Three-dimensional bump structure for a wing, control method and system

Abstract

The application provides a three-dimensional bulging structure, a control method and a system for an airfoil. The three-dimensional bulging structure comprises: a composite skin, first driving members and second driving members, the first driving members and the second driving members are arranged in a thickness direction of the composite skin in a laminated manner, wherein the first driving members and the second driving members each comprise a plurality of shape memory alloy (SMA) wires and a plurality of composite round rods, the SMA wires and the composite round rods of the first driving members are arranged in a first direction, and the SMA wires and the composite round rods of the second driving members are arranged in a second direction; and a temperature control assembly comprising heating wires connected with the SMA wires, the heating wires are used for heating the SMA wires to trigger the SMA wires to generate deformation, drive the composite round rods to be pressed and bent outward, and further drive the composite skin to form a three-dimensional bulging curved surface.

Description

Technical Field

[0001] This invention relates to the field of aviation, and more specifically, to a three-dimensional bulge structure, control method, and system for an aircraft wing. Background Technology

[0002] Drag reduction optimization for aircraft has long been a research hotspot in the field of aerodynamics. The current mainstream cruise condition for civil airliners is high subsonic flight. Under this condition, the local airflow velocity on the upper surface of the wing will exceed the speed of sound before the flight speed, forming a local supersonic flow region, which in turn generates local shock waves. When the flight speed reaches the critical Mach number, the shock wave drag will increase significantly compared to the low-speed condition, resulting in a significant decrease in the overall aerodynamic performance of the aircraft.

[0003] To reduce shock wave drag under high subsonic conditions, existing technologies widely employ shock wave control bulge drag reduction schemes. This scheme does not require significant modifications to the original airfoil structure, nor does it add additional viscous drag. By adjusting the flow field structure near the shock wave through bulges placed in the shock wave generation area of ​​the airfoil, the original normal shock wave with large energy loss is transformed into an oblique shock wave with lower energy loss, thereby achieving the effect of reducing shock wave drag.

[0004] Currently, most publicly available shock wave control bump schemes adopt fixed structures or two-dimensional deformable structures. These bumps can achieve good shock wave control effects near their design operating point. However, during actual flight, parameters such as Mach number, angle of attack, and flight altitude will change dynamically within a large range. Existing bump schemes are prone to generating additional aerodynamic losses under flight conditions that deviate from the design operating point, and may even lead to an unexpected increase in the overall drag of the aircraft. This limits the application scenarios and overall benefits of shock wave drag reduction technology to a certain extent. Summary of the Invention

[0005] This summary is provided to introduce, in a simplified form, some concepts that will be further described in the following detailed description. This summary is not intended to identify key or essential features of the claimed subject matter, nor is it intended to help determine the scope of the claimed subject matter.

[0006] One of the objectives of this invention is to provide a three-dimensional bulge structure, control method, and system for airfoils. Employing a three-dimensional adaptive bulge design based on shape memory alloys, and addressing the bulge configuration requirements under different flight conditions, the invention utilizes the two-way shape memory effect of the two-way shape memory alloy. Temperature control drives the SMA filaments to deform, correspondingly forming a three-dimensional bulge shape that matches the current operating conditions, thus achieving adaptive airfoil shock wave control over a wide operating range.

[0007] According to one aspect of the present invention, a three-dimensional bulge structure for an airfoil is provided, the three-dimensional bulge structure comprising: a composite material skin including a first driving member and a second driving member, the first driving member and the second driving member being stacked along the thickness direction of the composite material skin, wherein each of the first driving member and the second driving member includes multiple shape memory alloy SMA wires and multiple composite material rods, the SMA wires and composite material rods of the first driving member being arranged along a first direction, and the SMA wires and composite material rods of the second driving member being arranged along a second direction; and a temperature control assembly including heating wires connected to the SMA wires, the heating wires being used to heat the SMA wires to trigger deformation of the SMA wires, causing the composite material rods to bend outward under pressure, thereby driving the composite material skin to form a three-dimensional bulge surface.

[0008] In one embodiment of the present invention, the first direction is the chord direction of the wing, and the second direction is the span direction of the wing.

[0009] In one embodiment of the present invention, the first driving member and the second driving member are fixed and limited by a composite material matrix.

[0010] In one embodiment of the present invention, the SMA filaments in the first driving member are arranged alternately with the composite material rods, and the SMA filaments in the second driving member are also arranged alternately with the composite material rods.

[0011] In one embodiment of the present invention, the SMA wires and composite material rods in the first driving member are arranged in an equidistant, staggered pattern, and the SMA wires and composite material rods in the second driving member are also arranged in an equidistant, staggered pattern.

[0012] In one embodiment of the present invention, one or more of the following: the diameter of the SMA filament, the diameter of the composite material rod, and the spacing and number of the SMA filament and the composite material rod, are matched with the shock wave intensity and area size of the arrangement area of ​​the three-dimensional bulge structure.

[0013] In one embodiment of the present invention, in the arrangement area where the shock wave intensity is higher than a preset threshold, the spacing between the SMA wires and the composite material rods is reduced; the number of SMA wires and composite material rods is increased; and / or the diameter of the SMA wires and the diameter of the composite material rods are increased.

[0014] In one embodiment of the present invention, the SMA wire is a two-way shape memory alloy wire, and the initial temperature of the SMA wire's inverse martensitic transformation is higher than the highest structural temperature that the wing surface can reach due to changes in the external environment.

[0015] In one embodiment of the invention, a composite material skin is arranged on the wing surface, and the initial curvature of the composite material rod matches the curvature of the wing surface.

[0016] According to another aspect of the present invention, a wing structure is provided, the wing structure comprising: a wing base; and a three-dimensional bulge structure of the present invention, wherein a composite material skin in the three-dimensional bulge structure is disposed on the surface of the wing base, and the direction in which the composite material rod in the three-dimensional bulge structure buckles outward under pressure corresponds to the outer side of the wing base.

[0017] According to another aspect of the invention, an aircraft is provided, the aircraft comprising: the wing structure of the invention; and a three-dimensional bulge control unit configured to determine flight conditions based on the aircraft's flight parameters and to determine the heating power of the heating wires based on a target bulge configuration corresponding to the flight conditions.

[0018] According to another aspect of the present invention, a three-dimensional bulge control method for an airfoil is provided. This method is used in the three-dimensional bulge structure of the present invention. The three-dimensional bulge control method includes: acquiring flight parameters; determining flight conditions based on the flight parameters; and obtaining a target bulge configuration according to the determined flight conditions, wherein the target bulge configurations corresponding to different flight conditions are three-dimensional curved surfaces with different bulge heights; determining the heating power of a heating wire based on the target bulge configuration; and applying the determined heating power to the heating wire to heat the SMA filament, thereby triggering deformation of the SMA filament, causing the composite material rod to buckle outward under pressure, and thus driving the composite material skin to form the target bulge configuration.

[0019] In one embodiment of the present invention, determining the flight condition based on flight parameters includes determining the flight condition according to a flight parameter-flight condition mapping table; obtaining the target bulge configuration based on the determined flight condition includes obtaining the target bulge configuration according to a flight condition-bulge configuration mapping table; and determining the heating power of the heating wires based on the target bulge configuration includes determining the heating power of each heating wire in the heating wires according to a bulge configuration-heating power mapping table.

[0020] According to another aspect of the present invention, a three-dimensional bulge control system for an aircraft wing is provided, the three-dimensional bulge control system comprising: a memory; and at least one processor communicatively coupled to the memory, the at least one processor being configured to perform the three-dimensional bulge control method of the present invention.

[0021] These and other features and advantages will become apparent from the following detailed description and with reference to the accompanying drawings. It should be understood that the foregoing general description and the following detailed description are illustrative only and do not limit the scope of the claims. Attached Figure Description

[0022] To gain a more detailed understanding of the manner in which the features of the present invention are described above, reference can be made to various embodiments to provide a more specific description of the above-briefly summarized aspects, some of which are illustrated in the accompanying drawings. However, it should be noted that the drawings illustrate only certain typical aspects of the invention and should not be considered as limiting its scope, as this description may allow for other equivalent and effective aspects.

[0023] Figure 1a and Figure 1b This is a schematic diagram of a wing structure according to an embodiment of the present invention.

[0024] Figure 2a and Figure 2b This is a schematic diagram of a three-dimensional bulge structure for an aircraft wing according to an embodiment of the present invention.

[0025] Figure 3 This is a cross-sectional view of a three-dimensional bulge structure for an aircraft wing according to an embodiment of the present invention.

[0026] Figure 4 This is a schematic diagram of the driving principle of a three-dimensional bulge structure for an airfoil according to an embodiment of the present invention.

[0027] Figure 5 This is a deformation diagram of a three-dimensional bulge structure for an aircraft wing according to an embodiment of the present invention.

[0028] Figure 6 This is a flowchart of a three-dimensional bulge control method for an airfoil according to an embodiment of the present invention.

[0029] Figure 7 This is a schematic diagram of a three-dimensional bulge control system for an aircraft wing according to an embodiment of the present invention.

[0030] The accompanying drawings are not drawn to scale. Detailed Implementation

[0031] Embodiments of this disclosure will now be described in more detail with reference to the accompanying drawings. While some embodiments of this disclosure are shown in the drawings, it should be understood that this disclosure can be implemented in various forms and should not be construed as limited to the embodiments set forth herein. Rather, these embodiments are provided to provide a more thorough and complete understanding of this disclosure. It should be understood that the accompanying drawings and embodiments of this disclosure are for illustrative purposes only and are not intended to limit the scope of protection of this disclosure.

[0032] In the description of this disclosure, it should be noted that, unless otherwise stated, "a plurality of" means two or more (including two); the terms "upper," "lower," "left," "right," "inner," "outer," etc., indicating orientation or positional relationships are only for the convenience of describing this disclosure 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, and therefore should not be construed as a limitation of this disclosure. Furthermore, the terms "first," "second," "third," etc., are used for descriptive purposes only and should not be construed as indicating or implying relative importance. "Vertical" is not vertical in the strict sense, but within the allowable tolerance range. "Parallel" is not parallel in the strict sense, but within the allowable tolerance range.

[0033] The directional terms used in the following description refer to the directions shown in the figures and are not intended to limit the specific structure of this disclosure. It should also be noted that, unless otherwise explicitly specified and limited, the terms "installation," "connection," and "joining" should be interpreted broadly. For example, they can refer to a fixed connection, a detachable connection, or an integral connection; they can refer to a direct connection or an indirect connection through an intermediate medium. Those skilled in the art can understand the specific meaning of the above terms in this disclosure based on the specific circumstances.

[0034] In this document, the term "embodiment" means that a particular feature, structure, or characteristic described in connection with an embodiment may be included in at least one embodiment of this disclosure. The appearance of this phrase in various places throughout the specification does not necessarily refer to the same embodiment, nor is it a separate or alternative embodiment mutually exclusive with other embodiments. It will be explicitly and implicitly understood by those skilled in the art that the embodiments described herein can be combined with other embodiments.

[0035] In the description of the embodiments of this disclosure, the term "and / or" is merely a description of the relationship between related objects, indicating that three relationships can exist. For example, A and / or B can represent: A existing alone, A and B existing simultaneously, or B existing alone. Additionally, the character " / " in this document generally indicates that the preceding and following related objects have an "or" relationship.

[0036] Figure 1a and Figure 1b A schematic diagram of a wing structure 100 according to an embodiment of the present invention is shown. It should be understood that... Figure 1a and Figure 1b The wing structure 100 shown is merely exemplary and not limiting. The wing structure of the present invention may include, but is not limited to, other types of wing structures. Figure 1a and Figure 1b The wing structure shown has 100 or more components.

[0037] like Figure 1aand Figure 1b As shown, the wing structure 100 may include a wing base 110 and a three-dimensional bulge structure 120.

[0038] In this embodiment, the three-dimensional bulge structure 120 can be arranged on the surface of the wing base 110, for example, it can be set in the high-shock-occurrence area near the leading edge, corresponding to the location where local shock waves frequently occur under high subsonic cruise conditions. As an example and not a limitation, the three-dimensional bulge structure 120 can be arranged on the upper surface of the wing base 110. It should be understood that this embodiment only exemplifies a single three-dimensional bulge structure 120, but in actual applications, multiple three-dimensional bulge structures 120 can be arranged along the spanwise direction according to the wing's spanwise shock wave distribution characteristics, wing size, and other requirements, all of which fall within the scope of this invention.

[0039] The three-dimensional bulge structure 120 may include a composite material skin. In one embodiment, the three-dimensional bulge structure 120 may be mounted on the surface of an existing wing, for example, by bonding it to the wing surface. In another embodiment, the three-dimensional bulge structure 120 may be integrated as part of the wing skin, i.e., the three-dimensional bulge structure 120 replaces a portion of the original wing surface.

[0040] When the three-dimensional bulge structure 120 is deformed under pressure, it bulges outwards onto the wing, effectively interfering with shock wave formation. The three-dimensional bulge structure 120 can be driven to produce different degrees of deformation through temperature control, corresponding to form a three-dimensional bulge shape that matches the current operating conditions, thereby achieving adaptive wing shock wave control over a wide operating range.

[0041] In one embodiment, the wing structure 100 described above can be applied to aircraft, such as civil airliners, cargo planes, and other high-subsonic aircraft. The aircraft may also include a three-dimensional bulge control unit, which can reuse the computing power module of the aircraft's existing flight control computer or be set as an independent embedded unit dedicated to shock wave control. The three-dimensional bulge control unit can be configured to determine the current flight conditions based on the aircraft's flight parameters (such as Mach number, angle of attack, flight altitude, etc. collected by airborne sensors) and determine the heating power of the heating wires based on the target bulge configuration corresponding to the current flight conditions, so as to drive the three-dimensional bulge structure 120 to form a bulge surface that matches the current conditions.

[0042] The following combination Figures 2a to 3 The three-dimensional bulge structure is described in detail. Figure 2a and Figure 2b A schematic diagram of a three-dimensional bulge structure 120 for an aircraft wing according to an embodiment of the present invention is shown. Figure 3 A cross-sectional view of a three-dimensional bulge structure 120 for an aircraft wing according to an embodiment of the present invention is shown. It should be understood that... Figures 2a to 3 The three-dimensional bulge structure 120 shown is merely exemplary and not limiting. The three-dimensional bulge structure of the present invention may include structures larger than... Figures 2a to 3 The three-dimensional bulge structure shown has 120 or fewer components.

[0043] like Figures 2a to 3 As shown, the three-dimensional bulge structure 120 may include a composite material skin and a temperature control component.

[0044] The following is an explanation of composite material skin.

[0045] like Figure 2a As shown, the composite material skin may include a first drive member 210 and a second drive member 220. For example... Figure 3 As shown, the first driving member 210 and the second driving member 220 can be stacked along the thickness direction of the composite material skin, with the first driving member 210 arranged along a first direction and the second driving member 220 arranged along a second direction. The first direction and the second direction are different. Preferably, the first direction and the second direction are orthogonal or nearly orthogonal. This arrangement of dual driving members can achieve independent deformation control of the bulge along two directions, forming a continuous three-dimensional curved surface configuration.

[0046] In one example, referring to Figure 1, the first direction can be the chord direction of the wing, and the second direction can be the spanwise direction of the wing. It should be understood that the first and second directions can be adjusted to other directions according to the actual installation scenario of the bulge, adapting to the needs of different airfoils or non-wing-type shock wave control scenarios. Different angle arrangements can be used under different operating conditions, as long as independent deformation control in two dimensions can be achieved, all of which fall within the scope of this invention.

[0047] In one example, such as Figure 3 As shown, the first driving component 210 and the second driving component 220 can be fixed and positioned by a composite material matrix. This fixing method requires no additional connectors and ensures uniform force transmission between layers. Furthermore, the stacking order of the first driving component 210 and the second driving component 220 is interchangeable. For example, it can be adjusted according to deformation priority requirements. Specifically, placing the chordal driving component on the upper layer can improve the control response speed of the chordal curvature, and placing the spanwise driving component on the upper layer can improve the control accuracy of the spanwise bulge distribution, thus adapting to the shock wave control requirements of different models.

[0048] like Figure 2aAs shown, the first driving component 210 and the second driving component 220 may each include a shape memory alloy (SMA) wire 206 and a composite material rod 208. The SMA wire 206 and the composite material rod 208 of the first driving component 210 are arranged along a first direction, while the SMA wire 206 and the composite material rod 208 of the second driving component 220 are arranged along a second direction. They can work together to achieve a "drive-load" function, meaning the SMA wire 206 provides the deformation driving force, and the composite material rod 208 ensures the stiffness and direction of the deformation are controllable. For example, the composite material skin can be based on a flexible matrix material, with the composite embedded composite material rod 208 and SMA wire 206 integrally molded, eliminating the need for an additional independent transmission structure and resulting in higher overall integration. The composite material rod 208 serves as the load-bearing component, and the SMA wire 206 serves as the driving component, working collaboratively within the driving components in two orthogonal directions.

[0049] The direction in which the composite material rod in the three-dimensional bulge structure 120 buckles outward under pressure corresponds to the outer side of the wing matrix 110. This buckling direction design ensures that the three-dimensional bulge structure bulges outward when deformed under pressure, which can effectively interfere with shock wave formation without squeezing the original structures such as pipes and cables inside the wing, thus avoiding affecting the basic functions of the wing.

[0050] In one example, the SMA wires 206 and composite material rods 208 in the first drive component 210 are arranged in an alternating pattern, as are the SMA wires 206 and composite material rods 208 in the second drive component 220. For instance, taking an airfoil scenario, when the first drive component 210 is a chordal drive component and the second drive component 220 is a spanwise drive component, the chordal drive component is composed of chordal SMA wires 206 and chordal composite material rods 208 arranged in an alternating pattern along the chordal direction, and the spanwise drive component is composed of spanwise SMA wires 206 and spanwise composite material rods 208 arranged in an alternating pattern along the spanwise direction. The two are stacked to form an orthogonal drive frame. This orthogonal drive frame can achieve independent deformation control of the bulge in both the chordal and spanwise dimensions, and can flexibly adjust the chordal curvature and spanwise distribution of the bulge to form a three-dimensional curved surface that matches the shock wave gradient distribution on the airfoil surface, thereby improving the shock wave drag reduction effect under a wide range of operating conditions.

[0051] In further examples, such as Figure 2a As shown, the SMA wires 206 and composite material rods 208 in the first driving component 210 are arranged in an equidistant, staggered pattern, as are the SMA wires 206 and composite material rods 208 in the second driving component 220. This arrangement can achieve uniform deformation of the bulge surface, adapting to the requirements of conventional areas with uniform shock wave distribution.

[0052] In another example, one or more of the following factors—the diameter of the SMA wire 206, the diameter of the composite material rod 208, and the spacing and number of SMA wires 206 and composite material rods 208—are matched (e.g., positively correlated) with the shock wave intensity and region size of the arrangement area of ​​the three-dimensional bulge structure 120. For example, in a further example, in an arrangement area where the shock wave intensity is higher than a preset threshold, compared to an arrangement area where the shock wave intensity is lower than or equal to the preset threshold, the spacing between the SMA wires 206 and composite material rods 208 is reduced to improve deformation control accuracy; the number of SMA wires 206 and composite material rods 208 is increased to improve local driving force; and / or the diameter of the SMA wires 206 and the diameter of the composite material rods 208 are increased to improve structural load-bearing capacity, thereby adapting to the high aerodynamic load requirements of the strong shock wave region.

[0053] In one example, the SMA wire 206 can be a two-way shape memory alloy wire. However, it should be understood that the SMA wire 206 can also be made of other materials with temperature-triggered deformation characteristics. The composite material rod 208 can be a carbon fiber reinforced composite material. However, it should be understood that multiple composite material rods 208 can also be made of other composite materials that meet stiffness requirements, such as glass fiber and aramid fiber, or lightweight alloy materials with suitable stiffness. In addition, the initial temperature of the martensitic reverse phase transformation of the SMA wire 206 is higher than the highest structural temperature that the wing surface can reach due to changes in the external environment, in order to avoid unintended triggering of the phase transformation of the SMA wire 206, ensuring that the deformation is controlled only by active heating, and improving the reliability of regulation.

[0054] The following is an explanation of the temperature control components.

[0055] like Figure 2b As shown, the temperature control component may include a heating wire 230 connected to the SMA filament 206. The heating wire 230 is used to heat the SMA filament 206 to trigger a phase change and deformation of the SMA filament 206, causing the composite material rod 208 to buckle outward under pressure, thereby driving the composite material skin to form a three-dimensional bulge surface that matches the current working conditions.

[0056] In one example, such as Figure 2b As shown, each SMA wire 206 can be connected to a corresponding heating wire 230 at both ends, so that different heating wires 230 can independently adjust the heating power of the corresponding SMA wire 206 without interfering with each other. In one example, the heating wire 230 can be arranged in the same layer as the corresponding SMA wire 206. By being arranged in the same layer as the driving component, there is no need for cross-layer wiring, which can reduce heat conduction loss and improve heating efficiency.

[0057] In one example, referring to Figure 1, the composite material skin can be arranged on the wing surface, such as the shock wave generation area on the upper surface of the wing, and the initial curvature of the composite material rod 208 matches the curvature of the wing surface (such as matching the curvature of the wing surface of the arrangement area), so that multiple composite material rods 208 are pressed towards the outside of the wing and buckled, which ensures that the bulge can effectively interfere with the shock wave flow field, and will not squeeze the original pipelines and cable structures inside the wing.

[0058] It should be understood that, depending on the actual situation, the above examples can be implemented individually or in combination.

[0059] As can be seen, the three-dimensional bulge structure of the present invention can form a three-dimensional bulge surface that adapts to different flight conditions through the independent deformation control capability of the dual orthogonal direction driving components. This breaks through the limitations of existing fixed bulges and two-dimensional deformable bulges in terms of adaptability to different flight conditions. It can achieve a stable shock wave drag reduction effect in all flight phases, including takeoff, climb, cruise, and landing, effectively improving the aerodynamic benefits of high subsonic aircraft and reducing operating costs.

[0060] The following combination Figure 4 and Figure 5 The bulge driving and deformation process of one embodiment of the present invention will be described in detail. Figure 4 A schematic diagram illustrating the driving principle of a three-dimensional bulge structure for an airfoil according to an embodiment of the present invention is shown. Figure 5 A schematic diagram of the deformation of a three-dimensional bulge structure for an aircraft wing according to an embodiment of the present invention is shown.

[0061] It should be understood that this embodiment uses orthogonal driving in the chord and spanwise directions as an example for explanation. In practical applications, other driving arrangements in different directions can also be adopted according to the characteristics of shock wave distribution, all of which are within the scope of this invention.

[0062] Combination Figure 4 and Figure 5 During the driving process, when it is necessary to adjust the bulge configuration, current is passed through the heating wire 230 to heat the SMA wire 206 at the corresponding position. The SMA wire 206 is heated and triggers a phase change and generates shrinkage deformation, which applies axial pressure to the composite material round rod 208 arranged in the same direction. When the pressure exceeds the instability critical stress of the composite material round rod 208, the composite material round rod 208 buckles and deforms, which in turn drives the composite material skin to bulge outward to form a bulge surface.

[0063] Because the upper wing panel itself has inherent curvature, the initial curvature of the composite material rod 208 matches the wing curvature of the installation area. Therefore, when it is under pressure and becomes unstable, it will preferentially bend towards the outer surface of the wing, ensuring that the bulge protrusion direction meets the shock wave control requirements and will not deform and squeeze the original internal structure towards the inner side of the wing.

[0064] Because the composite material skin has SMA wires 206 and composite material rods 208 arranged in two orthogonal directions in the chord and span directions, the shape memory alloy wires 206 in different directions and positions can be independently heated and controlled to achieve independent buckling deformation in the chord and span directions. The independent buckling deformation of the chord and span drive components are superimposed on each other, and finally a three-dimensional bulge surface matching the shock wave distribution under the current flight conditions is formed on the skin surface, so as to achieve the effect of weakening the shock wave intensity and reducing the shock wave drag.

[0065] It should be understood that the above driving process is only an illustrative example. In actual applications, the heating sequence and power of each SMA wire can be adjusted according to the working conditions to adapt to different bulge deformation rates and configuration requirements, all of which fall within the scope of this invention.

[0066] Figure 6 A flowchart of a three-dimensional bulge control method 600 for an airfoil according to an embodiment of the present invention is shown. The three-dimensional bulge control method 600 can be used in the three-dimensional bulge structure 120 described above to achieve shock wave drag reduction under all flight conditions by adaptively adjusting the bulge configuration.

[0067] like Figure 6 As shown, the three-dimensional bulge control method 600 may include the following steps: In step 602, flight parameters can be acquired, flight conditions can be determined based on these parameters, and the target bulge configuration can be obtained according to the determined flight conditions. The target bulge configuration corresponding to different flight conditions is a three-dimensional curved surface with different bulge heights. Different three-dimensional curved surfaces can have different surface curvatures to accurately match the shock wave distribution characteristics of the current flight conditions.

[0068] In one example, the flight condition can be determined based on a flight parameter-flight condition mapping table. Specifically, flight parameters may include parameters such as flight Mach number, angle of attack, altitude, and atmospheric density. Other auxiliary parameters, such as the rate of change of Mach number, rate of change of angle of attack, fuel load, and wing vibration signals, can also be introduced according to control accuracy requirements to improve the accuracy of condition identification. The flight condition can be determined based on a preset flight parameter-flight condition mapping table, such as dividing it into typical conditions like takeoff and climb, subsonic cruise, transonic transition, and landing. It can also be refined into more condition ranges based on aerodynamic characteristics, or optimized only for the single cruise condition, all of which fall within the scope of this invention.

[0069] In one example, the target bulge configuration can be obtained based on a flight condition-bulge configuration mapping table. Specifically, the target bulge configuration can be obtained, for example, based on a flight condition-bulge configuration mapping table calibrated in advance through numerical simulation or wind tunnel testing. The flight condition-bulge configuration mapping table can store configuration parameters such as chord curvature, spanwise bulge height, and bulge coverage area for each flight condition.

[0070] In step 604, the heating power of the heating wires can be determined based on the target bulge configuration. By differentially controlling the power of the heating wires at different locations, fine-tuning of the bulge surface can be achieved.

[0071] In one example, the heating power of each heating wire in the heating conductor can be determined based on a bulge configuration-heating power mapping table. For instance, the bulge configuration-heating power mapping table can be pre-calibrated and can store the heating power thresholds required for different regions of the SMA wire to achieve the target deformation. The power of the heating wire at the corresponding position can be assigned according to the curvature distribution of the target configuration.

[0072] It should be understood that the three-level lookup table logic in this embodiment is a preferred solution. In practical applications, the use of each mapping table in the above example (such as the flight parameter-flight condition mapping table, the flight condition-bulge configuration mapping table, and the bulge configuration-heating power mapping table) can be flexibly combined according to actual needs (such as according to airborne computing power resources, condition response requirements, etc.), and it is not necessary to use all of them.

[0073] In step 606, a determined heating power can be applied to the heating wire to heat the SMA filament, triggering deformation of the SMA filament. This causes the composite material rod to buckle outward under pressure, thereby driving the composite material skin to form the target bulge configuration and achieving shock wave control. Specifically, the shock wave flow field can be disturbed by the three-dimensional bulge surface to achieve the control effect of weakening the shock wave intensity and reducing the shock wave drag.

[0074] It should be understood that the process in this embodiment is merely exemplary, and each step can be adjusted, modified, added, and / or deleted according to actual needs. For example, when computing power is sufficient, the intermediate table lookup step can be omitted, and the target bulge configuration or heating power can be directly calculated from flight parameters; when optimizing only for fixed cruise conditions, a fixed heating power threshold can be preset, omitting the condition matching and configuration retrieval steps; a real-time feedback correction step for the bulge configuration can also be added, using pressure sensors arranged on the skin surface to collect aerodynamic load data, dynamically fine-tuning the target bulge configuration, and further improving shock wave control accuracy, etc.

[0075] As can be seen, the three-dimensional bulge control method of the present invention can quickly match the optimal bulge configuration according to the real-time flight conditions, realize the adaptive dynamic control of the three-dimensional bulge, without the need for complex real-time flow field calculations, and has a fast response speed and high reliability. It is adapted to the computing power limitations and flight safety requirements of the airborne system, effectively broadens the operating condition coverage of the shock wave control bulge, and can achieve a stable shock wave drag reduction effect in the entire flight phase, thereby improving the operational economy of the aircraft.

[0076] Figure 7A schematic diagram of a three-dimensional bulge control system for an airfoil according to an embodiment of the present invention is shown. The hardware architecture of this embodiment can be directly adapted to the embedded operating environment of the airborne flight control computer, reuse the existing sensor, actuator and other hardware resources of fly-by-wire aircraft, without the need for additional dedicated control units, resulting in low integration costs; it can also adopt an independent shock wave control unit according to the retrofit requirements of older aircraft models, thus having wider adaptability.

[0077] The system may include components that are connected to or communicate with the bus 720. For example, the system may include the bus 720, a processor 705, and one or more memories 710, etc.

[0078] Processor 705 can be any type of processor and may include, but is not limited to, general-purpose processors and / or dedicated processors (e.g., special processing chips), intelligent hardware devices (e.g., general-purpose processors, DSPs, CPUs, microcontrollers, ASICs, FPGAs, programmable logic devices, discrete gate or transistor logic components, discrete hardware components, or any combination thereof). In some cases, processor 705 may be configured to use a memory controller to operate the memory array. In other cases, a memory controller (not shown) may be integrated into processor 705. Processor 705 may be responsible for managing bus 720 and general processing, including executing software 715 stored on memory 710. Processor 705 may also be configured to perform various functions described herein related to the three-dimensional bulge control method for wings. For example, the processor 705 can be configured to: acquire flight parameters, determine flight conditions based on the flight parameters, and obtain a target bulge configuration according to the determined flight conditions, wherein the target bulge configuration corresponding to different flight conditions is a three-dimensional curved surface with different bulge heights; determine the heating power of the heating wire based on the target bulge configuration; and apply the determined heating power to the heating wire, thereby heating one or more SMA filaments among multiple SMA filaments to trigger a phase transition and generate deformation, causing the composite material rod in the corresponding direction to be compressed and buckled outward, thereby driving the composite material skin to form the target bulge configuration and realizing shock wave control.

[0079] Memory 710 can be any storage device capable of storing data. Memory 710 may include, but is not limited to, disk drives, optical storage devices, solid-state storage, floppy disks, hard disks, magnetic tapes or any other magnetic media, optical discs or any other optical media, ROM (Read-Only Memory), RAM (Random Access Memory), cache memory and / or any other memory chip or cartridge, and / or any other medium from which a computer can read data, instructions and / or code. Memory 710 may store computer-executable software 715 including computer-readable instructions that, when executed, cause a processor to perform the various functions described herein. Memory 710 may have various data / instructions / code for implementing the various functions described herein related to the design of a three-dimensional bulge control system for an airfoil. In one example, memory 710 may also store preset flight parameter-flight condition mapping tables, flight condition-bulge configuration mapping tables, and / or bulge configuration-heating power mapping tables.

[0080] Software 715 may be stored in memory 710 and includes, but is not limited to, an operating system, one or more application programs, drivers, and / or other data and code. Instructions for performing the various functions described herein may be included in one or more application programs, and the components of the system may be implemented by processor 705 reading and executing the instructions of one or more application programs. In some cases, software 715 may not be directly executable by the processor, but may (e.g., when compiled and executed) enable the computer to perform the various functions described herein related to the three-dimensional bulge control method for wings.

[0081] It should be understood that the above hardware architecture is merely exemplary. In practical applications, the configuration and connection relationship of hardware components can be adjusted according to airborne safety level and performance requirements, as long as the three-dimensional bulge control function for wings of the present invention can be realized.

[0082] The above describes the solution related to the three-dimensional bulge of the present invention. The improvements of the present invention include at least: (1) Higher structural integration: The integrated design of composite material skin with orthogonal dual drive layer is adopted. The "drive-load" function is achieved by SMA wire and composite material round rod. There is no need to set up an independent transmission mechanism. The overall thickness is thin and the weight is light. It can be adapted and installed without making major changes to the original wing structure. The modification cost is low and the reliability is high.

[0083] (2) Wider range of operating conditions: Through the independent deformation control capability in both chord and span directions, a three-dimensional bulge surface that matches the shock wave distribution characteristics of different flight conditions can be formed. This breaks through the limitation that existing fixed bulges and two-dimensional deformable bulges can only adapt to a single design point. Stable shock wave drag reduction effect can be achieved in all flight stages such as takeoff, climb, cruise and landing, and the average drag reduction benefit under a wide range of operating conditions is significantly improved.

[0084] (3) Simple and reliable control logic: The control logic adopts a pre-calibrated three-mapped table lookup control logic, which does not require complex real-time flow field calculation. It has a fast response speed and low computing power requirements. It can be directly adapted to the existing airborne flight control hardware resources of civil aircraft without the need to add a dedicated control unit. It has strong engineering feasibility. At the same time, the control logic can flexibly adjust the calibration parameters according to the aerodynamic characteristics of different aircraft models, and has wide adaptability.

[0085] The above improvements can effectively reduce the cruise drag of high subsonic aircraft, reduce operational energy consumption and carbon emissions, and have high engineering application value and economic benefits.

[0086] The foregoing description includes examples of various aspects of the claimed subject matter. It is certainly impossible to describe every conceivable combination of components or methods for the purpose of depicting the claimed subject matter, but those skilled in the art will recognize that many further combinations and arrangements of the claimed subject matter are possible. Thus, the disclosed subject matter is intended to cover all such changes, modifications, and variations that fall within the spirit and scope of the appended claims.

Claims

1. A three-dimensional bulge structure for an aircraft wing, characterized in that, The three-dimensional bulge structure includes: A composite material skin includes a first driving member and a second driving member, which are stacked along the thickness direction of the composite material skin. Each of the first and second driving members includes multiple shape memory alloy SMA wires and multiple composite material rods. The SMA wires and composite material rods of the first driving member are arranged along a first direction, and the SMA wires and composite material rods of the second driving member are arranged along a second direction. The temperature control component includes a heating wire connected to the SMA filament. The heating wire is used to heat the SMA filament to trigger deformation, which causes the composite material rod to bend outward under pressure, thereby driving the composite material skin to form a three-dimensional bulging surface.

2. The three-dimensional bulge structure according to claim 1, characterized in that, The first direction is the chord direction of the wing, and the second direction is the span direction of the wing.

3. The three-dimensional bulge structure according to claim 1, characterized in that, The first driving component and the second driving component are fixed and limited by a composite material matrix.

4. The three-dimensional bulge structure according to claim 1, characterized in that, In the first driving component, SMA wires and composite material round rods are arranged alternately, and in the second driving component, SMA wires and composite material round rods are arranged alternately.

5. The three-dimensional bulge structure according to claim 4, characterized in that, In the first driving component, the SMA wires and composite material rods are arranged in an equidistant, staggered pattern, and in the second driving component, the SMA wires and composite material rods are arranged in an equidistant, staggered pattern.

6. The three-dimensional bulge structure according to claim 1, characterized in that, One or more of the following factors—the diameter of the SMA filament, the diameter of the composite material rod, and the spacing and number of SMA filaments and composite material rods—are matched with the shock wave intensity and area size of the arrangement region of the three-dimensional bulge structure.

7. The three-dimensional bulge structure according to claim 6, characterized in that, In the arrangement area where the shock wave intensity is higher than a preset threshold, The spacing between the SMA filaments and the composite material rod is reduced; The number of SMA wires and composite material rods is increased; and / or The diameter of the SMA wire and the diameter of the composite material rod are increased.

8. The three-dimensional bulge structure according to claim 1, characterized in that, The SMA wire is a two-way shape memory alloy wire, and the initial temperature of the martensitic reverse phase transformation of the SMA wire is higher than the highest structural temperature that the wing surface can reach due to changes in the external environment.

9. The three-dimensional bulge structure according to claim 1, characterized in that, The composite material skin is arranged on the wing surface, and the initial curvature of the composite material rod matches the curvature of the wing surface.

10. A wing structure, characterized in that, The wing structure includes: Wing base; and The three-dimensional bulge structure as described in any one of claims 1-9, In this structure, the composite material skin in the three-dimensional bulge structure is arranged on the surface of the wing substrate, and the direction in which the composite material rod in the three-dimensional bulge structure buckles outward under pressure corresponds to the outer side of the wing substrate.

11. An aircraft, characterized in that, The aircraft include: The wing structure as described in claim 10; and A three-dimensional bulge control unit is configured to determine flight conditions based on the aircraft's flight parameters and to determine the heating power of the heating wires based on the target bulge configuration corresponding to the flight conditions.

12. A three-dimensional bulge control method for an airfoil, characterized in that, The three-dimensional bulge control method is used for a three-dimensional bulge structure as described in any one of claims 1-9, and the three-dimensional bulge control method includes: The flight parameters are acquired, the flight conditions are determined based on the flight parameters, and the target bulge configuration is obtained according to the determined flight conditions. The target bulge configuration corresponding to different flight conditions is a three-dimensional curved surface with different bulge heights. The heating power of the heating wire is determined based on the target bulge configuration; and A determined heating power is applied to the heating wire to heat the SMA filament, triggering deformation of the SMA filament, which in turn causes the composite material rod to buckle outward under pressure, thereby driving the composite material skin to form the target bulge configuration.

13. The three-dimensional bulge control method according to claim 12, characterized in that, The flight conditions are determined based on the flight parameter-flight condition mapping table; The target bulge configuration is obtained based on a flight condition-bulge configuration mapping table; and / or The heating power of each heating wire in the heating wire is determined according to the bulge configuration-heating power mapping table.

14. A three-dimensional bulge control system for an aircraft wing, characterized in that, The three-dimensional bulge control system includes: Memory; and At least one processor communicatively coupled to the memory, the at least one processor being configured to perform the three-dimensional bulge control method as described in claim 12 or 13.

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