Aircraft energy and heat integrated management system

Abstract

The invention discloses an aircraft energy and heat integrated management system, and belongs to the technical field of control and adjustment. The system integrates a turbine-based combined cycle engine subsystem, a fuel oil heat management subsystem, an environment control subsystem and an aircraft accessory subsystem, air entraining and energy extraction are carried out from the TBCC engine subsystem, combined work of all the subsystems is coordinated, energy distribution and heat management regulation and control of the whole machine are achieved, and the energy utilization rate of the whole machine is improved. The system has stage adjustment and mode switching capabilities, can perform power distribution and thermal management regulation and control according to various flight working conditions including takeoff, climbing, stamping cruising and returning in a flight envelope, greatly enriches the selection of effective heat sinks, improves the utilization efficiency of the heat sinks, and improves the energy efficiency of the heat sinks. The temperatures of the heating sources at different positions are optimally regulated and controlled, weight compensation is effectively reduced, the maneuverability and the long cruising ability of the aircraft are improved, and the flight cost is reduced.

Description

Technical Field

[0001] The present invention relates to energy distribution, thermal energy management and environmental control technologies for hypersonic aircraft, discloses an aircraft energy and thermal integrated management system, and specifically discloses an integrated energy and thermal management system for hypersonic aircraft based on a turbine-based combined cycle (TBCC) engine, belonging to the technical field of control and regulation. Background Art

[0002] As aircraft develop towards high power density, long flight time and stealth, the load of energy and thermal management systems increases, while traditional decentralized management is difficult to achieve rapid response and efficient operation. National defense needs have put forward higher requirements for aircraft stealth, maneuverability, supersonic cruise, etc. In this context, aircraft energy / thermal management issues have become more prominent. Faced with the increasing flight Mach number, the use of composite materials, the reduction of effective heat sinks, and the many demands for the maneuverability, stealth, maintainability, and power consumption and heat dissipation of next-generation aircraft, traditional decentralized airborne electromechanical systems and single energy management methods are far from meeting them. At present, the main technologies used in the field of aircraft thermal control and energy regulation at home and abroad include air circulation cooling technology, liquid loop cooling technology, auxiliary power unit (APU) power generation technology, and ramjet turbine power generation technology.

[0003] Air circulation cooling technology is widely used in various types of aircraft environmental control systems, but due to its low refrigeration performance coefficient, the reliability of the system under ground conditions is poor. In addition, since it introduces external ram air, the use of aircraft at high Mach numbers is subject to certain restrictions. These are in conflict with the characteristics of hypersonic aircraft, which suffer from severe aerodynamic heating and high flight altitude and speed.

[0004] Liquid loop cooling technology is an indirect cooling technology, and the intermediate coolant is a liquid such as fuel. The thermal conductivity and mass heat capacity of liquid are much greater than those of air. Under the same equipment power, the use of liquid coolant can reduce the flow rate and pipe size to the equipment, which is more effective for cooling concentrated heat loads and long-distance heat loads, but the targeted cooling effect on a large number of dispersed heat sources is poor, which is easy to cause redundant cooling.

[0005] APU is a small gas turbine engine that runs independently of the main engine and is used to provide electrical energy for aircraft ground operations and emergency situations. APU does not rely on the main engine when working, and can operate independently on the ground or in flight, providing electrical energy flexibly on the ground or in the air, and can still provide a certain amount of electrical energy when the main engine fails. However, due to the small size of APU, its power generation capacity is limited and cannot meet the needs of all systems. In addition, due to its high fuel consumption, APU is usually used to provide electrical energy for the start-up phase of the aircraft.

[0006] Ramjet turbine generator technology extracts air from the main engine to drive a small turbine, thereby driving the generator to generate electricity. It has a high power generation efficiency and can utilize the waste heat or surplus energy of the main engine to improve energy utilization. Compared with traditional generators, turbine generators are more compact and lightweight, but require specially designed air extraction and management systems, which increases the design complexity and affects the performance of the main engine. The extracted air may have a certain impact on the performance and efficiency of the main engine.

[0007] Given that the defects of existing aircraft thermal control and energy regulation technologies cannot meet the thermal energy management needs of hypersonic aircraft, the energy-thermal management integrated system of hypersonic aircraft has gradually become an important research direction. However, the existing energy-thermal management integrated system of hypersonic aircraft still has the following shortcomings:

[0008] Increased system complexity: Due to the deep coupling of energy and thermal management, the complexity of system integration has increased significantly, which has put forward higher requirements for design and maintenance. The traditional hypersonic vehicle energy-thermal management integrated system is difficult to balance the requirements of system performance and structural compactness;

[0009] Insufficient dynamic adaptability: The existing hypersonic vehicle energy-thermal management integrated system has a weak dynamic response capability to environmental changes during flight missions, making it difficult to achieve efficient real-time adjustments. Especially under complex flight conditions, system overload or control lag is prone to occur;

[0010] Low efficiency in heat load distribution: When dealing with high heat flux conditions, the current hypersonic vehicle energy-thermal management integrated system has insufficient ability to distribute heat loads and optimize heat dissipation paths, resulting in local overheating and thermal stress concentration, affecting the overall performance and structural life of the vehicle.

[0011] Energy utilization limitation: In the energy supply and demand balance and multi-task switching, the existing hypersonic aircraft energy-thermal management integrated system cannot make full use of energy, which easily leads to waste or insufficient energy supply.

[0012] In addition, the existing hypersonic vehicle energy-thermal management integrated system often considers people or equipment, is relatively independent, and has no overall arrangement. The energy and thermal management system of a hypersonic vehicle should not only provide a good and comfortable working environment for personnel, but also provide energy for electronic equipment and provide a safe and reliable working environment for other equipment. This puts forward more stringent requirements for the design of the energy-thermal management integrated system of a hypersonic vehicle. Factors such as weight, performance, cost, reliability, and compensation loss should be considered, and the thermal management system, energy control system, and aircraft engine propulsion system need to be integrated for design.

[0013] Therefore, the energy-thermal management integrated system needs to further break through the existing technical bottlenecks, improve the intelligence, reliability and adaptability of the system, and provide strong support for the development of a new generation of aircraft. How to provide an integrated energy and thermal management system for hypersonic aircraft to achieve rated energy supply and solve the problem of distributed heat load heat dissipation is a technical problem that technicians in this field need to solve urgently. Summary of the invention

[0014] The purpose of the present invention is to provide an aircraft energy and thermal integrated management system in response to the deficiencies of the above-mentioned background technology. In view of the energy demand and thermal management demand of the flight mission envelope of the completion phase of hypersonic aircraft, an integrated energy-thermal management system is proposed to solve the two major difficult problems of waste heat treatment and effective energy utilization during flight, and to regulate the energy distribution and thermal management of the entire aircraft to cope with different engine operating states and energy efficiency requirements. Power distribution and thermal management are regulated according to various flight conditions in the flight envelope including take-off, climb, cruise, and return, so as to improve the utilization efficiency of heat sinks, achieve optimal regulation of the temperatures of heat sources at different positions, effectively reduce weight compensation, improve aircraft maneuverability and long endurance, and reduce flight costs.

[0015] The present invention adopts the following technical solutions to achieve the above-mentioned invention objectives:

[0016] An integrated energy and heat management system for an aircraft comprises: a TBCC engine subsystem, a fuel thermal management subsystem, an environmental control subsystem and an aircraft accessory subsystem; the TBCC engine subsystem is used to perform work on gas introduced from the outside to generate mechanical energy; the fuel thermal management subsystem is used to cool an onboard heat source, and at the same time, perform ram air cooling and / or liquid nitrogen cooling on the fuel after cooling the onboard heat source, and supply oil to the TBCC engine subsystem; the environmental control subsystem is used to provide auxiliary power after introducing high-pressure gas from the TBCC engine subsystem or operating gas from the outside, provide emergency power when the TBCC engine subsystem does not provide high-pressure gas, and introduce low-pressure gas from the TBCC engine subsystem to perform air circulation cooling on the engine compartment, landing gear compartment and cockpit; the aircraft accessory subsystem works under the action of auxiliary power or emergency power, or is driven by extracting mechanical energy output by the TBCC engine subsystem, or is driven by the action of operating gas introduced from the outside.

[0017] As a further optimization scheme for the integrated energy and thermal management system of an aircraft, the TBCC engine subsystem includes a turbofan engine and a ramjet engine. During the takeoff phase, neither the turbofan engine nor the ramjet engine works. During the climb working phase and the restarting phase, only the turbofan engine works, and the low-pressure gas at the outlet of the low-pressure compressor of the turbofan engine is introduced into the environmental control subsystem through the low-pressure bleed air pipeline, and the high-pressure gas at the outlet of the high-pressure compressor of the turbofan engine is introduced into the environmental control subsystem through the high-pressure bleed air pipeline. The mechanical energy generated by the high-pressure turbine of the turbofan engine is transmitted to the aircraft accessory subsystem through the mechanical transmission shaft. During the TBCC engine mode conversion phase, the turbofan engine is shut down and the ramjet engine is started. During the ramjet engine working alone, the working gas introduced from the outside is transmitted to the aircraft accessory subsystem through the ramjet bleed air pipeline. During the unpowered descent phase, neither the turbofan engine nor the ramjet engine works.

[0018] As a further optimization scheme for the integrated energy and thermal management system of an aircraft, the fuel thermal management subsystem: does not work during the take-off phase; cools the onboard heat source during the climb working phase, the TBCC engine mode conversion phase and the restart phase, and at the same time, uses ram air cooling and liquid nitrogen cooling for the fuel after cooling the onboard heat source; cools the onboard heat source during the ramjet engine single-operation phase and the unpowered descent phase, and at the same time, uses liquid nitrogen cooling for the fuel after cooling the onboard heat source.

[0019] As a further optimization scheme for the integrated energy and thermal management system of an aircraft, the fuel thermal management subsystem includes: a ram air cooling circuit and a liquid nitrogen cooling circuit; in the ram air cooling circuit, the fuel in the main fuel tank flows through the first heat exchanger, the second heat exchanger and the fourth heat exchanger in sequence for circulated heat exchange, the first heat exchanger and the second heat exchanger are liquid-liquid heat exchangers, and the fourth heat exchanger is a liquid-gas heat exchanger with ram air as the cooling medium; in the liquid nitrogen cooling circuit, the fuel in the main fuel tank flows through the first heat exchanger, the second heat exchanger and the third heat exchanger in sequence, and a part of the fuel is pumped to each combustion chamber in the TBCC engine subsystem, and a part of the fuel enters the fifth heat exchanger for cooling and is then pumped back to the main fuel tank through the sixth heat exchanger, the third heat exchanger is a liquid-liquid heat exchanger, the fifth heat exchanger is a liquid-liquid heat exchanger with liquid nitrogen as the cooling medium, and the sixth heat exchanger is a liquid-gas heat exchanger, and the engine heat load is connected to the third heat exchanger to form a liquid-liquid heat exchange circuit.

[0020] As a further optimization scheme for the integrated energy and thermal management system of an aircraft, the environmental control subsystem includes: an auxiliary power unit, an emergency power unit and a liquid nitrogen cascade utilization device; the auxiliary power unit provides auxiliary power after introducing working gas from the outside during the take-off phase; provides auxiliary power after introducing high-pressure gas from the TBCC engine subsystem during the climbing working phase and the restarting phase, and introduces low-pressure gas from the TBCC engine subsystem to perform air circulation cooling on the engine compartment, landing gear compartment and cockpit; the emergency power unit provides emergency power during the TBCC engine mode conversion phase and the unpowered descent phase; the liquid nitrogen cascade utilization device performs cascade liquid nitrogen cooling on the engine compartment, landing gear compartment and cockpit during the TBCC engine mode conversion phase, the ramjet engine single working phase and the unpowered descent phase.

[0021] As a further optimization solution for the integrated energy and thermal management system of an aircraft, the auxiliary power unit includes: a power turbine, a high-speed transmission shaft, a cooling turbine, an inspiration motor, a low-pressure fan, an auxiliary power unit combustion chamber and an auxiliary fuel tank; during the takeoff phase, the inspiration motor is in starter mode, the inspiration motor drives the low-pressure fan to work, the working gas introduced from the outside is transported to the auxiliary power unit combustion chamber through the low-pressure fan, the fuel in the auxiliary fuel tank is transported to the auxiliary power unit combustion chamber, the fuel burns in the auxiliary power unit combustion chamber, and the high-temperature and high-pressure gas generated drives the power turbine to rotate; during the climbing phase and the restart phase, the low-pressure fan is driven by the inspiration motor to drive the auxiliary power unit combustion chamber. The high-pressure gas introduced by the TBCC engine subsystem is transported to the auxiliary power unit combustion chamber, and the fuel in the auxiliary fuel tank is transported to the auxiliary power unit combustion chamber. The high-temperature and high-pressure gas generated by the combustion of the fuel in the auxiliary power unit combustion chamber drives the power turbine. The low-pressure gas introduced from the TBCC engine subsystem passes through the low-pressure fan, and is cooled by the sixth heat exchanger and the seventh heat exchanger in turn. The gas cooled by the seventh heat exchanger enters the cooling turbine and the power turbine. The cooling gas passing through the power turbine performs air circulation cooling on the engine compartment, landing gear compartment and cockpit. The sixth heat exchanger is a liquid-to-gas heat exchanger, and the seventh heat exchanger is an air-to-gas heat exchanger.

[0022] As a further optimization solution for the integrated energy and thermal management system of an aircraft, the emergency power unit includes: a power turbine, an auxiliary power unit combustion chamber and emergency fuel. During the TBCC engine mode conversion phase and the unpowered descent phase, the emergency fuel is transported to the auxiliary power unit combustion chamber, and the emergency fuel burns in the auxiliary power unit combustion chamber. The high-temperature and high-pressure gas generated drives the power turbine to rotate.

[0023] As a further optimization plan for the aircraft energy and thermal integrated management system, the liquid nitrogen cascade utilization device includes: a liquid nitrogen storage tank and a liquid nitrogen booster pump. The low-temperature liquid nitrogen in the liquid nitrogen storage tank is pressurized by the liquid nitrogen booster pump and pumped to the cockpit, and the heated nitrogen in the cockpit is sent to the landing gear compartment. A stream of low-temperature liquid nitrogen is diverted from the liquid nitrogen booster pump and pumped into the landing gear compartment, and a stream of low-temperature liquid nitrogen is diverted from the liquid nitrogen booster pump and pumped into the engine compartment.

[0024] As a further optimization scheme for the integrated energy and thermal management system of an aircraft, the aircraft accessory subsystem includes: an aircraft accessory casing and a ramjet turbine; during the take-off phase, the auxiliary power provided by the auxiliary power unit is transmitted to the aircraft accessory casing, and the aircraft accessory casing outputs electrical energy to start the main engine; during the climbing working phase and the restarting phase, the auxiliary power provided by the auxiliary power unit is transmitted to the aircraft accessory casing, and the mechanical energy extracted from the TBCC engine subsystem is transmitted to the aircraft accessory casing, and the aircraft accessory casing outputs electrical energy to start the main engine and drive the hydraulic system to work; during the TBCC engine mode conversion phase and the unpowered descent phase, the emergency power provided by the auxiliary power unit is transmitted to the aircraft accessory casing, and the aircraft accessory casing outputs electrical energy to start the main engine and drive the hydraulic system to work; during the ramjet engine working alone, the working gas introduced from the outside drives the ramjet turbine, and the mechanical energy generated by the ramjet turbine is transmitted to the aircraft accessory casing, and the aircraft accessory casing outputs electrical energy to start the main engine and drive the hydraulic system to work.

[0025] As a further optimization solution for the aircraft energy and thermal integrated management system, batteries are used to supply power to output electrical energy during the TBCC engine mode conversion phase and the unpowered descent phase.

[0026] The present invention adopts the above technical solution and has the following beneficial effects:

[0027] (1) The improved integrated energy and thermal management system for supersonic aircraft of the present invention aims to adapt to the flight envelope of hypersonic aircraft as the optimal design principle, deeply couple the energy management and thermal management functions, and improve the efficiency, response speed and reliability of the overall system through integrated design.

[0028] (2) Compared with the traditional aircraft thermal management method, the improved integrated energy and thermal management system of the supersonic aircraft provided by the present invention adds a combined cooling solution for distributed and large heat sources through the fuel thermal management subsystem and the environmental control subsystem. It adopts a combination of multiple heat sinks such as bleed air cooling, fuel cooling, and liquid nitrogen step cooling. The corresponding cooling method can be matched according to the change of thermal load in the flight envelope. The improved air / fluid circulation loop can greatly improve the cooling efficiency of the heat load generated by the aircraft at different stages, thereby improving the heat dissipation efficiency of the entire aircraft.

[0029] (3) The present invention improves the integrated energy and thermal management system of supersonic aircraft, proposes a new main engine shaft power extraction method - turbofan engine energy extraction, and proposes an integrated auxiliary-emergency power generation method. The main engine shaft power extraction and ramjet turbine power generation, APU power generation, and battery power output form an energy closed loop, which can realize power extraction at various stages of the aircraft flight according to the energy demand in the aircraft envelope, greatly reducing the flight curb weight while realizing more diverse and feasible power generation methods, meeting the needs of more high-power electrical equipment, and greatly improving power generation efficiency. BRIEF DESCRIPTION OF THE DRAWINGS

[0030] In order to more clearly illustrate the embodiments of the present invention or the technical solutions in the prior art, the drawings required for use in the embodiments will be briefly introduced below. Obviously, the drawings described below are only some embodiments of the present invention. For those skilled in the art, other drawings can be obtained based on these drawings without paying creative labor.

[0031] Figure 1 It is a schematic diagram of the composition of the integrated energy and thermal management system for a hypersonic aircraft of the present invention.

[0032] Figure 2 It is a schematic diagram of the working state of the integrated energy and thermal management system of a hypersonic aircraft during the take-off phase of the present invention.

[0033] Figure 3 It is a schematic diagram of the working state of the integrated energy and thermal management system of a hypersonic aircraft during the climbing phase of the present invention.

[0034] Figure 4 It is a schematic diagram of the working state of the integrated energy and thermal management system of a hypersonic aircraft in the mode conversion stage of the present invention.

[0035] Figure 5 It is a schematic diagram of the working state of the integrated energy and thermal management system of a hypersonic aircraft in the ramming stage of the present invention.

[0036] Figure 6 It is a schematic diagram of the working state of the integrated energy and thermal management system of a hypersonic aircraft in the unpowered stage of the present invention.

[0037] Figure 7 It is a schematic diagram of the working state of the integrated energy and thermal management system of a hypersonic aircraft in the restart phase of the present invention.

[0038] Explanation of the numbers in the figure: 1. Inlet, 2. Low-pressure compressor, 3. High-pressure compressor, 4. Turbofan engine combustion chamber, 5. Bypass, 6. High-pressure turbine, 7. Low-pressure turbine, 8. Gas mixing chamber, 9. Afterburner, 10. Low-speed nozzle, 11. Isolation section, 12. Ramjet engine combustion chamber, 13. High-speed nozzle, 14. Main fuel tank, 15. Aerodynamic skin heat transfer, 16. Engine compartment, 17. Landing gear compartment, 18. Cockpit, 19. Liquid nitrogen storage tank, 20. Low-pressure fan, 21. Initiator motor, 22. Cooling turbine, 23. , high-speed transmission shaft, 24, power turbine, 25, emergency fuel, 26, auxiliary power unit combustion chamber, 27, auxiliary fuel tank, 28, aircraft accessory casing, 29, external environment, 30, battery, 31, electrical energy, 32, hydraulic energy, 33, ramjet turbine, I1, low-pressure bleed air pipeline, I2, high-pressure bleed air pipeline, I3, mechanical transmission shaft, I4, ramjet bleed air pipeline, B1, fuel boost pump, B2, liquid nitrogen boost pump, Z1, electronic equipment thermal load, Z2, lubricating oil-hydraulic oil thermal load, Z3, engine thermal load, RC1, ramjet Air, RC2, liquid nitrogen, HR1, the first heat exchanger, HR2, the second heat exchanger, HR3, the third heat exchanger, HR4, the fourth heat exchanger, HR5, the fifth heat exchanger, HR6, the sixth heat exchanger, HR7, the seventh heat exchanger, V1, the first fuel flow regulating valve, V2, the second fuel flow regulating valve, V3, the third fuel flow regulating valve, V4, the electronic equipment regulating valve, V5, the lubricating oil-hydraulic oil regulating valve, V6, the engine thermal load flow regulating valve, V7, the ram air flow regulating valve, V8, the liquid nitrogen flow regulating valve, V9, the first The fourth fuel flow regulating valve, V10, the fifth fuel flow regulating valve, V11, the first gas regulating valve, V12, the second gas regulating valve, V13, the third gas regulating valve, V14, the fourth gas regulating valve, V15, the fifth gas regulating valve, V16, the sixth gas regulating valve, V17, the seventh gas regulating valve, V18, the eighth gas regulating valve, V19, the ninth gas regulating valve, V20, the sixth fuel flow regulating valve, V21, the seventh fuel flow regulating valve, V22, the eighth fuel flow regulating valve, V23, the ninth fuel flow regulating valve. DETAILED DESCRIPTION

[0039] In order to make the purpose, technical solution and advantages of the present invention clearer, the technical solution of the present invention will be clearly and completely described below in conjunction with the drawings in the embodiments of the present invention. Obviously, the described embodiments are part of the embodiments of the present invention, rather than all the embodiments. Based on the embodiments of the present invention, all other embodiments obtained by those skilled in the art without creative work are within the scope of protection of the present invention.

[0040] like Figure 1As shown, the present invention discloses an integrated energy and thermal management system for a hypersonic aircraft based on a TBCC engine, which is composed of a TBCC engine subsystem, a fuel thermal management subsystem, an environmental control subsystem, and an aircraft accessory subsystem. Air bleed and energy extraction are performed through the TBCC engine subsystem, and the joint work between the various subsystems is coordinated to achieve energy distribution and thermal management regulation of the aircraft during flight missions to meet the energy efficiency requirements of different engine operating states. During the flight, the various subsystems work in conjunction with each other through pipelines and mechanical shafts to complete gas transportation and mechanical energy output, bleed air and extract electrical energy from the TBCC engine subsystem to meet the thermal management needs of the aircraft's heat-generating components and the electrical energy needs during flight, achieve optimal regulation of electrical energy distribution and thermal energy management for hypersonic aircraft, and improve the aircraft's maneuverability and long endurance. Along the flight profile, it can cover the energy distribution and thermal management regulation needs of takeoff, climb phase, mode conversion phase, ramjet cruise independent working phase, unpowered taxi phase, and restart flight phase in the hypersonic flight envelope. Among them, the energy demand is reflected in the system as the power output of electrical energy and hydraulic energy, and the thermal management demand is reflected in the system as multi-heat source matching ram air cooling, fuel cooling, and liquid nitrogen cooling solutions.

[0041] The TBCC engine subsystem is the main power source of the entire energy management system. Figure 1As shown, the TBCC engine subsystem adopts a parallel ramjet engine, including: a turbofan engine consisting of an air inlet 1, a low-pressure compressor 2, a high-pressure compressor 3, a turbofan engine combustion chamber 4, a high-pressure turbine 6, a low-pressure turbine 7, an outer duct 5, a gas mixing chamber 8, an afterburner 9, and a low-speed nozzle 10; and a ramjet engine consisting of an air inlet 1, an isolation section 11, a ramjet engine combustion chamber 12, and a high-speed nozzle 13. When the turbofan engine is working, the external working condition gas enters the low-pressure compressor 2 through the air inlet 1 for pressurization, the low-pressure gas at the outlet of the low-pressure compressor 2 is sent to the high-pressure compressor 3 and the outer duct 5, the high-temperature and high-pressure gas at the outlet of the high-pressure compressor 3 is sent to the turbofan engine combustion chamber 4 for combustion, the high-temperature and high-pressure combustion gas generated by the turbofan engine combustion chamber 4 drives the high-pressure turbine 6 and the low-pressure turbine 7 to rotate, the gas expanded by the low-pressure turbine 7 and the low-pressure gas output by the outer duct 5 enter the gas mixing chamber 8 together, the mixed gas at the outlet of the gas mixing chamber 8 is sent to the afterburner 9 for further combustion, and the high-temperature combustion gas generated by the afterburner 9 is sent to the low-speed nozzle 10 to generate thrust. When the ramjet engine is working, the external working condition gas is compressed through the air inlet 1 and sent to the isolation section 11, the high-temperature and high-pressure gas at the outlet of the isolation section 11 is sent to the ramjet engine combustion chamber 12, and the high-temperature and high-pressure gas generated by the combustion in the ramjet engine combustion chamber 12 is sent to the high-speed nozzle 13 to generate thrust. During flight, the ramjet engine and the turbofan engine can be used separately or reused according to the flight conditions, and have the ability to quickly switch modes. The low-pressure bleed air pipeline I1 introduces the low-pressure gas from the outlet of the low-pressure compressor 2 into the environmental control subsystem; the high-pressure bleed air pipeline I2 introduces the high-pressure gas from the outlet of the high-pressure compressor 3 into the environmental control subsystem; the mechanical transmission shaft I3 transmits the mechanical energy generated by the high-pressure turbine 6 to the aircraft accessory subsystem; the ramjet bleed air pipeline I4 delivers the gas sent from the air inlet 1 to the aircraft accessory subsystem when the ramjet engine is working. Figure 1As shown, the fuel thermal management subsystem includes: a main fuel tank 14, an aerodynamic skin heat transfer 15, a fuel boost pump B1, an electronic equipment thermal load Z1, a lubricating oil-hydraulic oil thermal load Z2, an engine thermal load Z3, ram air RC1, liquid nitrogen RC2, a first heat exchanger HR1, a second heat exchanger HR2, a third heat exchanger HR3, a fourth heat exchanger HR4, a fifth heat exchanger HR5, a sixth heat exchanger HR6, an electronic equipment regulating valve V4, a lubricating oil-hydraulic oil regulating valve V5, an engine thermal load flow The first heat exchanger HR1, the second heat exchanger HR2, the third heat exchanger HR3, and the fifth heat exchanger HR5 are all liquid-liquid heat exchangers, and the fourth heat exchanger HR4 and the sixth heat exchanger HR6 are all liquid-gas heat exchangers. The main fuel tank 14 is connected to the aerodynamic skin heat transfer 15, which means that the aerodynamic heat during the flight is applied to the main fuel tank 14. After being pressurized by the boost pump B1, the fuel passes through the first heat exchanger HR1 and the second heat exchanger HR2 in sequence for heat exchange treatment. The electronic equipment heat load Z1 is connected to the first heat exchanger HR1, and the lubricating oil-hydraulic oil heat load Z2 is connected to the second heat exchanger HR2. After the heat exchange is completed, it is divided into two streams of fuel. One of the streams of fuel passes through the fourth heat exchanger HR4 to exchange heat with the ram air RC1 to complete cooling. The cooled fluid again participates in the heat exchange circuit of the first heat exchanger HR1 to form a heat exchange circuit of HR1-HR2-HR4-HR1. When the flow rate in the HR1-HR2-HR4-HR1 heat exchange circuit increases, the cooling speed and cooling effect will be greatly improved. It is calculated that when the fuel flow rate is 0.5kg / s, the cooling effect of the HR1-HR2-HR4-HR1 heat exchange circuit is improved by 6%; the other stream of fuel The fuel directly flows through the liquid-liquid heat exchange loop composed of the third heat exchanger HR3 and the engine heat load Z3 for heat exchange. A part of the fuel after heat exchange in the liquid-liquid heat exchange loop is directly delivered to the ramjet engine combustion chamber 12, the turbofan engine combustion chamber 4 and the afterburner 9 through the first fuel flow regulating valve V1, the second fuel flow regulating valve V2 and the third fuel flow regulating valve V3; the other part of the fuel after heat exchange in the liquid-liquid heat exchange loop returns to the fifth heat exchanger HR5 and is cooled by liquid nitrogen RC2. The fuel after low-temperature cooling can be regulated by the fourth fuel flow regulating valve V9 and the fifth fuel flow regulating valve V10 to control the fuel flow participating in the circulation and the flow returning to the main fuel tank 14. According to the different operating conditions of the TBCC engine subsystem from low Mach to high Mach, it can realize adaptive selection of ramjet air RC1 or liquid nitrogen RC2 as the optimal cooling medium, and it can also realize the use of combined cooling to improve the utilization efficiency of the heat sink.The system performs fuel diversion regulation on the fuel before the engine heat load Z3 through the eighth fuel flow regulating valve V22, and can effectively control the cooling flow of heat exchange with the engine through the ninth fuel flow regulating valve V23, so as to maintain the fuel outlet temperature after heat exchange with the engine within the effective safety temperature range of 110°C, thereby improving the combustion efficiency of the fuel entering the turbofan engine combustion chamber 4 and the afterburner 9. It is calculated that compared with the fuel temperature entering the inlet of the turbofan engine combustion chamber 4 and the afterburner 9 being controlled at 110°C, not only does the fuel have the highest cooling efficiency before reaching the coking temperature, but also compared with the fuel directly entering the combustion chamber at a basic temperature of 60°C, the combustion efficiency can be improved by 9.45%.

[0042] like Figure 1 As shown, the environmental control subsystem includes an auxiliary-emergency power unit and a liquid nitrogen cascade utilization device. The environmental control subsystem is connected to the low-pressure air bleed pipeline I1 and the high-pressure air bleed pipeline I2 respectively. The auxiliary-emergency power unit can provide the power requirements of the aircraft during take-off and emergency restart, which can greatly save the use of liquid nitrogen while ensuring the utilization efficiency of the heat sink and enhance the aircraft's endurance.

[0043] The liquid nitrogen cascade utilization device includes: an engine compartment 16, a landing gear compartment 17, a cockpit 18, a liquid nitrogen storage tank 19, and a liquid nitrogen booster pump B2. The low-temperature liquid nitrogen in the liquid nitrogen storage tank 19 is stored in an environment of -196°C. The cockpit 18, the landing gear compartment 17, and the engine compartment 16 are cooled in a cascade manner by utilizing the large latent heat of vaporization and excellent cooling capacity of the low-temperature liquid nitrogen. The low-temperature liquid nitrogen is pressurized by the liquid nitrogen booster pump B2 and pumped to the cockpit 18. Spray cooling is performed in the cockpit 18 to cool the cockpit to below 25°C. At this time, the outlet temperature of the cockpit 18 is 25°C. The heated gaseous liquid nitrogen can still be passed into the landing gear compartment 17. At this time, the heated gaseous liquid nitrogen Liquid nitrogen cooling efficiency is low, so a stream of low-temperature liquid nitrogen is introduced from the liquid nitrogen booster pump B2 to assist in cooling the landing gear compartment 17, and the purpose of saving liquid nitrogen is achieved in this part. The landing gear compartment 17 is cooled to 110°C and the high-temperature gas is discharged; another stream of low-temperature liquid nitrogen is introduced from the liquid nitrogen booster pump B2 to spray cool the engine compartment 16, and the high-temperature gas is discharged after the engine compartment temperature is cooled to 110°C. The fifth gas regulating valve V15 is used to control the outlet flow of low-temperature liquid nitrogen in the liquid nitrogen storage tank 19. According to calculations, compared with the scheme without step cooling, this scheme adopts step cooling to save 13% of liquid nitrogen consumption, which can effectively save liquid nitrogen consumption.

[0044] The auxiliary-emergency power unit includes: a low-pressure fan 20, an inspiration motor 21, a cooling turbine 22, a high-speed transmission shaft 23, a power turbine 24, an emergency fuel 25, an auxiliary power unit combustion chamber 26, and an auxiliary fuel tank 27. The power turbine 24, the high-speed transmission shaft 23, the cooling turbine 22, the inspiration motor 21, the low-pressure fan 20, the auxiliary power unit combustion chamber 26, and the auxiliary fuel tank 27 constitute the auxiliary power unit. The power turbine 24, the auxiliary power unit combustion chamber 26, and the emergency fuel 25 constitute the emergency power unit. During the aircraft startup phase, the starting motor 21 is in the starting motor state, and the starting motor 21 drives the low-pressure fan 20 to work. The gas sucked in from the outside is delivered to the auxiliary power unit combustion chamber 26 through the sixth gas regulating valve V16, the low-pressure fan 20, and the seventh gas regulating valve V17. At the same time, the auxiliary fuel tank 27 also delivers fuel to the auxiliary power unit combustion chamber 26 through the sixth fuel flow regulating valve V20. The high-temperature and high-pressure gas generated by the combustion of the fuel in the auxiliary power unit combustion chamber 26 drives the power turbine 24 to rotate, thereby driving the auxiliary power unit to work. Specifically, the mechanical properties generated by the rotation of the power turbine 24 are transmitted to the cooling turbine 22, the starting motor 21, and the low-pressure fan 20 through the high-speed transmission shaft 23. The various power-consuming components of the aircraft accessory subsystem also provide mechanical energy for the aircraft accessory casing 28 in the aircraft accessory subsystem; the high-pressure gas at the outlet of the high-pressure compressor 3 introduced directly through the high-pressure bleed air channel I2 enters the auxiliary power unit combustion chamber 26 through the seventh gas regulating valve V17, and the emergency fuel 25 is delivered to the auxiliary power unit combustion chamber 26 through the seventh fuel flow regulating valve V21. The high-temperature and high-pressure gas generated by the combustion of the emergency fuel 25 in the auxiliary power unit combustion chamber 26 drives the power turbine 24 to rotate and perform work, thereby driving the emergency power unit to work and providing mechanical energy for the aircraft accessory casing 28 in the aircraft accessory subsystem. The eighth gas regulating valve V18 is used to adjust the gas flow rate of the high-pressure gas entering the auxiliary power unit combustion chamber 26.

[0045] At the same time, in the auxiliary-emergency power device, the low-pressure bleed air pipeline I1 transports the low-pressure gas at the outlet of the low-pressure compressor 2 to the low-pressure fan 20 to drive it to rotate. The low-pressure gas at the outlet of the low-pressure compressor 2 passes through the low-pressure fan 20 and is cooled in turn through the sixth heat exchanger HR6 and the seventh heat exchanger HR7. The sixth heat exchanger HR6 is a liquid-gas heat exchanger, and the seventh heat exchanger HR7 is an air-gas heat exchanger. The gas cooled by the seventh heat exchanger HR7 enters the cooling turbine 22 and the power turbine 24. The gas passing through the power turbine 24 passes through the second gas regulating valve V12, The third gas regulating valve V13 and the fourth gas regulating valve V14 are connected in parallel to the engine compartment 16, the landing gear compartment 17 and the cockpit 18. After the heat exchange and temperature rise in the engine compartment 16, the landing gear compartment 17 and the cockpit 18, the gas passes through the seventh heat exchanger HR7 again to complete the closed heat exchange cycle. The bleed air flow participating in the closed heat exchange cycle is regulated and controlled by the first gas regulating valve V11. This process is carried out during the climbing phase of the aircraft at a low Mach number, providing the electrical energy required for the aircraft's take-off phase, and meeting the cooling requirements during the low-speed flight phase and the electrical energy requirements for emergency restart of the aircraft.

[0046] like Figure 1 As shown, the aircraft accessory subsystem includes: an aircraft accessory casing 28, a ramjet turbine 33, and a ninth gas regulating valve V19. When the ramjet engine is working, the working gas from the ramjet bleed pipe I4 drives the ramjet turbine 33 to rotate at a high speed, and the mechanical energy generated by the ramjet turbine 33 is delivered to the aircraft accessory casing 28. The gas expanded by the ramjet turbine 33 is delivered to the external environment 29 through the ninth gas regulating valve V19. The mechanical transmission shaft I3 transmits the mechanical energy generated by the high-pressure turbine 6 to the aircraft accessory casing 28. The mechanical energy from the mechanical transmission shaft I3 and the power turbine 24 also jointly drives the aircraft accessory casing 28 to work, completes the output of electric energy 31 and drives the aircraft hydraulic system to work, which is reflected in the output of hydraulic energy 32. The aircraft accessory subsystem can use the battery 30 to directly output electric energy 31 according to different flight conditions, so as to achieve efficient energy distribution effect.

[0047] The integrated energy and thermal management system based on the TBCC hypersonic aircraft engine can switch the system mode, extract energy or refrigeration cycle according to different flight conditions, and can keep the system within a controllable range by adjusting the valve opening. The entire system contains 21 regulating valves, each of which is used to realize the system mode switching and control the energy and thermal management system to meet the corresponding energy demand and cooling capacity demand in each mode. Among them, when the valve opening is 0, it is in the off state, and when the valve opening is 1, it is in the fully open state. According to the flight envelope and energy demand of the aircraft, along the flight profile, the energy distribution and thermal management control of the take-off, climb working stage, TBCC engine mode conversion stage, ramjet engine single working stage, unpowered descent stage, and turbine restart stage are completed in sequence. In response to different engine working conditions and energy efficiency requirements within the flight envelope, the energy control system should have adaptive adjustment and mode switching capabilities.

[0048] The working states and mode switching capabilities of the integrated energy and thermal management system of a hypersonic aircraft based on a TBCC engine at each stage are shown in Table 1.

[0049]

[0050] Table 1

[0051] like Figure 2 As shown, in the integrated energy and thermal management system of a hypersonic aircraft based on a TBCC engine, only the auxiliary power unit and the aircraft accessory subsystem work during the take-off phase. When the system is in the ground start mode, the inspiration motor 21 is in the starter mode and is powered by its own power supply. The inspiration motor 21 drives the low-pressure fan 20 to work. Under ground working conditions, the air drawn from the outside passes through the sixth gas regulating valve V16, the low-pressure fan 20, and the seventh gas regulating valve V17 in sequence, and then enters the auxiliary power unit combustion chamber 26 and mixes with the fuel from the auxiliary fuel tank 27 to burn. The high-temperature and high-pressure gas generated after the combustion enters the power turbine 24 and expands, driving the power turbine 24 to rotate at a high speed, and transmits the mechanical energy to the aircraft accessory casing 28 through the shaft power to output high-power electrical energy 31 to start the aircraft main engine. Using this method to start the main engine effectively avoids the inefficient and time-consuming aircraft engine start-up process of the ground support vehicle, shortening the take-off time. The valve configuration during the take-off phase is shown in Table 2:

[0052]

[0053] Table 2

[0054] like Figure 3As shown, the integrated energy and thermal management system of a hypersonic aircraft based on a TBCC engine goes through an air climbing stage and a horizontal acceleration stage in sequence during the climbing stage, and only the turbofan engine works in the TBCC engine subsystem; the low-pressure gas introduced by the low-pressure bleed air pipeline I1 is bleed-cooled in a closed heat exchange cycle in the environmental control subsystem; the high-pressure gas extracted by the high-pressure bleed air pipeline I2 is sent to the environmental control subsystem to start the auxiliary-emergency power unit, and the mechanical energy generated by the auxiliary-emergency power unit and the mechanical energy generated by the turbofan engine extracted by the mechanical transmission shaft I3 are transmitted to the aircraft accessory casing 28, driving the aircraft hydraulic system to work, which is reflected in the output of hydraulic energy 28, and at the same time, the shaft work is extracted and the electrical energy 31 is output in the aircraft accessory casing 28; the fuel thermal management system cools the fuel under the control of V4-V10, V22, and V23, and the cooled fuel is sent to the turbofan engine combustion chamber 4 and the afterburner combustion chamber 9 through the second fuel flow regulating valve V2 and the third fuel flow regulating valve V3. The valve configuration during the climbing stage is shown in Table 3:

[0055]

[0056]

[0057] Table 3

[0058] like Figure 4 As shown, the integrated energy and thermal management system of a hypersonic aircraft based on a TBCC engine undergoes a conversion process of shutting down the turbofan engine and starting the ramjet engine during the mode conversion stage. In this stage, the emergency power unit works, and the mechanical energy generated by the emergency power unit is transmitted to the aircraft accessory casing 28 through the transmission shaft, driving the aircraft hydraulic system to work, and at the same time, shaft power is extracted and output electric energy 31 is output in the aircraft accessory casing 28; in this stage, the aircraft accessory subsystem uses a battery 30 to output electric energy 31 to power the power consumption system; in the environmental control subsystem, due to the high flight altitude and Mach number, engine bleed air cooling is no longer applicable, so this system cools each cabin through a liquid nitrogen step cooling device;

[0059] The working condition of the fuel thermal management subsystem in this stage is the same as that of the turbofan working stage alone. All valves in the fuel subsystem are in the open state, and a combined cooling scheme of ram air and liquid nitrogen is adopted. The cooled fuel is sent to the ramjet engine combustion chamber 12, the turbofan engine combustion chamber 4, and the afterburner combustion chamber 9 through the first fuel flow regulating valve V1, the second fuel flow regulating valve V2, and the third fuel flow regulating valve V3. The valve configuration in the mode conversion stage is shown in Table 4:

[0060]

[0061]

[0062] Table 4

[0063] like Figure 5 As shown, in the integrated energy and thermal management system of a hypersonic aircraft based on a TBCC engine, during the ramjet powered climb, cruise, reconnaissance, weapon delivery, and cruise return phases, only the ramjet engine works in the TBCC engine subsystem, and the system uses the ramjet bleed air pipeline I4 to bleed air, and transport the high-temperature and high-pressure gas to the ramjet turbine 33 for expansion and high-speed rotation, providing mechanical energy for the aircraft accessory sub-casing 28, driving the aircraft hydraulic system to work, and at the same time extracting shaft power and outputting electrical energy 31 in the aircraft accessory casing 28; in the fuel thermal management subsystem, due to the high flight Mach number, the engine The bleed air cooling method is no longer applicable, so the liquid nitrogen cooling solution is used throughout this stage. The fuel is pressurized by the fuel booster pump B1 and then cooled by liquid nitrogen through the HR1-HR2-HR3-HR5 heat exchange circuit. The fuel cooled by liquid nitrogen is divided into two streams through the fourth fuel flow regulating valve V9 and the fifth fuel flow regulating valve V10 to participate in the cooling cycle or return to the main fuel tank; in this stage, the cockpit, landing gear compartment, and engine compartment are also cooled by spray cooling using a liquid nitrogen cascade cooling device. The cooling effect is the same as that in the mode conversion stage. The valve configuration in the stamping stage is shown in Table 5:

[0064]

[0065]

[0066] Table 5

[0067] like Figure 6 The integrated energy and thermal management system of a hypersonic aircraft based on a TBCC engine is shown. When the TBCC engine subsystem is working in the powerless stage, the TBCC engine subsystem does not work. In the system, only the fuel thermal management subsystem and the liquid nitrogen cascade utilization device are put into operation. In this stage, the emergency power device is put into operation, and the mechanical energy generated by the emergency power device is transmitted to the aircraft accessory casing 28 through the transmission shaft to drive the aircraft hydraulic system to work. In this stage, the battery 30 is used to output electric energy 31, and at the same time, the shaft power is extracted in the aircraft accessory casing 28 to output electric energy 31 to power the power system; in the fuel thermal management system, due to the high flight Mach number, the engine bleed air cooling method is also no longer applicable at this time, so the liquid nitrogen cooling solution is used throughout this stage; in this stage, the cooling of the cockpit, landing gear compartment, and engine compartment is also sprayed with a liquid nitrogen cascade cooling device, and the cooling effect is the same as that in the mode conversion stage. The valve configuration in the powerless stage is shown in Table 6:

[0068]

[0069] Table 6

[0070] like Figure 7 The integrated energy and thermal management system of a hypersonic aircraft based on a TBCC engine is shown in the restart phase, which goes through the air climb and horizontal acceleration phases in sequence. The system working status is similar to Figure 3 The same as the ramp phase shown, the valve configuration during the restart phase is shown in Table 3.

[0071] The specific settings are based on actual construction requirements and are not specifically limited in this application.

[0072] It should be noted that, in this application, relational terms such as "first" and "second" are only used to distinguish one entity or operation from another entity or operation, and do not necessarily require or imply any such actual relationship or order between these entities or operations. Moreover, the terms "include", "comprise" or any other variants thereof are intended to cover non-exclusive inclusion, so that a process, method, article or device including a series of elements includes not only those elements, but also other elements not explicitly listed, or also includes elements inherent to such process, method, article or device. In the absence of further restrictions, the elements defined by the sentence "comprise a ..." do not exclude the presence of other identical elements in the process, method, article or device including the elements.

[0073] The above is only a specific implementation of the present application, so that those skilled in the art can understand or implement the present application. Various modifications to these embodiments will be apparent to those skilled in the art, and the general principles defined in the present application can be implemented in other embodiments without departing from the spirit or scope of the present application. Therefore, the present application will not be limited to these embodiments shown herein, but will conform to the widest scope consistent with the principles and novel features of the present application.

Claims

1. An integrated energy and thermal management system for an aircraft, characterized in that: include: The TBCC engine subsystem is used to perform work on the gas introduced from the outside to generate mechanical energy; A fuel thermal management subsystem, used to cool the onboard heat source, and to perform ram air cooling and / or liquid nitrogen cooling on the fuel after cooling the onboard heat source, and supply fuel to the TBCC engine subsystem; An environmental control subsystem, used to provide auxiliary power by introducing high-pressure gas from the TBCC engine subsystem or operating gas from the outside, to provide emergency power when the TBCC engine subsystem does not provide high-pressure gas, and to introduce low-pressure gas from the TBCC engine subsystem to perform air circulation cooling on the engine compartment, landing gear compartment and cockpit; and, The aircraft accessory subsystem works under the auxiliary power or emergency power, or is driven by extracting the mechanical energy output by the TBCC engine subsystem, or is driven by the working gas introduced from the outside.

2. The aircraft energy and thermal integrated management system according to claim 1, characterized in that: The TBCC engine subsystem includes a turbofan engine and a ramjet engine; During the takeoff phase, both the turbofan and ramjet engines are inoperative; During the climbing operation phase and the restart phase, only the turbofan engine is working, the low-pressure gas at the outlet of the low-pressure compressor of the turbofan engine is introduced into the environmental control subsystem through the low-pressure bleed air pipeline, and the high-pressure gas at the outlet of the high-pressure compressor of the turbofan engine is introduced into the environmental control subsystem through the high-pressure bleed air pipeline. The mechanical energy generated by the high-pressure turbine of the turbofan engine is transmitted to the aircraft accessory subsystem through the mechanical transmission shaft; During the TBCC engine mode conversion phase, the turbofan engine is shut down and the ramjet engine is started; During the ramjet engine's independent operation phase, the operating gas introduced from the outside is transferred to the aircraft accessory subsystem via the ramjet bleed air pipeline; During the unpowered descent phase, both the turbofan engine and the ramjet engine do not work.

3. The aircraft energy and thermal integrated management system according to claim 2, characterized in that: The fuel thermal management subsystem does not work during the takeoff phase; during the climb operation phase, the TBCC engine mode conversion phase and the restart phase, the onboard heat source is cooled, and the fuel after the onboard heat source is cooled is cooled by ram air and liquid nitrogen; during the ramjet engine single operation phase and the unpowered descent phase, the onboard heat source is cooled, and the fuel after the onboard heat source is cooled is cooled by liquid nitrogen.

4. The aircraft energy and thermal integrated management system according to claim 3, characterized in that: The fuel thermal management subsystem includes: A ram air cooling circuit, in which the fuel in the main fuel tank flows through the first heat exchanger, the second heat exchanger and the fourth heat exchanger in sequence for circulated heat exchange, wherein the first heat exchanger and the second heat exchanger are liquid-liquid heat exchangers, and the fourth heat exchanger is a liquid-gas heat exchanger using ram air as a cooling medium; and, Liquid nitrogen cooling circuit, the fuel in the main fuel tank flows through the first heat exchanger, the second heat exchanger and the third heat exchanger in sequence, and a part of the fuel is pumped to each combustion chamber in the TBCC engine subsystem, and a part of the fuel enters the fifth heat exchanger for cooling and is pumped back to the main fuel tank through the sixth heat exchanger. The third heat exchanger is a liquid-liquid heat exchanger, the fifth heat exchanger is a liquid-liquid heat exchanger with liquid nitrogen as the cooling medium, and the sixth heat exchanger is a liquid-gas heat exchanger. The engine heat load is connected to the third heat exchanger to form a liquid-liquid heat exchange circuit.

5. The aircraft energy and thermal integrated management system according to claim 4, characterized in that: The environmental control subsystem includes: an auxiliary power unit, which provides auxiliary power after introducing working gas from the outside during the take-off phase; provides auxiliary power after introducing high-pressure gas from the TBCC engine subsystem during the climbing working phase and the restart phase, and introduces low-pressure gas from the TBCC engine subsystem to perform air circulation cooling on the engine compartment, the landing gear compartment and the cockpit; Emergency power unit, providing emergency power during TBCC engine mode conversion and power-off descent; and, The liquid nitrogen cascade utilization device is used to perform cascade liquid nitrogen cooling on the engine compartment, landing gear compartment and cockpit during the TBCC engine mode conversion stage, the ramjet engine independent operation stage and the unpowered descent stage.

6. The aircraft energy and thermal integrated management system according to claim 5, characterized in that: The auxiliary power unit includes: a power turbine, a high-speed transmission shaft, a cooling turbine, an inspiration motor, a low-pressure fan, an auxiliary power unit combustion chamber and an auxiliary fuel tank; During the takeoff phase, the initiator motor is in the starter mode, the initiator motor drives the low-pressure fan to work, the working gas introduced from the outside is transported to the auxiliary power unit combustion chamber through the low-pressure fan, the fuel in the auxiliary fuel tank is transported to the auxiliary power unit combustion chamber, the fuel is burned in the auxiliary power unit combustion chamber, and the high-temperature and high-pressure gas generated drives the power turbine to rotate; During the climbing operation phase and the restart phase, the high-pressure gas introduced from the TBCC engine subsystem is transported to the auxiliary power unit combustion chamber, and the fuel in the auxiliary fuel tank is transported to the auxiliary power unit combustion chamber. The high-temperature and high-pressure gas generated by the combustion of the fuel in the auxiliary power unit combustion chamber drives the power turbine. The low-pressure gas introduced from the TBCC engine subsystem passes through the low-pressure fan and is cooled by the sixth heat exchanger and the seventh heat exchanger in turn. The gas cooled by the seventh heat exchanger enters the cooling turbine and the power turbine. The cooling gas passing through the power turbine performs air circulation cooling on the engine compartment, the landing gear compartment and the cockpit. The sixth heat exchanger is a liquid-to-gas heat exchanger, and the seventh heat exchanger is an air-to-gas heat exchanger.

7. The aircraft energy and thermal integrated management system according to claim 5, characterized in that: The emergency power unit includes: a power turbine, an auxiliary power unit combustion chamber and emergency fuel. During the TBCC engine mode conversion stage and the unpowered descent stage, the emergency fuel is transported to the auxiliary power unit combustion chamber, and the emergency fuel is burned in the auxiliary power unit combustion chamber. The high-temperature and high-pressure gas generated drives the power turbine to rotate.

8. The aircraft energy and thermal integrated management system according to claim 5, characterized in that: The liquid nitrogen cascade utilization device comprises: a liquid nitrogen storage tank and a liquid nitrogen booster pump. The low-temperature liquid nitrogen in the liquid nitrogen storage tank is pressurized by the liquid nitrogen booster pump and pumped to the cockpit, and the nitrogen in the cockpit after heating is sent to the landing gear compartment. A stream of low-temperature liquid nitrogen is diverted from the liquid nitrogen booster pump and pumped into the landing gear compartment, and a stream of low-temperature liquid nitrogen is diverted from the liquid nitrogen booster pump and pumped into the engine compartment.

9. An aircraft energy and thermal integrated management system according to any one of claims 1 to 8, characterized in that: The aircraft accessory subsystem includes: an aircraft accessory casing and a ramjet turbine, During the take-off phase, the auxiliary power provided by the auxiliary power unit is transmitted to the aircraft accessory casing, and the aircraft accessory casing outputs electrical energy to start the main engine; During the climbing operation phase and the restart phase, the auxiliary power provided by the auxiliary power unit is transmitted to the aircraft accessory casing, the mechanical energy extracted from the TBCC engine subsystem is transmitted to the aircraft accessory casing, the aircraft accessory casing outputs electrical energy, starts the main engine and drives the hydraulic system to work; During the TBCC engine mode conversion stage and the unpowered descent stage, the emergency power provided by the auxiliary power unit is transmitted to the aircraft accessory casing, and the aircraft accessory casing outputs electrical energy to start the main engine and drive the hydraulic system to work; During the ramjet engine's independent working stage, the working gas introduced from the outside drives the ramjet turbine. The mechanical energy generated by the ramjet turbine is transmitted to the aircraft accessory casing. The aircraft accessory casing outputs electrical energy, starts the main engine and drives the hydraulic system to work.

10. An aircraft energy and thermal integrated management system according to claim 9, characterized in that: During the TBCC engine mode conversion stage and the unpowered descent stage, batteries are used to supply power to output electrical energy.

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