A thermal management system for hydrogen-fueled aircraft engines with coordinated multi-thermodynamic processes

Abstract

The present invention discloses a thermal management system for a hydrogen-fueled aviation engine that coordinates multiple thermodynamic processes. The system includes a high-bypass-ratio turbofan engine structure and an intercooling and regenerative cycle structure. The intercooling and regenerative cycle structure includes a hydrogen storage tank, a boost pump, a metering valve, a regenerator, an intercooler, a first-stage radiator for bleed air, and a second-stage radiator. The system coordinates and optimizes multiple thermodynamic processes such as high-temperature fuel gas reheating, compressor interstage air cooling, turbine blade cooling air precooling, and hydrogen fuel preheating. It rationally utilizes low-temperature hydrogen fuel and the cooled combustion product - water as a heat sink, overcomes the problems of excessive pressure and relatively complex system caused by the complex bleed air structure in the traditional intercooling and regenerative layout, effectively reduces compressor power consumption, fuel consumption and overall machine waste heat loss, improves the cooling quality of turbine blade cooling air, and meets the low-energy consumption and high-reliability thermal management requirements of civil aviation engines.

Description

Technical Field

[0001] The present invention belongs to the technical field of aviation engines, and in particular relates to a thermal management system for hydrogen fuel aviation engines with coordinated multi-thermodynamic processes. Background Art

[0002] Compared with traditional aviation kerosene, hydrogen fuel has the advantages of being carbon-free and having a high calorific value, which can meet the industry's needs in reducing carbon emissions. In addition, due to the characteristics of large reserves, low temperature, and high specific heat capacity, liquid hydrogen is also a high-quality heat sink with huge heat absorption potential in the engine. It has the potential to be used in the thermal management system of the entire machine and improve the thermal cycle process of the entire machine.

[0003] In traditional aircraft engine thermal management systems, the intercooled heat recovery cycle is considered an important thermal management solution that can effectively reduce engine fuel consumption and minimize overall heat loss. The traditional intercooled heat recovery cycle cools the compressor inlet air through an intercooler, thereby reducing the compressor inlet temperature and reducing compression work. At the same time, the high-temperature combustion gas in front of the tail nozzle heats the compressor outlet airflow, increasing the combustion chamber inlet air temperature and reducing fuel consumption. Although this technology can theoretically effectively reduce compressor power consumption and fuel consumption, in actual application, due to the heavy system mass, high structural complexity, and large gas losses along the way, it is currently more widely used in ground-based gas turbines, but has yet to be widely promoted in aircraft engines.

[0004] Therefore, it is urgent to combine the usage characteristics of hydrogen fuel heat sinks, explore the important role of water, the combustion product of hydrogen fuel in the air, in the energy allocation process of the whole machine, and establish an aviation engine thermal management system based on the intercooling heat recovery cycle that comprehensively considers the comprehensive utilization of multiple heat sinks and the coordinated layout of multiple thermal processes. Summary of the Invention

[0005] The technical problem to be solved by the present invention is to overcome the shortcomings of the existing technology and provide a hydrogen-fueled aircraft engine thermal management system that coordinates multiple thermodynamic processes. The system can avoid the pressure loss caused by the complex air bleed structure in the traditional intercooling and heat recovery layout, and coordinately optimize multiple thermodynamic processes such as high-temperature fuel gas heat recovery, compressor interstage air cooling, turbine blade cooling air precooling, and hydrogen fuel preheating. It efficiently utilizes the heat sink potential of low-temperature hydrogen fuel and condensed water, cools multiple heat sources in the main channel of the engine, and simultaneously achieves multiple thermal management goals such as reducing fuel consumption, reducing compressor power consumption, reducing overall waste heat loss, and optimizing the thermal protection process of high-temperature components.

[0006] In order to solve the above technical problems, the basic concept of the technical solution adopted by the present invention is:

[0007] A multi-thermodynamic process coordinated thermal management system for hydrogen-fueled aviation engines, including a high-bypass-ratio turbofan engine structure and an intercooler heat regeneration cycle structure;

[0008] The high bypass ratio turbofan engine structure includes an air inlet, a fan, a compressor, a combustion chamber, a high-pressure turbine, a low-pressure turbine, and a tail nozzle connected in sequence;

[0009] The intercooling heat regeneration cycle structure includes a hydrogen storage tank, a booster pump, a metering valve, a regenerator, an intercooler, a first-stage radiator for bleed air, and a second-stage radiator for bleed air;

[0010] The hydrogen storage tank, boost pump, and metering valve are connected in sequence. The hot end of the regenerator is connected in series between the low-pressure turbine and the tail nozzle. The cold end inlet and outlet of the regenerator are respectively connected to the metering valve and the cold end inlet of the bleed air first-stage radiator. The cold end outlet of the bleed air first-stage radiator is connected to the combustion chamber.

[0011] The hot end of the intercooler is connected in series between the fan and the compressor, the cold end inlet and outlet of the intercooler are connected to the hot end outlet of the regenerator and the cold end inlet of the bleed air secondary radiator respectively, and the cold end outlet of the bleed air secondary radiator is connected to the combustion chamber;

[0012] The hot end inlet and outlet of the first-stage bleed air radiator are respectively connected to the last stage of the compressor and the hot end inlet of the second-stage bleed air radiator, and the hot end outlet of the second-stage bleed air radiator is connected to the internal cooling channel of the high-pressure turbine blade.

[0013] The low-temperature liquid hydrogen fuel passes through the regenerator and the first-stage radiator of bleed air in turn, cooling the high-temperature combustion gas and the turbine blade cooling gas in turn, and liquefies the water vapor in the high-temperature combustion gas. The low-temperature liquid water passes through the intercooler and the second-stage radiator of bleed air in turn, cooling the compressor inlet air and the turbine blade cooling gas after being cooled by the hydrogen fuel.

[0014] The system of the present invention uses low-temperature hydrogen fuel and the cooled combustion product - water as a heat sink to carry out collaborative optimization of multiple thermodynamic processes such as high-temperature gas heat recovery, compressor interstage air cooling, turbine blade cooling air pre-cooling, and hydrogen fuel preheating;

[0015] The high-temperature gas reheating process utilizes low-temperature hydrogen fuel to absorb the heat of the gas after the low-pressure turbine, solving the problems of increased pipe weight and excessive gas pressure loss caused by using the compressor outlet air as a cooling source in the traditional reheating process. At the same time, it effectively reduces the temperature of the gas at the outlet of the aircraft engine and reduces waste heat loss.

[0016] The compressor interstage air cooling process utilizes the condensed water obtained by cooling the gas during the heat recovery process to absorb the heat of the fan outlet air, thereby reducing the compressor inlet air temperature and reducing system power consumption;

[0017] The turbine blade cooling gas pre-cooling process utilizes hydrogen fuel after cooling the combustion gas and water after cooling the fan outlet air as a heat source to absorb the compressor bleed air used as turbine blade cooling gas, so that the heat sink is fully utilized before entering the combustion chamber, while improving the cooling quality of the turbine blade cooling gas;

[0018] The hydrogen fuel preheating process uses high-temperature combustion gas and turbine blade cooling gas as heat sources to preheat the hydrogen fuel in sequence, so that it reaches a higher temperature level before entering the combustion chamber, thereby reducing the heat expended when the fuel is heated and burned, and reducing fuel consumption.

[0019] Furthermore, the regenerator, intercooler, first-stage bleed air radiator, and second-stage bleed air radiator are all made of high-temperature alloy materials, and use a serpentine, plate-fin, or shell-and-tube structure, and can meet the requirements of long-term stable operation under 1000K temperature and 3MPa pressure conditions.

[0020] A hydrogen-fueled aircraft engine thermal management system is established based on the principle of an intercooled heat regeneration cycle and a highly efficient and compact heat exchanger device. This enables energy transfer between the various working fluids within the system, utilizing low-temperature hydrogen fuel to cool the engine turbine exhaust and turbine blade cooling air. The hydrogen fuel is preheated before being injected into the combustion chamber to participate in combustion, effectively reducing fuel consumption. Furthermore, water generated during the heat regeneration process is used to cool the compressor intake air and further absorb heat from the turbine blade cooling air, thereby improving the cooling quality of the turbine blade cooling air, reducing compressor power consumption, and increasing the core engine's thermal cycle efficiency. The heated water vapor is also injected into the combustion chamber to mix with the main stream, continuing to participate in the thermal cycle and increasing engine thrust.

[0021] After adopting the above technical solution, the present invention has the following beneficial effects compared with the prior art:

[0022] The system of the present invention can avoid the pressure loss caused by the complex air bleed structure in the traditional intercooling and heat recovery layout, and fully utilize the advantages of large onboard capacity and strong heat dissipation potential of hydrogen fuel aircraft to carry out coordinated optimization of multiple thermodynamic processes such as high-temperature fuel gas heat recovery, compressor interstage air cooling, turbine blade cooling air precooling, and hydrogen fuel preheating, thereby improving the thermodynamic cycle process of aircraft engines, reducing engine fuel consumption, reducing compressor power consumption, improving the thermodynamic cycle efficiency of the entire machine, and ensuring efficient and reliable operation of hot end components. Ultimately, a thermal management technology solution suitable for hydrogen fuel high bypass ratio turbofan engines is formed to support the future research and development and application of new energy aircraft in the field of civil aviation.

[0023] The specific embodiments of the present invention will be described in further detail below with reference to the accompanying drawings. BRIEF DESCRIPTION OF THE DRAWINGS

[0024] The accompanying drawings are part of this application and are used to provide a further understanding of the present invention. The exemplary embodiments of the present invention and their descriptions are used to explain the present invention, but do not constitute an undue limitation of the present invention. Obviously, the drawings described below are only some embodiments. For those of ordinary skill in the art, other drawings can be obtained based on these drawings without inventive effort. In the accompanying drawings:

[0025] Figure 1 It is a schematic diagram of the connection of the system of the present invention.

[0026] In the figure: 1-intake duct; 2-fan; 3-compressor; 4-combustion chamber; 5-high-pressure turbine; 6-low-pressure turbine; 7-tail nozzle; 8-hydrogen storage tank; 9-boost pump; 10-metering valve; 11-regenerator; 12-intercooler; 13-bleed air first-stage radiator; 14-bleed air second-stage radiator.

[0027] It should be noted that these drawings and textual descriptions are not intended to limit the conceptual scope of the present invention in any way, but rather to illustrate the concept of the present invention for those skilled in the art by referring to specific embodiments. DETAILED DESCRIPTION

[0028] In order to make the purpose, technical solutions and advantages of the embodiments of the present invention clearer, the technical solutions in the embodiments will be clearly and completely described below in conjunction with the drawings in the embodiments of the present invention. The following embodiments are used to illustrate the present invention but are not used to limit the scope of the present invention.

[0029] In the description of the present invention, it should be noted that the terms "upper", "lower", "front", "back", "left", "right", "vertical", "inside", "outside", etc., indicating directions or positional relationships, are based on the directions or positional relationships shown in the accompanying drawings. They are only for the convenience of describing the present invention and simplifying the description, and do not indicate or imply that the device or element referred to must have a specific direction, be constructed and operated in a specific direction. Therefore, they cannot be understood as limiting the present invention.

[0030] In the description of the present invention, it should be noted that, unless otherwise expressly specified or limited, the terms "mounted," "connected," and "connected" should be understood broadly. For example, they may refer to fixed, detachable, or integral connections; mechanical or electrical connections; and direct or indirect connections through an intermediary. Those skilled in the art will understand the specific meanings of the above terms in the present invention based on the specific circumstances.

[0031] Example

[0032] like Figure 1As shown, the thermal management system for a hydrogen-fueled aircraft engine with coordinated multiple thermodynamic processes described in this embodiment includes a high-bypass-ratio turbofan engine structure and an intercooler heat regeneration cycle structure;

[0033] The high bypass ratio turbofan engine structure includes an air inlet 1, a fan 2, a compressor 3, a combustion chamber 4, a high-pressure turbine 5, a low-pressure turbine 6, and a tail nozzle 7 connected in sequence; the intercooling heat regeneration cycle structure includes a hydrogen storage tank 8, a boost pump 9, a metering valve 10, a regenerator 11, an intercooler 12, a bleed air first-stage radiator 13, and a bleed air second-stage radiator 14;

[0034] The hydrogen storage tank 8, boost pump 9, and metering valve 10 are connected in sequence and located upstream of the regenerator 11. They are used to provide low-temperature hydrogen fuel at an appropriate flow rate and pressure for the turbofan engine. The typical temperature of the hydrogen fuel is approximately 40K. The cold inlet and outlet of the regenerator 11 are respectively connected to the metering valve 10 and the cold inlet of the first-stage bleed air radiator 13, while the hot inlet and outlet are respectively connected to the low-pressure turbine 6 and the tail nozzle 7. The low-temperature hydrogen fuel cools the exhaust gas of the low-pressure turbine 6 in the regenerator 11. After passing through the regenerator 11, the gas is significantly cooled before being discharged from the engine body, reducing waste heat loss during the engine exhaust process. A portion of the water vapor in the gas is cooled and liquefied and then directed to the intercooler 12 to cool the compressor inlet air and turbine blade cooling gas. After being preheated in the regenerator, the hydrogen fuel can be heated to approximately 500K.

[0035] The cold-end inlet and outlet of the intercooler 12 are connected to the hot-end outlet of the regenerator 11 and the cold-end inlet of the bleed air secondary radiator 14, respectively. The hot-end inlet and outlet are connected to the fan 2 and the compressor 3, respectively. Liquid water cools the inlet air of the compressor 3 in the intercooler 12, reducing the inlet air temperature of the compressor 3, thereby reducing the power consumption of the compressor during the isentropic compression process and improving the energy utilization of the engine. The cold-side liquid water absorbs a large amount of heat, heats up, and then vaporizes to cool the turbine blade cooling air.

[0036] The cold end inlet and outlet of the first-stage bleed air radiator 13 are connected to the cold end outlet of the regenerator 11 and the combustion chamber 4, respectively. The hot end inlet air is drawn from the last stage of the compressor 3, and the outlet is connected to the hot end inlet of the second-stage bleed air radiator 14. The hydrogen fuel cools the turbine blade cooling air drawn from the compressor 3 in the first-stage bleed air radiator 13, thereby improving the cooling quality of the cold air. The hydrogen fuel is further preheated and then injected into the combustion chamber 4 to participate in combustion, which helps to reduce the heat required for the fuel during the temperature rise process of the combustion chamber, and can effectively reduce fuel consumption. The water vapor generated during the combustion process can be liquefied during the reheating process.

[0037] The cold end inlet and outlet of the bleed air secondary radiator 14 are respectively connected to the cold end outlet of the intercooler 12 and the combustion chamber 4, and the hot end inlet is connected to the hot end outlet of the bleed air primary radiator 13. The outlet air will be introduced into the internal cooling channel of the high-pressure turbine 5 blades. The water vapor further cools the turbine blade cooling air in the bleed air secondary radiator 14, and then is injected into the combustion chamber 4 to mix with the mainstream and continue to participate in the thermodynamic cycle process of the engine system. The cooling quality of the turbine blade cooling air is further improved, and can then be used for the air film cooling process of the turbine blades.

[0038] The regenerator 11, intercooler 12, first-stage bleed air radiator 13 and second-stage bleed air radiator 14 are made of high-temperature alloy materials and have a coiled tube, plate-fin or shell-and-tube structure, and can operate stably for a long time under the temperature conditions of 1000K and the pressure conditions of 3MPa.

[0039] The system structure of the present invention improves the shortcomings of the traditional intercooling heat recovery cycle in practical application, utilizes low-temperature hydrogen fuel instead of compressor air to participate in the heat recovery process, fully utilizes the advantages of large temperature difference heat exchange between the cold and hot side fluids, avoids the airflow pressure loss caused by the complex bleed structure, and comprehensively considers the energy conservation and emission reduction requirements and thermal protection requirements of hydrogen-fired engines, integrates multiple thermodynamic processes such as high-temperature fuel gas heat recovery, compressor interstage air cooling, turbine blade cooling air precooling, and hydrogen fuel preheating, and through the efficient use of the heat sink potential of low-temperature hydrogen fuel, achieves multiple thermal management goals such as reducing fuel consumption, reducing compressor power consumption, reducing overall waste heat loss, and optimizing the thermal protection process of high-temperature components, supporting the future research and development and application of new energy aircraft in the field of civil aviation.

[0040] The above description is merely a preferred embodiment of the present invention and does not constitute any form of limitation to the present invention. Although the present invention has been disclosed as above in terms of a preferred embodiment, it is not intended to limit the present invention. Any technician familiar with this patent can make slight changes or modifications to equivalent embodiments using the above technical content without departing from the scope of the technical solution of the present invention. However, any simple modifications, equivalent changes, and modifications made to the above embodiments based on the technical essence of the present invention without departing from the content of the technical solution of the present invention are still within the scope of the solution of the present invention.

Claims

1. A multi-thermodynamic process coordinated hydrogen fuel aircraft engine thermal management system, characterized by: Including high bypass ratio turbofan engine structure and intercooling heat recovery cycle structure; The high bypass ratio turbofan engine structure comprises an air inlet (1), a fan (2), a compressor (3), a combustion chamber (4), a high-pressure turbine (5), a low-pressure turbine (6), and a tail nozzle (7) connected in sequence; The intercooling and heat regeneration cycle structure comprises a hydrogen storage tank (8), a booster pump (9), a metering valve (10), a regenerator (11), an intercooler (12), a first-stage bleed air radiator (13), and a second-stage bleed air radiator (14); The hydrogen storage tank (8), the boost pump (9), and the metering valve (10) are connected in sequence, the hot end of the regenerator (11) is connected in series between the low-pressure turbine (6) and the tail nozzle (7), the cold end inlet and outlet of the regenerator (11) are respectively connected to the metering valve (10) and the cold end inlet of the bleed air first-stage radiator (13), and the cold end outlet of the bleed air first-stage radiator (13) is connected to the combustion chamber (4); The hot end of the intercooler (12) is connected in series between the fan (2) and the compressor (3), the cold end inlet and outlet of the intercooler (12) are respectively connected to the hot end outlet of the regenerator (11) and the cold end inlet of the bleed air secondary radiator (14), and the cold end outlet of the bleed air secondary radiator (14) is connected to the combustion chamber (4); The hot end inlet and outlet of the first-stage bleed air radiator (13) are respectively connected to the last stage of the compressor (3) and the hot end inlet of the second-stage bleed air radiator (14), and the hot end outlet of the second-stage bleed air radiator (14) is connected to the internal cooling channel of the high-pressure turbine (5) blade.

2. The multi-thermodynamic process coordinated hydrogen fuel aircraft engine thermal management system according to claim 1, characterized in that: The regenerator (11), intercooler (12), primary air radiator (13), and secondary air radiator (14) are all made of high-temperature alloy materials, and are of a coiled tube, plate-fin, or shell-and-tube structure, and are capable of long-term stable operation under 1000K temperature and 3MPa pressure conditions.

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