Hydrogen fuel aero-engine thermal management system with multiple thermal processes cooperating

Abstract

The invention discloses a multi-thermal process collaborative hydrogen fuel aero-engine thermal management system which comprises a large bypass ratio turbofan engine structure and an intercooling regenerative cycle structure, and the intercooling regenerative cycle structure comprises a hydrogen storage tank, a booster pump, a metering valve, a regenerator, an intercooler, a bleed air primary radiator and a bleed air secondary radiator. According to the system, multiple thermal processes such as high-temperature fuel gas heat regeneration, compressor interstage air cooling, turbine blade cooling gas pre-cooling and hydrogen fuel preheating are collaboratively optimized, and low-temperature hydrogen fuel and a cooled combustion product, namely water, are reasonably utilized as heat sinks; the problems that pressure is too large and a system is complex due to a complex air entraining structure in a traditional indirect cooling regenerative layout are solved, the power consumption of an air compressor, fuel consumption and waste heat loss of the whole machine are effectively reduced, the cooling quality of turbine blade cooling air is improved, and the heat management requirements of low energy consumption and high reliability of a civil aviation engine are met.

Description

Technical Field

[0001] The present invention belongs to the technical field of aero-engines, and in particular, relates to a thermal management system for a hydrogen-fueled aero-engine 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 intercooling heat recovery cycle is considered to be an important thermal management solution that can effectively reduce the engine fuel consumption rate and reduce the waste heat loss of the entire machine. The traditional intercooling heat recovery cycle cools the compressor inlet air through the intercooler, thereby reducing the compressor inlet temperature to reduce the compression work; at the same time, the high-temperature combustion gas in front of the tail nozzle is used to heat the compressor outlet airflow, which plays a role in increasing the combustion chamber inlet air temperature and reducing fuel consumption. Although this technology can effectively reduce compressor power consumption and fuel consumption in theory, in actual application, due to the heavy system mass, high structural complexity, and large gas loss along the way, it is currently widely used in ground gas turbines, but it is still difficult 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 prior art 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 can efficiently utilize the heat sink potential of low-temperature hydrogen fuel and condensed water, cool multiple heat sources in the main channel of the engine, and simultaneously achieve 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: A multi-thermodynamic process coordinated thermal management system for hydrogen fuel aircraft engines, including a high bypass ratio turbofan engine structure and an intercooler heat recovery cycle structure; The high bypass ratio turbofan engine structure comprises 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; The intercooling heat recovery cycle structure includes a hydrogen storage tank, a booster pump, a metering valve, a heat regenerator, an intercooler, a primary radiator for bleed air, and a secondary radiator for bleed air; The hydrogen storage tank, booster 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 connected to the metering valve and the cold end inlet of the first-stage radiator of the bleed air respectively, and the cold end outlet of the first-stage radiator of the bleed air is connected to the combustion chamber; 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 respectively connected to the hot end outlet of the regenerator and the cold end inlet of the bleed air secondary radiator, and the cold end outlet of the bleed air secondary radiator is connected to the combustion chamber; 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.

[0007] The low-temperature liquid hydrogen fuel passes through the regenerator and the first-stage radiator of bleed air in turn, successively cooling the high-temperature combustion gas and the turbine blade cooling air, 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, successively cooling the compressor inlet air and the turbine blade cooling air cooled by the hydrogen fuel.

[0008] The system of the present invention uses low-temperature hydrogen fuel and the cooled combustion product - water as a heat sink to 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; The high-temperature gas reheating process utilizes low-temperature hydrogen fuel to absorb the heat of the gas after the low-pressure turbine, thereby solving the problems of increased pipe weight and excessive gas pressure loss caused by using the compressor outlet air as a cold source in the traditional reheating process, and at the same time effectively reduces the gas temperature at the outlet of the aircraft engine and reduces waste heat loss; 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, which can reduce the compressor inlet air temperature and reduce system power consumption; The turbine blade cooling gas precooling process uses the hydrogen fuel after cooling the combustion gas and the water after cooling the fan outlet air as a cold 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; 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 up and burned, thereby reducing fuel consumption.

[0009] Furthermore, the regenerator, intercooler, primary air radiator and secondary air radiator are all made of high-temperature alloy materials, with a serpentine, plate-fin or shell-and-tube structure, and can work stably for a long time under the temperature conditions of 1000K and the pressure conditions of 3MPa.

[0010] A hydrogen-fueled aircraft engine thermal management system is established based on the principle of intercooling heat recovery cycle and an efficient and compact heat exchanger device to achieve energy transfer between the various working fluids in the system. Low-temperature hydrogen fuel is used to cool the engine turbine exhaust and turbine blade cooling air. The hydrogen fuel is preheated and then injected into the combustion chamber to participate in combustion, thereby effectively reducing fuel consumption. In addition, the water generated during the heat recovery process is used to cool the compressor intake and further absorb the heat of the turbine blade cooling air, thereby improving the cooling quality of the turbine blade cooling air, reducing the power consumption of the compressor, and improving the thermal cycle efficiency of the core engine. The heated water vapor will also be injected into the combustion chamber to mix with the mainstream, continue to participate in the thermal cycle process, and improve the engine thrust.

[0011] After adopting the above technical solution, the present invention has the following beneficial effects compared with the prior art: 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 utilizes the advantages of large onboard capacity and strong heat dissipation potential of hydrogen fuel aircraft. It can coordinate and optimize multiple thermal processes such as high-temperature fuel gas heat recovery, compressor interstage air cooling, turbine blade cooling air precooling, and hydrogen fuel preheating, improve the thermal cycle process of aircraft engines, reduce engine fuel consumption, reduce compressor power consumption, improve the thermal cycle efficiency of the entire machine, ensure efficient and reliable operation of hot end components, and finally form a thermal management technology solution suitable for hydrogen fuel high bypass ratio turbofan engines to support the research and development and application of future new energy aircraft in the field of civil aviation.

[0012] The specific implementation modes of the present invention are further described in detail below in conjunction with the accompanying drawings. BRIEF DESCRIPTION OF THE DRAWINGS

[0013] The accompanying drawings are part of this application and are used to provide a further understanding of the present invention. The illustrative embodiments of the present invention and their descriptions are used to explain the present invention, but do not constitute an improper limitation of the present invention. Obviously, the drawings described below are only some embodiments. For ordinary technicians in this field, other drawings can be obtained based on these drawings without creative work. In the drawings: Figure 1 It is a schematic diagram of system connection of the present invention.

[0014] 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.

[0015] 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 are intended to illustrate the concept of the present invention for those skilled in the art by referring to specific embodiments. DETAILED DESCRIPTION

[0016] 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.

[0017] In the description of the present invention, it should be noted that the directions or positional relationships indicated by terms such as “upper”, “lower”, “front”, “back”, “left”, “right”, “vertical”, “inside” and “outside” are based on the directions or positional relationships shown in the accompanying drawings, and are only for the convenience of describing the present invention and simplifying the description, rather than indicating or implying that the device or element referred to must have a specific direction, be constructed and operated in a specific direction, and therefore cannot be understood as a limitation on the present invention.

[0018] In the description of the present invention, it should be noted that, unless otherwise clearly specified and limited, the terms "installed", "connected", and "connected" should be understood in a broad sense, for example, it can be a fixed connection, a detachable connection, or an integral connection; it can be a mechanical connection or an electrical connection; it can be a direct connection or an indirect connection through an intermediate medium. For ordinary technicians in this field, the specific meanings of the above terms in the present invention can be understood according to specific circumstances. Example

[0019] like Figure 1 As shown, a multi-thermodynamic process coordinated hydrogen fuel aircraft engine thermal management system described in this embodiment includes a high bypass ratio turbofan engine structure and an intercooler heat recovery cycle structure; 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 recovery 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 primary radiator 13, and a bleed air secondary radiator 14; The hydrogen storage tank 8, the boost pump 9, and the metering valve 10 are connected in sequence and are located upstream of the regenerator 11, and are used to provide low-temperature hydrogen fuel with suitable flow and pressure for the turbofan engine. At this time, the typical temperature value of the hydrogen fuel is about 40K. 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 first-stage radiator 13 of the bleed air, and the hot end inlet and outlet are respectively connected to the low-pressure turbine 6 and the tail nozzle 7. The low-temperature hydrogen fuel cools the combustion gas after the low-pressure turbine 6 in the regenerator 11. The combustion gas is greatly cooled after passing through the regenerator 11 and then discharged from the engine body, reducing the waste heat loss in the exhaust process of the engine. A part of the water vapor in the combustion gas is cooled and liquefied and then led to the intercooler 12 to cool the compressor inlet air and the turbine blade cooling gas. The hydrogen fuel can be heated to about 500K after being preheated by the regenerator.

[0020] 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 hot end inlet and outlet are respectively connected to the fan 2 and the compressor 3. Liquid water cools the inlet air of the compressor 3 in the intercooler 12 to reduce 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 rate of the engine. The cold side liquid water absorbs a large amount of heat and then vaporizes after heating up, and is used to cool the turbine blade cooling air. The cold end inlet and outlet of the first-stage radiator 13 of bleed air 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 radiator 14 of bleed air. The hydrogen fuel cools the turbine blade cooling air drawn from the compressor 3 in the first-stage radiator 13 of bleed air, 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 the combustion, which is beneficial to reduce the heat required for the fuel in the process of heating up the combustion chamber, and can effectively reduce the fuel consumption. The water vapor generated during the combustion process can be liquefied during the heat recovery process. 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 then it can be used for the air film cooling process of the turbine blades.

[0021] The regenerator 11, intercooler 12, primary bleed air radiator 13 and secondary 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 work stably for a long time under the temperature conditions of 1000K and the pressure conditions of 3MPa.

[0022] 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 of cold and hot side fluids, avoids the air flow pressure loss caused by complex air 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 achieves multiple thermal management goals such as reducing fuel consumption, reducing compressor power consumption, reducing whole machine waste heat loss, and optimizing high-temperature component thermal protection processes by efficiently utilizing the heat sink potential of low-temperature hydrogen fuel, thereby supporting the research and development and application of future new energy aircraft in the field of civil aviation.

[0023] The above is only a preferred embodiment of the present invention, and does not limit the present invention in any form. Although the present invention has been disclosed as a preferred embodiment, it is not used to limit the present invention. Any technician familiar with this patent can make some changes or modify the technical contents suggested above into equivalent embodiments without departing from the scope of the technical solution of the present invention. However, any simple modification, equivalent change and modification made to the above embodiments according to the technical essence of the present invention without departing from the content of the technical solution of the present invention still fall 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 intercooler 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) which are connected in sequence; The intercooling heat recovery cycle structure comprises a hydrogen storage tank (8), a booster pump (9), a metering valve (10), a heat regenerator (11), an intercooler (12), a primary radiator (13) for bleed air, and a secondary 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 first-stage radiator (13) of the bleed air; and the cold end outlet of the first-stage radiator (13) of the bleed air 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 bleed air radiator (13) are respectively connected to the last stage of the compressor (3) and the hot end inlet of the second bleed air radiator (14), and the hot end outlet of the second bleed air radiator (14) is connected to the internal cooling channel of the blades of the high-pressure turbine (5).

2. A 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 serpentine, plate-fin, or shell-and-tube structure, and are capable of long-term stable operation under 1000K temperature conditions and 3MPa pressure conditions.

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