Aircraft rudder thermal protection device based on internal-external composite cooling structure and aircraft

Abstract

The application provides an aircraft wing rudder thermal protection device based on an internal-external composite cooling structure and an aircraft, which is composed of a cold gas source, a cold gas transportation device and an internal-external composite cooling structure. The cold gas source provides high-pressure low-temperature cold gas, and the cold gas transportation device contains a cold gas pipeline, a control valve and a cold gas buffer cavity. The internal-external composite cooling structure includes an internal cooling channel and a surface mass injection channel: the internal cooling channel integrates a turbulence rib column, a curved channel, a ribbed channel and a coupled layer plate cooling structure, strengthens convective heat exchange and improves structural strength; the surface mass injection channel is a discrete gas film hole, and the cold gas heated and warmed by the internal cooling channel is sprayed out through the discrete gas film hole to form a protective gas film, which isolates the high-temperature airflow. The application has the following advantages: 1) multiple internal heat exchange strengthening structures are integrated to improve the heat absorption capacity of the coolant; 2) the thermal protection capacity is strong; 3) the thermal protection performance can be dynamically adjusted to adapt to the changing flight thermal environment; and 4) the wing rudder can be widely cooled.

Description

Technical Field

[0001] This invention belongs to the field of aerodynamic heat and thermal protection technology for aircraft, specifically relating to a thermal protection device for aircraft wings and rudders based on an internal and external composite cooling structure, and an aircraft. Background Technology

[0002] The wing rudder is a critical component of hypersonic vehicles, used to provide lift and control flight attitude. During flight, the wing rudder directly faces the high-speed incoming airflow, forming a detached shock wave at its leading edge. This shock wave compresses the incoming air, significantly increasing the gas temperature and causing intense aerodynamic heating of the wing rudder, even ablating its surface. This affects the hypersonic vehicle's lift and flight attitude, and can even lead to a crash. To ensure the safe and reliable operation of hypersonic vehicles, advanced aerodynamic thermal protection technologies are urgently needed.

[0003] Currently, the technical solutions for aerodynamic thermal protection of hypersonic aircraft, especially for wing and rudder components, can be divided into the following four types.

[0004] The first technical solution is a passive thermal protection technology based on heat-insulating materials. Specifically, for wing rudders, this involves wrapping a heat-insulating jacket around the metal rudder core, with this material layer directly bearing the aerodynamic thermal load. [1] Its advantage is that its thermal protection capability can cover the entire wing and rudder surface. However, this solution also has obvious limitations: 1) the temperature resistance of the wall is limited by the upper limit of the temperature resistance of the heat shield material (usually not higher than 1100℃); 2) it cannot actively adjust the thermal protection performance and is difficult to adapt to the variable flight thermal environment.

[0005] The second technical solution is a semi-passive thermal protection technology based on heat pipes. Specifically for wing rudders, this involves: arranging heat pipes in the high heat flux region of the wing rudder's leading edge; or integrating the wing rudder with a high-temperature alkali metal heat pipe, making the wing rudder structure a shell for the heat pipe. [2] The working principle of this technical solution is as follows: the working fluid absorbs heat and evaporates at the hot end, the vapor flows to the cold end and releases heat to condense, and then returns to the hot end through the capillary structure, forming a continuous cycle. The advantage of the second technical solution is that it has the ability to reduce wall temperature in a wide range of high heat load areas of the wing and rudder. However, its disadvantages are also prominent: 1) its thermal protection capability is limited, constrained by the sonic limit, condensation limit, and capillary limit; 2) it cannot actively adjust its thermal protection capability to adapt to varying thermal loads.

[0006] The third technical solution is active thermal protection technology based on coolant flow channels. Specifically for wing rudders, it can be further divided into two categories based on the relative spatial positions of the coolant and the surface to be protected: internal forced convection cooling and surface quality ejector cooling. Internal forced convection cooling involves arranging coolant flow channels within the wing rudder's leading edge wall. Forced convection heat transfer between the coolant and the channel wall removes the heat generated by aerodynamic heating, achieving the goal of cooling the wing rudder's leading edge wall. Surface quality ejector cooling can be further divided into two sub-categories based on the location of the coolant flow channels: first, ejector cooling of the fuselage surface upstream of the wing rudder. [3] Second, reverse jet cooling at the leading edge of the wing and rudder. [4] The first sub-category of technical solutions involves arranging surface mass ejection channels on the fuselage surface upstream of the wing rudder, allowing the actively ejected gas to form an "air wall" at the leading edge of the wing rudder, blocking the high-temperature incoming flow and achieving the goal of thermal protection. The second sub-category of technical solutions also involves arranging surface mass ejection channels at the leading edge of the wing rudder to introduce cool air into the high-heat-load area, forming a protective air film to isolate the high-temperature incoming flow. The advantages of the third technical solution are: 1) strong thermal protection capability, i.e., a high upper limit on the total temperature of the incoming flow that it can withstand; 2) dynamically adjustable thermal protection performance. However, its disadvantages are: 1) it only utilizes the heat absorption or insulation capacity of the coolant, without simultaneously utilizing both heat absorption and insulation capabilities, thus not fully exploiting the thermal protection potential of the cold source; 2) the cooling area is limited, with the cooling effect confined to the area near the coolant flow channel (such as the wing rudder root or leading edge).

[0007] The fourth technical solution is an active thermal protection technology based on an internal and external composite cooling structure. Specifically for the wing rudder, this involves a sandwich structure with ribbed channels inside the wing rudder, allowing the coolant to absorb heat from the wall surface through convection. The heated coolant is then ejected through mass ejection channels on the wing rudder surface, forming a protective gas film on the wall surface, isolating the high-temperature mains from the wing rudder wall and achieving the goal of cooling the wall surface. The advantages of this technical solution are: 1) strong thermal protection capability, i.e., a high upper limit for the total incoming flow temperature; 2) dynamically adjustable thermal protection performance; and 3) ability to achieve wide-area cooling of the wing rudder. However, its disadvantage is that the internal heat exchange enhancement structure of the wing rudder is singular, with ribbed channels only arranged inside the wing rudder, resulting in limited heat exchange enhancement effect and insufficient utilization of the coolant's heat absorption capacity.

[0008] In summary, existing heat shield and heat pipe solutions suffer from limited thermal protection capabilities and the inability to actively adjust thermal protection capabilities. Traditional active thermal protection technologies rely solely on the heat absorption or insulation capabilities of the coolant, and their cooling area is limited, failing to meet the needs of large-scale cooling. Furthermore, active thermal protection technologies based on internal and external composite cooling structures currently employ a single internal enhanced heat exchange structure, simply using ribbed channels, resulting in insufficient utilization of the coolant's heat absorption capacity. To meet the future demands of high Mach numbers, reusability, and variable operating conditions in flight, there is an urgent need to develop an aerodynamic thermal protection technology that combines the following advantages: 1) Integrating multiple internal enhanced heat exchange structures to improve the coolant's heat absorption capacity, while employing surface-quality ejector channel structures to utilize the coolant's insulation capabilities; 2) Strong thermal protection capabilities, i.e., a high upper limit for the total incoming flow temperature; 3) Dynamically adjustable thermal protection performance to adapt to varying flight thermal environments; 4) Ability to achieve large-scale cooling of wings and rudders.

[0009] References

[0010] [1] Zhou Changcheng. Study on low-temperature preparation of C / SiC composite material and its air rudder heat shield [D]. National University of Defense Technology, 2008.

[0011] [2] Wang Changji. Study on heat transfer mechanism and performance of cyclic phase change thermal protection system for hypersonic vehicles [D]. Nanjing University of Aeronautics and Astronautics, 2022. DOI:10.27239 / d.cnki.gnhhu.2022.001447.

[0012] [3] Yuan Ye, Cao Zhanwei, Ma Wei, et al. Experimental study on the influence of active ejector cooling on the thermal environment of air rudder [J]. Missiles and Space Launch Vehicle Technology, 2021, (06): 48-51.

[0013] [4] Feng Yuan, Guo Qingyang, Gang Dundian, et al. Study on drag reduction and heat protection characteristics of reverse jet at the leading edge of high-speed air rudder [J]. Aerospace Technology, 2025, (04): 13-22+48. DOI: 10.16338 / j.issn.2097-0714.20230276. Summary of the Invention

[0014] To address the aforementioned needs, this invention proposes a thermal protection device for aircraft wing rudders based on an internal and external composite cooling structure. Combining various cooling channel configurations and surface quality ejection channels within the wing rudders, it enhances the heat absorption capacity of the coolant while utilizing its heat insulation capabilities. Furthermore, it can adjust the coolant supply pressure in real-time to match the thermal protection requirements under different flight conditions. The specific solution is as follows:

[0015] A thermal protection device for aircraft wing rudders based on an internal and external composite cooling structure includes a cold air source, a cold air transport device, and an internal and external composite cooling structure connected in sequence. The cold air source and the cold air transport device are located inside the aircraft, and the internal and external composite cooling structure is located inside the wing rudders.

[0016] The internal and external composite cooling structure includes internal cooling channels for the wing and rudder and surface mass ejection channels for the wing and rudder; among which,

[0017] The internal cooling channel of the wing rudder is a meandering channel arranged inside the wing rudder. The meandering channel includes a first channel, a second channel, and a third channel connected in sequence along the gas flow direction. The first channel is connected to the cold air transport device. The third channel is parallel to the outer surface of the leading edge of the wing rudder and is connected to the surface mass ejection channel of the wing rudder. The second channel is connected between the first channel and the third channel. The first channel is provided with an array of turbulence ribs. The second channel is provided with bends and several ribs along the gas flow direction.

[0018] The surface mass ejection channel of the wing rudder is a discrete air film hole array set on the outer surface of the leading edge of the wing rudder;

[0019] The cold air output from the cold air source is transported to the internal and external composite cooling structure through the cold air transport device. The cold air flows through the internal cooling channel of the wing rudder and is then discharged through the discrete air film hole array to form a covering air film on the outer surface of the wing rudder to protect it.

[0020] Furthermore, the cold air source is one of the following or a combination thereof: a cold air source formed by cooling the incoming air through expansion, a cold air source formed by precooling the incoming air using fuel, or a cold air source output from a high-pressure air cylinder carried inside the aircraft.

[0021] Furthermore, the cold air transport device includes cold air piping, control valves, and a cold air buffer chamber;

[0022] The air conditioning ducts are pipes with circular, rectangular or triangular cross-sections, and the air conditioning ducts connect the air conditioning source and the internal and external composite cooling structure;

[0023] In the air conditioning duct, a control valve and an air conditioning buffer chamber are installed sequentially along the gas flow direction.

[0024] The control valve is used to dynamically adjust the pressure and mass flow rate of the cooling air in real time.

[0025] The cold air buffer chamber is used to rectify and decelerate the cold air to create near-stagnant flow conditions. Temperature sensors and pressure sensors are installed in the cold air buffer chamber to monitor the temperature and pressure of the cold air, respectively.

[0026] Furthermore, the first channel is a rectangular cavity distributed along the wing rudder height direction, which is assumed to be the longitudinal direction.

[0027] The turbulence ribs in the first channel are arranged in an array in the longitudinal direction, and the turbulence ribs are arranged side by side or in a staggered arrangement in the gas flow direction.

[0028] Each baffle rib is fixed laterally within the first channel, with one or both ends of the baffle rib fixedly connected to the inner wall of the first channel.

[0029] The cross-section of the spoiler ribs is circular, rhomboid, or elliptical, with the tip of the rhomboid or the short axis of the ellipse facing the air inlet side.

[0030] Furthermore, the second channel is a narrow strip channel with a sweep angle, and curves are arranged at both ends of the second channel to connect the first channel and the third channel.

[0031] A row of ribs is provided on both sides of the inner wall of the second channel along the gas flow direction. The ribs on both sides of the inner wall of the second channel are arranged side by side or in a staggered arrangement.

[0032] The ribs are one or more combinations of continuous V-ribs, discontinuous V-ribs, straight ribs, discontinuous straight ribs or W-ribs, with the tips of continuous V-ribs, discontinuous V-ribs or W-ribs facing the air inlet side.

[0033] Furthermore, the third channel is parallel to the outer surface of the wing rudder leading edge, and a coupling plate cooling structure is provided between the third channel and the outer surface of the wing rudder leading edge.

[0034] The cooling structure of the coupling plate has a hollow cavity inside, and there are some turbulence ribs in the hollow cavity. The turbulence ribs are cylindrical or rhomboid columns that are arranged side by side or in a staggered manner in the hollow cavity.

[0035] One side of the hollow chamber is connected to the third channel via an array of impact holes;

[0036] The other side of the hollow chamber is connected to the atmosphere through a discrete air film hole array. Each discrete air film hole is set at an angle or vertically between the hollow chamber and the outer surface of the wing leading edge.

[0037] Furthermore, the angle between each discrete film air hole and the leading edge of the wing rudder is an acute angle.

[0038] Furthermore, the discrete air film pores can be round, scoop-shaped, fan-shaped, or dumbbell-shaped.

[0039] An aircraft equipped with any of the above-mentioned aircraft wing and rudder thermal protection devices.

[0040] Compared with the prior art, the present invention has the following beneficial effects:

[0041] 1. This invention integrates multiple internal heat transfer enhancement structures to improve the heat absorption capacity of the cold air source, while employing a surface mass ejection channel structure to utilize the heat insulation capacity of the coolant. Currently, active thermal protection schemes based on internal and external composite cooling structures rely solely on ribbed channels for enhancing heat transfer within the wing rudder, resulting in limited heat transfer enhancement effects and insufficient utilization of the coolant's heat absorption capacity. This invention integrates multiple heat transfer enhancement structures within the wing rudder's internal cooling channels, including turbulence ribs, curved structures, ribbed second channels, and a laminated cooling system coupling impact holes, ribs, and film cooling holes, thereby increasing the convective heat transfer coefficient between the coolant and the wall surface and unlocking its heat absorption potential. Simultaneously, mass ejection channels are arranged on the wing rudder surface to extract the coolant from the internal cooling channels, forming a film on the wing rudder surface to isolate the high-temperature mainstream. Therefore, this invention significantly improves the coolant's heat absorption capacity while utilizing the film insulation capacity.

[0042] 2. Strong thermal protection capability, meaning a high upper limit for the total incoming flow temperature it can withstand. With the wing and rudder materials remaining unchanged, the highest temperature that the thermal jacket solution can withstand is 1373 K; after adopting an integrated design of the wing and rudder with an alkali metal high-temperature heat pipe, this upper limit is increased to 1500 K. The aircraft wing and rudder thermal protection device based on an internal and external composite cooling structure proposed in this invention further breaks through the upper limit of the total incoming flow temperature to 1853 K, supporting the extreme thermal protection requirements at higher flight speeds.

[0043] 3. Dynamically adjustable thermal protection performance to adapt to varying flight thermal environments. Traditional heat shields and heat pipe structures lack active thermal protection adjustment capabilities, making it difficult to adapt to drastically changing thermal loads across a wide speed range and multiple flight conditions. The aircraft wing and rudder thermal protection device proposed in this invention, based on an internal and external composite cooling structure, can actively and flexibly adjust the cooling gas pressure or mass flow rate by real-time adjustment of the control valve's set parameters, achieving optimal thermal protection under different flight speeds, angles of attack, and rudder deflection angles.

[0044] 4. Enables wide-area cooling of wing rudders. Traditional active thermal protection technologies only provide cooling effects near the coolant flow channels (such as the wing rudder root or leading edge), offering insufficient thermal protection for other areas. The aircraft wing rudder thermal protection device proposed in this invention, based on an internal and external composite cooling structure, utilizes the heat absorption and insulation capabilities of the coolant to expand the cooling range of the wing rudder surface, achieving full-area cooling of the wing rudder. Attached Figure Description

[0045] To more clearly illustrate the technical solutions in the embodiments of the present invention, the drawings used in the description of the embodiments or the prior art will be briefly introduced below. Obviously, the drawings described below are only some embodiments of the present invention. For those skilled in the art, other drawings can be obtained based on these drawings without creative effort.

[0046] Figure 1 A schematic diagram of an aircraft equipped with the wing-rudder thermal protection device of the present invention;

[0047] Figure 2 This is a diagram of the internal structure of the wing rudder thermal protection device;

[0048] Figure 3 A cross-sectional view of the internal structure of the wing rudder thermal protection device;

[0049] Figure 4 Schematic diagram of three ways to provide cold air sources for the wing rudder thermal protection device;

[0050] Figure 5 The diagram shows the distribution of three different types of turbulence ribs within the first channel;

[0051] Figure 6 This is a schematic diagram of the second channel and the ribs on both sides inside it, in one embodiment;

[0052] Figure 7 A schematic diagram showing four different rib distribution patterns inside the second channel;

[0053] Figure 8 This is a schematic diagram of the wing rudder internal cooling channel, rudder leading edge impact hole-rib-air film coupling layered cooling structure of the present invention.

[0054] Figure 9 This is a schematic diagram showing different hole shapes for the wing rudder surface quality ejection channel of the present invention;

[0055] Figure 10 The numerical simulation computational domain for the embodiment;

[0056] Figure 11 The distribution of net heat flux density reduction at the rudder leading edge caused by different orifice shapes under the same incoming flow conditions;

[0057] Figure 12 The average circumferential temperature distribution and its difference at the leading edge of the wing and rudder before and after cold air injection are obtained through flow-heat conjugate heat transfer calculation simulation.

[0058] Explanation of reference numerals in the attached figures:

[0059] Air source 1;

[0060] 2. Cold air transport device; 2.1. Cold air pipeline; 2.2. Control valve; 2.3. Cold air buffer chamber;

[0061] 3. Internal and external composite cooling structure; 3.1. Internal cooling channel of wing rudder; 3.2. Mass ejection channel of wing rudder surface;

[0062] First channel 3.1.1, spoiler rib 3.1.1.1;

[0063] Second channel 3.1.2, bend 3.1.2.1, rib 3.1.2.2;

[0064] Third channel 3.1.3;

[0065] Coupled layer cooling structure 3.1.4, impact hole array 3.1.4.1, hollow cavity 3.1.4.2, turbulence ribs 3.1.4.3;

[0066] Discrete air film pore array 3.2. Detailed Implementation

[0067] In the following description, numerous specific details are set forth in order to provide a more thorough understanding of the invention. However, it will be apparent to those skilled in the art that the invention can be practiced without one or more of these details. In other instances, certain technical features well-known in the art have not been described in order to avoid obscuring the invention.

[0068] To fully understand this invention, detailed steps and structures will be presented in the following description to illustrate the technical solution of this invention. Preferred embodiments of the invention are described in detail below; however, in addition to these detailed descriptions, the invention may have other embodiments.

[0069] To address the thermal protection challenges faced by hypersonic aircraft wing rudders in extreme thermal environments, this invention discloses a wing rudder thermal protection device based on an internal and external composite cooling structure. This device aims to achieve the following design objectives: 1) Integrating multiple internal enhanced heat exchange structures to improve the heat absorption capacity of the cold air, while employing a surface quality ejector channel structure to utilize the heat insulation capacity of the cold air; 2) Strong thermal protection capability, i.e., a high upper limit for the total incoming flow temperature; 3) Dynamically adjustable thermal protection performance to adapt to varying flight thermal environments; 4) Achieving wide-range cooling of the wing rudder.

[0070] To achieve the above objectives, the present invention proposes a method as follows: Figure 1 and Figure 2 The thermal protection device for the aircraft wing rudder shown has the following core components: 1. a cold air source, 2. a cold air transport device, and 3. an internal and external composite cooling structure connected in sequence.

[0071] The operating principle and core functions of this invention are as follows (see Figure 2-3 High-pressure cold air is drawn from cold air source 1 (which can be incoming air after expansion cooling, incoming air pre-cooled by fuel, or high-pressure air cylinders carried by the aircraft, see...). Figure 4The cooled air enters the control valve 2.2 via the cooled air pipe 2.1. The control valve 2.2 adjusts the total pressure or mass flow rate of the discharged cooled air according to real-time thermal protection requirements. It then flows through the cooled air pipe 2.1 into the cooled air buffer chamber 2.3. The buffer chamber 2.3 is a cavity with a certain volume, where the cooled air forms near-stagnant flow conditions, and the total temperature and pressure of the cooled air are measured simultaneously, providing key input parameters for dynamically adjusting thermal protection performance. The cooled air in the buffer chamber 2.3 is then discharged into the wing rudder's internal cooling channel 3.1 via the cooled air pipe 2.1, additionally cooling the rudder shaft. Temperature sensors and pressure sensors are installed in the cooled air buffer chamber 2.3 to monitor the cooled air temperature and pressure. By using the parameters collected in real-time by the temperature and pressure sensors, combined with the thermal load requirements of the current flight conditions, the control valve can dynamically correct the cooled air delivery parameters to ensure that the cooling effect matches the actual thermal protection requirements.

[0072] Within the wing rudder's internal cooling channel 3.1, cool air absorbs and carries away heat from the wing rudder structure through various enhanced heat exchange structures. The specific process is as follows:

[0073] First, the cold air flows through the first channel 3.1.1. Within the first channel 3.1.1, there is an array of baffles 3.1.1.1 arranged in a full-height / half-height structure, either side-by-side or staggered. The cross-sectional shape of the baffles 3.1.1.1 includes, but is not limited to, circular, rhomboid, etc. (e.g., Figure 5 As shown in the figure, convective heat transfer between the cold air and the wall is enhanced by increasing the disturbance to the flow and the heat exchange contact area.

[0074] Then, the cold airflow enters the second channel 3.1.2 through the bend 3.1.2.1, where flow separation and reattachment occur, enhancing convective heat transfer. Specifically, a row of ribs 3.1.2.2 is provided on both sides of the inner wall of the second channel 3.1.2 along the gas flow direction (e.g., ...). Figure 6 As shown), the ribs 3.1.2.2 on both sides of the inner wall of the second channel 3.1.2 are arranged side by side or in a staggered pattern. The ribs 3.1.2.2 are one or more combinations of continuous V-ribs, discontinuous V-ribs, straight ribs, discontinuous straight ribs, or W-ribs. The tips of the continuous V-ribs, discontinuous V-ribs, or W-ribs face the air inlet side and can be optimized by adjusting parameters such as rib spacing, tilt angle, and discontinuity pattern (see...). Figure 7 ).

[0075] Then, the cold airflow exiting from the second channel 3.1.2 enters the third channel 3.1.3 via the bend 3.1.2.1, and then enters the coupling plate cooling structure 3.1.4 (as shown). Figure 8 (As shown). The coupling plate cooling structure 3.1.4 has a hollow cavity 3.1.4.2 inside, and a turbulence rib 3.1.4.3 is provided inside the hollow cavity 3.1.4.2 (as shown). Figure 3The two ends of the turbulence ribs 3.1.4.3 are connected to the two ends of the inner wall of the hollow cavity 3.1.4.2, serving as support and turbulence. The turbulence ribs 3.1.4.3 are arranged side by side or in a staggered manner inside the hollow cavity 3.1.4.2, and their cross-sectional shape is cylindrical or rhomboid.

[0076] An impact hole array 3.1.4.1 connects one side of the hollow chamber 3.1.4.2 to the third channel 3.1.3. The axes of each impact hole are vertically or obliquely positioned on the connecting wall between the third channel 3.1.3 and the hollow chamber 3.1.4.2. The other side of the hollow chamber 3.1.4.2 is connected to the atmosphere through a discrete film perforation array 3.2. Each discrete film perforation is obliquely or vertically positioned between the hollow chamber 3.1.4.2 and the outer surface of the wing's leading edge. The cold air passes through the impact hole array 3.1.4.1 to form an impact jet, which significantly enhances convective heat transfer between the cold air and the internal wall. Subsequently, the cold air flows through the turbulence ribs 3.1.4.3, increasing flow disturbance and expanding the heat transfer contact area, further enhancing convective heat transfer. Furthermore, the turbulence ribs 3.1.4.3 also improve the structural strength of the wing.

[0077] Finally, the cooled air, having absorbed heat from the wall, is ejected from the leading edge surface of the wing rudder through the surface mass ejection channel 3.2 (discrete film perforation array), forming a protective film that isolates the high-temperature incoming flow, thereby reducing the heat flux density on the leading edge surface of the wing rudder. Figure 9 The adhesion of cool air can be improved and the heat flux on the leading edge surface of the wing rudder reduced by optimizing the tilt angle of the discrete film cooling holes and the shape of the cooling holes (including but not limited to round holes, scoop-shaped holes, fan-shaped holes, and dumbbell-shaped holes). Preferably, the angle between the discrete film cooling holes and the front end of the outer surface of the wing rudder leading edge is an acute angle to achieve better protection.

[0078] Example 1

[0079] To evaluate the effectiveness of this invention, computational fluid dynamics simulations were used to investigate the thermal protection performance of the internal and external composite cooling thermal protection structure on the wing leading edge. First, the boundary conditions given by computational fluid dynamics are introduced. The entire computational domain is visible. Figure 10 Mesh was applied only to the fluid domain. The computational domain cross-section was a 260*260mm square, and the wing model was placed on a 400mm long plate. The leading edge of the plate was 200mm from the computational domain inlet, and the trailing edge was 200mm from the computational domain outlet. The film cooling holes on the leading edge of the wing had a diameter of 1mm, an angle of 30°, and were either circular or scoop-shaped, totaling 11 holes.

[0080] The inlet is given a total pressure of 0.15 MPa and a total temperature of 1500 K, while the outlet static pressure is 287 Pa. This setting ensures that the mainstream Mach number reaches 5, i.e., five times the speed of sound. The wall is a no-slip isothermal wall with a wall temperature set to 300 K. The total pressure at the film gas inlet is set to 7.6 kPa, and the total temperature is set to 300 K. The computational fluid dynamics solver has been verified through mesh independence checks and experimental validation using the turbulence model, and will not be shown here.

[0081] Before introducing the thermal protection effect of this invention on the wing and rudder, we will first introduce the thermal protection performance evaluation index of this invention—Net Heat Flux Reduction (NHFR), which is expressed as follows:

[0082]

[0083] Where, q c q is the wall heat flux density during coolant injection. uc This represents the heat flux density of the wall without cooling holes. Its physical meaning is the proportion of the reduction in heat flux density of the wall after coolant injection to the heat flux density of the wall without cooling holes.

[0084] The thermal protection effect of the internal and external composite cooling structure on the leading edge of the wing and rudder under different jet angles and orifice shapes is as follows: Figure 11 As shown in the figure, the area enclosed by the black solid line represents the cold air coverage area where the net heat flux density reduction (NHFR) is greater than 0.5. The figure demonstrates that film cooling has a significant heat reduction effect on the wing leading edge. Within the cold air coverage area (enclosed by the black solid line), the net heat flux density reduction exceeds 50%, and the area-average net heat flux density reduction for all areas in the figure is above 31%.

[0085] To screen for structures with excellent thermal protection performance, this embodiment conducts a parametric study, evaluating the impact of circular and frigate-shaped holes on the thermal protection performance of the wing and rudder leading edge region under a jet angle of 30°. Evaluation indicators for thermal protection performance include the coolant coverage area on the wall, the reduction in surface-average net heat flux density, and the mass flow rate of coolant consumed. Figure 11 As shown, the coolant coverage area is sensitive to the orifice shape. The coolant coverage area produced by a circular orifice is 0.18 cm², while that produced by a scoop-shaped orifice is only 0.12 cm², meaning that the coolant coverage area of ​​a circular orifice is 1.5 times that of a scoop-shaped orifice.

[0086] Figure 11 The reduction in average net heat flux density is also sensitive to the orifice shape. The reduction in average net heat flux density produced by a circular orifice is 0.35, while that of a swirl-shaped orifice drops to 0.31, meaning that the reduction in average net heat flux density produced by a circular orifice is 1.13 times that of a swirl-shaped orifice.

[0087] Analysis of coolant consumption shows that the coolant mass flow rate consumed by the circular orifice is slightly lower than that of the scoop orifice. The coolant mass flow rate consumed by a single circular orifice is 0.0063 g / s, while that of a single scoop orifice is 0.0064 g / s, meaning that the coolant mass flow rate consumed by the circular orifice is 98% of that of the scoop orifice.

[0088] In summary, a circular orifice, consuming 98% of the coolant mass flow rate of a scoop-shaped orifice, still achieves 1.5 times the coolant coverage area and 1.13 times the reduction in average net heat flux density. This indicates that in hypersonic flows, the thermal protection performance of film cooling is sensitive to orifice shape, and optimizing structural parameters has engineering value.

[0089] It is worth emphasizing that in subsonic flows, scoop-shaped orifices offer superior thermal protection compared to circular orifices, a well-established fact. However, in hypersonic flows, the thermal protection performance of scoop-shaped orifices is inferior to that of circular orifices, contrary to the trend observed in subsonic flows. Therefore, optimizing orifice shapes in hypersonic flows is not common knowledge and requires creative effort.

[0090] Example 2

[0091] In practical engineering, the effectiveness of thermal protection is typically evaluated by the reduction in temperature of the wing and rudder walls after the injection of cold gas. Based on this indicator, the thermal protection performance of this invention was assessed through fluid-structure-thermal coupling numerical simulations. The computational domain is as follows: Figure 10 As shown, its size is three times that of the computational domain of Example 1, and the solver settings are consistent with those of Example 1. Unlike Example 1, Example 2 introduces a solid wall with a thickness of 5 mm. Fifteen cylindrical film vents with a diameter of 2 mm and an angle of 30° are arranged on the leading edge of the wing rudder.

[0092] The computational domain inlet total pressure was set at 22.89 MPa, the total temperature at 2332.8 K, and the outlet static pressure at 5529 Pa, thus ensuring an incoming Mach number of 7, or seven times the speed of sound. These boundary conditions reproduced the hypersonic flight condition at an altitude of 20 km and a Mach number of 7. The total temperature of the cooling gas was 300 K, and the total mass flow rate of the 15 film cooling orifices was 11.6 g / s.

[0093] Figure 12The heat reduction effect of the internal and external composite cooling structure on the leading edge region of the wing rudder is demonstrated. It can be observed that the internal and external composite cooling structure has an excellent heat reduction effect on the wing rudder leading edge. Without the internal and external composite cooling structure, the lowest temperature of the wing rudder leading edge wall reaches 1853 K. After introducing the composite cooling structure, the lowest temperature of the wing rudder leading edge wall drops to 653 K, a reduction of 1200 K (i.e., 65%). Furthermore, under no cooling conditions, the average temperature of the entire wing rudder leading edge wall is 1853 K; with the internal and external composite cooling, the average temperature of the entire wing rudder leading edge wall decreases to 813 K, a reduction of 1040 K (i.e., 56%).

[0094] Based on the analysis of the need for additional thermal protection measures, the wing and rudder thermal protection device based on an internal and external composite cooling structure proposed in this invention has met the thermal protection requirements in most areas, eliminating the need for additional thermal protection measures. This further confirms its superior heat reduction performance. Figure 10 As shown, within the range of Y = 100–450 mm, the wall temperature remains consistently below 900 K, which is lower than the allowable operating temperature of nickel-based superalloys (973 K). Therefore, no additional thermal protection technology is required in this region, and the area requiring no additional thermal protection technology accounts for 78% of the total area of ​​the wing leading edge.

[0095] The above are merely preferred embodiments of the present invention and are not intended to limit the invention. For example, the positions and shapes of the ribs, flow bends, impact holes, and film gas holes can be modified and varied in various ways by those skilled in the art. Any modifications, equivalent substitutions, or improvements made within the spirit and principles of the present invention should be included within the scope of protection of the present invention.

[0096] The thermal protection device based on an internal and external composite cooling structure described in this invention has been applied to turbine blades and combustion chamber walls in the field of aero-engines. However, turbine blades and combustion chamber walls operate in subsonic / transonic mainstreams, and related technological applications and research are concentrated under subsonic / transonic conditions. The technology for applying this device to the thermal protection of hypersonic vehicle wing rudders has not been publicly reported, and its working principle differs fundamentally from that in the aero-engine field: in hypersonic flow, the compressibility of the gas is significantly enhanced, leading to the generation of shock waves. The emergence of shock waves causes the wall heat transfer characteristics and the physical mechanism of the interaction between the coolant and the mainstream to be fundamentally different from those in subsonic / transonic flow. Therefore, extending the internal and external composite cooling structure to hypersonic wing rudder thermal protection is not a simple technology transfer, but involves the exploration and innovation of a series of new physical mechanisms, providing a new technical solution for the aerodynamic thermal protection design of hypersonic vehicles.

[0097] Compared to existing hypersonic wing fin thermal protection technologies, the core difference of this invention lies in the fact that it not only incorporates mass ejection channels on the wing fin surface but also integrates multiple cooling channels with enhanced heat transfer structures within the wing fin, including spoiler ribs, bends, ribbed channels, and coupling plate cooling structures. This design utilizes the heat insulation effect of the coolant on the surface while significantly improving the enhanced heat transfer effect inside the wing fin and increasing the coolant's heat absorption capacity.

[0098] The preferred embodiments of the present invention have been described above. It should be understood that the present invention is not limited to the specific embodiments described above, and the devices and structures not described in detail should be understood as being implemented in a conventional manner in the art. Any person skilled in the art can make many possible variations and modifications to the technical solutions of the present invention using the methods and techniques disclosed above, or modify them into equivalent embodiments with equivalent changes, without departing from the scope of the present invention. This does not affect the essential content of the present invention. Therefore, 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 present invention's technical solutions still fall within the protection scope of the present invention.

Claims

1. A thermal protection device for aircraft wing rudders based on an internal and external composite cooling structure, characterized in that, It includes a cold air source (1), a cold air transport device (2) and an internal and external composite cooling structure (3) connected in sequence. The cold air source (1) and the cold air transport device (2) are located inside the aircraft, and the internal and external composite cooling structure (3) is located inside the wing rudder. The internal and external composite cooling structure (3) includes an internal cooling channel (3.1) for the wing and rudder and a surface mass ejection channel (3.2) for the wing and rudder; wherein, The internal cooling channel (3.1) of the wing rudder is a meandering channel arranged inside the wing rudder. The meandering channel includes a first channel (3.1.1), a second channel (3.1.2), and a third channel (3.1.3) connected in sequence along the gas flow direction. The first channel (3.1.1) is connected to the cold air transport device (2). The third channel (3.1.3) is parallel to the outer surface of the leading edge of the wing rudder and is connected to the surface mass ejection channel (3.2) of the wing rudder. The second channel (3.1.2) is connected between the first channel (3.1.1) and the third channel (3.1.3). The first channel (3.1.1) is provided with an array of distributed turbulence ribs (3.1.1.1). The second channel (3.1.2) is provided with a bend (3.1.2.1) and several ribs (3.1.2.2) along the gas flow direction. The surface mass ejection channel (3.2) of the wing rudder is a discrete air film hole array set on the outer surface of the leading edge of the wing rudder; The cold air output from the cold air source (1) is transported to the inner and outer composite cooling structure (3) through the cold air transport device (2). The cold air flows through the internal cooling channel (3.1) of the wing rudder and is then discharged through the discrete air film hole array (3.2) to form a covering air film on the outer surface of the wing rudder to protect the wing rudder.

2. The aircraft wing and rudder thermal protection device according to claim 1, characterized in that, The cold air source (1) is one of the following or a combination thereof: a cold air source formed by cooling the incoming air through expansion, a cold air source formed by precooling the incoming air using fuel, or a cold air source output from an air cylinder carried inside the aircraft.

3. The aircraft wing and rudder thermal protection device according to claim 1, characterized in that, The cold air transport device (2) includes a cold air pipeline (2.1), a control valve (2.2), and a cold air buffer chamber (2.3). The cold air duct (2.1) is a pipe with a circular, rectangular or triangular cross section. The cold air duct (2.1) is connected to the cold air source (1) and the internal and external composite cooling structure (3). On the cold air duct (2.1), a control valve (2.2) and a cold air buffer chamber (2.3) are sequentially installed along the gas flow direction. The control valve (2.2) is used to dynamically adjust the pressure and mass flow rate of the cooling air in real time. The cold air buffer chamber (2.3) is used to rectify and decelerate the cold air to form an approximate stagnant flow condition. Temperature sensors and pressure sensors are installed in the cold air buffer chamber (2.3) to monitor the temperature and pressure of the cold air, respectively.

4. The aircraft wing and rudder thermal protection device according to claim 1, characterized in that, The first channel (3.1.1) is a rectangular cavity distributed along the rudder height direction, assuming the rudder height direction is longitudinal. The turbulence ribs in the first channel (3.1.1) 3.1.1.1) are arranged in an array in the longitudinal direction, with turbulence ribs ( 3.1.1.1) Distributed side-by-side or staggered along the gas flow direction; Each baffle rib (3.1.1.1) is horizontally fixed to the first channel ( Within 3.1.1), one or both ends of the turbulence rib (3.1.1.1) are connected to the first channel ( 3.1.1) Fixed connection to the inner wall, The cross-section of the turbulence rib (3.1.1.1) is circular, rhomboid or elliptical, with the tip of the rhomboid or the short axis of the ellipse facing the air inlet side.

5. The aircraft wing and rudder thermal protection device according to claim 4, characterized in that, The second channel (3.1.2) is a narrow strip channel with a sweep angle. The two ends of the second channel (3.1.2) are respectively arranged with bends (3.1.2.1) to connect the first channel (3.1.1) and the third channel (3.1.3). A row of ribs (3.1.2.2) is provided on both sides of the inner wall of the second channel (3.1.2) along the gas flow direction. The ribs (3.1.2.2) on both sides of the inner wall of the second channel (3.1.2) are arranged side by side or in a staggered arrangement. The rib (3.1.2.2) is one or more combinations of continuous V-ribs, discontinuous V-ribs, straight ribs, discontinuous straight ribs or W-ribs, with the tips of the continuous V-ribs, discontinuous V-ribs or W-ribs facing the air inlet side.

6. The aircraft wing and rudder thermal protection device according to claim 4, characterized in that, The third channel (3.1.3) is parallel to the outer surface of the wing rudder leading edge, and a coupling plate cooling structure (3.1.4) is provided between the third channel (3.1.3) and the outer surface of the wing rudder leading edge. The coupling plate cooling structure (3.1.4) has a hollow cavity (3.1.4.2) inside, and a few turbulence ribs (3.1.4.3) are provided in the hollow cavity (3.1.4.2). The turbulence ribs (3.1.4.3) are cylindrical or rhomboid columns arranged side by side or in a staggered arrangement in the hollow cavity (3.1.4.2). The hollow chamber (3.1.4.2) is connected to the third channel (3.1.3) by an array of impact holes. 3.1.4.1) Connected; The other side of the hollow chamber (3.1.4.2) is connected to the atmosphere through a discrete air film hole array (3.2), and each discrete air film hole is set at an angle or vertically between the hollow chamber (3.1.4.2) and the outer surface of the wing leading edge.

7. The aircraft wing and rudder thermal protection device according to claim 6, characterized in that, The angle between each discrete film air hole and the leading edge of the wing rudder is an acute angle.

8. The aircraft wing and rudder thermal protection device according to claim 6, characterized in that, Discrete air film pores can be round, scoop-shaped, fan-shaped, or dumbbell-shaped.

9. An aircraft, characterized in that, The aircraft is equipped with a wing and rudder thermal protection device according to any one of claims 1-8.

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