Comprehensive heat exchanger of aircraft

Abstract

The invention relates to the technical field of heat exchangers, and discloses an aircraft comprehensive heat exchanger which comprises a shell, a longitudinal vortex generator fin set and a micro-channel liquid cooling circulation system, a flow guide cavity for containing cooling liquid is formed in the shell, and the shell and an aircraft belly fairing are integrally installed; the longitudinal vortex generator fin set comprises a plurality of fins of a three-dimensional twisted structure, every two fins symmetrically arranged in the horizontal direction form a fin unit, the fin units are fixed to the face, away from an aircraft belly fairing, of the shell in rows, the twisting angle theta of the fins ranges from 30 degrees to 60 degrees, the ratio of the twisting wavelength to the height of the fins ranges from 0.5 to 2.5, and the width of the fins ranges from 0.5 to 2.5. The micro-channel liquid cooling circulation system is arranged in the shell and comprises a first-stage main channel arranged in the arrangement direction of the fin units, a second-stage branch channel comprises a plurality of rhombic channels penetrating through the first-stage main channel, and a third-stage micro-channel is parallel to the first-stage main channel and penetrates through the second-stage branch channel and can optimize the heat exchange characteristic.

Description

TECHNICAL FIELD

[0001] The present application relates to the technical field of heat exchangers, and particularly relates to a comprehensive heat exchanger for an aircraft. BACKGROUND

[0002] With the transformation of the global aviation industry towards green and low carbon, hydrogen fuel cell power systems have become an important research direction for civil aviation aircraft power devices due to their zero carbon emissions and high energy conversion efficiency. However, a large amount of waste heat will be generated during the operation of the hydrogen fuel cell stack, and if the heat dissipation efficiency is insufficient, the temperature of the hydrogen fuel cell stack will be out of control, which will seriously affect the system life and safety. The weight, aerodynamic resistance and reliability of the heat dissipation system of the civil aviation aircraft are extremely sensitive, and it is necessary to realize the multi-objective collaborative optimization of high-efficiency heat dissipation, low resistance, light weight and low cost in a limited space. Therefore, the development of a high-efficiency, light-weight and low-aerodynamic-resistance heat exchanger has become one of the core technical challenges for the commercial application of civil aviation fuel cell power systems.

[0003] The performance requirements of the heat exchanger for civil aviation aircraft are much higher than those for ground or low-speed aircraft, mainly in the demand for aerodynamic characteristics and lightweight of civil aviation aircraft. The skin friction resistance of the civil aviation aircraft in the cruising stage (Ma 0.8~0.85) accounts for more than 40% of the total resistance, and any additional aerodynamic loss needs to be strictly limited. Studies have shown that if the local resistance of the heat exchanger increases by 1%, the overall fuel efficiency will decrease by 0.3%~0.5%. According to the standard of the International Air Transport Association (IATA), every kilogram of additional weight will increase the life cycle cost of the aircraft by about 1000~1500 US dollars. Therefore, the ratio of heat dissipation power to weight of the heat exchanger should be higher than 1.0kW / kg.

[0004] The traditional fuel cell stack cooling system has problems of insufficient heat exchange capacity, low heat exchange capacity to power ratio of the heat exchanger and great influence on the aerodynamic design of the aircraft when applied to the working environment of high altitude, high speed and limited heat dissipation area of the aircraft. The traditional fin heat exchanger mainly adopts flat fin, corrugated fin or louver fin structure, and there is a significant contradiction between the heat dissipation performance and the aerodynamic characteristics. CN116565255A proposes to use corrugated fins, which are arranged at an angle by arranging adjacent first fin heat sinks and second fin heat sinks, changing the air flow direction, improving the turbulence effect, and facilitating full heat exchange between the air and the fin heat sink. However, in the actual use environment of the aircraft, although the design increases the heat dissipation of the fin, the flow separation zone is easy to cause the boundary layer to thicken, resulting in a nonlinear increase in aerodynamic resistance, which cannot meet the demand of the fuselage integrated heat dissipation of ultra-low resistance. Although the flat fin has simple structure and low manufacturing cost, the fluid in the flow channel is mainly in laminar flow, and the convective heat transfer coefficient is usually less than 200 W / (m²·K), which is difficult to meet the high heat dissipation demand of the fuel cell. In order to improve the heat exchange capacity, CN111867322A increases the fin density or height to expand the heat dissipation area, but it will cause the aerodynamic resistance to rise exponentially. Therefore, it is necessary to propose a new aircraft comprehensive heat exchanger. SUMMARY

[0005] In order to solve the above technical problems, the present application provides an aircraft comprehensive heat exchanger which can continuously strengthen the longitudinal vortex under the condition of low aerodynamic resistance and optimize the heat exchange characteristics.

[0006] The present application provides an aircraft comprehensive heat exchanger, comprising: a shell, a longitudinal vortex generator fin group and a micro-channel liquid cooling circulation system, the shell is internally provided with a flow guide cavity containing cooling liquid, the shell is integrally installed with the aircraft belly fairing, the longitudinal vortex generator fin group comprises a plurality of three-dimensional twisted fins, two fins symmetrically arranged along the horizontal direction form a fin unit, the fin units are fixed in a row on one side of the shell away from the aircraft belly fairing, the twist angle θ of the fin is 30°-60°, and the ratio of the twist wavelength to the height of the fin is 0.5-2.5, the micro-channel liquid cooling circulation system is arranged inside the shell and comprises: a primary main channel, a secondary branch channel and a tertiary micro-channel, the primary main channel is arranged along the arrangement direction of the fin unit, the secondary branch channel comprises a plurality of diamond-shaped channels penetrating the primary main channel, and the tertiary micro-channel is parallel to the primary main channel and penetrates the secondary branch channel.

[0007] Optionally, the height H of the fin is 2mm-10mm, the thickness d of the fin is 0.3mm-1mm, and the spacing S between the two adjacent fin units is 1mm-5mm.

[0008] Optionally, the single-cycle twist angle θ of the fin is 45°±5°, the ratio of the twist wavelength to the height of the fin is 1.84±0.05, and the ratio of the spacing S between two adjacent fin units to the height H of the fin is 0.2~0.5.

[0009] Optionally, the shell and fins can be integrally formed by 3D printing.

[0010] Optionally, the fins are made of aluminum alloy with a thermal conductivity ≥160 W / (m·K) and a tensile strength ≥350 MPa. The fin surface has an AlO ceramic layer with a thickness of 5μm~15μm.

[0011] Optionally, the microchannel liquid cooling circulation system forms an angle of 75° to 85° with the direction of airflow.

[0012] Optionally, the cross-sectional shape of the primary main channel is trapezoidal, with the upper base width being 0.8mm~1.2mm and the height being 0.4mm~0.6mm.

[0013] Optionally, the cross-sectional width of the secondary branch channel is 40% to 50% of that of the primary main channel. The secondary branch channel is at the same height and flush with the primary main channel. The top of the tertiary micro channel is lower than the top of the secondary branch channel. The tertiary micro channel covers at least 80% of the fin surface area.

[0014] Optionally, the height of the three-level microchannel is 0.1mm~0.2mm, and the channel spacing between two adjacent three-level microchannels is 0.3mm~0.5mm.

[0015] Optionally, both the inlet and outlet ends of the primary main channel are connected to the coolant circulation pipeline. The coolant circulation pipeline is connected to the inlet of the primary main channel through a tapered inlet section. The ratio of the cross-sectional area of ​​the tapered inlet section to the cross-sectional area of ​​the primary main channel is 3:1. The closing angle of the tapered inlet section 3-4 is 15°~25°. A backflow suppression protrusion is provided at the bifurcation node between the secondary branch channel and the primary main channel.

[0016] The technical solution provided by the embodiments of the present invention has the following advantages compared with the prior art: The integrated heat exchanger for aircraft provided in this invention designs the longitudinal vortex generator fin assembly as a three-dimensional twisted fin configuration, which causes the cold air flowing through the longitudinal vortex to generate an axial velocity component, prolonging the residence time of the cold air and reducing the lateral velocity gradient. This reduces pressure loss caused by flow separation, while the periodic twisting creates a local flow instability zone, which reduces the thickness of the thermal boundary layer. By causing the cold air flowing through the longitudinal vortex to generate an axial velocity component and disrupting the boundary layer, the fins significantly improve heat transfer without excessively increasing aerodynamic drag. Meanwhile, the centrifugal force generated by the fin twisting can accelerate the shedding of the condensate film, avoiding the risk of icing and exhibiting good self-cleaning properties. By symmetrically designing two fins into a fin unit and fixing the fin units in rows inside the shell, the flow pattern of gas flowing through the shell can be effectively changed from laminar flow to turbulent flow, significantly improving the convective heat transfer coefficient. Furthermore, the grouped fins can effectively guide the airflow direction, forming the required longitudinal vortex, rather than lateral or other directions. Under low aerodynamic resistance conditions, the longitudinal vortex is continuously enhanced, optimizing the heat transfer characteristics. At the same time, a microchannel liquid cooling circulation pipeline is integrated inside the shell, i.e., at the top of the longitudinal vortex generator fin group. Through a primary main channel, a secondary branch channel, and a tertiary microchannel, a multi-level branching structure of a fractal tree-like microchannel network is formed, achieving synergistic optimization of flow uniformity and heat transfer area. Under the premise that the aerodynamic resistance remains basically unchanged, the use of embedded microchannels significantly increases the heat transfer, achieving deep coupling between the coolant flow channel and the air-cooled fins, forming a dual circulation mode, and improving the synergistic efficiency of the cooling medium. Attached Figure Description

[0017] Figure 1 This is a schematic diagram of the overall structure of an integrated heat exchanger for an aircraft, provided as an embodiment of the present invention. Figure 2 A schematic diagram of the connection state between the fins and the shell of the tortuous longitudinal vortex generator provided in an embodiment of the present invention; Figure 3 This is a schematic diagram of the fin structure provided in an embodiment of the present invention; Figure 4 This is a schematic diagram of the structure of the arrayed finned units provided in an embodiment of the present invention; Figure 5 A top view of the microchannel liquid cooling circulation system provided in an embodiment of the present invention; Figure 6 This is a partial structural schematic diagram of the microchannel liquid cooling circulation system provided in an embodiment of the present invention.

[0018] Explanation of reference numerals in the attached figures: 1. Shell; 2. Fins; 3. Microchannel liquid cooling circulation system; 3-1. Primary main channel; 3-2. Secondary branch channel; 3-3. Tertiary microchannel; 4. Coolant circulation pipeline; 5. Centrifugal pump; 6. Throttling valve. Detailed Implementation

[0019] The following detailed description of a specific embodiment of the present invention is provided in conjunction with the accompanying drawings. However, it should be understood that the scope of protection of the present invention is not limited to the specific embodiment.

[0020] In the description of this invention, it should be understood that the terms "center," "longitudinal," "lateral," "length," "width," "thickness," "upper," "lower," "front," "rear," "left," "right," "vertical," "horizontal," "top," "bottom," "inner," "outer," "axial," "radial," and "circumferential" indicate the orientation or positional relationship based on the orientation or positional relationship shown in the accompanying drawings. They are only for the convenience of describing the technical solution of this invention and simplifying the description, and do not indicate or imply that the device or element referred to must have a specific orientation, or be constructed and operated in a specific orientation. Therefore, they should not be construed as limitations on this invention.

[0021] The present invention will be described below through several specific embodiments. To keep the following description of the embodiments clear and concise, detailed descriptions of known functions and components may be omitted. When any component of an embodiment of the present invention appears in more than one drawing, the component may be represented by the same reference numerals in each drawing.

[0022] refer to Figure 1 , Figure 2 , Figure 3 , Figure 4 and Figure 5 ,in, Figure 1 This is a schematic diagram of the overall structure of an integrated heat exchanger for an aircraft, provided by an embodiment of the present invention. Figure 2 This is a structural schematic diagram of the connection state between the fins and the shell of the tortuous longitudinal vortex generator provided in an embodiment of the present invention. Figure 3 This is a schematic diagram of the fin structure provided in an embodiment of the present invention. Figure 4 This is a schematic diagram of the structure of the arrayed finned units provided in an embodiment of the present invention. Figure 5 A top view of the microchannel liquid cooling circulation system provided in an embodiment of the present invention, as shown below. Figure 1 , Figure 2 , Figure 3 , Figure 4 and Figure 5As shown, this embodiment of the invention provides an integrated heat exchanger for an aircraft, comprising: a shell 1, a longitudinal vortex generator fin assembly, and a microchannel liquid-cooled circulation system 3. The shell 1 has a flow guide cavity for holding coolant. The shell 1 is integrated with the aircraft's belly fairing. The air-cooling system achieves forced convection heat transfer through ram air at the leading edge of the belly. The thermal resistance ratio (Rliquid / Rair) between the liquid-cooled and air-cooled sides is maintained at 0.6~0.8. The overall system heat transfer coefficient K ≥ 550 W / (m²·K). The coolant is an ethylene glycol aqueous solution, wherein the volume fraction of ethylene glycol is 40%~60%. The thermal conductivity of the coolant at 120℃ is ≥ 0.75 W / (m·K), and the viscosity is ≤ 3.5%. The longitudinal vortex generator fin assembly comprises several three-dimensional twisted fins 2. Two fins 2 arranged symmetrically in the horizontal direction form a fin unit. The fin units are fixed in a row on the side of the shell 1 facing away from the aircraft's belly fairing. The twist angle θ of the fins 2 is 30°. The twisted wavelength to height ratio of fin 2 is 0.5~2.5. A microchannel liquid cooling circulation system 3 is disposed inside the shell 1, including: a primary main channel 3-1, a secondary branch channel 3-2, and a tertiary microchannel 3-3. The primary main channel 3-1 is arranged along the arrangement direction of the fin units. The secondary branch channel 3-2 includes multiple diamond-shaped channels penetrating the primary main channel 3-1 and extending at a 60° bifurcation angle. The tertiary microchannel 3-3 is parallel to the primary main channel 3-1 and penetrates the secondary branch channel 3-2. The tertiary microchannel 3-3 covers at least 80% of the surface area of ​​fin 2. The microchannel liquid cooling circulation system 3 forms a fractal tree-like microchannel network. The microchannel liquid cooling circulation system 3 is integrally formed by SLM 3D printing with a forming accuracy of ±0.05. The inner wall surface roughness Ra≤6.3μm is mm. The microchannel liquid cooling circulation system is prepared using a 3D printing strategy, which effectively avoids the thermal stress dissipation and cracking caused by the difference in thermal expansion coefficients at the interface of dissimilar materials in traditional heat exchangers.

[0023] The fins 2, with their three-dimensional twisted configuration, achieve vortex-flow synergy to generate vortices. When the air velocity exceeds a certain level, a shear layer instability effect occurs between adjacent vortex pairs, triggering the generation of secondary vortices and achieving a balanced optimization between low flow resistance and high heat transfer. The ratio of the lateral spacing of the longitudinal vortex pairs to the fin spacing is set to 1:2 to ensure effective interference between adjacent vortex systems and avoid flow separation.

[0024] This invention provides an integrated heat exchanger for aircraft. By designing the longitudinal vortex generator fin assembly as a three-dimensional twisted fin configuration, it generates an axial velocity component in the cold air, reducing flow separation pressure loss, disrupting the boundary layer, and increasing heat transfer. This extends the flow time of the airflow on the heat exchanger surface without excessively affecting aerodynamic drag. By symmetrically designing two fins as a fin unit and fixing these fin units in rows within the shell, the flow pattern of the gas flowing through the shell can be effectively changed from laminar to turbulent, significantly improving the convective heat transfer coefficient. Furthermore, the grouped fins can effectively guide the airflow direction, forming a... The required longitudinal vortex, rather than the lateral or other directions, is continuously enhanced under low aerodynamic drag conditions to optimize heat transfer characteristics. Simultaneously, a microchannel liquid-cooled circulation pipeline is integrated inside the shell, at the top of the longitudinal vortex generator fin assembly. Through a primary main channel, a secondary branch channel, and a tertiary microchannel, a multi-level branching structure of a fractal tree-like microchannel network is formed, achieving synergistic optimization of flow uniformity and heat transfer area. Under the premise of basically unchanged aerodynamic drag, the use of embedded microchannels significantly increases the heat transfer, realizes deep coupling between the coolant flow channel and the air-cooled fins, forms a dual circulation mode, and improves the synergistic efficiency of the cooling medium.

[0025] Refer again Figure 2 and Figure 3 The height H of fin 2 is 2mm to 10mm, the thickness d is 0.3mm to 1mm, and the spacing S between two adjacent fin units is 1mm to 5mm.

[0026] To comprehensively consider the aircraft's heat dissipation power and aerodynamic drag, and to meet its heat dissipation requirements without excessively increasing aerodynamic drag, in the integrated heat exchanger for the aircraft provided in this embodiment, the single-cycle twist angle θ of fin 2 is 45°±5°, the ratio of the twist wavelength to the height of fin 2 is 1.84±0.05, the ratio of the spacing S between two adjacent sets of fin units to the height H of fin 2 is 0.2~0.5, and at a Reynolds number Re=5×10⁻⁶... 3 ~2×10 5 Within the range.

[0027] Optionally, the shell 1 and the fins 2 are integrally formed by 3D printing. The integrally formed 3D printed parts have high structural strength, which is conducive to meeting the needs of different flight environments of the aircraft. The integrally formed 3D printed parts are designed with topology optimization and use a lattice structure, which is conducive to achieving lightweight and improving the power-to-weight ratio.

[0028] In this embodiment, fin 2 is made of aluminum alloy with a thermal conductivity ≥160 W / (m·K) and tensile strength ≥350 MPa. The surface of fin 2 has a 5μm~15μm thick AlO ceramic layer. The microhardness of the AlO ceramic layer on the surface of fin 2 is ≥1500 HV, and the surface roughness Ra=1.2±0.3μm, to enhance the component's hardness, making it wear-resistant, heat-resistant, and electrically insulating. The aluminum alloy material chosen for fin 2, with a thermal conductivity ≥160 W / (m·K), provides good temperature uniformity at the heat sink, resulting in superior heat dissipation performance. The tensile strength ≥350 MPa further enhances its structural stability. According to ASTM B117 standards, its microhardness reaches 1500 HV, its salt spray corrosion resistance exceeds 2000 hours, and its surface roughness Ra=1.2μm promotes efficient heat exchange in the gas-liquid two-phase flow. The shell structure employs a topology optimization strategy, using algorithms to calculate the optimal material distribution under stress, removing redundant material and retaining key load-bearing structures, thus meeting lightweight requirements while ensuring reliable structural strength. 3D printing technology enables the topology optimization and integrated design of the liquid cooling plate structure, eliminating the need for welding or brazing. Therefore, a seamless, integral component without weld joints can be directly achieved, preventing leaks.

[0029] Specifically, the microchannel liquid cooling circulation system 3 forms an angle of 75°~85° with the airflow direction, creating an orthogonal heat exchange design. By matching the liquid cooling and air cooling flow directions, the channels of the microchannel liquid cooling circulation system 3 are arranged along the twisting direction of the fins 3, forming an angle of 75°~85° with the airflow direction. This allows the heat from the liquid cooling pipes to be conducted laterally to the air side through the fin substrate, avoiding local hot spots caused by heat flow superposition in traditional designs. The orthogonal pipe layout can effectively improve the temperature uniformity of the radiator. Compared with counter-flow or co-flow circulation pipes, the overall flow of the liquid cooling pipes is shorter, the pressure drop is smaller, the requirements for the pump are reduced, and it is beneficial for lightweight design.

[0030] Optionally, the cross-sectional shape of the primary main channel 3-1 is trapezoidal, which can improve the structural strength to a certain extent and enhance the heat exchange performance between the coolant and the heat exchange plate. The upper bottom width of the primary main channel 3-1 is 0.8mm~1.2mm and the height is 0.4mm~0.6mm.

[0031] refer to Figure 6 , Figure 6 This is a partial structural diagram of the microchannel liquid cooling circulation system provided in an embodiment of the present invention, as shown below. Figure 6As shown, the cross-sectional width of the secondary branch channel 3-2 is 40% to 50% of that of the primary main channel 3-1, ensuring a reasonable coolant flow rate, heat exchange time, and heat exchange capacity while maintaining a reasonable inlet and outlet temperature difference, ensuring that the coolant operates within the efficient operating temperature range of the fuel cell stack. The secondary branch channel 3-2 is at the same height and flush with the primary main channel 3-1. The top of the tertiary microchannel 3-3 is lower than the top of the secondary branch channel 3-2. Three to five primary main channels 3-1 are arranged on one side of the fin, with a spacing of 10% to 30% of the width of the fin. The secondary branch channels 3-2 form a symmetrically distributed "tree-like" fractal structure.

[0032] The secondary branch channel 3-2 is at the same height and flush with the primary main channel 3-1. The top of the tertiary microchannel 3-3 is lower than the top of the secondary branch channel 3-2, which can ensure the flow uniformity effect and control the cooling time of the coolant in the channel to ensure that the inlet and outlet temperatures of the coolant are within the high-efficiency operating temperature range of the hydrogen fuel cell stack.

[0033] Optionally, the height of the three-stage microchannel 3-3 is 0.1mm~0.2mm. The three-stage microchannel 3-3 is formed by laser etching, which can control the flow rate and velocity to control the heat exchange and heat exchange time. The channel spacing between two adjacent three-stage microchannels 3-3 is 0.3mm~0.5mm, which can ensure the heat uniformity of the heat exchange plate and improve the heat exchange performance.

[0034] Refer again Figure 1 The inlet and outlet ends of the primary main channel 3-1 are both connected to the coolant circulation pipe 4, forming a dual-circulation cooling architecture. The coolant circulation pipe 4 is connected to the inlet of the primary main channel 3-1 through a tapered inlet section. A centrifugal pump 5 and a throttle valve 6 are connected to the coolant circulation pipe 4. The pressure resistance of the coolant circulation pipe 4 is ≥12 MPa, and the leakage rate is ≤1×10 Pa·m³ / s. The ratio of the cross-sectional area of ​​the tapered inlet section to the cross-sectional area of ​​the primary main channel 3-1 is 3:1 to ensure that the coolant is evenly distributed to each main channel. The closing angle of the tapered inlet section 3-4 is 15°~25°. At the bifurcation node between the secondary branch channel 3-2 and the primary main channel 3-1, there is a hemispherical backflow suppression protrusion with a diameter of 0.2mm~0.4mm to prevent backflow. The local acceleration effect suppresses the generation of eddies and reduces pressure loss. The coolant circulation pipe 4 is connected to a dynamic control system to control the coolant flow rate Q in the coolant circulation pipe 4. The opening of the throttle valve 6 is adjusted in real time according to the flight altitude h and the fuel cell load power P to control the coolant flow rate Q.

[0035] The liquid cooling system connects the longitudinal vortex generator fin assembly to the hydrogen fuel cell stack via coolant circulation pipe 4 and microchannel liquid cooling circulation system 3. This allows the highly thermally conductive coolant, which has absorbed a large amount of heat in the hydrogen fuel cell stack, to efficiently exchange heat with the air through the microchannel liquid cooling circulation system 3, the shell 1, and the longitudinal vortex generator fin assembly. Simultaneously, the liquid cooling system employs an intelligent dynamic control strategy, adjusting the throttle valve opening in real time based on the aircraft's flight altitude h and the load power P of the hydrogen fuel cell stack to control the coolant flow rate Q. The air cooling system utilizes ram air at the leading edge of the fuselage to achieve forced convection heat transfer.

[0036] The dynamic control system regulates coolant flow through the following logic: (a) When the flight altitude h ≥ 10 km, the adjustment relationship between the coolant flow rate Q and the fuel cell load power P is as follows: Q=k 1* P 0.75 *e −0.02h Where k1 is the correction coefficient, with a value ranging from 0.8 to 1.2; (b) When the ambient temperature T≥40℃ and the humidity RH≥80%, the overload mode is activated, and the flow rate Q is increased to 1.2~1.5 times the baseline value; (c) The response time of throttle valve 6 is ≤50 ms and the control accuracy error is ≤±2%.

[0037] The integrated heat exchanger for aircraft provided in this embodiment of the invention, under the following test conditions: (a) incoming air velocity 50 m / s, angle of attack 0°~10°; (b) coolant inlet temperature 120℃, flow rate 1.2L / min~2.5 L / min; (c) ambient temperature -50℃~45℃, humidity 0%RH~90%RH; the performance of the integrated heat exchanger for aircraft provided in this embodiment of the invention meets the following requirements: (i) heat dissipation ≥80 kW / m², pressure drop ≤100 Pa; (ii) power density ratio ≥1.5kW / kg.

[0038] The installation angle between the shell 1 and the belly fairing is 3°~8°, the windward area ratio is ≤15%, the overall weight of the heat exchanger is ≤20 kg / m², the increase in the aircraft's cruise drag is ≤0.15%, and the heat dissipation power fluctuation is ≤±3% under Ma 0.6~0.85 cruise conditions.

[0039] Heat exchanger wind tunnel testing In this embodiment, the heat exchanger is installed on the belly fairing of a 4:1 scale model of a hydrogen-electric aircraft. Its heat exchange capacity is tested in a wind tunnel environment with a temperature of 25°C, an incoming wind speed of 50 m / s, and an angle of attack of 0°~10°. During the test, the coolant is circulated and heated to 120°C by an external heat source, rather than the hydrogen fuel cell stack. This invention is applicable to heat dissipation in aircraft thermal environment management systems and electronic equipment thermal management systems, therefore, no restrictions are placed on the heat source.

[0040] Heat exchanger ground extreme condition test In this embodiment, the heat exchanger is installed on the belly fairing of the hydrogen-electric aircraft. Its heat exchange capacity is tested in a test environment with a temperature of 25°C, natural convection, and an angle of attack of 0°. During the test, the coolant is circulated and heated to 120°C by an external onboard hydrogen fuel cell stack.

[0041] Aircraft assembly heat exchanger testing In this embodiment, the heat exchanger is installed on the belly fairing of the hydrogen-electric aircraft, and its heat exchange capacity is tested under typical takeoff and landing conditions and cruise conditions of the hydrogen-electric aircraft.

[0042] The above inventions are merely a few specific embodiments of the present invention. However, the embodiments of the present invention are not limited thereto, and any variations that can be conceived by those skilled in the art should fall within the protection scope of the present invention.

Claims

1. An aircraft integrated heat exchanger, comprising: The application relates to a shell (1) internally provided with a flow guide cavity containing cooling liquid, which is integrally arranged on an aircraft belly fairing; a longitudinal vortex generator fin group comprising a plurality of three-dimensional twisted fins (2), two horizontally symmetrically arranged fins (2) forming a fin unit, the fin unit being fixed in an array on one side of the shell (1) away from the aircraft belly fairing, the twist angle theta of the fin (2) being 30-60 DEG, and the ratio of the twist wavelength to the height of the fin (2) being 0.5-2.5; and a micro-channel liquid cooling circulation system (3) arranged in the shell (1) and comprising a primary main channel (3-1), a secondary branch channel (3-2) and a tertiary micro-channel (3-3), the primary main channel (3-1) being arranged along the arrangement direction of the fin unit, the secondary branch channel (3-2) comprising a plurality of diamond-shaped channels penetrating the primary main channel (3-1), and the tertiary micro-channel (3-3) being parallel to the primary main channel (3-1) and penetrating the secondary branch channel (3-2). The height H of the fin (2) is 2-10 mm, the thickness d is 0.3-1 mm, and the spacing S between two adjacent fin units is 1-5 mm. The twist angle theta of a single period of the fin (2) is 45 DEG + / - 5 DEG, the ratio of the twist wavelength to the height of the fin (2) is 1.84 + / - 0.05, and the ratio of the spacing S between two adjacent fin units to the height H of the fin (2) is 0.2-0.

5. The shell (1) and the fin (2) are integrally formed by 3D printing.

2. The aircraft recuperator as recited in claim 1, wherein, The material of the fin (2) is aluminum alloy, the thermal conductivity coefficient is greater than or equal to 160 W / (m*K), the tensile strength is greater than or equal to 350 MPa, and the surface of the fin (2) is provided with an AlO ceramic layer with a thickness of 5-15 mu m.

3. The aircraft recuperator as recited in claim 2, wherein, The micro-channel liquid cooling circulation system (3) forms an included angle of 75-85 DEG with the air flow direction.

4. The aircraft recuperator as recited in claim 1, wherein, The cross-sectional shape of the primary main channel (3-1) is trapezoidal, the upper base width of the primary main channel (3-1) is 0.8-1.2 mm, and the height is 0.4-0.6 mm.

5. The aircraft recuperator as recited in claim 1 or 4, wherein The cross-sectional width of the secondary branch channel (3-2) is 40-50% of the primary main channel (3-1), the secondary branch channel (3-2) is equal in height to the primary main channel (3-1) and is flush with the primary main channel (3-1), the top of the tertiary micro-channel (3-3) is lower than the top of the secondary branch channel (3-2), and the tertiary micro-channel (3-3) covers at least 80% of the surface of the fin (2).

6. The aircraft recuperator as recited in claim 1, wherein, The height of the tertiary micro-channel (3-3) is 0.1-0.2 mm, and the channel spacing between two adjacent tertiary micro-channels (3-3) is 0.3-0.5 mm.

7. The aircraft recuperator as recited in claim 1, wherein ​ 8. The aircraft recuperator as recited in claim 1, wherein ​ 9. The aircraft recuperator as recited in claim 1, wherein ​ 10. The aircraft recuperator as recited in claim 1, wherein, The inlet end and the outlet end of the primary main channel (3-1) are connected with the cooling liquid circulating pipeline (4), the cooling liquid circulating pipeline (4) is connected with the inlet of the primary main channel (3-1) through a tapered inlet section, the ratio of the cross-sectional area of the tapered inlet section to the cross-sectional area of the primary main channel (3-1) is 3:1, the closing angle of the tapered inlet section is 15°~25°, and the secondary branch channel (3-2) is provided with a backflow suppression protrusion at the branching node of the primary main channel (3-1).

Images