists of multiple high-performance solid-state power controllers, employing a modular design for flexible expansion based on actual needs. Its core function is to precisely control the power distribution on the DC bus according to the integrated controller's commands, allocating power as needed to the electric pump unit, auxiliary function units, or controlling the charging and discharging operations of the energy storage unit, ensuring a stable and suitable power supply for all equipment. The sensor layer, deployed at corresponding measurement points on the second-stage rocket, includes heat flux sensors, temperature sensors, pressure sensors, voltage sensors, current sensors, and a battery management unit. This sensor layer collects real-time operating parameters such as reentry heat flux density, rocket body structure temperature, cryogenic propellant pressure and temperature, DC bus voltage and current, and remaining power in the energy storage unit, transmitting these parameters to the integrated controller.

[0094] The power execution and application module 103 includes an electric pump unit and auxiliary function units. The electric pump unit consists of a high-performance motor and a centrifugal pump. The motor is connected to the power distribution unit via a cable and is stably powered by the DC bus. The speed of the electric pump unit can be precisely adjusted via commands from the integrated controller, thereby independently controlling the flow rate of cryogenic propellant in the active cooling circuit, accurately adapting to thermal protection requirements, and completely eliminating the coupling limitation with turbine speed. The auxiliary function units cover various electrical equipment required for subsequent missions after rocket reentry, mainly including electrically heated vaporizers for pressurizing the propellant tanks before secondary ignition, onboard navigation equipment, communication equipment, and attitude control systems. These devices draw power from the DC bus through the power distribution unit, achieving energy reuse across stages and improving the success rate and reliability of rocket missions.

[0095] To achieve collaborative operation, precise control, and dynamic optimization among system modules, this system adopts a layered control hardware and software architecture, dividing the control function into four layers: sensor layer, control layer, execution layer, and physical layer. Each layer interacts with data and transmits commands through standardized signal interfaces, ensuring clear and scalable control logic. The specific architecture is as follows: Figure 3 As shown.

[0096] like Figure 3 As shown, the sensor layer, located at the bottom of the architecture, serves as the system's sensing organ. It comprises various high-precision sensors and a battery management unit, collecting key parameters during the rocket's reentry process in real time and feeding the status back to the control layer. These key parameters include reentry heat flux density, rocket body temperature, coolant pressure and temperature, electric pump speed, DC bus voltage and current, and remaining power in the energy storage unit. The data collected by the sensor layer is preprocessed before being transmitted to the control layer, providing precise data support for control decisions.

[0097] The control layer, situated above the sensor layer, is the core control layer of the system, consisting of an integrated controller and a power distribution unit. The integrated controller runs a model-predictive dynamic energy management algorithm, calculating the optimal power distribution scheme based on real-time data transmitted from the sensor layer and its built-in physical model, and outputting drive and control commands. The power distribution unit, acting as the execution carrier of these control commands, performs precise energy distribution and charging / discharging control according to the integrated controller's instructions, allocating power to the execution layer.

[0098] The execution layer, located above the control layer, consists of dedicated drivers for various actuators, including electric pump drivers, generator excitation / rectifier controllers, energy storage unit charge / discharge controllers, and auxiliary load power distribution controllers. The core function of the execution layer is to receive control commands from the control layer, convert these commands into specific actions, and precisely control the electric pump speed, generator excitation current, battery charge / discharge power, etc., ensuring that each device operates stably according to preset logic.

[0099] The physical layer, located at the top of the architecture, includes physical devices such as electric pumps, generators, battery packs, and tank boosters. Driven by the execution layer, the physical layer performs specific physical actions and feeds back the status to the control layer via the sensor layer, forming a complete closed-loop control system.

[0100] The model-based predictive dynamic energy balance strategy is a core innovation that distinguishes it from existing technologies. Unlike existing passive response control, this strategy actively predicts heat flow changes and energy demands during reentry, optimizes energy allocation in advance, achieves dynamic energy balance and efficient utilization, and ensures stable and reliable system operation under various conditions. The algorithm flow is as follows: Figure 4 As shown.

[0101] like Figure 4 As shown, at the start of each new control cycle, the integrated controller first collects real-time data, including reentry heat flux density, rocket body structural temperature, remaining energy in the energy storage unit, and electric pump power. The integrated controller uses a built-in physical model as a prediction model, inputting real-time parameters such as current heat flux density, flight speed, and coolant flow rate to predict the turbine power generation curve within the next Δt, where Δt is the future time domain length. The prediction model is a simplified lumped-parameter physical model validated through extensive simulations and experiments, integrating theoretical and experimental data from multiple fields such as rocket reentry flight mechanics, aerodynamic heat transfer, and turbine power generation characteristics, enabling it to accurately output the predicted turbine power generation curve within the short future time domain. This provides a lead time for subsequent optimization decisions, among which It represents time, with the shortest future time domain being, for example, 5 seconds.

[0102] The integrated controller acquires the current system status, including the power demand of the electric drive pump. And the remaining battery power constraint, among which This includes the electric pump's own losses and the power required to drive the flow of the cooling medium. The integrated controller also obtains the remaining charge level of the energy storage unit through the battery management unit, clearly defining the battery's charge and discharge capacity constraints and providing boundary conditions for optimization solutions.

[0103] The integrated controller incorporates parameters such as predicted turbine power generation, current electric pump power demand, and remaining battery capacity constraints into a multi-objective optimization model to solve a multi-objective optimization problem. The decision variable is the battery charging and discharging power, and the objective is peak shaving and valley filling or maximizing battery recovery. The optimization objective is to minimize a preset cost function. This function can be flexibly adjusted according to task requirements. For example, the cost function can be set to... ,in The objective function, assuming average power generation, aims to stabilize power generation, achieving peak shaving and valley filling. The cost function can also be set to maximize energy recovery. The aim is to maximize the recovery of excess energy during reentry and improve energy utilization efficiency. Power to charge the energy storage unit.

[0104] To ensure the safe and stable operation of the system, the optimization solution must satisfy the following constraints. The power balance constraint is... ,in The charging and discharging power of the energy storage unit is defined, with positive values ​​indicating discharging and negative values ​​indicating charging. This constraint ensures a balance between energy supply and demand in the system. Battery safety constraints are... ,generally Set to 20%, Setting it to 80% avoids overcharging and over-discharging the battery, extending its lifespan. This represents the remaining electrical charge of the energy storage unit over time. The pump power limiting constraint is... ,in To ensure the minimum pump power required for basic thermal protection, This constraint, representing the maximum rated power of the electric pump, ensures that the pump operates within a safe range.

[0105] If the optimal solution is obtained, the battery charge / discharge command of the first control step is executed, that is, the rolling optimization strategy is adopted, and only the optimal battery charge / discharge power sequence is executed. The first control step instruction in the process involves re-collecting real-time data, predicting power generation, and solving for the optimal solution upon the arrival of the next control cycle to ensure the dynamic adaptability of the control strategy. If the optimal solution is not obtained, a preset safety strategy is adopted, such as keeping the battery inactive or prioritizing cooling, while waiting for the next control cycle.

[0106] During the second-stage reentry of the rocket, the operating conditions, including aerodynamic heat flux, flight attitude, and energy demand, change drastically. To ensure the controller can reliably handle various complex operating conditions and enable the system to achieve optimal operation at different stages, a complete operating mode and state transition logic covering the entire flight profile from second-stage separation to the end of reentry were designed. The specific state transition logic is as follows: Figure 5 As shown, each mode is described in detail in Table 1.

[0107] like Figure 5 As shown, the system enters standby self-test mode (mode 0) from the initial state. In mode 0, the system completes the self-test and waits for the re-entry trigger condition. When the heat flux exceeds the threshold 1 (100kW / m²), the system will enter standby self-test mode. 2 When the heat flux exceeds the threshold value, the system switches to re-entry startup mode, i.e., mode 1. To avoid frequent mode switching caused by heat flux fluctuations around the threshold, hysteresis control logic is introduced for mode switching: when the heat flux is greater than the sum of threshold 1 and the hysteresis, the system switches from mode 0 to mode 1; when the heat flux is less than the difference between threshold 1 and the hysteresis, the system returns from mode 1 to mode 0, with the hysteresis set to 10 kW / m³. 2In Mode 1, the electric pump starts and establishes minimum cooling flow. When the generator power exceeds the electric pump power consumption, the system switches to energy self-sufficiency mode (Mode 2), achieving energy self-sufficiency. During Mode 2 operation, if a heat flux peak is encountered, the system switches to peak shaving mode (Mode 3), where the energy storage unit absorbs excess energy. After the heat flux subsides, the system returns from Mode 3 to Mode 2. During Mode 2 operation, if a decrease in heat flux or disturbance causes a drop in power generation, the system switches to energy compensation mode (Mode 4), where the energy storage unit discharges to compensate for the power shortfall. After power generation recovers, the system returns from Mode 4 to Mode 2. After reentry, the system switches from Mode 2 to energy distribution mode (Mode 5), distributing stored energy to auxiliary functional units. In any of Modes 0 to 5, if a system failure occurs, the system immediately switches to fault degradation mode under fault handling mode (Mode 6), where the battery provides power alone to ensure minimum cooling flow. After reentry or crash, the system terminates operation.

[0108] Table 1. State Transition Table for Full Flight Profile Mode By converting the mechanical energy output by the turbine into electrical energy, the rigid mechanical coupling between the turbine and the cooling pump in existing technologies is completely broken. This allows cooling flow regulation to be independent of turbine speed, enabling precise and rapid adjustment based on real-time heat flow changes. This significantly improves the response speed and control accuracy of the thermal protection system, effectively addressing sudden heat flow shocks. The introduction of an energy storage unit can efficiently absorb excess electrical energy during the peak reentry heat flow phase, avoiding energy waste. Simultaneously, during the heat flow trough phase or power disturbances, the stored electrical energy can be released to supplement power gaps, smoothing DC bus power fluctuations, improving system operational stability, and enhancing energy utilization efficiency. Based on model prediction-driven dynamic energy management algorithms, while meeting the primary objective of rocket thermal protection, the charging and discharging strategies of the energy storage unit can be dynamically optimized to achieve multiple objectives such as maximizing energy recovery, extending battery life, and stabilizing bus voltage. Compared to existing passive response control, this significantly improves the system's intelligence level and overall performance. The aerodynamic thermal energy captured during the reentry phase, which would otherwise be wasted, can be stored as electrical energy and applied to subsequent missions after reentry, enabling cross-stage energy reuse in the propulsion system. This reduces the rocket's own power consumption and improves mission success rate. The well-designed fault degradation mode allows the energy storage unit to act as an emergency power source when the system experiences problems such as sensor anomalies or generator failures. This ensures the continuous operation of the core cooling function, significantly improving the system's robustness and the flight survival probability of the reusable rocket, and reducing the risk of rocket recovery failure.

[0109] Figure 6This diagram illustrates a flowchart of an embodiment of a reusable second-stage rocket energy integrated management method provided by the present invention. This method employs a reusable second-stage rocket energy integrated management system 100 as provided by the present invention. Figure 6 As shown, the method includes the following steps: S1. The heat capture and energy conversion module 101, which is located in the reentry heat protection zone of the second-stage rocket, uses the cryogenic propellant of the second-stage rocket to absorb aerodynamic heat to generate a high-temperature and high-pressure working fluid, and converts the thermal energy of the high-temperature and high-pressure working fluid into electrical energy output. S2. The power distribution and management module 102 converts the electrical energy output by the heat capture and energy conversion module 101 into stable DC power, stores the stable DC power, and distributes the stable DC power according to control commands. S3. The power execution and application module 103 drives the cryogenic propellant to circulate in the heat capture and energy conversion module 101 according to the electrical energy allocated by the power distribution and management module 102, and provides electrical energy for the subsequent mission of the second-stage rocket.

[0110] It should be noted that the beneficial effects of the reusable second-stage rocket energy integrated management method provided in the above embodiments are the same as those of the reusable second-stage rocket energy integrated management system 100 described above, and will not be repeated here. Furthermore, the method and system embodiments provided in the above embodiments belong to the same concept, and their specific implementation process can be found in the method embodiments, which will not be repeated here.

[0111] The above description is merely a preferred embodiment of the present invention and an explanation of the technical principles employed. Those skilled in the art should understand that the scope of disclosure in this invention is not limited to technical solutions formed by specific combinations of the above-described technical features, but should also cover other technical solutions formed by arbitrary combinations of the above-described technical features or their equivalents without departing from the above-disclosed concept. For example, technical solutions formed by substituting the above features with (but not limited to) technical features with similar functions disclosed in this invention.

[0112] It should be noted that the terms "first," "second," etc., used in the specification and claims of this application are used to distinguish similar objects and represent a limitation on a specific order or sequence. Where appropriate, the order of use for similar objects can be interchanged so that the embodiments of this application described herein can be implemented in an order other than that shown or described.

[0113] Although embodiments of the present invention have been shown and described above, it is understood that the above embodiments are exemplary and should not be construed as limiting the present invention. Those skilled in the art can make changes, modifications, substitutions and variations to the above embodiments within the scope of the present invention.

Claims

1. A reusable second-stage rocket energy integrated management system, characterized in that, include: The heat capture and energy conversion module is installed in the reentry heat protection zone of the second-stage rocket. It uses the cryogenic propellant carried by the second-stage rocket to absorb aerodynamic heat to generate a high-temperature and high-pressure working fluid, and converts the thermal energy of the high-temperature and high-pressure working fluid into electrical energy output. The power distribution and management module is used to convert the electrical energy output by the heat capture and energy conversion module into stable DC power, store the stable DC power, and distribute it according to control commands. The power execution and application module is used to drive the cryogenic propellant to circulate in the heat capture and energy conversion module according to the electrical energy allocated by the power distribution and management module, and to provide electrical energy for the subsequent mission of the second-stage rocket.

2. The reusable second-stage rocket energy integrated management system according to claim 1, characterized in that, The heat capture and energy conversion module includes: an active cooling circuit and a gas turbine power generation unit; The active cooling circuit is located in the reentry thermal protection zone of the second-stage rocket and is used to absorb the aerodynamic heat and generate the high-temperature and high-pressure working fluid using the cryogenic propellant carried by the second-stage rocket. The gas turbine power generation unit is connected to the active cooling circuit and is used to receive the high-temperature and high-pressure working fluid and convert the thermal energy of the high-temperature and high-pressure working fluid into electrical energy output.

3. The reusable second-stage rocket energy integrated management system according to claim 2, characterized in that, The power distribution and management module includes a rectifier / voltage regulation unit, an energy storage unit, an integrated controller, and a power distribution unit; The rectifier / voltage regulator unit is connected to the gas turbine power generation unit and is used to convert the electrical energy output by the gas turbine power generation unit into the stable DC power and feed it into the DC bus. The energy storage unit is connected to the DC bus via a bidirectional DC / DC converter for storing the stable DC power. The integrated controller is connected to the sensor layer, which is deployed at the corresponding measurement points of the second-stage rocket. The integrated controller is used to acquire real-time operating parameters and generate control commands based on the real-time operating parameters. The power distribution unit is connected to the DC bus and the integrated controller, and is used to distribute the stable DC power on the DC bus according to the control command.

4. The reusable second-stage rocket energy integrated management system according to claim 3, characterized in that, The power execution and application module includes an electric pump unit and an auxiliary function unit; The electric pump unit is connected to the power distribution unit and is used to drive the cryogenic propellant to circulate in the active cooling loop according to the stable DC power distributed by the power distribution unit. The auxiliary function unit is connected to the power distribution unit and is used to provide power for the subsequent missions of the second-stage rocket according to the stable DC power distributed by the power distribution unit.

5. The reusable second-stage rocket energy integrated management system according to claim 3, characterized in that, The energy storage unit includes a battery pack and a supercapacitor, and the battery pack and the supercapacitor are connected to the DC bus through the bidirectional DC / DC converter.

6. The reusable second-stage rocket energy integrated management system according to claim 3, characterized in that, The integrated controller has a built-in prediction model, and the integrated controller is specifically used for: The turbine power generation prediction curve for the future time domain is output based on the real-time operating parameters and the prediction model, and the control command is generated based on the turbine power generation prediction curve and the real-time operating parameters.

7. The reusable second-stage rocket energy integrated management system according to claim 6, characterized in that, The real-time operating parameters include reentry heat flux density, rocket body structure temperature, pressure and temperature of the cryogenic propellant, rotational speed of the electric pump unit, voltage and current of the DC bus, and remaining power of the energy storage unit; the integrated controller is specifically used for: The predicted turbine power generation curve, the reentry heat flux density, the rocket body structure temperature, the pressure and temperature of the cryogenic propellant, the rotational speed of the electric pump unit, the voltage and current of the DC bus, and the remaining power of the energy storage unit are input into a multi-objective optimization model. The control command is generated by solving the multi-objective optimization model.

8. The reusable second-stage rocket energy integrated management system according to claim 4, characterized in that, The electric pump unit includes a motor and a centrifugal pump. The motor is connected to the power distribution unit, and the centrifugal pump is connected to the motor. The motor is used to drive the centrifugal pump according to the stable DC power distributed by the power distribution unit, so that the centrifugal pump drives the cryogenic propellant to circulate in the active cooling loop.

9. The reusable second-stage rocket energy integrated management system according to claim 4, characterized in that, The auxiliary functional unit includes a tank booster, a navigation device, a communication device, and an attitude control system. The tank booster, the navigation device, the communication device, and the attitude control system are respectively connected to the power distribution unit. The tank pressurizer is used to provide electrical energy to pressurize the tank of the second-stage rocket according to the stable DC power allocated by the power distribution unit. The navigation equipment, the communication equipment, and the attitude control system are used to provide operating power for the subsequent missions of the second-stage rocket according to the stable DC power allocated by the power distribution unit.

10. A method for integrated energy management of a reusable second-stage rocket, employing the integrated energy management system for a reusable second-stage rocket as described in any one of claims 1 to 9, characterized in that, include: The heat capture and energy conversion module, located in the reentry heat protection zone of the second-stage rocket, utilizes the cryogenic propellant of the second-stage rocket to absorb aerodynamic heat and generate a high-temperature, high-pressure working fluid, which then converts the thermal energy of the high-temperature, high-pressure working fluid into electrical energy output. The power distribution and management module converts the electrical energy output by the heat capture and energy conversion module into stable DC power, stores the stable DC power, and distributes the stable DC power according to control commands. The power execution and application module drives the cryogenic propellant to circulate in the heat capture and energy conversion module according to the electrical energy allocated by the power distribution and management module, and provides electrical energy for the subsequent missions of the second-stage rocket.

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