Aerodynamic layout of sub-span and super-span speed domain of patrolling bomb

Abstract

The sub-span and super-span speed domain aerodynamic layout comprises double-shaft variable-sweep-wing mechanisms arranged at the bottom of a fuselage, the sweep angle can be freely changed and locked within the range of 0-90 degrees, each double-shaft variable-sweep-wing mechanism comprises an inner wing section and an outer wing section, the wing root of each inner wing section is in pivot connection with the fuselage, and the wing root of each outer wing section is in pivot connection with the fuselage. The sweep-forward angle of the inner wing section is adjusted within any angle of 0-90 degrees; the wing root of the outer wing section is pivotally connected with the wing tip of the inner wing section, and the sweepback angle of the outer wing section is adjusted at any angle of 0-90 degrees; the machine body is in a flat shuttle shape, and the machine head is in a pointed-end shape. According to the implementation mode, in the stretching, folding and flying processes, the influence trends of the double-shaft variable-swept-wing mechanism on the pressure center focus of the patrolling bomb are opposite, so that the pressure center focus change range of cross-speed domain deformation flight is controlled within an extremely small interval, and the pneumatic control stability is remarkably improved. In addition, when being stored, the double-shaft variable-swept-wing mechanism can be completely folded and parallel to the fuselage and is adaptive to narrow spaces such as a launch canister or a transport case, so that the storage and deployment capabilities are improved.

Description

Technical Field

[0001] This invention relates to the field of loitering munitions, and in particular to an aerodynamic configuration of a loitering munition in the sub-trans- and super-trans-velocity ranges. Background Technology

[0002] Modern and future flight missions, whether for high-speed penetration military aircraft, long-endurance reconnaissance drones, or aerospace vehicles aiming for rapid global reach, all demand efficient, stable, and controllable flight across a wide speed range—from low-speed takeoff and landing, subsonic cruise (Mach number < 0.8), transonic acceleration (Mach number 0.8-1.4) to supersonic (Mach number > 1.4) and even hypersonic flight. This requirement is known as "wide-range" or "full-speed-range" flight capability.

[0003] However, subsonic flight, which aims for high lift and low drag, typically requires wings with a high aspect ratio (i.e., slender wings), straight wings, or wings with a small sweep angle, to generate sufficient lift and reduce induced drag. Supersonic flight, on the other hand, first needs to overcome shock wave drag, requiring wings with a low aspect ratio, a large sweep angle, or even a triangular shape, to delay shock wave generation and reduce shock wave drag. Traditional fixed-wing configurations can only achieve optimal performance within a specific speed range, failing to consider performance across the entire speed spectrum, thus severely limiting the aircraft's mission adaptability and combat effectiveness.

[0004] Related loitering munitions or drones adopt a single-axis variable sweep design to give them a wide speed range, but single-axis variable sweep loitering munitions still have the following problems: 1. Dramatic Changes in Aerodynamic Center of Pressure and Focal Point: When the wing sweep angle changes, the center of pressure and focal point of the entire loitering munition shift significantly. Considering cost and feasibility, the center of gravity of a loitering munition is usually fixed, which leads to drastic changes in the aircraft's longitudinal stability (pitch attitude). To balance the aircraft, the flight control system needs to constantly trim, increasing the complexity of control and the risk of instability.

[0005] 2. Considering the matching of deformation handling characteristics, wing mounting position and center of gravity, the wing span and aspect ratio are limited: the lift-to-drag ratio performance at both low and high speeds is significantly affected. This restricts the overall aerodynamic efficiency in the sub-, trans-, and super-species speed ranges.

[0006] 3. Occupies a large amount of fuselage space: The single-axis variable sweep wing encroaches on valuable internal space of the fuselage, affecting the layout of other subsystems such as the loitering munition air intake and mission payload.

[0007] 4. Insufficient storage efficiency: Although the wings can be folded back, the overall profile is still large, and its storage size is still not ideal for drones that need to be launched from compact launch devices (such as shipborne vertical launch systems or land-based launch tubes). Summary of the Invention

[0008] The summary section of this invention provides a brief overview of the concepts, which will be described in detail in the detailed description section that follows. This summary section is not intended to identify key or essential features of the claimed invention, nor is it intended to limit the scope of the claimed invention.

[0009] Some embodiments of the present invention provide aerodynamic configurations for loitering munitions in the sub-, trans-, and super-speed ranges to address the technical problems mentioned in the background section above.

[0010] Some embodiments of the present invention provide an aerodynamic layout for a loitering munition in the sub-trans- and super-trans-speed ranges, including a dual-axis variable-sweep wing mechanism disposed at the bottom of the fuselage, forming a low-wing configuration, capable of arbitrarily changing and locking the sweep angle within the range of 0° to 90°, wherein... Each of the aforementioned dual-axis variable sweep wing mechanisms includes an inner wing section and an outer wing section. The wing root of the inner wing section is pivotally connected to the fuselage and can rotate relative to the fuselage, thereby adjusting the forward sweep angle of the inner wing section within any angle from 0° to 90°. The outer wing root is pivotally connected to the inner wing tip, allowing it to rotate relative to the inner wing section, thus adjusting the sweep angle of the outer wing section at any angle from 0° to 90°. The length of the outer wing section is greater than or equal to twice the length of the inner wing section; The fuselage is flat and spindle-shaped, and the head is pointed.

[0011] The above embodiments of the present invention have the following beneficial effects: During the extension, folding, and sweeping processes, the inner and outer wing sections of the dual-axis variable-sweep wing mechanism have opposite effects on the pressure center of the loitering munition, thus compensating for each other. This controls the range of pressure center changes during cross-speed range deformation flight within a very small range, fundamentally improving the stability and maneuverability problems caused by the drastic pressure center changes of traditional single-axis variable-sweep wings, and significantly enhancing aerodynamic control stability.

[0012] In addition, the dual-section folding design allows the inner and outer wing sections of the dual-axis variable sweep wing mechanism to be completely folded and parallel to the fuselage when stored, forming an extremely compact envelope shape. This perfectly fits into small spaces such as weapon launch tubes or transport containers, without occupying valuable internal space. It solves the problem that the lateral dimensions of traditional variable sweep wings are still too large after storage and affect the arrangement of loitering munition air intakes, mission payloads, and other mechanisms, thus improving storage and deployment capabilities. Attached Figure Description

[0013] To more clearly illustrate the specific embodiments of the present invention or the technical solutions in the prior art, the drawings used in the description of the specific embodiments or the prior art will be briefly introduced below. Obviously, the drawings described below are some embodiments of the present invention. For those skilled in the art, other drawings can be obtained from these drawings without creative effort.

[0014] Figure 1 This is a top view of an embodiment of the sub- and super-speed-range aerodynamic layout of the loitering munition of the present invention in its extended state. Figure 2 This is a front view of an embodiment of the sub-trans- and super-trans-velocity aerodynamic layout of the loitering munition of the present invention in an extended state. Figure 3 A top view of a swept-back state of an embodiment of the sub- and super-speed-range aerodynamic layout of the loitering munition of the present invention. Figure 4 A bottom view of the retracted state of an embodiment of the sub- and super-velocity aerodynamic layout of the loitering munition of the present invention. Figure 5 This is a front view of the retracted state of an embodiment of the sub-trans- and super-trans-velocity aerodynamic layout of the loitering munition of the present invention. Figure 6 This is a side view of the retracted state of an embodiment of the sub-trans- and super-trans-velocity aerodynamic layout of the loitering munition of the present invention. Figure 7 This is a schematic diagram showing how the stability of a single-axis variable sweep wing and a dual-axis variable sweep wing mechanism changes with the deployment angle at transonic speeds (e.g., Ma=0.8). Figure 8 Schematic diagrams of the shape of NACA SC(2)-0710 supercritical airfoil, NACA0006 supersonic airfoil and target airfoil; Figure 9 This is a side view of an embodiment of the headstock profile formed according to the present invention; Figure 10 This is a schematic diagram of one embodiment of the fuselage of the present invention; Figure 11 This is a perspective view of an embodiment of the headstock of the present invention; Figure 12 This is a schematic diagram of an embodiment of the von Kármán curves on both sides of the fuselage symmetry plane of the present invention; Figure 13 This is a schematic diagram of an embodiment of the von Kármán curve group of the present invention; Figure 14 This is a schematic diagram of an embodiment of the multi-section curved surface formed by the present invention.

[0015] Figure 15This is a schematic diagram of the loitering munition entity formed according to the present invention.

[0016] Explanation of reference numerals in the attached figures: 11. Nose section; 111. Nose section generatrix; 112. Nose section vertex; 113. Von Kármán curve group; 114. Multi-section curved surface; 12. Fuselage; 121. Fuselage cross-section; 13. Tail; 131. Y-shaped tail fin; 2. Dual-axis variable sweep wing mechanism; 21. Inner wing section; 211. Inner wing section root; 212. Inner wing section tip; 22. Outer wing section; 221. Outer wing section root; 222. Outer wing section tip. Detailed Implementation

[0017] The technical solution of the present invention will be clearly and completely described below with reference to the embodiments. Obviously, the described embodiments are only some embodiments of the present invention, and not all embodiments. Based on the embodiments of the present invention, all other embodiments obtained by those skilled in the art without creative effort are within the scope of protection of the present invention.

[0018] In the description of this invention, it should be understood that the terms "center," "longitudinal," "lateral," "length," "width," "thickness," "upper," "lower," "front," "rear," "left," "right," "vertical," "horizontal," "top," "bottom," "inner," "outer," "clockwise," and "counterclockwise," etc., indicate the orientation or positional relationship based on the orientation or positional relationship shown in the accompanying drawings. They are used only for the convenience of describing this invention and simplifying the description, and do not indicate or imply that the device or element referred to must have a specific orientation, or be constructed and operated in a specific orientation. Therefore, they should not be construed as limiting this invention.

[0019] Furthermore, the terms "first" and "second" are used for descriptive purposes only and should not be construed as indicating or implying relative importance or implicitly specifying the number of indicated technical features. Thus, a feature defined as "first" or "second" may explicitly or implicitly include one or more of the stated features. In the description of this invention, "a plurality of" means two or more, unless otherwise explicitly specified. Furthermore, the terms "installed," "connected," and "linked" should be interpreted broadly; for example, they may refer to a fixed connection, a detachable connection, or an integral connection; they may refer to a mechanical connection or an electrical connection; they may refer to a direct connection or an indirect connection through an intermediate medium; and they may refer to the internal connection of two components. Those skilled in the art can understand the specific meaning of the above terms in this invention based on the specific circumstances.

[0020] This disclosure will now be described in detail with reference to the accompanying drawings and embodiments.

[0021] Please see Figures 1 to 6 The aerodynamic layout of the loitering munition of the present invention in the sub-trans- and super-trans-speed range includes a dual-axis variable-sweep wing mechanism 2 disposed at the bottom of the fuselage 12. Each dual-axis variable-sweep wing mechanism 2 includes an inner wing section 21, an inner wing pivot, an outer wing section 22, and an outer wing pivot.

[0022] The fixed end of the inner wing pivot is connected to the bottom of the fuselage 12, and the driving end of the inner wing pivot is connected to the wing root 211 of the inner wing section. The aforementioned inner wing pivot is used to adjust the forward sweep angle of the inner wing section 21, so that the inner wing section 21 is parallel or perpendicular to the longitudinal axis of the fuselage 12.

[0023] The fixed end of the outer wing pivot is connected to the wingtip 212 of the inner wing section, and the driving end of the outer wing pivot is connected to the wing root 221 of the outer wing section. The aforementioned outer wing pivot is used to adjust the sweep angle of the outer wing section 22, so that the outer wing section 22 is parallel or perpendicular to the longitudinal axis of the fuselage 12.

[0024] By driving the inner wing pivot and the outer wing pivot to rotate respectively, the extension and folding of the dual-axis variable sweep wing mechanism 2 can be realized.

[0025] Specifically, during extension, the inner wing pivot drives the inner wing section tip 212 to rotate around the inner wing section root 211. The inner wing section 21 can rotate from being parallel to the longitudinal axis of the fuselage 12 to being perpendicular to the longitudinal axis of the fuselage 12, and the forward sweep angle can be adjusted from 90° to 0°. Simultaneously, the outer wing pivot drives the outer wing section tip 222 to rotate around the outer wing section root 221. The outer wing section 22 can rotate from being parallel to the longitudinal axis of the fuselage 12 to being perpendicular to the longitudinal axis of the fuselage 12, and the backward sweep angle can be adjusted from 90° to 0°. The inner wing section 21 and the outer wing section 22 transform from a folded state to an inverted V-shaped state, and the dual-axis variable sweep wing mechanism 2 forms an M-shaped overall configuration, finally transforming into the extended state. In this way, the relative deployment angle between the inner wing section 21 and the outer wing section 22 increases from 0° to 180°.

[0026] It should be noted that the inner wing section 21 and the outer wing section 22 can independently change their angles and have asymmetrical deformation capabilities on both sides, so as to change the overall aerodynamic configuration to adapt to different flight speeds from low subsonic to supersonic.

[0027] The inner wing pivot and the outer wing pivot can be a servo or motor with a self-locking function, thereby locking the inner wing section 21 and the outer wing section 22 after adjusting the forward and backward sweep angles.

[0028] After the dual-axis variable-sweep wing mechanism 2 is fully extended, it significantly increases the wing's aspect ratio, thereby improving the aircraft's lift and drag performance. During the extension process, the centroid of the inner wing section 21 shifts rearward, while the centroid of the outer wing section 22 shifts forward. The changes in the wing's center of pressure and focal point are relatively small, resulting in good handling and stability characteristics during the wing's extension process.

[0029] During folding, the inner wing pivot drives the inner wing tip 212 to rotate around the inner wing root 211, causing the inner wing section 21 to rotate from perpendicular to the longitudinal axis of the fuselage 12 to parallel to the longitudinal axis of the fuselage 12, and completely retract into the bottom of the fuselage, with the forward sweep angle varying from 0° to 90°. Simultaneously, the outer wing pivot drives the outer wing tip 222 to rotate around the outer wing root 221, causing the outer wing section 22 to rotate from perpendicular to the longitudinal axis of the fuselage 12 to parallel to the longitudinal axis of the fuselage 12, with the backward sweep angle varying from 0° to 90°. The inner wing section 21 and the outer wing section 22 transform from an extended state to an inverted V-shaped state, and the dual-axis variable sweep wing mechanism 2 forms an M-shaped overall configuration, finally transforming into a folded state. In this way, the inner wing section 21 and the outer wing section 22 can achieve any relative deployment angle from 180° to 0°.

[0030] After folding, the dual-axis variable-sweep wing mechanism 2 is parallel to the fuselage, facilitating its storage in weapon system launch tubes, launch racks, and other devices. It also facilitates transportation, allowing the main support point or launch slider to be directly positioned on the robust underside of the fuselage 12. The structure is simple, reliable, and easy to install into the launch tube. During folding, the centroid of the inner wing section 21 moves forward, while the centroid of the outer wing section 22 moves backward, resulting in minimal changes to the wing's center of pressure and focal point. This gives the wing excellent handling characteristics and flight stability during the variable-sweep process.

[0031] During flight, by sweeping the inner wing section 21 forward (sweep angle range of 0° to 75°) and simultaneously sweeping the outer wing section 22 backward (sweep angle range of 0° to 75°), a stable M-shaped aerodynamic shape is formed. The key aerodynamic benefit of this configuration is that the forward sweep of the inner wing section 21 causes the nose-down moment generated by its lift to compensate for the nose-up moment generated by the backward sweep of the outer wing section 22, thereby controlling the range of pressure center variation within a very small range during full-speed flight and greatly improving the inherent stability of the aircraft.

[0032] The length of the outer wing section 22 can be greater than or equal to twice the length of the inner wing section 21, thereby reducing the change in the center of gravity of the entire missile during the folding and unfolding process, which helps to reduce aerodynamic disturbances and improve the stability of the entire missile during unfolding.

[0033] The aerodynamic layout of the loitering munition of this invention, encompassing sub- and super-speed ranges, features a dual-axis variable-sweep wing mechanism 2. During deployment, the inner wing section 21 and the outer wing section 22 exert opposite influences on the loitering munition's pressure center focal point, mutually compensating for each other. This effectively controls the range of pressure center focal point variation during cross-speed range deformation flight within an extremely small interval, fundamentally improving the stability and maneuverability problems caused by abrupt pressure center changes in traditional single-axis variable-sweep wings, resulting in a significant improvement in aerodynamic control stability. Please refer to [link / reference]. Figure 7 , Figure 7 This is a schematic diagram showing how the stability of a single-axis variable-sweep wing and a dual-axis variable-sweep wing mechanism changes with the deployment angle at transonic speeds (e.g., Ma=0.8).

[0034] In addition, the unique dual-section folding design allows the inner wing section 21 and the outer wing section 22 to be completely folded and parallel to the fuselage when the dual-axis variable sweep wing mechanism 2 is stored, forming an extremely compact envelope shape. This perfectly fits into small spaces such as weapon launch tubes or transport containers, and does not occupy valuable space inside the fuselage. This solves the problem that the lateral size of traditional variable sweep wings is still too large after storage and affects the arrangement of loitering munition air intakes, mission payloads and other mechanisms, thus improving storage and deployment capabilities.

[0035] like Figure 8 As shown, in order to take into account both sub-, trans-, and supersonic flight states, the target airfoil for the inner wing section 21 and the outer wing section 22 is a high lift-to-drag ratio wide-speed-range airfoil obtained by weighted averaging of the NACA SC(2)-0710 supercritical airfoil and the NACA0006 supersonic airfoil. The integrated airfoil is relatively thin overall, effectively reducing shock wave drag. It has the ability to balance lift, low wave drag, and stall characteristics when combined with small camber and post-loading. The maximum camber is 0.47%, the maximum camber position is 84%, the maximum relative thickness is 7.95%, and the maximum thickness position is 39%.

[0036] Under Mach 2.0 conditions, the aerodynamic performance of this target airfoil is mainly dominated by linearized supersonic aerodynamic theory. Its near-symmetrical thin configuration design allows lift generation to rely primarily on small angle-of-attack deflection, avoiding additional wave drag caused by large camber. This characteristic gives the airfoil a significant advantage during supersonic cruise and acceleration phases.

[0037] In the transonic range (approximately 0.8–1.2 Mach), the thin profile significantly increases its critical Mach number. Even with sweep, the ultrasonic region on the upper surface of the target airfoil remains controlled, resulting in relatively low shock wave intensity and reducing the likelihood of severe boundary layer separation. Simultaneously, it retains the characteristics of a supercritical airfoil: a flat upper surface in the middle and a concave lower surface at the tail, effectively delaying the shock wave's appearance and significantly reducing its intensity. Compared to traditional transonic supercritical airfoils, while this target airfoil may not have the optimal lift-to-drag ratio in this range, its aerodynamic characteristics change continuously and stably, without sudden drag transitions.

[0038] At low subsonic speeds such as Mach 0.4, this target airfoil, due to its relatively small thickness and low lift curve slope, does not pursue extreme lift-to-drag ratio performance. However, with a swept-back configuration and a larger aspect ratio, the required lift level can be achieved by appropriately increasing the angle of attack and wing area. For UAV platforms that prioritize high-speed performance, such low-speed efficiency losses are acceptable.

[0039] Therefore, the aforementioned target airfoil embodies an overall design philosophy that prioritizes supersonic performance while also considering transonic stability and low-speed flyability. This airfoil exhibits clear aerodynamic mechanisms and continuous performance variations throughout its entire airfoil envelope, avoiding multiple airfoil switching or complex variable geometry designs, thus demonstrating outstanding engineering feasibility.

[0040] Furthermore, the loitering munition is designed for subsonic-transonic-sonic wide envelope flight missions within the range of 0.4–2.0 Ma. Through a combined design of high-speed aerodynamics and overall layout, the length of the inner wing section 21 is approximately 26% of the total length of the loitering munition, and the length of the outer wing section 22 is approximately 52% of the total length of the vehicle. The length ratio of the inner wing section 21 and the outer wing section 22 is primarily based on maintaining the lift-to-drag ratio and controlling wave drag during the transonic to supersonic flight phases, while also taking into account low-speed flyability.

[0041] Under transonic and supersonic flight conditions, the drag composition of loitering munitions undergoes fundamental changes, with wave drag becoming the decisive factor affecting the lift-to-drag ratio. Wave drag is closely related to airfoil thickness and equivalent normal Mach number, and is also extremely sensitive to wing planform. The significantly increased length proportion of the outer wing section 22 allows it to form the main equivalent sweep plane in a high-sweep state, effectively reducing the incoming flow normal Mach number, thereby weakening shock wave intensity and delaying shock wave formation. This characteristic is particularly critical in the supersonic range of Ma>1.2 and is a necessary condition for maintaining an acceptable lift-to-drag ratio.

[0042] The inner wing section 21, accounting for approximately 26% of the aircraft's total length, is primarily designed to provide a stable and continuous lift source at high speeds. Under supersonic conditions, lift is linearly related to wing area, while its dependence on aspect ratio is significantly reduced. A longer inner wing section helps concentrate lift distribution in the vicinity of the fuselage, reducing shock wave interference and additional wave drag at the wingtips, while also helping to meet area rule requirements and lower the overall wave drag level.

[0043] From a lift-to-drag ratio perspective, this configuration ensures excellent lift efficiency in the subsonic range of Mach 0.6 by providing sufficient lift area through the longer outer wing section 22, enabling the loitering munition to achieve stable flight within a reasonable angle of attack range. In the transonic to supersonic range, the combination of the long outer wing section 22 and the medium-length inner wing section 21 unifies lift generation and wave drag control within the same planar configuration, avoiding performance degradation caused by a single wing section excessively bearing the function of lift or wave drag adjustment.

[0044] In summary, the design scheme with the inner wing section 21 accounting for 26% of the length and the outer wing section 22 accounting for 52% of the length is a reasonable aerodynamic layout choice within the subsonic-transonic-supersonic flight envelope, with the supersonic lift-to-drag ratio as the core and taking into account transonic stability and subsonic flyability. It reflects the engineering design principle of "high-speed performance as the main factor and low-speed performance as the whole" in the aerodynamic design of wide-speed-range UAVs.

[0045] The fuselage 12 is a flattened spindle shape, and the fuselage cross-section 121 is a trapezoid with rounded corners and curved edges, with the vertical dimension smaller than the horizontal dimension. The curvature of the fuselage cross-section 121 and the position of the nose apex 2 are determined according to the flight speed.

[0046] like Figure 10 As shown, the non-rotating integral fuselage is determined according to the flight speed requirements. The fuselage section 121 is a transverse section of the fuselage perpendicular to the longitudinal central axis of the fuselage. In this design, the fuselage section 121 is the reference section at the junction of the nose and the fuselage. The edge of the fuselage section 121 is the positioning reference at the end of the von Kármán curve group 113.

[0047] like Figure 9 , 11 The nose apex 112 shown is the convergence endpoint of the foremost part of the nose profile, the common starting point of the von Kármán curve group 113, and also the origin reference for the longitudinal dimension of the nose 11. The position of the nose apex 112 directly determines the taper and bluntness of the nose 11, which are directly related to the shock wave drag of the nose 11 during flight. In this embodiment, the position of the nose apex 112 should be selected to ensure that the projection of the nose apex 112 onto the fuselage section 121 falls at the intersection of the fuselage section 121 and the plane of symmetry of the fuselage 12.

[0048] The fuselage 12 and the nose 11 are determined according to the following steps: Step S1: First, based on the flight speed requirements and the shock wave drag during flight, determine the basic parameters of the fuselage 12 and the position of the nose apex 112.

[0049] Step S2: Select key points at the edge of fuselage section 121. To prevent shock waves from being generated by turbulence during flight, the selection of key points should follow the arc length distance between two adjacent key points. The contour fitting error is determined based on the curvature Ks at the edge of the fuselage section where the key point is located and the preset contour fitting error δ1, so that the contour fitting error between adjacent key points is less than δ1. ≤ The contour fitting error δ1 is determined based on the flight speed.

[0050] In this embodiment, as Figure 9 , 11 As shown, in step S3, the von Kármán curve is generated on the fuselage symmetry plane.

[0051] The specific method for the nose section busbar 111: S 31 Project the nose vertex 112 onto the fuselage section 121, and connect the nose vertex 112 with the projection point. The length of this line segment is L.

[0052] S 32 : Connect any intersection point of the fuselage symmetry plane and the corresponding fuselage section 121 with the projection point. The length of this line segment is R.

[0053] S 33Taking the nose apex 112 as the origin, the direction of the fuselage axis 12 as the x-direction, and the direction perpendicular to the fuselage axis upwards as the y-direction, using the formula: The von Kármán curve at the 12 symmetry plane of the fuselage was determined.

[0054] In this embodiment, as Figure 12 As shown, in step S3, the method for generating the von Kármán curve group 113 connecting the apex 112 of the machine head and the key points of the cross section is as follows: S 34 Select several key points along the edge of fuselage section 121 on one side of the fuselage's symmetry plane, and connect each key point to the projection point of the nose on the section. The length of this line segment is... The angle between the line segment of length R and the plane of symmetry is .

[0055] S 35 Taking the nose apex 112 as the origin, the x-direction is along the fuselage axis, and the y-direction is upward perpendicular to the fuselage axis. Using the formula: Determine S 34 The von Kármán curve at key points.

[0056] S 36 : S 35 The von Kármán curves of the key points obtained in the process are rotated around the line connecting the nose and the projection point of the nose. and- The angle extends to the edges of the fuselage section 121 on both sides of the fuselage symmetry plane, forming the von Kármán curve group 113.

[0057] In this embodiment, as Figure 13 As shown, once the von Kármán curve group 113 of the nose cone 11's outer contour is determined, the skeleton of the nose cone 11's outer contour has been formed, thus forming the solid outer contour of the nose cone 11. Figure 12 As shown, in step S4, several planes perpendicular to the fuselage symmetry plane are created to truncate the von Kármán curve group 113.

[0058] The method for creating several planes perpendicular to the fuselage's plane of symmetry is as follows: S 41 Several key points are selected at the nose section generatrix 111. A plane is generated that passes through the key points, is perpendicular to the fuselage symmetry plane, and is parallel to the fuselage section 121. The von Kármán curve group 113 is then cut off.

[0059] S 42The points generated by the truncated von Kármán curve group 113 are connected using spline curves. Since the intersection points of the plane perpendicular to the fuselage symmetry plane with the von Kármán curve group 113 are discretely distributed, directly connecting them with straight lines or arcs would create broken lines and sharp-angled profiles, failing to meet the smoothness requirements of the nose's aerodynamic shape. In this embodiment, the spline curve is a cubic spline curve. The discrete intersection points formed by the intersection of each truncated plane with the von Kármán curve group 113 are used as control points. Through cubic spline curve fitting calculations, a closed cross-sectional profile that fits all control points is generated. During the fitting process, the profile deviation of the curve is strictly controlled within the preset nose outline fitting error threshold δ1 to ensure that the fitted cross-sectional profile matches the aerodynamic drag reduction characteristics of the von Kármán curve.

[0060] In this embodiment, S 41 During the process, to ensure the smoothness of the nose surface and prevent shock waves from being generated on the nose surface during flight, the plane spacing between two adjacent key points at the nose shape generatrix 111 is... The contour fitting error is determined based on the curvature Ks at the key point and the preset contour fitting error δ2, ensuring that the contour fitting error between adjacent key points is less than δ2. ≤ The contour fitting error δ2 is determined based on the flight speed.

[0061] In this embodiment, as Figure 14 As shown, in step S4, the method for generating the multi-section surface 14 is to use the truncated von Kármán curve group 113 as the surface section and the spline generated by connecting the truncated points as the guide line to generate the multi-section surface 14.

[0062] In this embodiment, as Figure 15 As shown, the multi-section curved surface 14 is joined with the fuselage section 121, and the solid outline of the nose 11 is generated by using the closed curved surface.

[0063] Revisit Figure 1 The fuselage 12 and tail 13 taper with a smooth, curved geometric contraction, with the engine exhaust nozzle integrated at the end of this contraction section, making its exit cross-section flush with the tail section of the fuselage 12. This geometrically integrated contraction design guides the tail airflow smoothly, effectively reducing airflow separation and low-pressure wake regions, and significantly reducing bottom drag.

[0064] Revisit Figure 1 and Figure 2 The fuselage 12 is equipped with a Y-shaped tail 131 at its end. This invention adopts a low-wing aerodynamic configuration, which is beneficial to improving the roll stability and ground effect of the aircraft. The Y-shaped tail 131 layout is a highly efficient aerodynamic shape with low drag and low radar cross-section.

[0065] In summary, the dual-axis V-shaped variable-sweep wing, flattened spindle-shaped fuselage, integrated retractable tail section, and Y-shaped tail of this invention together constitute a synergistic drag reduction and synergistic stability aerodynamic system. Both the inner wing section 21 and the outer wing section 22 adopt transonic airfoils. The geometric characteristics of this airfoil determine that it can maintain high aerodynamic efficiency even at near-sonic speeds. Combined with the variable-sweep function, it ensures that the aircraft can achieve its optimal performance in any target speed range by adjusting the sweep angle, thereby achieving a high lift-to-drag ratio across the entire speed range from subsonic to supersonic (Ma<3.0).

[0066] Finally, a streamlined fairing with an outer profile fixed to the wing is designed in the connection area between the fuselage 12-inner wing section 21 and the inner wing 21-outer wing section 22. The geometric surface of this fairing smoothly transitions with the adjacent wing and fuselage 12 surfaces. Its function is to eliminate aerodynamic disadvantages such as gaps and cavities that are inevitably generated due to the presence of the rotating mechanism, ensuring the continuity and smoothness of the aerodynamic shape, thereby suppressing the generation of interference drag.

[0067] Finally, it should be noted that the above embodiments are only used to illustrate the technical solutions of the present invention, and not to limit them; although the present invention has been described in detail with reference to the foregoing embodiments, those skilled in the art should understand that modifications can still be made to the technical solutions described in the foregoing embodiments, or equivalent substitutions can be made to some or all of the technical features; and these modifications or substitutions do not cause the essence of the corresponding technical solutions to deviate from the scope of the technical solutions of the embodiments of the present invention.

Claims

1. An aerodynamic configuration for a loitering munition in the sub-trans- and super-trans-velocity ranges, characterized in that, This includes a dual-axis variable-sweep wing mechanism located at the bottom of the fuselage, forming a low-wing configuration that can arbitrarily change and lock the sweep angle within a range of 0° to 90°. Each of the aforementioned dual-axis variable sweep wing mechanisms includes an inner wing section and an outer wing section. The wing root of the inner wing section is pivotally connected to the fuselage and can rotate relative to the fuselage, thereby adjusting the forward sweep angle of the inner wing section within any angle from 0° to 90°. The outer wing root is pivotally connected to the inner wing tip, allowing it to rotate relative to the inner wing section, thus adjusting the sweep angle of the outer wing section at any angle from 0° to 90°. The length of the outer wing section is greater than or equal to twice the length of the inner wing section; The fuselage is flat and spindle-shaped; the head is pointed.

2. The aerodynamic layout of the loitering munition in the sub- and super-speed ranges according to claim 1, characterized in that, The dual-axis variable sweep wing mechanism also includes an inner wing pivot and an outer wing pivot. The fixed end of the inner wing pivot is connected to the bottom of the fuselage, and the driving end of the inner wing pivot is connected to the wing root of the inner wing section. The fixed end of the outer wing pivot is connected to the wingtip of the inner wing section, and the driving end of the outer wing pivot is connected to the wing root of the outer wing section; the inner and outer wing sections can independently change their sweep angle.

3. The aerodynamic layout of the loitering munition in the sub-trans- and super-trans-velocity ranges according to claim 2, characterized in that, The inner wing pivot drives the inner wing section to rotate, and the forward sweep angle of the inner wing section can be adjusted from 90° to 0°. The outer wing pivot drives the outer wing section to rotate, and the sweep angle of the outer wing section can be adjusted from 90° to 0°. The relative deployment angle between the inner wing section and the outer wing section can be 0° to 180°.

4. The aerodynamic layout of the loitering munition in the sub-trans- and super-trans-velocity ranges according to claim 1, characterized in that, After the dual-axis variable-sweep wing mechanism is folded, it is completely retracted to the bottom of the fuselage and parallel to the fuselage.

5. The aerodynamic layout of the loitering munition in the sub- and super-speed ranges according to claim 1, characterized in that, The target airfoils for the inner and outer wing sections are high lift-to-drag ratio, wide speed range airfoils obtained by optimizing supercritical and supersonic airfoils.

6. The aerodynamic layout of the loitering munition in the sub-trans- and super-trans-velocity ranges according to claim 1, characterized in that, The fuselage and tail section are integrated with a smooth, curved geometric contraction.

7. The aerodynamic layout of the loitering munition in the sub-trans- and super-trans-velocity ranges according to claim 1, characterized in that, The fuselage is equipped with a Y-shaped tail fin at the end.

8. The aerodynamic layout of the loitering munition in the sub-trans- and super-trans-velocity ranges according to claim 1, characterized in that, The fuselage and nose section are determined according to the following steps: S1: Determine the non-rotating fuselage, fuselage cross-section, and nose apex based on the flight speed. The projection point of the nose apex onto the fuselage cross-section falls at the intersection of the fuselage cross-section and the fuselage symmetry plane. S2: Select key points, which are located at the edge of the fuselage section; S3: Generate a von Kármán curve on the fuselage symmetry plane as the nose shape generatrix, obtain the parameters required for the von Kármán curve of the key points of the fuselage-nose connection section, and generate a group of von Kármán curves connecting the nose apex and the key points of the section. S4: Create several planes perpendicular to the fuselage symmetry plane, cut off the von Kármán curve group, generate multi-section curved surfaces, and close several multi-section curved surfaces to form the solid nose outline.

9. The aerodynamic layout of the cruise aircraft in the sub-transit and super-transit speed ranges according to claim 8, characterized in that, In step S2, the arc length distance between two adjacent key points is determined based on the curvature Ks at the edge of the fuselage section where the key point is located and the preset contour fitting error δ1, so that the contour fitting error between adjacent key points is less than δ1.

10. The aerodynamic layout of the cruise aircraft in the sub-transit and super-transit speed ranges according to claim 7, characterized in that, The fuselage is a trapezoid with rounded corners and curved edges, with its vertical dimension smaller than its horizontal dimension.

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