An aircraft energy and thermal integrated management system

Abstract

The application discloses a kind of aircraft energy and heat comprehensive management system, belong to control, adjusting technical field.The system integrates turbine-based combined cycle engine subsystem, fuel heat management subsystem, environmental control subsystem, aircraft accessory subsystem, through from TBCC engine subsystem carry out air bleeding and energy extraction, coordinate the joint work between each subsystem, realize whole machine energy distribution and heat management regulation and control, to respond different engine working state and energy efficiency demand, system has stage regulation and mode switching capability, can be according to the multiple flight conditions in flight envelope including take-off, climb, ram cruise, return carry out power distribution and heat management regulation and control, greatly enrich the selection of effective heat sink and improve heat sink utilization efficiency, realize the temperature of optimal regulation and control different position heat source, effectively reduce weight compensation, improve the maneuverability and long endurance capability of aircraft and reduce flight cost.

Description

TECHNICAL FIELD

[0001] The present application relates to the energy distribution, thermal energy management and environmental control technology of hypersonic aircraft, and discloses an aircraft energy and heat comprehensive management system, in particular to an integrated energy and heat management system of a hypersonic aircraft based on a turbine-based combined cycle (TBCC) engine, and belongs to the technical field of control and regulation. BACKGROUND

[0002] With the development of aircraft towards high power density, long endurance and stealth, the load of energy and heat management system increases, and the traditional distributed management is difficult to realize rapid response and efficient operation. The national defense demand puts forward higher requirements for aircraft stealth, maneuverability and supersonic cruise, etc. Under this background, the aircraft energy / heat management problem is increasingly prominent. Facing the continuous increase of flight Mach number, the use of composite materials, the reduction of effective heat sink, and the many demands of next-generation aircraft maneuverability, stealth, maintainability, and power equipment power consumption and heat dissipation, the traditional distributed airborne electrical system and single energy management method cannot meet the requirements. At present, the main technologies applied in the field of aircraft heat control and energy regulation at home and abroad include air cycle cooling technology, liquid loop cooling technology, auxiliary power unit (APU) power generation technology, and ram turbine power generation technology.

[0003] Air cycle cooling technology is widely used in various types of aircraft environmental control systems, but due to its low refrigeration performance coefficient, the system has poor reliability in ground working conditions. In addition, the introduction of external ram air leads to certain limitations on the use of aircraft at high Mach numbers, which contradicts the characteristics of high hypersonic aircraft aerodynamic heating, high flight altitude and speed.

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

[0005] An APU is a small gas turbine engine that operates independently of the main engine and is used to provide electrical power for ground operations and emergency situations. The APU can operate independently on the ground or in flight, providing flexible electrical power on the ground or in the air, and can still provide some electrical power when the main engine fails. However, due to the small size of the APU, its power generation capacity is limited and cannot meet the needs of all systems, and the high fuel consumption characteristic makes the APU usually used to provide electrical power for the starting stage of the aircraft.

[0006] The ram turbine generator technology drives a small turbine to drive a generator by extracting air from the main engine, which has high power generation efficiency and can utilize the waste heat or residual energy of the main engine to improve energy utilization. Compared with traditional generators, turbine generators are more compact and lightweight, but they require a special air extraction and management system, increasing design complexity and affecting the performance of the main engine, and the extracted air may have some impact on the performance and efficiency of the main engine.

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

[0008] Increased system complexity: Due to the deep coupling of energy and thermal management, the complexity of system integration increases significantly, requiring higher design and maintenance requirements, and traditional energy-thermal management integrated systems of hypersonic vehicles cannot balance the needs of system performance and structural compactness;

[0009] Insufficient dynamic adaptability: The existing energy-thermal management integrated system of hypersonic vehicles has weak dynamic response capability to environmental changes during flight missions, making it difficult to achieve efficient real-time adjustment, especially in complex flight conditions, which can easily cause system overload or control lag;

[0010] Low efficiency of thermal load distribution: The current energy-thermal management integrated system of hypersonic vehicles has insufficient ability to distribute and optimize the heat dissipation path of thermal loads in high heat flux conditions, leading to local overheating and thermal stress concentration, affecting the overall performance and structural life of the aircraft;

[0011] Energy utilization rate limitation: In energy supply and demand balance and multi-task switching, the existing energy-thermal management integrated system of hypersonic vehicles cannot fully utilize energy, which can easily cause waste or insufficient power supply problems.

[0012] Except for the above, the existing hypersonic aircraft energy-thermal management integrated system often considers people or equipment, is relatively independent, and has not been arranged as a whole. The energy and thermal management system of the hypersonic aircraft not only needs to provide a good and comfortable working environment for personnel, but also needs to provide power for electronic equipment and a safe and reliable working environment for other equipment, which puts forward more stringent requirements for the design of the energy-thermal management integrated system of the hypersonic aircraft. The weight, performance, cost, reliability, compensation loss and other factors should be considered, and the thermal management system, energy regulation system and aircraft engine propulsion system should be designed together.

[0013] Therefore, the energy-thermal management integrated system needs to further break through the existing technical bottleneck, improve the intelligence, reliability and adaptability of the system, and provide strong support for the development of the new generation of aircraft. How to provide a comprehensive energy and thermal management system for the hypersonic aircraft to realize the rated energy supply and solve the distributed heat load heat dissipation is a technical problem to be solved by the technical personnel in the field. SUMMARY

[0014] The present application aims to overcome the above-mentioned problems in the prior art, and provides an aircraft energy and thermal comprehensive management system. The energy-thermal management integrated system is proposed to meet the energy demand and thermal management demand of the hypersonic aircraft during the completion of the flight mission envelope. The two difficult problems of waste heat treatment and effective energy utilization during flight are solved. The whole machine energy distribution and thermal management are regulated to cope with different engine working states and energy efficiency demands. The power distribution and thermal management regulation are carried out according to the multiple flight working conditions in the flight envelope including take-off, climbing, cruising and returning. The utilization efficiency of the heat sink is improved, the temperature of the heat source at different positions is optimally regulated, the weight compensation is effectively reduced, the aircraft maneuverability and long endurance capability are improved, and the flight cost is reduced.

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

[0016] The aircraft energy and heat comprehensive management system comprises a TBCC engine subsystem, a fuel heat management subsystem, an environmental control subsystem and an aircraft accessory subsystem; the TBCC engine subsystem is used for working on the gas introduced from outside to generate mechanical energy; the fuel heat management subsystem is used for cooling the on-board heat source and cooling the fuel after cooling the on-board heat source by means of ram air cooling and / or liquid nitrogen cooling to supply the fuel to the TBCC engine subsystem; the environmental control subsystem is used for providing auxiliary power after introducing the high-pressure gas from the TBCC engine subsystem or introducing the working condition gas from outside, providing emergency power when the TBCC engine subsystem does not provide the high-pressure gas, and introducing the low-pressure gas from the TBCC engine subsystem to perform air circulation cooling on the engine cabin, the landing gear cabin and the cockpit; and the aircraft accessory subsystem works under the action of the auxiliary power or the emergency power, or is driven by extracting the mechanical energy output by the TBCC engine subsystem, or is driven by the working condition gas introduced from outside.

[0017] As a further optimization scheme of the aircraft energy and heat comprehensive management system, the TBCC engine subsystem comprises a turbofan engine and a ramjet engine; in the take-off stage, neither the turbofan engine nor the ramjet engine works; in the climb working stage and the restart stage, only the turbofan engine works, the low-pressure gas at the low-pressure gas outlet of the low-pressure compressor in the turbofan engine is introduced into the environmental control subsystem through a low-pressure bleed air pipeline, the high-pressure gas at the high-pressure gas outlet of the high-pressure compressor in the turbofan engine is introduced into the environmental control subsystem through a high-pressure bleed air pipeline, and the mechanical energy generated by the high-pressure turbine in the turbofan engine is transmitted to the aircraft accessory subsystem through a mechanical transmission shaft; in the TBCC engine mode conversion stage, the turbofan engine is shut down and the ramjet engine is started; in the ramjet engine alone working stage, the working condition gas introduced from outside is transmitted to the aircraft accessory subsystem through a ramjet bleed air pipeline; and in the unpowered glide stage, neither the turbofan engine nor the ramjet engine works.

[0018] As a further optimization scheme of the aircraft energy and heat comprehensive management system, the fuel heat management subsystem: does not work in the take-off stage; cools the on-board heat source and cools the fuel after cooling the on-board heat source by means of ram air cooling and liquid nitrogen cooling in the climb working stage, the TBCC engine mode conversion stage and the restart stage; cools the on-board heat source and cools the fuel after cooling the on-board heat source by means of liquid nitrogen cooling in the ramjet engine alone working stage and the unpowered glide stage.

[0019] As a further optimization scheme of the aircraft energy and heat integrated management system, the fuel heat management subsystem comprises: a ram air cooling circuit and a liquid nitrogen cooling circuit; the ram air cooling circuit, fuel in the main fuel tank is circulated and heat-exchanged after flowing through the first heat exchanger, the second heat exchanger and the fourth heat exchanger in turn, 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; the liquid nitrogen cooling circuit, fuel in the main fuel tank is circulated and heat-exchanged after flowing through the first heat exchanger, the second heat exchanger and the third heat exchanger in turn, a part of the fuel is pumped to each combustion chamber in the TBCC engine subsystem, and a part of the fuel is cooled in the fifth heat exchanger and 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, 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 of the aircraft energy and heat integrated management system, the environmental control subsystem comprises: an auxiliary power device, an emergency power device and a liquid nitrogen cascade utilization device; the auxiliary power device provides auxiliary power after introducing working condition gas from the outside in the take-off stage, provides auxiliary power after introducing high-pressure gas from the TBCC engine subsystem in the climbing working stage and the restart stage, and circulates and cools the engine cabin, the landing gear cabin and the cockpit by introducing low-pressure gas from the TBCC engine subsystem; the emergency power device provides emergency power in the TBCC engine mode conversion stage and the unpowered glide stage; and the liquid nitrogen cascade utilization device performs cascade liquid nitrogen cooling on the engine cabin, the landing gear cabin and the cockpit in the TBCC engine mode conversion stage, the ramjet single working stage and the unpowered glide stage.

[0021] As a further optimization scheme of the aircraft energy and heat integrated management system, the auxiliary power device comprises a power turbine, a high-speed transmission shaft, a cooling turbine, an auxiliary generator, a low-pressure fan, an auxiliary power device combustion chamber and a sub-fuel tank; in the take-off stage, the auxiliary generator is in the starting mode, the auxiliary generator drives the low-pressure fan to work, the working condition gas introduced from the outside is transported to the auxiliary power device combustion chamber through the low-pressure fan, the fuel in the sub-fuel tank is transported to the auxiliary power device combustion chamber, the fuel is burned in the auxiliary power device combustion chamber, and the high-temperature and high-pressure gas generated drives the power turbine to rotate; in the climbing working stage and the restart stage, the high-pressure gas introduced from the TBCC engine subsystem is transported to the auxiliary power device combustion chamber, the fuel in the sub-fuel tank is transported to the auxiliary power device combustion chamber, the fuel is burned in the auxiliary power device combustion chamber to generate high-temperature and high-pressure gas to drive the power turbine, and the low-pressure gas introduced from the TBCC engine subsystem passes through the low-pressure fan, then passes through the sixth heat exchanger and the seventh heat exchanger in sequence and is cooled, the gas cooled by the seventh heat exchanger enters the cooling turbine and the power turbine, the cooling gas passing through the power turbine is used for air circulation cooling of the engine cabin, the landing gear cabin and the cockpit, the sixth heat exchanger is a liquid-gas heat exchanger, and the seventh heat exchanger is a gas-gas heat exchanger.

[0022] As a further optimization scheme of the aircraft energy and heat integrated management system, the emergency power device comprises a power turbine, an auxiliary power device combustion chamber and emergency fuel, in the TBCC engine mode conversion stage and the unpowered glide stage, the emergency fuel is transported to the auxiliary power device combustion chamber, the emergency fuel is burned in the auxiliary power device combustion chamber, and the high-temperature and high-pressure gas generated drives the power turbine to rotate.

[0023] As a further optimization scheme of the aircraft energy and heat integrated management system, the liquid nitrogen step 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 pumped to the cockpit after being pressurized by the liquid nitrogen booster pump, the nitrogen gas in the cockpit after being warmed is sent into the landing gear cabin, a part of the low-temperature liquid nitrogen is pumped into the landing gear cabin from the liquid nitrogen booster pump, and a part of the low-temperature liquid nitrogen is pumped into the engine cabin from the liquid nitrogen booster pump.

[0024] As a further optimization scheme of an aircraft energy and heat integrated management system, the aircraft accessory subsystem comprises: an aircraft accessory gearbox, a ram turbine; in the take-off stage, auxiliary power provided by an auxiliary power device is transmitted to the aircraft accessory gearbox, the aircraft accessory gearbox outputs electric energy, and the main engine is started; in the climbing working stage and the restart stage, auxiliary power provided by the auxiliary power device is transmitted to the aircraft accessory gearbox, mechanical energy extracted from the TBCC engine subsystem is transmitted to the aircraft accessory gearbox, the aircraft accessory gearbox outputs electric energy, the main engine is started, and the hydraulic system is driven to work; in the TBCC engine mode conversion stage and the unpowered glide stage, emergency power provided by the auxiliary power device is transmitted to the aircraft accessory gearbox, the aircraft accessory gearbox outputs electric energy, the main engine is started, and the hydraulic system is driven to work; in the ramjet engine single working stage, working condition gas introduced from outside drives the ram turbine, mechanical energy generated by the ram turbine is transmitted to the aircraft accessory gearbox, the aircraft accessory gearbox outputs electric energy, the main engine is started, and the hydraulic system is driven to work.

[0025] As a further optimization scheme of an aircraft energy and heat integrated management system, in the TBCC engine mode conversion stage and the unpowered glide stage, a storage battery is used to supply and output electric energy.

[0026] The application has the following beneficial effects by adopting the above technical scheme:

[0027] (1) The improved ultrasonic speed aircraft integrated energy and heat management system aims to adapt to the flight envelope of a hypersonic aircraft as the optimal design principle, deeply couples energy management and heat management functions, and improves the efficiency, response speed and reliability of the overall system through integrated design.

[0028] (2) Compared with the traditional aircraft heat management mode, the improved ultrasonic speed aircraft integrated energy and heat management system increases the combined cooling scheme for distributed large heat sources through the fuel heat management subsystem and the environmental control subsystem, adopts a combination of multiple heat sinks such as bleed air cooling, fuel cooling and liquid nitrogen cascade cooling, can match the corresponding cooling mode according to the change of the thermal load in the flight envelope, and the improved air / fluid circulation loop can greatly improve the cooling efficiency of the aircraft in different stages of generating thermal load and improve the overall heat dissipation efficiency.

[0029] (3) The improved hypersonic aircraft integrated energy and thermal management system of the application proposes a new main shaft power extraction method, i.e. turbofan engine energy extraction, and an integrated auxiliary-emergency power generation method, and the main shaft power extraction, ram turbine power generation, APU power generation and battery power output form an energy closed loop, so that the extraction of electric energy can be realized in each flight stage according to the energy demand in the flight envelope, the flight preparation mass is greatly reduced, more diverse and feasible power generation methods are realized, the demand of more high-power electrical equipment is met, and the power generation efficiency is greatly improved. BRIEF DESCRIPTION OF DRAWINGS

[0030] In order to more clearly illustrate the technical solutions in the embodiments of the application or the prior art, the drawings needed to be used in the embodiments will be briefly introduced below. Obviously, the drawings in the following description only constitute some embodiments of the application, and for those skilled in the art, other drawings can also be obtained from these drawings without creative labor.

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

[0032] Figure 2 It is a schematic diagram of the working state of the hypersonic aircraft integrated energy and thermal management system of the application in the take-off stage.

[0033] Figure 3 It is a schematic diagram of the working state of the hypersonic aircraft integrated energy and thermal management system of the application in the climb stage.

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

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

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

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

[0038] Explanations of the figure: 1, air inlet, 2, low pressure compressor, 3, high pressure compressor, 4, turbofan engine combustion chamber, 5, bypass duct, 6, high pressure turbine, 7, low pressure turbine, 8, gas mixing chamber, 9, afterburner, 10, low speed nozzle, 11, isolator, 12, ramjet combustion chamber, 13, high speed nozzle, 14, main fuel tank, 15, aerodynamic skin heat transfer, 16, engine bay, 17, landing gear bay, 18, cockpit, 19, liquid nitrogen tank, 20, low pressure fan, 21, starter generator, 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 gearbox, 29, external environment, 30, battery, 31, electrical energy, 32, hydraulic energy, 33, ram turbine, I1, low pressure bleed air line, I2, high pressure bleed air line, I3, mechanical transmission shaft, I4, ram bleed air line, B1, fuel boost pump, B2, liquid nitrogen boost pump, Z1, electronic equipment thermal load, Z2, oil-hydraulic oil thermal load, Z3, engine thermal load, RC1, ram air, RC2, liquid nitrogen, HR1, first heat exchanger, HR2, second heat exchanger, HR3, third heat exchanger, HR4, fourth heat exchanger, HR5, fifth heat exchanger, HR6, sixth heat exchanger, HR7, seventh heat exchanger, V1, first fuel flow regulating valve, V2, second fuel flow regulating valve, V3, third fuel flow regulating valve, V4, electronic equipment regulating valve, V5, oil-hydraulic oil regulating valve, V6, engine thermal load flow regulating valve, V7, ram air flow regulating valve, V8, liquid nitrogen flow regulating valve, V9, fourth fuel flow regulating valve, V10, fifth fuel flow regulating valve, V11, first gas regulating valve, V12, second gas regulating valve, V13, third gas regulating valve, V14, fourth gas regulating valve, V15, fifth gas regulating valve, V16, sixth gas regulating valve, V17, seventh gas regulating valve, V18, eighth gas regulating valve, V19, ninth gas regulating valve, V20, sixth fuel flow regulating valve, V21, seventh fuel flow regulating valve, V22, eighth fuel flow regulating valve, V23, ninth fuel flow regulating valve. DETAILED DESCRIPTION

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

[0040] As Figure 1As shown, the present invention discloses an integrated energy and thermal management system for hypersonic aircraft based on a TBCC engine, which consists of a TBCC engine subsystem, a fuel thermal management subsystem, an environmental control subsystem, and an aircraft accessory subsystem. The TBCC engine subsystem performs bleed air and energy extraction, while coordinating the joint work between the various subsystems to realize the energy distribution and thermal management regulation of the aircraft during flight missions, so as to meet the energy efficiency requirements of different engine operating states. During flight, the various subsystems work together through pipelines and mechanical shafts to complete gas delivery and mechanical energy output. They extract bleed air and electrical energy from the TBCC engine subsystem to meet the thermal management requirements of the aircraft's heat-generating components and the electrical energy requirements during flight. This achieves optimal control of electrical energy distribution and thermal energy management for hypersonic aircraft, improving the aircraft's maneuverability and long-endurance capability. Along the flight profile, it can cover the energy distribution and thermal management control requirements of the hypersonic flight envelope during takeoff, climb, mode transition, ramjet cruise, unpowered taxiing, and restart flight phases. Among these, the energy requirements are reflected in the power output of electrical and hydraulic energy, while the thermal management requirements are reflected in the multi-heat source matching ramjet air cooling, fuel cooling, and liquid nitrogen cooling scheme.

[0041] The TBCC engine subsystem is the primary power source for the entire energy management system. For example... Figure 1As shown, the TBCC engine subsystem adopts a parallel ramjet engine, including: a fanjet engine composed of an air inlet 1, a low-pressure compressor 2, a high-pressure compressor 3, a fanjet engine combustion chamber 4, a high-pressure turbine 6, a low-pressure turbine 7, an outer bypass 5, a gas mixing chamber 8, an afterburner 9, and a low-speed nozzle 10; and a ramjet engine composed of an air inlet 1, a separation section 11, a ramjet engine combustion chamber 12, and a high-speed nozzle 13. When the fanjet 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 bypass 5, the high-temperature and high-pressure gas at the outlet of the high-pressure compressor 3 is sent to the fanjet engine combustion chamber 4 for combustion, the high-temperature and high-pressure gas generated by the fanjet engine combustion chamber 4 drives the high-pressure turbine 6 and the low-pressure turbine 7 to rotate, the gas after expansion through the low-pressure turbine 7 and the low-pressure gas output by the outer bypass 5 are jointly sent to the gas mixing chamber 8, 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 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 after entering the air inlet 1 and is sent to the separation section 11, the high-temperature and high-pressure gas at the outlet of the separation section 11 is sent to the ramjet engine combustion chamber 12, the high-temperature and high-pressure gas generated by the ramjet engine combustion chamber 12 is sent to the high-speed nozzle 13 to generate thrust. During flight, the ramjet engine and the fanjet engine can be used alone or reused according to the flight condition, and have the ability to quickly realize mode switching, wherein the low-pressure bleed air pipeline I1 introduces the low-pressure gas at the gas outlet of the low-pressure compressor 2 to the environmental control subsystem; the high-pressure bleed air pipeline I2 introduces the high-pressure gas at the gas outlet of the high-pressure compressor 3 to 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; and the ramjet bleed air pipeline I4 sends 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, a pneumatic skin heat transfer 15, a fuel booster pump B1, an electronic device thermal load Z1, an 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 device regulating valve V4, an oil-hydraulic oil regulating valve V5, an engine thermal load flow regulating valve V6, a ram air flow regulating valve V7, a liquid nitrogen flow regulating valve V8, a fourth fuel flow regulating valve V9, a fifth fuel flow regulating valve V10, a seventh fuel flow regulating valve V21, an eighth fuel flow regulating valve V22, a ninth fuel flow regulating valve V23, wherein 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 both liquid-gas heat exchangers. The main fuel tank 14 is connected to the pneumatic skin heat transfer 15, indicating that the pneumatic heat during flight acts on the main fuel tank 14. The fuel is pressurized by the booster pump B1 and then sequentially passes through the first heat exchanger HR1 and the second heat exchanger HR2 for heat exchange treatment. The electronic device thermal load Z1 is connected to the first heat exchanger HR1, and the oil-hydraulic oil thermal load Z2 is connected to the second heat exchanger HR2. After heat exchange is completed, the fuel is divided into two streams. One stream of fuel passes through the fourth heat exchanger HR4 and the ram air RC1 for heat exchange to complete cooling. The cooled fluid is again involved in the heat exchange circuit of the first heat exchanger HR1 to form a HR1-HR2-HR4-HR1 heat exchange circuit. When the flow in the HR1-HR2-HR4-HR1 heat exchange circuit increases, the cooling speed and cooling effect are greatly improved. Calculation shows that when the fuel flow is 0.5 kg / s, the cooling effect of the HR1-HR2-HR4-HR1 heat exchange circuit is improved by 6%. The other stream of fuel directly flows through the third heat exchanger HR3 and the engine thermal load Z3 to form a liquid-liquid heat exchange circuit for heat exchange. Part of the fuel after heat exchange in the liquid-liquid heat exchange circuit is directly delivered to the ramjet 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. Another part of the fuel after heat exchange in the liquid-liquid heat exchange circuit is returned to the fifth heat exchanger HR5 for cooling with liquid nitrogen RC2. The low-temperature cooled fuel can be regulated by the fourth fuel flow regulating valve V9 and the fifth fuel flow regulating valve V10 to control the fuel flow involved in the circulation and the flow returned to the main fuel tank 14. According to different operating conditions of the TBCC engine subsystem from low Mach to high Mach, it can be realized that ram air RC1 or liquid nitrogen RC2 is adaptively selected as the optimal cooling medium, and the combined cooling method can also be used to improve the utilization efficiency of the heat sink.The system adjusts the fuel shunt through the eighth fuel flow regulating valve V22 before the engine heat load Z3, and through the ninth fuel flow regulating valve V23, the cooling flow exchanged with the engine can be effectively controlled, the fuel outlet temperature after the heat exchange with the engine is maintained in the effective safety temperature range of 110℃, the combustion efficiency of the fuel entering the turbofan engine combustion chamber 4 and the afterburner 9 can be improved, and after calculation, compared with the fuel temperature controlled at 110℃ entering the turbofan engine combustion chamber 4 inlet and the afterburner 9 inlet, not only the fuel has the highest cooling efficiency before reaching the coking temperature, but also the combustion efficiency can be improved by 9.45% compared with the fuel directly entering the combustion chamber at a basic temperature of 60℃.

[0042] As shown in Figure 1 The ring control subsystem includes an auxiliary-emergency power device and a liquid nitrogen cascade utilization device, and the ring control subsystem is connected with the low-pressure bleed air pipeline I1 and the high-pressure bleed air pipeline I2 respectively. The auxiliary-emergency power device can provide the power demand of the aircraft during the take-off stage and the emergency restart, can greatly save the liquid nitrogen consumption under the condition of ensuring the utilization efficiency of the heat sink, and enhances the endurance time of the aircraft.

[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℃, and the low-temperature liquid nitrogen is used to cascade cool the cockpit 18, the landing gear compartment 17 and the engine compartment 16 in turn by using the large latent heat of vaporization and excellent cooling capacity of the low-temperature liquid nitrogen. The low-temperature liquid nitrogen is pumped to the cockpit 18 after being pressurized by the liquid nitrogen booster pump B2, and the cockpit is cooled to below 25℃ by spray cooling in the cockpit 18. At this time, the outlet temperature of the cockpit 18 is 25℃, and the gaseous liquid nitrogen after being warmed up can still be introduced into the landing gear compartment 17. At this time, the cooling efficiency of the gaseous liquid nitrogen after being warmed up is low, so a low-temperature liquid nitrogen is introduced from the liquid nitrogen booster pump B2 to assist in cooling the landing gear compartment 17. The landing gear compartment 17 is cooled to 110℃, and the high-temperature gas is discharged. A low-temperature liquid nitrogen is introduced from the liquid nitrogen booster pump B2 to spray cool the engine compartment 16, and the engine compartment temperature is cooled to 110℃, and the high-temperature gas is discharged. The fifth gas regulating valve V15 is used to control the outlet flow of the low-temperature liquid nitrogen in the liquid nitrogen storage tank 19. According to the calculation, compared with the scheme without cascade cooling, the scheme using cascade cooling can save 13% of the liquid nitrogen consumption, and can effectively save the liquid nitrogen consumption.

[0044] The auxiliary-emergency power device comprises: a low-pressure fan 20, an exciting generator 21, a cooling turbine 22, a high-speed transmission shaft 23, a power turbine 24, emergency fuel 25, an auxiliary power device combustion chamber 26, a sub-fuel tank 27, the power turbine 24, the high-speed transmission shaft 23, the cooling turbine 22, the exciting generator 21, the low-pressure fan 20, the auxiliary power device combustion chamber 26, and the sub-fuel tank 27 constitute an auxiliary power device, and the power turbine 24, the auxiliary power device combustion chamber 26, and the emergency fuel 25 constitute an emergency power device. During the starting stage of the aircraft, the exciting generator 21 is in an exciting generator state, the exciting generator 21 drives the low-pressure fan 20 to work, and the gas sucked from the outside is transported into the auxiliary power device combustion chamber 26 through the sixth gas regulating valve V16, the low-pressure fan 20, and the seventh gas regulating valve V17, and at the same time, the sub-fuel tank 27 also transports fuel into the auxiliary power device 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 device combustion chamber 26 drives the power turbine 24 to rotate, thereby driving the auxiliary power device to work, and specifically: the mechanical properties generated by the rotation of the power turbine 24 are transmitted to each power-consuming component including the cooling turbine 22, the exciting generator 21, and the low-pressure fan 20 through the high-speed transmission shaft 23, and at the same time, mechanical energy is provided for the aircraft accessory casing 28 in the aircraft accessory subsystem; the high-pressure gas directly introduced into the high-pressure compressor 3 outlet through the high-pressure bleed air channel I2 enters the auxiliary power device combustion chamber 26 through the seventh gas regulating valve V17, the emergency fuel 25 is transported into the auxiliary power device 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 device combustion chamber 26 drives the power turbine 24 to rotate to do work, drives the emergency power device to work, and at the same time, provides mechanical energy for the aircraft accessory casing 28 in the aircraft accessory subsystem, and the eighth gas regulating valve V18 is used for adjusting the gas flow of the high-pressure gas entering the auxiliary power device combustion chamber 26.

[0045] Meanwhile, in the auxiliary-emergency power device, the low-pressure bleed air pipeline I1 delivers the low-pressure gas at the outlet of the low-pressure compressor 2 to the low-pressure fan 20 to drive it to rotate, and the low-pressure gas at the outlet of the low-pressure compressor 2 is sequentially cooled by the sixth heat exchanger HR6 and the seventh heat exchanger HR7 after passing through the low-pressure fan 20, the sixth heat exchanger HR6 is a liquid-gas heat exchanger, and the seventh heat exchanger HR7 is a gas-gas heat exchanger, the gas cooled by the seventh heat exchanger HR7 enters the cooling turbine 22 and the power turbine 24, and the gas passing through the power turbine 24 is delivered to the engine compartment 16, the landing gear compartment 17 and the cockpit 18 in a parallel manner through the second gas regulating valve V12, the third gas regulating valve V13 and the fourth gas regulating valve V14, and the gas is heated by the engine compartment 16, the landing gear compartment 17 and the cockpit 18, and then passes through the seventh heat exchanger HR7 to complete the closed heat exchange cycle, and the bleed air flow participating in the closed heat exchange cycle is adjusted and controlled by the first gas regulating valve V11, and this process is carried out in the climbing stage of the aircraft at a low Mach number, providing the required electric energy in the take-off stage of the aircraft, meeting the refrigeration demand in the low-speed flight stage and the electric energy demand in the emergency restart of the aircraft.

[0046] As shown in Figure 1 The aircraft accessory subsystem includes an aircraft accessory casing 28, a ram turbine 33 and a ninth gas regulating valve V19. When the ram engine is working, the working condition gas from the ram bleed air pipeline I4 drives the ram turbine 33 to rotate at high speed, and the mechanical energy generated by the ram turbine 33 is delivered to the aircraft accessory casing 28, and the gas expanded by the ram 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, and the mechanical energy from the mechanical transmission shaft I3 and the power turbine 24 also 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 embodied as the output of hydraulic energy 32. The aircraft accessory subsystem can directly output electric energy 31 by using the storage battery 30 to supply power according to different flight conditions, achieving efficient energy distribution effect.

[0047] The integrated energy and thermal management system based on the TBCC hypersonic aircraft engine can switch system modes, extract energy, or perform cooling cycles according to different flight conditions. All of these can be kept within a controllable range by adjusting the valve openings. The entire system contains 21 regulating valves, each used to switch system modes and control the energy and thermal management system to meet the corresponding energy and cooling requirements in each mode. A valve opening of 0 indicates a closed state, while an opening of 1 indicates a fully open state. Based on the aircraft's flight envelope and energy requirements, the system sequentially completes energy distribution and thermal management control along the flight profile during takeoff, climb, TBCC engine mode transition, ramjet engine standalone operation, unpowered descent, and turbine restart. Within the flight envelope, the energy management system should have adaptive adjustment and mode switching capabilities to cope with different engine operating states and energy efficiency requirements.

[0048] Table 1 shows the operating status and mode switching capabilities of the integrated energy and thermal management system for hypersonic aircraft based on TBCC engines at each stage.

[0049]

[0050] Table 1

[0051] like Figure 2 As shown, the integrated energy and thermal management system for hypersonic aircraft based on the TBCC engine operates only in the auxiliary power unit and aircraft accessory subsystems during takeoff. When the system is in ground start-up mode, the starter motor 21 is in starter mode and powered by its own power supply. The starter motor 21 drives the low-pressure fan 20. Under ground conditions, external air is drawn in sequentially through the sixth gas regulating valve V16, the low-pressure fan 20, and the seventh gas regulating valve V17 before entering the combustion chamber 26 of the auxiliary power unit and mixing with fuel from the auxiliary fuel tank 27 for combustion. The high-temperature and high-pressure gas generated after combustion enters the power turbine 24 and expands, driving the power turbine 24 to rotate at high speed. Through shaft power, mechanical energy is transferred to the aircraft accessory casing 28 to output high-power electrical energy 31 to start the main engine of the aircraft. This method of starting the main engine effectively avoids the inefficient and time-consuming aircraft engine start-up process of the ground support vehicle, shortening the takeoff time. The valve configuration during the takeoff phase is shown in Table 2.

[0052]

[0053] Table 2

[0054] like Figure 3As shown, the integrated energy and thermal management system for hypersonic aircraft based on the TBCC engine sequentially undergoes an in-flight climb phase and a horizontal acceleration phase during the climb phase. Only the turbofan engine operates within the TBCC engine subsystem. Low-pressure gas introduced through the low-pressure bleed air line I1 undergoes bleed air cooling via a closed-loop heat exchange cycle in the environmental control subsystem. High-pressure gas extracted through the high-pressure bleed air line I2 is sent to the environmental control subsystem to activate the auxiliary-emergency power unit. The mechanical energy generated by the auxiliary-emergency power unit and the mechanical energy extracted from the turbofan engine by the mechanical drive shaft I3 are transferred to the aircraft accessory housing 28, driving the aircraft hydraulic system, resulting in hydraulic energy output 28. Simultaneously, shaft work is extracted and electrical energy 31 is output within the aircraft accessory housing 28. The fuel thermal management system cools the fuel under the control of valves V4-V10, V22, and V23, and the cooled fuel is sent to the turbofan engine combustion chamber 4 and afterburner 9 via the second fuel flow regulating valve V2 and the third fuel flow regulating valve V3. The valve configuration during the climb phase is shown in Table 3.

[0055]

[0056]

[0057] Table 3

[0058] like Figure 4 As shown, the integrated energy and thermal management system for hypersonic aircraft based on the TBCC engine undergoes a transition process during the mode transition phase, involving the shutdown of the turbofan engine and the start-up of the ramjet engine. During this phase, the emergency power unit operates, and the mechanical energy generated by the emergency power unit is transmitted to the aircraft accessory housing 28 through the drive shaft, driving the aircraft hydraulic system. Simultaneously, the shaft work is extracted and output as electrical energy 31 in the aircraft accessory housing 28. During this phase, the aircraft accessory subsystem uses the battery 30 to output electrical energy 31 to power the electrical 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 uses a liquid nitrogen cascade cooling device to cool each compartment.

[0059] During this phase, the fuel thermal management subsystem operates in the same manner as when the turbofan engine operates alone. All valves in the fuel subsystem are open, and a combined cooling scheme of ram air and liquid nitrogen is used. The cooled fuel is then delivered to the ramjet engine combustion chamber 12, the turbofan engine combustion chamber 4, and the afterburner chamber 9 via 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 during the mode transition phase 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 the TBCC engine, during the ramjet-powered climb, cruise, reconnaissance, weapon delivery, and cruise return phases, only the ramjet engine operates in the TBCC engine subsystem. The system uses ramjet bleed air line I4 to deliver high-temperature, high-pressure gas to the ramjet turbine 33 for expansion and high-speed rotation, providing mechanical energy to the aircraft accessory sub-casing 28, driving the aircraft hydraulic system, and simultaneously extracting shaft power and outputting electrical energy 31 within the accessory sub-casing 28. In the fuel thermal management subsystem, due to the high flight Mach number, the engine... Since bleed air cooling is no longer applicable, liquid nitrogen cooling is used throughout this stage. Fuel is pressurized by fuel booster pump B1 and then cooled by liquid nitrogen through the HR1-HR2-HR3-HR5 heat exchange circuit. After being cooled by liquid nitrogen, the fuel is divided into two streams by 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. During this stage, the cockpit, landing gear bay, and engine bay are also cooled using a liquid nitrogen cascade cooling system with spray cooling. The cooling effect is the same as in the mode transition stage. The valve configuration during the ramjet stage is shown in Table 5.

[0064]

[0065]

[0066] Table 5

[0067] like Figure 6 The integrated energy and thermal management system for hypersonic aircraft based on the TBCC engine, as shown, operates in the no-power phase. During this phase, the TBCC engine subsystem is inactive, and only the fuel thermal management subsystem and the liquid nitrogen cascade utilization device are operational. The emergency power unit is activated during this phase, transmitting the mechanical energy generated to the aircraft accessory housing 28 via a drive shaft, driving the aircraft hydraulic system. During this phase, the battery 30 outputs electrical energy 31, and simultaneously, shaft power extraction in the aircraft accessory housing 28 outputs electrical energy 31 to power the electrical system. In the fuel thermal management system, due to the high flight Mach number, engine bleed air cooling is no longer applicable; therefore, liquid nitrogen cooling is used throughout this phase. During this phase, the cockpit, landing gear bay, and engine bay are also cooled using a liquid nitrogen cascade cooling device, with the same cooling effect as in the mode transition phase. The valve configuration during the no-power phase is shown in Table 6.

[0068]

[0069] Table 6

[0070] likeFigure 7 The hypersonic aircraft integrated energy and thermal management system based on the TBCC engine shown, in the restart phase, sequentially experiences air climbing, horizontal acceleration phase, the system working state and Figure 3 The valve configuration in the restart phase is shown in Table 3.

[0071] According to the actual construction requirements, it is specifically set, which is not limited in the present application.

[0072] It should be noted that in the present application, relationship 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 the entities or operations. Moreover, the terms "include", "contain" or any other variants thereof are intended to cover non-exclusive inclusion, so that the process, method, article or equipment including a series of elements not only includes those elements, but also includes other elements not explicitly listed or inherent to such process, method, article or equipment. Without more limitations, the element defined by the statement "including a" does not exclude the presence of other identical elements in the process, method, article or equipment including the element.

[0073] The above is only a specific embodiment of the present application, which enables those skilled in the art to 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 range consistent with the principles and novel features of the present application.

Claims

1. An aircraft energy and thermal integrated management system, characterized in that, The system comprises a TBCC engine subsystem for working on gas introduced from outside to generate mechanical energy; a fuel thermal management subsystem for cooling on-board heat sources and simultaneously performing ram air cooling and / or liquid nitrogen cooling on fuel cooled after the on-board heat sources to supply oil to the TBCC engine subsystem; a cabin control subsystem for providing auxiliary power after introducing high-pressure gas from the TBCC engine subsystem or introducing working condition gas from outside, providing emergency power when the TBCC engine subsystem does not provide high-pressure gas, and performing air circulation cooling on the engine cabin, landing gear cabin and cockpit by introducing low-pressure gas from the TBCC engine subsystem; and an aircraft accessory subsystem operating under the action of the auxiliary power or emergency power, or being driven by extracting mechanical energy output by the TBCC engine subsystem, or being driven by working condition gas introduced from outside; The TBCC engine subsystem comprises a turbofan engine and a ramjet engine; in the take-off stage, neither the turbofan engine nor the ramjet engine works; in the climb working stage and the restart stage, only the turbofan engine works, low-pressure gas at the low-pressure gas outlet of the low-pressure compressor in the turbofan engine is introduced into the cabin control subsystem through a low-pressure bleed air pipeline, high-pressure gas at the high-pressure gas outlet of the high-pressure compressor in the turbofan engine is introduced into the cabin control subsystem through a high-pressure bleed air pipeline, and mechanical energy generated by the high-pressure turbine in the turbofan engine is transmitted to the aircraft accessory subsystem through a mechanical transmission shaft; in the TBCC engine mode conversion stage, the turbofan engine is shut down and the ramjet engine is started; in the ramjet engine only working stage, working condition gas introduced from outside is transmitted to the aircraft accessory subsystem through a ramjet bleed air pipeline; in the unpowered glide stage, neither the turbofan engine nor the ramjet engine works; The fuel thermal management subsystem: does not work in the take-off stage; cools on-board heat sources and simultaneously performs ram air cooling and liquid nitrogen cooling on fuel cooled after the on-board heat sources in the climb working stage, the TBCC engine mode conversion stage and the restart stage; cools on-board heat sources and simultaneously performs liquid nitrogen cooling on fuel cooled after the on-board heat sources in the ramjet engine only working stage and the unpowered glide stage; The fuel thermal management subsystem comprises: a ram air cooling circuit, fuel in a main fuel tank is circulated and heat-exchanged after sequentially flowing through a first heat exchanger, a second heat exchanger and a fourth heat exchanger, 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 a cooling medium; and a liquid nitrogen cooling circuit, fuel in the main fuel tank is sequentially flowed through a first heat exchanger, a second heat exchanger and a third heat exchanger, a part of the fuel is pumped to each combustion chamber in the TBCC engine subsystem, a part of the fuel is cooled after entering a fifth heat exchanger and is pumped back to the main fuel tank through a 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 a cooling medium, the sixth heat exchanger is a liquid-gas heat exchanger, and engine thermal load is connected to the third heat exchanger to form a liquid-liquid heat exchange circuit. ​ The environmental control subsystem comprises: an auxiliary power device, which provides auxiliary power after introducing working condition gas from outside in the take-off stage, provides auxiliary power after introducing high pressure gas from the TBCC engine subsystem in the climbing working stage and restart stage, and circulates air from the TBCC engine subsystem to cool the engine cabin, landing gear cabin and cockpit by introducing low pressure gas; an emergency power device, which provides emergency power in the TBCC engine mode conversion stage and unpowered glide stage; and a liquid nitrogen step utilization device, which performs step liquid nitrogen cooling on the engine cabin, landing gear cabin and cockpit in the TBCC engine mode conversion stage, ramjet-only working stage and unpowered glide stage.

2. The aircraft energy and thermal management system of claim 1, wherein, The auxiliary power device comprises: a power turbine, a high-speed transmission shaft, a cooling turbine, an auxiliary generator, a low pressure fan, an auxiliary power device combustion chamber and a sub-fuel tank. In the take-off stage, the auxiliary generator is in the starter mode, the auxiliary generator drives the low pressure fan to work, the working condition gas introduced from outside is delivered to the auxiliary power device combustion chamber through the low pressure fan, the fuel in the sub-fuel tank is delivered to the auxiliary power device combustion chamber, the fuel is burned in the auxiliary power device combustion chamber, and the high temperature and high pressure gas generated drives the power turbine to rotate. In the climbing working stage and restart stage, the high pressure gas introduced from the TBCC engine subsystem is delivered to the auxiliary power device combustion chamber, the fuel in the sub-fuel tank is delivered to the auxiliary power device combustion chamber, the fuel is burned in the auxiliary power device combustion chamber to generate high temperature and high pressure gas which drives the power turbine, and the low pressure gas introduced from the TBCC engine subsystem is cooled by the sixth heat exchanger and the seventh heat exchanger in sequence after passing through the low pressure fan, the gas cooled by the seventh heat exchanger enters the cooling turbine and the power turbine, and the cooling gas passing through the power turbine circulates air to cool the engine cabin, landing gear cabin and cockpit, the sixth heat exchanger is a liquid-gas heat exchanger, and the seventh heat exchanger is a gas-gas heat exchanger.

3. The integrated aircraft energy and thermal management system of claim 1, wherein, The emergency power device comprises: a power turbine, an auxiliary power device combustion chamber and emergency fuel, in the TBCC engine mode conversion stage and unpowered glide stage, the emergency fuel is delivered to the auxiliary power device combustion chamber, the emergency fuel is burned in the auxiliary power device combustion chamber, and the high temperature and high pressure gas generated drives the power turbine to rotate.

4. The integrated aircraft energy and thermal management system of claim 1, wherein, The liquid nitrogen step 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, the nitrogen gas heated in the cockpit is sent to the landing gear cabin, a part of the low temperature liquid nitrogen is branched from the liquid nitrogen booster pump and pumped to the landing gear cabin, and another part of the low temperature liquid nitrogen is branched from the liquid nitrogen booster pump and pumped to the engine cabin.

5. The integrated aircraft energy and thermal management system of any of claims 1 to 4, wherein, The aircraft accessory subsystem comprises: an aircraft accessory nacelle and a ram turbine, In the take-off stage, the auxiliary power provided by the auxiliary power device is transmitted to the aircraft accessory nacelle, the aircraft accessory nacelle outputs electric energy, and the main engine is started. In the climbing work phase and the restart phase, the auxiliary power provided by the auxiliary power device is transmitted to the aircraft accessory gearbox, the mechanical energy extracted from the TBCC engine subsystem is transmitted to the aircraft accessory gearbox, the aircraft accessory gearbox outputs electric energy, the main engine is started and the hydraulic system is operated; In the TBCC engine mode conversion phase and the unpowered glide phase, the emergency power provided by the auxiliary power device is transmitted to the aircraft accessory gearbox, the aircraft accessory gearbox outputs electric energy, the main engine is started and the hydraulic system is operated; In the ramjet engine single work phase, the working condition gas introduced from the outside drives the ram turbine, the mechanical energy generated by the ram turbine is transmitted to the aircraft accessory gearbox, the aircraft accessory gearbox outputs electric energy, the main engine is started and the hydraulic system is operated.

6. The integrated aircraft energy and thermal management system of claim 5, wherein, In the TBCC engine mode conversion phase and the unpowered glide phase, the battery is used to supply electric energy.

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