A new type of heavy-load short-wing lift-type aircraft

Abstract

This invention discloses a novel high-payload short-wing lift fuselage aircraft, comprising an airfoil-shaped lift fuselage, forward short wings smoothly transitioning to both sides of the nose of the lift fuselage, rear short wings smoothly transitioning to both sides of the tail of the lift fuselage, a propeller-driven propulsion system located at the nose of the lift fuselage, a V-tail located at the tail of the lift fuselage, and ventral fins located on the belly of the lift fuselage. The wingspan of this novel high-payload short-wing lift fuselage aircraft is ≤2m, with the wingspan of the forward short wings being smaller than that of the rear short wings, and the angle of installation of the forward short wings being larger than that of the rear short wings. With this design, the lift fuselage can provide significant lift, and the short wings, in addition to providing some lift, can also achieve longitudinal trim of the aircraft by adjusting the angle of installation, meeting the static stability requirements of the aircraft. Furthermore, this design retains high-speed cruise performance and reduces the wingspan, thereby enabling the UAV to possess maneuverable and flexible capabilities, and to be applied in confined and complex environments.

Description

Technical Field

[0001] This invention relates to the technical field of aircraft, and in particular to a novel heavy-load short-wing lift-powered aircraft. Background Technology

[0002] With the rapid development of logistics services, the application of drones for point-to-point "last mile" transportation has become a trend, reducing labor costs while enabling timely delivery. In urban environments, the ability to quickly and flexibly transport packages is a demand, requiring drones that possess both high-speed cruising and maneuverability to achieve fast, efficient, and safe delivery.

[0003] Fixed-wing drones, due to their large wingspan, struggle to achieve point-to-point flights in urban environments with towering buildings. Therefore, the mainstream logistics industry is increasingly adopting rotary-wing drones. For example, one type of autorotary logistics drone is primarily used for short-to-medium distance logistics transportation, effectively handling order fulfillment in remote, mountainous, and island regions with limited transportation, as well as the transport of high-value-added products and emergency supplies, significantly improving logistics efficiency. Other companies are planning a three-tiered air transport network consisting of large manned transport aircraft, large branch-line drones, and small terminal drones. Currently, they are primarily developing vertical takeoff and landing (VTOL) cargo drones for applications in logistics, emergency rescue, and government use. Even better, some small multi-rotor drones utilize a six-axis design, with a maximum payload of 2.5 kg and a maximum delivery distance of 10 km when fully loaded.

[0004] With the technological advancements in drones, their applications are becoming increasingly widespread, and their payload requirements are becoming more complex. Currently, drones commonly used for express delivery and logistics mainly fall into two categories: rotary-wing drones and hybrid vertical takeoff and landing (VTOL) fixed-wing drones. While rotary-wing drones are small and possess advantages in low-speed maneuverability, their hovering and forward flight power are greater than those of fixed-wing drones; the heavier the payload, the higher the power consumption. Therefore, rotary-wing drones have shorter ranges and shorter endurance, making them unsuitable for medium-distance logistics transportation. Furthermore, their relatively slow cruising speed hinders their ability to deliver on time. Hybrid VTOL drones offer the advantage of vertical takeoff and landing using a rotary-wing mode and high-speed cruising using a fixed-wing mode. However, their large fixed-wing wingspan makes it difficult to maneuver flexibly in complex urban environments with towering buildings. Additionally, their large wingspan makes them unsuitable for confined spaces.

[0005] In response to the above-mentioned drawbacks, there is an urgent need for a new type of aircraft that can achieve high-speed and flexible flight while possessing a large payload capacity. Summary of the Invention

[0006] The purpose of this invention is to provide a novel high-payload short-wing lift airframe aircraft, so as to enable the aircraft to fly at high speed and flexibly while having a large payload capacity.

[0007] To address the aforementioned technical problems, this invention provides a novel heavy-load short-wing lift fuselage aircraft, comprising an airfoil-shaped lift fuselage, forward short wings smoothly transitioning to both sides of the nose of the lift fuselage, rear short wings smoothly transitioning to both sides of the tail of the lift fuselage, a propeller power unit located at the nose of the lift fuselage, a V-tail fin located at the tail of the lift fuselage, and a ventral fin located on the belly of the lift fuselage; wherein the wingspan of the novel heavy-load short-wing lift fuselage aircraft is ≤2m, the wingspan of the forward short wings is smaller than the wingspan of the rear short wings, and the angle of installation of the forward short wings is larger than the angle of installation of the rear short wings.

[0008] In one embodiment, the difference in wingspan between the fore stub wing and the rear stub wing is 0–15 cm.

[0009] In one embodiment, the difference between the mounting angle of the fore stub wing and the mounting angle of the aft stub wing is 3° to 15°.

[0010] In one embodiment, the length of the lifting fuselage is 0.8 to 1.4 m, and the width of the lifting fuselage is 0.3 to 0.6 m; the wingspan of the fore stub wing and the lifting fuselage is 0.8 to 1.2 m; and the wingspan of the rear stub wing and the lifting fuselage is 1.1 to 1.5 m.

[0011] In one embodiment, the airfoil of the lifting fuselage is NACA2418, NACA 633-018, NACA0018, or S1046.

[0012] In one embodiment, the airfoil of the fore stub wing and the aft stub wing is MH114, MH115, NACA2412, NACA23102, or E195.

[0013] In one embodiment, the root angle of the fore stub wing is 6° to 12°, the aerodynamic twist angle of the fore stub wing is -2° to -8°, the root chord length of the fore stub wing is 0.235 to 0.295 m, and the wingtip chord length of the fore stub wing is 0.224 to 0.284 m; the root angle of the rear stub wing is -1° to -3°, the aerodynamic twist angle of the rear stub wing is 2° to 4°, the root chord length of the rear stub wing is 0.235 to 0.295 m, and the wingtip chord length of the rear stub wing is 0.224 to 0.284 m.

[0014] In one embodiment, the smooth transition between the lifting fuselage and the fore stub wing and the aft stub wing is a transition structure designed with a B-spline curve.

[0015] In one embodiment, the lifting fuselage is provided with a plurality of empty compartments in sequence along its length; the empty compartment closest to the nose of the lifting fuselage is used to place batteries; and the empty compartment near the tail of the lifting fuselage is used to load cargo.

[0016] In one embodiment, the lifting fuselage is generally flat, and the lifting fuselage has integrally formed fuselage side wings on both sides. The fuselage side wings extend from the nose of the lifting fuselage to the tail, and the thickness of the fuselage side wings gradually decreases from the middle of the lifting fuselage to both sides, so that the fuselage side wings are connected to the front wing and the rear wing in a smooth transition.

[0017] The beneficial effects of this invention are as follows:

[0018] This invention employs a blended wing-body design with a lifting fuselage and stub wings. The lifting fuselage utilizes an airfoil design to provide significant lift. The stub wings contribute some lift and, by adjusting the angle of attack, achieve longitudinal trim, meeting the aircraft's static stability requirements. Compared to traditional fixed-wing UAVs, this stub-wing fixed-wing UAV retains high-speed cruise performance while significantly reducing wingspan by utilizing the lift generated by the lifting fuselage, resulting in highly maneuverable capabilities suitable for confined and complex environments. Attached Figure Description

[0019] To more clearly illustrate the technical solution of the present invention, the drawings used in the embodiments will be briefly introduced below. Obviously, the drawings described below are only some embodiments of the present invention. For those skilled in the art, other drawings can be obtained based on these drawings without creative effort.

[0020] Figure 1 This is a schematic diagram of the structure provided in an embodiment of the present invention;

[0021] Figure 2 yes Figure 1 A schematic diagram of the structure viewed from below;

[0022] Figure 3 yes Figure 1 A top-view structural diagram;

[0023] Figure 4 It is a quadratic Bézier curve graph;

[0024] Figure 5 It is a cubic Bézier curve;

[0025] Figure 6 It is a quadratic B-spline curve plot with three-point control;

[0026] Figure 7 It is a quadratic B-spline curve plot with five-point control;

[0027] Figure 8 It is a curve showing the lift and drag coefficients of an aircraft under unpowered conditions;

[0028] Figure 9 It is a pressure coefficient cloud map of an aircraft under unpowered conditions;

[0029] Figure 10 This is a schematic diagram of the velocity field at the cross-section of the fuselage under unpowered conditions;

[0030] Figure 11 This is a curve showing the lift and drag coefficients of an aircraft under powered conditions.

[0031] Figure 12 It is a pressure coefficient cloud map of an aircraft under powered conditions;

[0032] Figure 13 This is a schematic diagram of the velocity field at the cross-section of the fuselage when it is powered.

[0033] Figure 14 It is a graph of the moment coefficients in the pitch, yaw, and roll directions of an aircraft.

[0034] Figure 15 This is a graph showing the changes in acceleration during actual flight;

[0035] Figure 16 This is a graph showing the changes in angular velocity during actual flight.

[0036] Figure 17 This is a graph showing energy changes during actual flight.

[0037] The attached figures are labeled as follows:

[0038] 10. Lifting fuselage; 11. Fuselage wings; 12. Empty cabin;

[0039] 20. Short front wing;

[0040] 30. Short rear wing;

[0041] 40. Propeller power mechanism;

[0042] 50. V-shaped tail wing;

[0043] 60. Pelvic fins. Detailed Implementation

[0044] The technical solutions of the embodiments of the present invention will be clearly and completely described below with reference to the accompanying drawings.

[0045] To address the current needs for inter-building aerial transport and food delivery via staircases in cities, this invention designs a novel high-payload, short-wing, lift-powered aircraft to solve this inter-building aerial transport problem. To achieve inter-building aerial transport, the small distances between buildings in some residential areas need to be considered. To ensure the aircraft's safe flight, its size needs to be limited. However, it still needs to meet the functions of aerial transport or high-rise food delivery, and there are also requirements for the aircraft's payload. Therefore, for such special transport tasks, some constraints are proposed for the aircraft design:

[0046] a) Wingspan of the aircraft < 2m; b) Payload > 1.5kg; c) Takeoff weight between 5-10kg.

[0047] At this point, the implementation of the present invention is as follows: Figures 1 to 3 As shown, the aircraft includes an airfoil-shaped lifting fuselage 10, forward stub wings 20 smoothly connected to both sides of the nose of the lifting fuselage 10, rear stub wings 30 smoothly connected to both sides of the tail of the lifting fuselage 10, a propeller power unit 40 located at the nose of the lifting fuselage 10, a V-tail 50 located at the tail of the lifting fuselage 10, and a ventral fin 60 located on the belly of the lifting fuselage 10. Moreover, with this configuration, in addition to ensuring that the wingspan of the new heavy-load stub wing lifting fuselage aircraft is ≤2m, it is also recommended that the wingspan of the forward stub wings 20 be smaller than the wingspan of the rear stub wings 30, and that the installation angle of the forward stub wings 20 be greater than the installation angle of the rear stub wings 30.

[0048] It should be noted that the function of the fuselage of a traditional fixed-wing UAV is to connect the wings and other components, as well as to carry cargo. The fuselage does not generate lift, but it still generates drag. In view of this characteristic, the present invention proposes the above-mentioned novel heavy-load short-wing lift fuselage aircraft, which belongs to the integrated wing-body design of the lift fuselage 10 and the short wing.

[0049] The lifting fuselage 10 refers to a configuration where, in the direction of the airflow, the fuselage cross-section and airfoil lift characteristics are similar, exhibiting a relatively thick aerodynamic shape. Employing the lifting fuselage 10 enables the fuselage to generate lift. In the blended wing-body design of the lifting fuselage 10 and the stub wing, the fuselage can carry cargo while providing most of the lift; the stub wing provides some lift and can flexibly adjust its angle of attack to achieve longitudinal trim of the UAV, meeting static stability requirements. The transition between the lifting fuselage 10 and the stub wing is achieved through blended wing-body technology. The advantage of this blended wing-body design is that it retains the high-speed cruise characteristics of fixed-wing UAVs while reducing the wingspan, enabling maneuverable and flexible flight and adapting to confined and complex environments.

[0050] To provide a more detailed description of the new heavy-load, short-wing lift-powered aircraft, the following section will explain each major component in detail:

[0051] Regarding the aforementioned lifting fuselage 10, as Figure 1 As shown, the lifting fuselage 10 is generally flat, and the lifting fuselage 10 has integrally formed fuselage side wings 11 on both sides. The fuselage side wings 11 extend from the nose of the lifting fuselage 10 to the tail. From the middle of the lifting fuselage 10 to both sides, the thickness of the fuselage side wings 11 decreases from thick to thin, so that the fuselage side wings 11 are connected to the front stub 20 and the rear stub 30 in a smooth transition.

[0052] Regarding the size design of the lifting fuselage 10, it is recommended that the length of the lifting fuselage 10 be 0.8 to 1.4 m and the width of the lifting fuselage 10 be 0.3 to 0.6 m. For example, in this embodiment, the length of the lifting fuselage 10 is 1.1 m and the width of the lifting fuselage 10 is 0.45 m.

[0053] In addition, there are many options for the airfoil of the lifting fuselage 10. For example, the airfoil of the lifting fuselage 10 can be set to NACA2418, NACA 633-018, NACA0018, S1046, etc. In this embodiment, NACA2418 is selected as the airfoil of the fuselage side wing 11 to realize the large volume function of the lifting fuselage 10.

[0054] Moreover, as Figure 1 As shown, in this embodiment, multiple empty compartments 12 are sequentially arranged along the length of the lifting fuselage 10. The empty compartment 12 closest to the nose of the lifting fuselage 10 is used to place batteries, while the empty compartment 12 closest to the tail of the lifting fuselage 10 is used to load cargo.

[0055] Specifically, the first compartment 12 at the nose can be used to connect the carbon fiber tube to the forewing 20, and can also be used as a battery compartment to store lithium batteries. Because this is a tandem layout, the focus of the aircraft is close to the nose. Therefore, in order to better balance the center of gravity, this embodiment places two 4S lithium batteries in the first compartment 12 on the left and right sides. The second compartment 12 is designed to house the flight control board, install GPS and other equipment. The third compartment 12 is located near the tail of the fuselage, so the third compartment 12 can be used to load cargo.

[0056] Given that the direct combination of the lifting fuselage 10 and the stub wing would result in significant aerodynamic drag at the connection point, this invention employs a blended wing-body design to improve the aircraft's aerodynamic performance and reduce drag during flight. This design incorporates a transition section between the lifting fuselage 10 and the stub wing. Specifically, the design process utilizes a B-spline curve design method, where the smooth transitions between the lifting fuselage 10 and the forward stub wing 20, as well as the rear stub wing 30, are designed using B-spline curves.

[0057] It should be noted that B-spline curves, based on Bézier curves, retain the advantages of Bézier curves such as intuitiveness and convex hull properties, while overcoming the disadvantages of Bézier curves, and have the characteristics of automatic smooth splicing and flexible local modification.

[0058] The B-spline curve is formed by P0, P1, ..., P... n (n>3) A k-order spline curve is defined by a total of n+1 control points and a set of non-decreasing, continuously varying vector nodes ui. Its curve expression is as follows:

[0059]

[0060] In the formula, P i Let u represent control points, n be the number of control points, k be the order of the B-spline curve, and i represent the vector of the i-th node. i，k (u) represents the i-th k-th order B-spline basis function.

[0061] basis function B i，k (u) is derived from the de Boer-Cox recurrence formula, and its expression is:

[0062] B i，k (u)=U1B i，k-1 (u)+U2B i+1，k-1 (u)

[0063]

[0064]

[0065] Where U1 and U2 are the coefficients of the basis functions, and k represents the smoothness of the curve.

[0066] To facilitate understanding of Bézier curves and B-spline curves Figure 4 and Figure 5 The quadratic and cubic Bézier curves are given respectively. Figure 6 and Figure 7 The quadratic B-spline curves for different control points are then given. Figure 4This is a quadratic Bézier curve, controlled by three control points P0, P1, and P2. The curve begins at P0 and ends at P2, and its first derivative at P0 and P2 is 2(P1-P0) and 2(P2-P1), respectively. That is, the lines P0P1 and P1P2 are the slopes of the tangent lines at P0 and P2, respectively. A cubic Bézier curve also passes through the first and last control points, and is tangent to the first and last edges of the control polygon at its start and end points, respectively, while approximating the remaining control points as closely as possible. Other higher-order Bézier curves follow the same principle. Clearly, Bézier curves cannot be modified locally; changing the position of a single control point affects the entire curve shape. B-spline curves, on the other hand, can be defined piecewise. Given m+n+1 vertices, m+1 segments of degree n B-spline curves can be defined. These m+1 segments of degree n B-spline curves are connected end-to-end to form a single curve, where each segment of degree n B-spline curve has n+1 control vertices. Figure 6 It is a quadratic B-spline curve with 3 control points, where m = 0 and n = 2, i.e., one quadratic curve segment. Figure 7 This is a quadratic curve with 5 control points, where m=2 and n=2, meaning it is formed by connecting 3 quadratic curve segments end to end. Unlike Bézier curves, the start and end points of a quadratic B-spline curve are located at the midpoints of the control edges. The tangents at the start and end points are similar to those of Bézier curves. The first and last edges connecting the control points are tangent to the start and end points. However, at the same time, the points where multiple curve segments connect are tangent to the remaining control edges.

[0067] B-spline curve design is used for the transition section between the fuselage and wing of an aircraft. First, the fuselage profile and the wing root profile are obtained by breaking the lines to create corresponding spline curves. Then, B-spline curve design is used to obtain two B-spline curves connecting the partial fuselage profile and the wing root profile. These multiple spline curves form a closed surface, resulting in the surface designed using B-spline curves. The surface effect is visible. Figure 1 Obviously, the transition surface after the wing-body blending design can make the curvature transition between the fuselage and the wing more natural and smooth. For example, in this embodiment, the nose of the lifting fuselage 10 is connected to the fore stub wing 20 through the B-spline curve arc surface B1, and the tail of the lifting fuselage 10 is connected to the rear stub wing 30 through the B-spline curve arc surface B2.

[0068] It should be noted that the main factors that generate lift on an airfoil include the incoming airflow velocity, the airfoil's reference area, and the airfoil's lift coefficient, which can be specifically expressed as:

[0069]

[0070] L represents lift, air density, and incoming air velocity. S is the reference area of ​​the wing, which is represented by the lift coefficient.

[0071] Among them, at small angles of attack, the lift coefficient is linearly positively correlated with the angle of attack of the incoming flow (the same applies to the wing installation angle), that is, as the angle of attack of the incoming flow increases, the lift coefficient will also increase.

[0072] Because the airfoil used in the lifting fuselage 10 is a positive lift airfoil, while the fuselage can provide greater lift, it also generates an additional nose-down moment, resulting in insufficient longitudinal stability and difficulty in aerodynamic trim. Therefore, this invention addresses the aircraft trim problem by changing the installation angles and aerodynamic twist angles of the fore and aft winglets. A larger fore wing installation angle than a rear wing installation angle allows the fore wing to generate greater lift, resulting in a larger nose-down moment relative to the aircraft's center of gravity. Conversely, a smaller rear wing installation angle results in relatively less lift generated by the rear wing, leading to a smaller nose-down moment relative to the aircraft's center of gravity. This ensures that the resultant moment of the fore and aft wings on the aircraft is a nose-down moment, overcoming the nose-down moment caused by the fuselage.

[0073] Furthermore, because the aircraft's center of gravity is relatively close to the nose, the lift torque of the canard about the center of gravity is relatively small, while the torque of the aft wing is relatively large. However, for static stability reasons, the center of gravity needs to be located in front of the aerodynamic focus. Therefore, on the one hand, efforts need to be made to shift the overall aerodynamic focus of the aircraft backward to ensure the center of gravity can be moved backward. From this perspective, the wingspan of the aft wing needs to be appropriately increased to achieve this aerodynamic focus shift. On the other hand, in order to ensure that the combined torque of the canard and aft wings about the aircraft is a pitching moment, the angle of attack of the canard needs to be increased, and the angle of attack of the aft wing needs to be decreased, i.e., increasing the difference in angle of attack between the canard and aft wings.

[0074] Based on the above considerations, after multiple design attempts, the optimal wingspan settings for the front stub 20 and rear stub 30 were finally determined. At the same time, considering that the aircraft's anti-stall performance might be poor due to an excessively large installation angle, negative aerodynamic twist angles were designed for both the front and rear stub 30 to ensure that the aircraft has good anti-stall performance.

[0075] Specifically, regarding the aforementioned forewing 20, such as Figure 1 and Figure 3 As shown, the airfoil can be selected in many ways. For example, the airfoil of the fore stub 20 can be set to MH114, NACA2412, NACA23102, E195, etc. In this embodiment, the airfoil of the fore stub 20 is set to MH114 to improve the aerodynamic performance of the aircraft.

[0076] In addition, there are corresponding preferred setting ranges for various setting parameters of the fore stub 20. For example, the wingspan of the fore stub 20 and the lifting fuselage 10 can be set to 0.8 to 1.2 m, the installation angle α1 at the wing root of the fore stub 20 can be 6° to 12°, the aerodynamic twist angle α2 of the fore stub 20 can be -2° to -8°, the wing root chord length of the fore stub 20 can be 0.235 to 0.295 m, and the wingtip chord length of the fore stub 20 can be 0.224 to 0.284 m.

[0077] In this embodiment, the wingspan of the fore stub 20 and the lifting fuselage 10 is set to 1m, the installation angle α1 at the wing root of the fore stub 20 is 9°, the aerodynamic twist angle α2 of the fore stub 20 is -5°, the wing root chord length of the fore stub 20 is 0.265m, and the wingtip chord length of the fore stub 20 is 0.254m.

[0078] Specifically, the installation angle a1 at the wing root of the fore stub 20 refers to the angle between the chord line of the wing root airfoil section and the axis of the lifting fuselage 10 when the fore stub 20 is installed on the lifting fuselage 10, and the aerodynamic twist angle a2 of the fore stub 20 refers to the angle from the wing root to the wingtip of the fore stub 20 where the wing angle of attack gradually changes.

[0079] Regarding the aforementioned rear stub wing 30, as Figure 1 and Figure 3 As shown, there are many options for the airfoil. For example, the airfoil of the rear stub 30 can be set to MH114, NACA2412, NACA23102, E195, etc. In this embodiment, the airfoil of the rear stub 30 is set to MH114 to improve the aerodynamic performance of the aircraft.

[0080] In addition, there are corresponding preferred setting ranges for various setting parameters of the rear stub wing 30. For example, the wingspan of the rear stub wing 30 and the lifting fuselage 10 can be set to 1.1 to 1.5 m, the installation angle b1 at the wing root of the rear stub wing 30 can be -1° to -3°, the aerodynamic twist angle b2 of the rear stub wing 30 can be 2° to 4°, the wing root chord length of the rear stub wing 30 can be 0.235 to 0.295 m, and the wingtip chord length of the rear stub wing 30 can be 0.224 to 0.284 m.

[0081] In this embodiment, the wingspan of the rear stub wing 30 and the lifting fuselage 10 is set to 1.3m, the installation angle b1 at the wing root of the rear stub wing 30 is -2°, the aerodynamic twist angle b2 of the rear stub wing 30 is 3°, the wing root chord length of the rear stub wing 30 is 0.265m, and the wingtip chord length of the rear stub wing 30 is 0.254m.

[0082] Specifically, the installation angle b1 at the wing root of the rear stub 30 refers to the angle between the chord line of the wing root airfoil section and the axis of the lifting fuselage 10 when the rear stub 30 is installed on the lifting fuselage 10, and the aerodynamic twist angle b2 of the rear stub 30 refers to the angle at which the wing angle of attack gradually changes from the wing root to the wingtip of the front stub 20.

[0083] It should be noted that since the matching between the forewing 20 and the aft wing 30 will result in better performance, when setting the relevant parameters of the forewing 20 and the aft wing 30, it is recommended to preferably set the difference between the wingspan of the forewing 20 and the wingspan of the aft wing 30 to be 0-15cm, and the difference between the mounting angle a1 of the forewing 20 and the mounting angle b1 of the aft wing 30 to be 3°-15°.

[0084] Regarding the aforementioned propeller power mechanism 40, as Figure 1 As shown, it uses dual motors and propellers on the left and right sides of the leading edge of the lifting fuselage 10. The advantage of using a forward-pull propeller is that it can make full use of the propeller's slip flow effect to increase the airflow velocity on the surface of the lifting fuselage 10 and the forewing 20, thereby increasing the dynamic pressure and improving the lift of the UAV.

[0085] It should be noted that the specific design parameters of the aircraft in this embodiment include the following configuration: the total takeoff weight is approximately 6.1 kg, of which the airframe structure weighs approximately 1.9 kg, the power system weighs approximately 0.39 kg, the lithium battery weighs approximately 0.94 kg, the flight control and other equipment weighs approximately 1 kg, and the payload weighs approximately 1.87 kg, with the payload accounting for approximately 30.6% of the total takeoff weight.

[0086] The total takeoff weight of an aircraft can be divided into the following parts:

[0087] W t =W ε +W m +W b +W p

[0088] In the formula, W t This is the takeoff weight, which remains unchanged during flight; W ε For structural weight, W m For the weight of the power and control devices, W b For battery weight, W p This refers to the weight of the mission payload.

[0089] Based on these design considerations, this embodiment is designed with a maximum thrust of 26.33N for a single motor and a total thrust of approximately 52.66N, resulting in a thrust-to-weight ratio of 0.863. Specifically, the propeller power unit 40 uses a Shuangtian ECO2826720kV brushless motor, which has a maximum thrust of 26.33N under static measurement. Considering the total takeoff weight of approximately 6.1kg, a dual-engine power system is used, achieving a total thrust of 52.66N. Therefore, the thrust-to-weight ratio of this aircraft is:

[0090]

[0091] The propeller is a 13-inch blade that matches the motor. Considering the blade diameter and the requirement that the fuselage components will not come into contact with the ground during takeoff, the landing gear height is designed to be 17cm.

[0092] It should be noted that, since this embodiment adopts a combination design of lifting fuselage 10 and short wings, the wingspan is greatly reduced, which will result in a small lateral damping coefficient of the UAV, making it difficult to converge lateral stability during flight and resulting in poor dynamic stability. Therefore, to solve this problem, this embodiment adds a V-tail fin 50 and a ventral fin 60.

[0093] Specifically, regarding the V-tail fin 50, such as Figure 1 As shown, the V-tail 50 is located on the upper surface of the tail of the lifting fuselage 10 and extends diagonally upward to both sides of the lifting fuselage 10. After the V-tail 50 is installed, the V-tail 50 can take on the functions of both the horizontal tail and the vertical tail, and play a role in longitudinal and lateral stability at the same time.

[0094] Regarding the pelvic fin 60, as Figure 2 As shown, the ventral fin 60 is located on the lower surface of the tail of the lifting fuselage 10 and extends diagonally downward to both sides of the lifting fuselage 10. After the ventral fin 60 is installed, it is equivalent to a vertical tail fin, which can increase the directional stability of the UAV when it sideslips.

[0095] To better verify the rationality of the above design scheme, the present invention performed CFD aerodynamic simulation calculations on the prototype model of the design to better analyze its flight performance.

[0096] (1) Non-powered aerodynamic performance

[0097] Without considering the front propeller (no power), the lift-drag coefficient curve of this aircraft is shown. Figure 8 At a 0° angle of attack, the lift coefficient is greater than 0.1. At a 10° angle of attack, the lift coefficient is between 0.8 and 0.9. Within the small angle of attack range, the lift coefficient changes linearly with the angle of attack. The polar curves show that the lift-to-drag ratio reaches its maximum at a 6° angle of attack, approximately 7.

[0098] Figure 9 The pressure coefficient contour map of the aircraft under unpowered conditions is presented. The pressure coefficient is relatively low at the leading edge of the lifting fuselage 10 and the fore stub wing 20, while the pressure difference between the upper and lower surfaces is large, indicating significant lift. The velocity distribution at section 10 of the lifting fuselage is visible when the incoming flow velocity is 18 m / s. Figure 10 When airflow passes over the surface of the lift fuselage 10, it is similar to airflow over an aircraft wing, with the airflow velocity on both the upper and lower surfaces experiencing a process of first increasing and then decreasing. Among them, the airflow velocity is faster and the pressure is lower on the upper surface, and there is a pressure difference between the upper and lower surfaces, so the fuselage can also generate lift.

[0099] (2) Possesses dynamic aerodynamic performance

[0100] The advantage of placing the propeller power unit 40 at the front is that it can make full use of the propeller's slipstream effect. The impact of a front-mounted propeller on the aerodynamic performance of the aircraft is evident. Figure 11 Comparison Figure 8 and Figure 11 It can be observed that the nose rotor can increase the lift coefficient of the aircraft. Although it also introduces drag, the overall maximum lift-to-drag ratio remains essentially unchanged at approximately 7. The slipstream behind the nose rotor disk can accelerate over the surfaces of the wing and lifting fuselage 10, increasing the dynamic pressure on these surfaces and thus improving the aircraft's lift. This means that during takeoff, the aircraft can achieve a greater lift coefficient, reduce the minimum takeoff speed, and decrease the takeoff distance, enabling short takeoff and landing.

[0101] Figure 12 The image shows the pressure cloud map of the surface of the lift fuselage 10 when powered by a propeller. (Comparison) Figure 9 and Figure 12 It can be observed that with propeller power, the pressure coefficient on the upper surface of the forward stub 20 behind the propeller disk is smaller, and the pressure coefficient on the upper surface of the rear stub 30 is also relatively smaller, indicating that the wake behind the propeller disk has a higher velocity when flowing over the fuselage surface. The propeller slipflow increases the dynamic pressure on the aircraft surface, which can improve the overall lift coefficient of the aircraft. The velocity field at the propeller disk cross-section is visible when propeller power is applied. Figure 13 When the incoming flow velocity was 18 m / s, the airflow velocity behind the propeller disk increased significantly, with the airflow velocity exceeding 20 m / s in most areas. (Comparison) Figure 10 and Figure 13 This indicates that the propeller can accelerate the airflow behind the propeller disk, creating a slipstream effect on the aircraft surface. This effect can increase the lift coefficient of the aircraft.

[0102] (3) Stability analysis

[0103] Figure 14The curves for the aircraft's roll moment coefficient versus sideslip angle, pitch moment coefficient versus angle of attack, and yaw moment coefficient versus sideslip angle are presented. As can be seen from the figures, when the aircraft rolls to the right, resulting in a rightward slide, it generates a negative roll moment—a leftward roll—allowing it to recover its initial state. This indicates that the aircraft has stability in the roll direction. When the aircraft pitches up, it generates a negative pitch moment—a nose-down moment—causing it to tend to nose-down, thus possessing the ability to recover its initial state. This indicates that the aircraft has stability in the pitch direction. When the aircraft slides to the right, it generates a positive yaw moment—a rightward yaw—causing it to tend to yaw to the right. This demonstrates stability in the yaw direction.

[0104] Assuming the aircraft maintains static stability in pitch, yaw, and roll, the focal point position can be calculated—0.4m from the nose. Considering static stability margin, the design center of gravity is 0.38m, resulting in a static stability margin of approximately 8%.

[0105] Furthermore, the invention underwent actual flight testing. Figure 15 These are acceleration change data during actual flight of the aircraft. Figure 16 These are data on the change in angular velocity during the actual flight of the aircraft. Figure 17 These are data on energy changes such as lithium battery voltage and current during flight. Flight test results show that the short-wing lift fuselage aircraft designed in this invention possesses short takeoff and landing capabilities and exhibits good flight stability. Utilizing the lift fuselage 10 and propeller slipstream effect, it can achieve high-lift, high-payload flight targets.

[0106] In summary, this invention generates primary lift through the lifting fuselage 10, enabling the aircraft to achieve high lift and heavy payload capacity. The stub wings, besides helping the aircraft address stability issues, also enable high-speed cruise. Compared to traditional fixed-wing aircraft, this invention has a smaller wingspan, a smaller turning radius in the air, and greater maneuverability in confined spaces. Within the same class, this invention offers advantages such as larger volume and greater payload capacity.

[0107] The above description represents the preferred embodiments of the present invention. It should be noted that those skilled in the art can make various improvements and modifications without departing from the principles of the present invention, and these improvements and modifications are also considered to be within the scope of protection of the present invention.

Claims

1. A novel large-load short-wing lifting fuselage aircraft, characterized in that, it comprises a lifting fuselage comprising an airfoil, front short wings smoothly connected to both sides of the head of the lifting fuselage, rear short wings smoothly connected to both sides of the tail of the lifting fuselage, a propeller power mechanism arranged at the head of the lifting fuselage, a V-tail arranged at the tail of the lifting fuselage, and a belly fin arranged at the belly of the lifting fuselage; the lifting fuselage is flat as a whole, and body side wings are integrally formed on both sides of the lifting fuselage, the body side wings extend from the head to the tail of the lifting fuselage, and the thickness of the body side wings gradually decreases from the middle of the lifting fuselage to both sides, and the body side wings are smoothly connected to the front short wings and the rear short wings; the smooth transition between the lifting fuselage and the front short wings is a transition structure designed by B-spline curve, and the smooth transition between the lifting fuselage and the rear short wings is a transition structure designed by B-spline curve; the installation angle of the root of the front short wing is 6°-12°, the aerodynamic twist angle of the front short wing is -2°--8°, the root chord length of the front short wing is 0.235-0.295 m, and the tip chord length of the front short wing is 0.224-0.284 m; the installation angle of the root of the rear short wing is -1°--3°, the aerodynamic twist angle of the rear short wing is 2°-4°, the root chord length of the rear short wing is 0.235-0.295 m, and the tip chord length of the rear short wing is 0.224-0.284 m; and the wingspan of the novel large-load short-wing lifting fuselage aircraft is ≤2 m, the wingspan of the front short wing is smaller than that of the rear short wing, and the installation angle of the front short wing is larger than that of the rear short wing. 2.The novel large-load short-wing lifting fuselage aircraft according to claim 1, characterized in that, the difference between the wingspan of the front short wing and the wingspan of the rear short wing is 0-15 cm. 3.The novel large-load short-wing lifting fuselage aircraft according to claim 1, characterized in that, the difference between the installation angle of the front short wing and the installation angle of the rear short wing is 3°-15°. 4.The novel large-load short-wing lifting fuselage aircraft according to claim 1, characterized in that, the length of the lifting fuselage is 0.8-1.4 m, and the width of the lifting fuselage is 0.3-0.6 m; the wingspan formed by the front short wing and the lifting fuselage is 0.8-1.2 m; the wingspan formed by the rear short wing and the lifting fuselage is 1.1-1.5 m. 5.The novel large-load short-wing lifting fuselage aircraft according to claim 1, characterized in that, the airfoil of the lifting fuselage is NACA2418 airfoil, NACA 633-018 airfoil, NACA0018 airfoil, or S1046 airfoil. 6.The novel large-load short-wing lifting fuselage aircraft according to claim 1, characterized in that, the airfoil of the front short wing and the rear short wing is MH114 airfoil, MH115 airfoil, NACA2412 airfoil, NACA23102, or E195 airfoil.

7. The novel large load short wing lifting body aircraft according to claim 1, characterized in that, a plurality of empty cabins are arranged along the length direction of the lifting body in sequence; the empty cabin closest to the nose of the lifting body is used to place the battery; the empty cabin closest to the tail of the lifting body is used to load goods.

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