A composite high-speed aircraft and method of maneuvering the same

Abstract

The application provides a composite high-speed aircraft and a control method thereof, which comprises a fuselage, a main rotor automatic telescopic mechanism, wings, canards and propellers; two wings are respectively arranged on the two sides of the fuselage close to the tail end of the fuselage; two canards are respectively arranged on the two sides of the fuselage close to the front end of the fuselage; one propeller is respectively arranged on the edge of each of the two canards; the two propellers are connected through a coordination shaft, the rotation speeds of the two propellers are same, and the rotation directions of the two propellers are opposite; the coordination shaft transversely penetrates through the two canards; a single main rotor automatic telescopic mechanism is arranged on the top of the fuselage, is arranged along the longitudinal direction of the fuselage and is located between the wings and the canards; and the folding and unfolding of the main rotor are realized when the flight mode is switched. The application is based on a flight mechanics model, the structure of the folding and unfolding main rotor and the tiltable propeller is used on the composite aircraft, and the speed bottleneck is broken through the switching of the helicopter mode and the fixed-wing mode, so that the composite high-speed flight is realized.

Description

Technical Field

[0001] The present invention relates to the field of composite rotorcraft and flight mechanics and control technology, and in particular to a composite high-speed aircraft and a control method thereof. Background Art

[0002] Conventional helicopters currently rely on high-speed rotors to provide thrust. At high speeds, this inevitably leads to shock waves on the leading blades and stalls on the trailing blades, limiting their speed. Conventional helicopters can reach a maximum speed of over 300 km / h, with a maximum dive speed of nearly 400 km / h. However, both military and civilian applications inevitably demand faster and more maneuverable helicopters, and high speed and maneuverability have become the future trend in helicopter development. In the military, the advantages of high-speed helicopters, combined with their high speed and maneuverability, have the potential to directly transform combat operations. They can simultaneously meet diverse mission requirements, including vertical takeoff and landing, hovering, short-range support, and military supply transport, serving as the primary force for short- to medium-range air support. In the civilian sector, high-speed helicopters can not only replace traditional transport helicopters but also enable a certain level of grid-based services. In the future, they are expected to become the backbone of intelligent urban transportation, making people's lives more convenient.

[0003] A compound helicopter is a concept for implementing a high-speed helicopter. It's defined as any type of helicopter with lift and thrust devices attached to a conventional helicopter. The general idea is to add some fixed-wing aircraft components to a traditional helicopter, allowing it to fly in helicopter mode at low speeds. As speed increases, the rotors are gradually unloaded, lifting power gradually shifted to the wings, while thrust is carried by the propulsion system. At high speeds, the helicopter operates entirely in fixed-wing mode. A compound helicopter is a fusion of a helicopter and a fixed-wing aircraft. Its advantages include: unlike fixed-wing aircraft, it doesn't require an airport runway, making takeoff and landing easier and offering excellent low-speed flight performance; and it offers significant advantages over traditional helicopters in terms of speed, maneuverability, and range.

[0004] However, in the existing high-speed compound helicopter design, one problem that has not yet been properly resolved is how to break through the speed bottleneck of the compound helicopter's steady linear motion. Summary of the Invention

[0005] In view of this, the present invention provides a composite high-speed aircraft and a control method thereof.

[0006] The proposed aircraft design includes: a fuselage, a main rotor automatic retractable mechanism, wings, canards, and propellers; the two wings are positioned on either side of the fuselage near the rear end; the two canards are positioned on either side of the fuselage near the front end; a propeller is mounted on the edge of each canard; the two propellers are connected by a coordination shaft, rotating at the same speed but in opposite directions; the coordination shaft passes transversely through the two canards; a single main rotor automatic retractable mechanism is mounted on top of the fuselage, along the longitudinal direction of the fuselage and between the wings and canards; and the main rotor is folded and unfolded when switching flight modes. Based on a flight mechanics model, the present invention utilizes the folding and unfolding main rotor and tilting propeller structures in a composite aircraft. By switching between helicopter and fixed-wing modes, the aircraft overcomes the speed bottleneck and achieves composite high-speed flight.

[0007] Based on a flight mechanics model, this solution applies the control strategies of folding and unfolding the main rotor and tilting propellers to a compound helicopter, breaking through the helicopter's speed bottleneck. It is particularly suitable for the development of compound high-speed UAVs.

[0008] The present invention specifically adopts the following technical solutions:

[0009] A composite high-speed aircraft, wherein:

[0010] The two wings (30) are respectively arranged on both sides of the fuselage (10) and close to the tail end of the fuselage;

[0011] The two canards (40) are respectively arranged on both sides of the fuselage (10) and close to the front end of the fuselage; a propeller is respectively installed at the wingtip of the two canards; the two propellers (50) are connected by a coordination shaft, and have the same rotation speed and opposite rotation directions; the coordination shaft passes through the two canards (40) transversely;

[0012] The main rotor automatic retractable mechanism (20) is installed on the top of the fuselage and is located between the wing (30) and the canard (40) in the longitudinal direction of the fuselage; when the flight mode is switched, the main rotor of the main rotor automatic retractable mechanism (20) is folded or unfolded.

[0013] Furthermore, the flight modes of the aircraft include: helicopter mode and fixed-wing mode;

[0014] In the helicopter mode, the main rotor in the main rotor automatic retractable mechanism (20) is in an extended state, and during vertical take-off and landing, hovering, and low-speed forward movement, the propeller of the main rotor is adjusted to be parallel to the fuselage by tilting to provide partial lift, and the pitch angle of one of the two propellers (50) is positive and the other is negative, so as to obtain a yaw moment to balance the anti-torque of the main rotor;

[0015] In the fixed-wing mode, the main rotor in the main rotor automatic retractable mechanism is in a folded state, and the propeller (50) is tilted to a direction perpendicular to the fuselage in a high-speed forward state to serve as a thrust propeller.

[0016] Furthermore, the main rotor automatic retractable mechanism (20) comprises: a main rotor, a main rotor box and a hydraulic press;

[0017] The main rotor is connected to the main rotor box through a rotating main shaft; the main rotor box is fixedly connected to the hydraulic press through ears on both sides of the hydraulic press; the hydraulic press is fixedly installed inside the fuselage, and drives the main rotor box and the main rotor to extend out of the fuselage or retract into the fuselage through telescopic movement.

[0018] Furthermore, the installation angle of the wing (30) is 3.5°.

[0019] Furthermore, when the aircraft switches between the helicopter mode and the fixed-wing mode, the aircraft's angle of attack varies between 0° and 13°.

[0020] Furthermore, ailerons are respectively installed at the trailing edges of the two wings (30), and vertical tails (31) are respectively installed at the wing tips.

[0021] Furthermore, the aileron area is 5% to 7% of the wing area;

[0022] The chord length of the aileron is 20% to 25% of the chord length of the wing.

[0023] Furthermore, the inclination angle of the two propellers during vertical flight is expressed as α, the total weight of the entire aircraft is W, the total thrust of the two propellers is T1, the moment arm length relative to the center of gravity is L1, the total lift of the main rotor is T2, and the moment arm length relative to the center of gravity is L2. Then, the balance of the longitudinal pitching moment needs to satisfy T1×cosα×=T2×L2, and the vertical force balance needs to satisfy T1×cosα+T2=W.

[0024] The installation position of the main rotor can be determined by calculation based on the above conditions.

[0025] Furthermore, the aircraft takes off vertically in helicopter mode and reaches the minimum stall speed V in fixed-wing mode when the forward flight speed reaches V min The angle of attack is maintained at 0 before; during the acceleration phase, the main rotor provides the thrust required for the helicopter to move forward;

[0026] In the aircraft from V min Accelerate to 1.2V min During the flight, the angle of attack is first adjusted from 0 to 13°, and then the main rotor is retracted and the propeller is tilted to convert the mode to fixed-wing mode.

[0027] Furthermore, the aircraft matches the optimal speed and cruising state with an angle of attack of 3.5°, and adjusts the angle of attack to 13° when deceleration is required. min When the mode is switched to helicopter mode, the speed is kept at V min to adjust the angle of attack to 0 for further deceleration and / or landing in helicopter mode.

[0028] Compared with the existing technology, the present invention and its preferred embodiment are based on the flight mechanics model and design an automatic retractable main rotor mechanism. The structure of folding and unfolding the main rotor and the tiltable propeller is used on the compound helicopter. The speed of the steady linear motion of the compound helicopter is improved by switching between helicopter mode and fixed-wing mode, realizing a compound high-speed aircraft, which is of great significance to the field of flight mechanics and control of compound rotorcraft. BRIEF DESCRIPTION OF THE DRAWINGS

[0029] The present invention is further described in detail below with reference to the accompanying drawings and specific embodiments:

[0030] Figure 1 A structural diagram of the helicopter mode of the compound high-speed aircraft provided in an embodiment of the present invention.

[0031] Figure 2 A structural diagram of the fixed-wing mode of a composite high-speed aircraft provided in an embodiment of the present invention.

[0032] Figure 3 Schematic diagram of the main rotor automatic retractable mechanism in helicopter mode provided by an embodiment of the present invention.

[0033] Figure 4 Schematic diagram of the main rotor automatic retraction mechanism in fixed-wing mode provided by an embodiment of the present invention.

[0034] Figure 5 A schematic diagram of a complete flight take-off and landing process provided by an embodiment of the present invention.

[0035] Figure 6 A mapping diagram of lift-to-drag ratio and angle of attack α provided in an embodiment of the present invention.

[0036] Figure 7 This is a Cm-α diagram provided by an embodiment of the present invention.

[0037] Figure 8 The Cm-C provided in the embodiment of the present invention L picture. DETAILED DESCRIPTION

[0038] Hereinafter, specific embodiments of the present application will be described in detail with reference to the accompanying drawings. Based on these detailed descriptions, those skilled in the art will be able to clearly understand the present application and implement the present application. Without violating the principles of the present application, the features of different embodiments may be combined to obtain new implementations, or certain features of certain embodiments may be substituted to obtain other preferred implementations.

[0039] It should be noted that the terms used herein are only for describing specific embodiments and are not intended to limit the exemplary embodiments according to the present application. As used herein, unless the context clearly indicates otherwise, the singular form is also intended to include the plural form. In addition, it should be understood that when the terms "comprise" and / or "include" are used in this specification, they indicate the presence of features, steps, operations, devices, components and / or combinations thereof.

[0040] To make the features and advantages of this patent more clearly understood, the following embodiments are specifically described in detail as follows:

[0041] It should be noted that, unless otherwise defined, the technical or scientific terms used in the present invention should have the usual meanings understood by people with ordinary skills in the field to which the present invention belongs. The "first", "second" and similar words used in the present invention do not indicate any order, quantity or importance, but are only used to distinguish different components. "Include" or "comprise" and similar words mean that the elements or objects appearing before the word include the elements or objects listed after the word and their equivalents, without excluding other elements or objects. "Connect" or "connected" and similar words are not limited to physical or mechanical connections, but may include electrical connections, whether direct or indirect. "Up", "down", "left", "right" and the like are only used to indicate relative positional relationships. When the absolute position of the object being described changes, the relative positional relationship may also change accordingly.

[0042] like Figure 1 、 Figure 2 As shown, the embodiment of the present invention first provides a design of the most basic structure of a composite high-speed aircraft, including: a fuselage 10, a main rotor automatic retractable mechanism 20, wings 30, canards 40 and propellers 50;

[0043] The two wings 30 are respectively placed on both sides of the fuselage 10 and close to the tail end of the fuselage 10;

[0044] The two canards 40 are respectively placed on both sides of the fuselage 10 and close to the front end of the fuselage 10; a propeller 50 is installed on the edge of each canard 40; the two propellers 50 are connected by a coordination shaft, and the two propellers 50 have the same speed and opposite directions; the coordination shaft passes through the canards 40 horizontally; a single main rotor automatic retractable mechanism 20 is installed on the top of the fuselage 10, along the longitudinal direction of the fuselage 10 and between the wing 30 and the canard 40; when the flight mode is switched, the main rotor in the main rotor automatic retractable mechanism 20 is folded and unfolded.

[0045] In the basic design of the composite high-speed aircraft provided above in the embodiment of the present invention, the aerodynamic efficiency of the aircraft can be improved and the maneuverability can be increased through the combination of the fuselage, wings, canards and propellers, making the aircraft more suitable for high-speed flight. The main rotor automatic retractable mechanism can realize the folding and unfolding of the main rotor when the flight mode is switched, thereby adapting to different flight modes and scenarios, and enhancing the adaptability and versatility of the aircraft. In addition, by installing propellers and coordinating shafts, the stability and maneuverability of the aircraft can be improved, while the flight noise can be reduced and the flight environment can be improved. The composite high-speed aircraft design has higher flight efficiency, adaptability and flight performance, and can quickly and flexibly perform various tasks and flight operations in the air.

[0046] The following describes in detail various preferred designs, design processes, and control methods of the composite high-speed aircraft according to embodiments of the present invention:

[0047] in, Figure 1 Corresponding to the basic structural diagram of the composite high-speed aircraft in the helicopter mode of this embodiment, Figure 2 Corresponding to the basic structure diagram of the compound high-speed aircraft in fixed-wing mode, the specific design process includes:

[0048] 1. Determine the layout of the composite high-speed aircraft:

[0049] The fuselage 10 adopts a theme layout including a single main rotor automatic retractable mechanism 20 , a wing 30 , a canard 40 and a propeller 50 .

[0050] The wing 30 is one of the key components of the aircraft and is mounted on the fuselage. Its primary function is to generate lift. The two wings 30 are placed on either side of the fuselage, close to the tail end. An aileron is mounted on each side of the trailing edge of each wing 30, and a tail fin 31 is mounted on each edge of each wing 30. The twin tail fins allow the rudder to be positioned away from the aircraft's centerline, making it less likely to be blocked by the fuselage and wings at large angles of attack. The twin tail fin design provides a larger stabilizer and rudder area, effectively reducing the height of the tail fins. The twin tail fins without a horizontal tail improve the aircraft's stability and maneuverability, reducing sideslip and steering resistance during flight and enhancing overall performance.

[0051] In one embodiment, two canards 40 are positioned on either side of the fuselage 10, near the front end. Two tiltable propellers 50 are mounted on the edges of the canards. During vertical takeoff and landing, hovering, and low-speed travel, the small propellers tilt to align with the fuselage to provide partial lift. The two propellers are designed with a positive and negative pitch angle, respectively, to generate a yaw moment to balance the main rotor's counter-torque. During high-speed travel, the propellers tilt perpendicular to the fuselage, becoming thrust propellers. A coordinating shaft passes transversely through the canards, connecting the two propellers to ensure they rotate at the same speed and in opposite directions.

[0052] The advantage of using canard layout is that it can provide greater lift than conventional layout under the same wing area. The flight of this aircraft is divided into two parts: helicopter mode and fixed-wing mode. The minimum level flight speed of the aircraft is Where W is the weight of the aircraft itself, ρ is the air density; S is the wingspan area; c Lmax is the maximum lift coefficient. The conversion speed between the two modes needs to satisfy v 转换 ≥v min To prevent the aircraft from stalling, greater lift under the same wing area means a smaller conversion speed, which can reduce the flight process of helicopter mode before mode conversion and improve flight efficiency; secondly, it is to make the balancing of the entire aircraft in the two modes adopt the same idea, that is, both are balancing models with the center of gravity between the two pulling forces, which is convenient for further design, analysis and production.

[0053] A single main rotor automatic retractable mechanism is installed on the top of the fuselage, along the longitudinal direction of the fuselage and between the wing and the canard.

[0054] In one embodiment, reference Figure 3 、 Figure 4 The main rotor automatic retractable mechanism is respectively a helicopter mode and a fixed-wing mode, and the main rotor automatic retractable mechanism includes a main rotor 1, a main rotor box 2, and a hydraulic press 3.

[0055] The two blades of the main rotor are independent of each other, and the main rotor can be raised and lowered by a hydraulic press inside the fuselage to retract into the fuselage for the purpose of engaging fixed-wing mode. The main rotor housing 2 is connected to the hydraulic press 3 via tabs on both sides. When the aircraft needs to switch to helicopter mode, the hydraulic press 3 operates to lift the main rotor housing 2 upward, allowing the main rotor 1 to extend out of the fuselage and then begin operation. When the aircraft needs to switch to fixed-wing mode, the main rotor 1 stops operating, and the hydraulic press 3 operates to retract the main rotor housing 2 downward, allowing the main rotor 1 to retract into the fuselage. To facilitate stowage and deployment, the main rotor can refer to the design of existing folding rotors.

[0056] 2. Select the external parameters of the main components of the composite high-speed aircraft, mainly involving the external parameters and installation positions of the main components such as wings, canards, fuselage, propellers, and main rotors.

[0057] In one embodiment, the overall parameters of the prototype helicopter are designed according to Table 1, and the maximum take-off weight of the prototype is calculated to be 500 kg, the wingspan is 7 m, the root chord length of the wings is 1 m, the canard span is 5 m, the root chord length of the canard is 0.6 m, the vertical tail span is 1 m, the root chord length of the vertical tail is 0.5 m, the main rotor diameter is 5.48 m, the propeller diameter is 1.332 m, the fuselage length is 6 m, the maximum width of the fuselage is 1.2 m, and the maximum height of the fuselage is 1.4 m. The overall parameters of this embodiment represent the modes of the composite high-speed aircraft helicopter as follows: Figure 1 As shown, the fixed-wing mode of the compound high-speed aircraft is as follows Figure 2 shown.

[0058] Table 1 Overall parameters of the prototype helicopter

[0059]

[0060]

[0061] In one embodiment, the aircraft's shape parameters and installation position are divided into two aspects: helicopter mode and fixed-wing mode, and the overall parameters are analyzed to analyze how they are weighed, selected, and iteratively calculated;

[0062] 1) Aerodynamic analysis of fixed-wing mode

[0063] In one embodiment, after the shape design is given, Xflr5 can provide an aerodynamic characteristics analysis of the entire aircraft based on the vortex lattice method (VLM).

[0064] The design points of mode conversion speed flight, cruising flight, and maximum speed flight are selected. These design points are all considered to be steady linear motion. Let the total lift be L, the total drag be D, the total weight of the aircraft be W, and the thrust of the propulsion system be T. Then, each state of the fixed-wing aircraft's steady level flight must satisfy: L = W, T = D. The atmospheric density is ρ, the flight speed is v, and the lift coefficient of the entire aircraft is C. L The drag coefficient of the whole machine is recorded as C D , the wing area is recorded as S, then the aircraft needs to meet the lift requirements when it is in steady level flight resistance

[0065] In one embodiment, Xflr5 is used to perform aircraft shape design and aerodynamic analysis, and the parameters involved in the design process are recorded as shown in Table 2:

[0066] Table 2 Aerodynamic parameters during the design process

[0067]

[0068]

[0069] Therefore, in this embodiment, the installation angle of the wing is set to 3.5°; when the aircraft switches between helicopter mode and fixed-wing mode, the aircraft angle of attack changes between 0° and 13°; the specific analysis process is as follows

[0070] 1.1) First, we need to introduce a concept: the minimum level flight speed v of the aircraft. min The minimum level flight speed refers to the minimum speed at which the aircraft can maintain steady straight and level flight at a certain altitude. The actual minimum level flight speed v min It will be restricted by many factors, such as lift restriction, vibration restriction and maximum horizontal tail deviation angle restriction. This embodiment considers the lift restriction during low-speed and high-angle flight. It can be calculated that the minimum level flight speed of the aircraft in fixed-wing mode at 0m is 39.3m / s. The conversion speed v between the two modes should be maintained during design. 转换 ≥v min To prevent the aircraft from stalling during mode switching, the mission profile has been taken as follows:

[0071] v 转换 =1.2v min =47.16m / s

[0072] Xflr5 shows that the maximum lift coefficient is at 13° angle of attack, C Lmax =0.987. Among them, the mission profile is shown as follows Figure 5 shown.

[0073] 1.2) The cruising state of an aircraft is the most economical flight state. The ratio of the aircraft's lift to drag is defined as the lift-to-drag ratio, which is expressed as K. The aircraft has the largest lift-to-drag ratio in the cruising state, and the corresponding angle of attack is called the favorable angle of attack α opt , the speed is called the favorable speed V opt Xflr5 can give the K-α graph of the lift-to-drag coefficient ratio of the whole aircraft and the angle of attack, see Figure 6 , we can obtain the maximum lift-to-drag ratio state, that is, the cruise state angle of attack α opt and lift coefficient C Lopt ,Depend on The aircraft cruising speed can be calculated.

[0074] Favorable angle of attack α opt =3.5°, the maximum lift-to-drag ratio Kmax = 17.655. In actual use, the aircraft generally maintains a constant altitude and constant speed cruise, that is, a constant straight flight at a certain altitude. After calculation, at sea level (i.e.: sea level 0m) vopt =64.525m / s; v at 1000m opt =67.733m / s; v at 3000m opt =74.897m / s.

[0075] The aircraft maintains a horizontal flight during cruising, so the wing mounting angle is set at 3.5°.

[0076] 1.3) Resistance in Steady Straight Flight The lift coefficient of the aircraft is C L and the drag coefficient C D Therefore, when flying at maximum speed, keep the angle of attack small to reduce the drag.

[0077] Choose 0 angle of attack for calculation, then C L =0.129,

[0078] 2) Helicopter mode parameters

[0079] In this embodiment, the design of the helicopter mode is mainly based on the slipstream theory (i.e., momentum method) during hovering and forward flight, and the main content is the aerodynamic analysis during hovering and forward flight, so as to determine the parameters of the main rotor and propeller.

[0080] Let the pulling force be T, the total weight be W, and the available power be N. 可用 , the required power is P 需用 When flying in helicopter mode, T≥W, N must be satisfied. 可用 ≥P 需用 .

[0081] The atmospheric density is recorded as ρ, the flight speed is recorded as v, and the lift coefficient of the entire aircraft is recorded as C L The wing area is recorded as S. The design point of the helicopter is also a steady level flight state, which needs to meet the lift resistance

[0082] In this embodiment, the helicopter mode parameters of the aircraft are determined, specifically the parameters and installation positions of the main rotor propeller and other major components, as shown in Table 3:

[0083] Table 3 Helicopter mode parameters

[0084]

[0085] The specific calculation process is as follows:

[0086] 2.1) Propeller radius calculation

[0087] Consider the total aircraft resistance at 3000m and 126.156m / s (maximum speed) SC D =417.828N. At this time, the total thrust of the aircraft is provided by two propellers. Each state of the aircraft in steady level flight must meet the following requirements: L = W, T = D. Then the propeller thrust T 小 =0.5D=208.914N.

[0088] Define a parameter, the rotor load P, where the total weight of the helicopter is W and the rotor radius is R, then P = πR 2 , the rotor load of a conventional helicopter is P = 150N / m 2 —450N / m 2 Generally speaking, the greater the propeller disc load P, the greater the power required during hovering. In order to reduce the required power, a smaller propeller disc load P is used. 小 =150N / m 2 .

[0089] The blade tip loss coefficient k=0.92.

[0090] According to the slipstream theory during vertical flight, the relative airflow velocity at the propeller disc is defined as v1, and the induced velocity is calculated as The propeller radius

[0091] 2.2) Calculation of main rotor radius

[0092] Consider the lift situation when hovering at 0m. The slipstream theory for vertical flight is still used, and the propeller disc load is still taken as 150N / m. 2 The formula is the same as above.

[0093] At this time, v1=8.158m / s, T 小 =2ρπv1 2 kr 2 =209.035N. Then T 主 =W-2T 小 =4481.93N.

[0094] The greater the propeller load, the smaller the main rotor radius. In fixed-wing mode, the main rotor needs to be retracted into the fuselage, which is restricted by the fuselage length. At the same time, the greater the propeller load P, the greater the power required, which is also restricted by the available power of the engine. The propeller load is taken as 190N / m 2 .

[0095] The induced speed of the main rotor is v1 = 9.181 m / s. 主 =2.740m.

[0096] In this embodiment, the center of gravity of the whole aircraft is used as the reference point, the total lift of the two propellers is T1, the length of the moment arm relative to the center of gravity is L1, the total lift of the main rotor is T2, and the length of the moment arm relative to the center of gravity is L2, then it is necessary to satisfy T1×L1=T2×L2. In this embodiment, the stability analysis is performed, and the center of gravity position is determined as X in the calculation of longitudinal static stability. G =2.746m, the propeller x-position is 0.8m, and the main rotor x-position is 2.93m, it can be verified that the pitching moment about the center of gravity can be balanced.

[0097] 2.3) Lift calculation for forward flight at sea level

[0098] Consider the lift conditions in forward flight at 0 m and transition speed. Use the slipstream theory in forward flight.

[0099] Define “characteristic induction velocity” v 10 , which is the average induced speed at the propeller disk in the hovering state while maintaining the same drag coefficient as in forward flight. Forward flight speed v0 = 39.3 m / s. And v1 is the induced speed at the main rotor blade during forward flight, which is given by It can be obtained that v1 = 2.145m / s.

[0100] The tension formula is T = 2ρπR 2 (v0+v1)v1k, note that v0+v1 refers to the vector sum of the two, because the inclination angle of the blade disk is very small, it is approximately equal to v0. We can get T 主 =4481.514N. The working condition of the propeller in hovering state is T 小 = 209.035 N. It can be seen that the rotor system alone can provide 500 kg of lift when flying forward at the conversion speed.

[0101] 2.4) Balanced counter-torque calculation

[0102] The left and right propellers have a positive and negative pitch angle respectively, and the components of the propeller thrust are in opposite directions, so a yaw moment can be obtained to balance the anti-torque of the main rotor.

[0103] In this embodiment, the main rotor's counter torque can be Calculate M k =117.64NM, the propeller is installed in a small cabin at the canard wingtip, and the pulling force of a single propeller is expressed as T 小 , the inclination angle of the two propellers in vertical flight is expressed as α, and the canard wingspan L = 5m, then the directional moment balance of helicopter mode flight needs to satisfy T 小 × sinα × L=M k, calculating α = 6.5°. This method of balancing counter-torque leads to a coupling problem in balancing the pitching moment and the yaw moment. As the propeller tilts, its vertical thrust component, i.e., the propeller's lift, decreases, resulting in an unbalanced pitching moment. Furthermore, if the main rotor thrust remains constant, the propeller thrust must be increased to maintain vertical force balance. This increases the horizontal thrust component, which in turn leads to an unbalanced yaw moment. To solve this problem, the thrust and mounting position of the main rotor must be adjusted. Let the total weight of the aircraft be W, the total thrust of the two propellers be T1, the moment arm length relative to the center of gravity be L1, the total lift of the main rotor be T2, and the moment arm length relative to the center of gravity be L2. The longitudinal pitching moment balance must satisfy T1 × cos α × = T2 × L2, and the vertical force balance must satisfy T1 × cos α + T2 = W. It can be calculated that the main rotor thrust T2 = 4484.62N, L2 = 0.18m, and the actual rotor load when the main rotor is hovering is The actual installation position is 2.926m.

[0104] 3. Design the wings, canards, and other parts of the fuselage, including ailerons, lift-enhancing devices, landing gear, and engines.

[0105] The design in this embodiment is mainly based on selection and appearance design, so the parameters of related components are only temporary values ​​in the preliminary design process and can be determined through other calculations and experiments.

[0106] 1) Aileron design

[0107] The aileron is one of the main control components for the lateral heading of the aircraft. In this embodiment, the shape and installation position of the aileron are preliminarily selected. The selection of specific parameters requires further precise aerodynamic analysis.

[0108] The principles for selecting the ailerons in this embodiment are as follows: ① The ailerons are generally installed on both sides of the wing's trailing edge. ② The aileron area is generally 5%-7% of the wing area. ③ The aileron chord length is generally 20%-25% of the wing chord length. ④ While meeting the aircraft's lateral maneuverability requirements, the aileron span should be as small as possible to maximize the flap span. Since this embodiment does not include trailing edge flaps, the aileron span range is relatively wide. The aileron parameters are shown in Table 4:

[0109] Table 4 Aileron parameters

[0110] <![CDATA[副翼面积(m 2 )]]> 0.288 Aileron chord length (m) 0.16 Aileron span (m) 1.8 Aileron installation y-axis position (m) ±1.4

[0111] 2) Design of lift-enhancing device

[0112] The main function of an aircraft's high-lift device is to provide additional lift at low speeds and delay airflow separation to improve the aircraft's stall performance. By utilizing helicopter-like vertical takeoff and landing, the hybrid aircraft avoids the takeoff and landing phases and low-speed forward flight of fixed-wing aircraft. This results in better low-speed performance and, compared to fixed-wing aircraft, is less dependent on the additional lift provided by the high-lift device. Therefore, this embodiment does not incorporate a high-lift device.

[0113] 3) Landing gear design

[0114] The landing gear is an attachment on the underside of an aircraft that supports the aircraft during takeoff, landing, or taxiing, and allows for ground mobility. The landing gear is the only component that supports the entire aircraft and is therefore an integral part of the aircraft; without it, the aircraft would be unable to take off, land, or move on the ground. In this embodiment, after takeoff, the landing gear can be retracted depending on the aircraft's performance.

[0115] 4) Engine design

[0116] Aircraft engines mainly include piston and jet engines. Piston engines can be divided into liquid-cooled and air-cooled types; jet engines can be divided into air jet engines and rocket engines. In this embodiment, the engine is selected as the ROTAX914 piston engine.

[0117] 4. Estimate the weight and center of gravity of the composite high-speed aircraft and verify and calculate its stability.

[0118] The estimation process of the weight and center of gravity of the entire aircraft and the verification and calculation of its stability are related to whether the entire aircraft can fly.

[0119] 1) Weight center of gravity estimation

[0120] The weight of an aircraft is divided into three parts: empty weight ME, fuel weight MF and payload weight MPL. In this embodiment, the weight estimation only considers the empty weight part. The fuel weight needs to be determined according to the requirements of the range and flight time. The maximum take-off weight is W, and the payload weight can be calculated by MPL = W-ME-MF. The empty weight only considers the weight of the wings, canards, vertical tail, fuselage, propellers, main rotors, propeller cabins, electromechanical devices, landing gear, and power systems. The weight and position of other major components such as the transmission system, fuel tanks, cargo compartments, etc. must meet the center of gravity requirements during design. The center of gravity position is mainly determined by the requirements of longitudinal static stability. In the preliminary design stage, it is necessary to first give a center of gravity position that meets the stability, and then adjust the weight and layout of each component to make the actual center of gravity coincide with the theoretical center of gravity. The weight and installation position of each component in this embodiment are shown in Table 5:

[0121] Table 5 Weight and installation position of some components

[0122]

[0123] In this embodiment, the center of gravity position is calculated using Xflr5 software as X_C G =2.746m,Y_C G =0.000m,Z_C G =0.059m.

[0124] 2) Stability analysis

[0125] In this embodiment, the longitudinal static stability and longitudinal dynamic stability of the composite aircraft are calculated and verified, and it is proved that the aircraft has good longitudinal and lateral stability characteristics.

[0126] 2.1) Longitudinal static stability

[0127] The static stability of an aircraft refers to its tendency to return to its original equilibrium position after being disturbed. The reference point for the longitudinal moment of the aircraft is the center of gravity C. G . Depends on the mass distribution of the aircraft; the point where the resultant force of the aircraft's aerodynamic force acts is called the pressure center, and the pressure center changes with the angle of attack during flight; but when the angle of attack changes, the point of action of the aircraft's aerodynamic force increment is not the pressure center, but the focus (also called the aerodynamic center), which has nothing to do with the angle of attack and changes with the Mach number Ma. The longitudinal static stability of the aircraft requires that the focus of the entire aircraft is located behind the center of gravity. The parameters and installation positions of the wings, fuselage and horizontal tail (or canard) will affect the longitudinal static stability of the entire aircraft. Generally, positively cambered wings and fuselage are statically unstable components, while horizontal tails are statically stable components. There are two indicators to measure the longitudinal static stability of an aircraft. One is that the slope of the Cm-α graph is negative, that is, the coefficient Cm-α is guaranteed to be <0; the other is that Cm-C L The vertical intercept of the graph is positive to ensure that the aircraft has longitudinal static stability and can be trimmed at a positive angle of attack. This example uses Xflr5 for analysis. Figure 7 The Cm-α plot of this embodiment given by Xflr5 is: Figure 8 The Cm-C of this example given by Xflr5 L The graph proves that this embodiment has longitudinal static stability.

[0128] 2.2) Longitudinal dynamic stability

[0129] The dynamic stability of an aircraft refers to the overall characteristics of the aircraft during the recovery process after being disturbed. The longitudinal motion of the aircraft is described by four variables: velocity v, angle of attack α, pitch angular velocity q, and track inclination θ.

[0130] The characteristic equation of the longitudinal motion equation is a quartic equation of one variable. Its corresponding characteristic roots can be divided into two pairs. The two pairs of characteristic roots describe two longitudinal motion modes. That is, the original fourth-order differential system corresponds to two second-order oscillatory systems. The characteristic roots of a second-order oscillatory system reflect the characteristics of the motion mode corresponding to this system. Each real characteristic root or each pair of conjugate complex roots represents a basic modal motion:

[0131] ① If the characteristic root is a real root, then this motion mode is a non-periodic exponential motion;

[0132] ② The characteristic roots are conjugate complex roots, and this motion is a periodic oscillation motion;

[0133] ③ If the real part of the characteristic root is negative, then the motion mode converges;

[0134] ④ If the real part of the characteristic root is positive, then the motion mode diverges;

[0135] ⑤ If the real part of the characteristic root is 0, then this motion mode is critically stable.

[0136] The larger the absolute value of the real part of the characteristic root, that is, the farther it is from the imaginary axis in the complex plane, the faster the mode converges. Conversely, the larger the absolute value of the imaginary part of the characteristic root, that is, the farther it is from the real axis in the complex plane, the faster the mode oscillates and the shorter its period. The characteristic roots of the longitudinal motion equation of a conventional aircraft consist of two sets of conjugate complex roots. The corresponding motion mode is called a "short-period mode" if the real part has a large absolute value and is always negative. The primary motion variables are velocity v and aircraft attitude angle, where the pitch rate changes ahead of the angle of attack. This motion mode has a very short period and high vibration frequency, making it difficult for the pilot to intervene. Therefore, aircraft require high performance in short-period modes. The corresponding motion mode is called a "long-period mode" if the real part has a small absolute value and is usually negative. The primary motion variables are velocity v and track inclination angle θ, where the velocity changes ahead of the track angle. This motion mode has a long period, high vibration frequency, and slow motion, hence the name "sinking and floating" motion, making it easy for the pilot to control. Long-period modes of aircraft can be unstable.

[0137] By simulating the characteristic root trajectory of the longitudinal motion equation of the composite high-speed aircraft in this embodiment, it is found that there are two sets of characteristic roots, namely: λ1 = -4.5296 ± 6.7240i, whose real part is negative and the absolute value is large, corresponding to the short-period mode; λ2 = -0.0071 ± 0.2146i, whose real part is negative and the absolute value is small, corresponding to the long-period mode.

[0138] 2.3) Lateral dynamic stability

[0139] The dynamic stability of an aircraft refers to its overall performance during the recovery process after a disturbance. The lateral motion of an aircraft is described by four variables: yaw angle β, roll angular velocity p, yaw angular velocity r, and roll angle φ. The characteristic equation of the lateral motion equation is a quartic equation. For a normally configured aircraft, the roots of this characteristic equation consist of three components: a negative real root with a large absolute value. This corresponds to the "roll mode," whose primary motion variables are roll angular velocity p and roll angle φ. This is a non-periodic motion with rapid decay. A pair of conjugate complex roots with negative real parts corresponds to the "Dutch roll mode," whose primary motion variables are roll angle φ, sideslip angle β, and yaw angle ψ. This is a periodic motion with a period of several to more than ten seconds and a high oscillation frequency. Aircraft have stringent requirements for the characteristics of the Dutch roll mode. A real root, which can be positive or negative, has a small absolute value of the real part. This corresponds to a "helical mode" motion, with the primary motion variables being the yaw angle ψ and the roll angle φ. This is a non-periodic motion, with both convergence and divergence occurring very slowly, on the order of hundreds of seconds, making corrections easy. Aircraft do not require high-quality helical mode characteristics.

[0140] By simulating the characteristic root trajectories of the lateral and yaw motion equations of the composite high-speed aircraft in this embodiment, three groups of characteristic roots are obtained, namely: λ1 = -10.8595, corresponding to the roll mode; λ2 = -0.6874±4.2516i, corresponding to the Dutch roll mode; λ3 = 0.0181, which is a real root with a positive real part and a small absolute value, corresponding to the spiral mode.

[0141] Based on the relevant concepts and theories of aircraft stability, the calculation and verification process of static stability and dynamic stability in this embodiment shows that the aircraft already has good longitudinal and lateral stability characteristics. In short, the aircraft of this embodiment can fly and fly smoothly.

[0142] This invention, based on flight dynamics calculations and simulations, considers parameters from all aspects of aircraft design, including the specific design parameters of the rotors, wings, and fuselage. All design parameters are reflected in the equilibrium equations and software simulations. Therefore, calculations and simulations can be fed back into the overall design process, providing a reference for optimizing overall aircraft parameters. Furthermore, this embodiment also verifies and calculates aircraft stability.

[0143] According to the above aircraft body design, the embodiment of the present invention further provides a control method of a composite high-speed aircraft, referring to Figure 5 A task diagram of the flight process and mode conversion process of a composite high-speed aircraft provided in an embodiment of the present invention.

[0144] In this embodiment, the aircraft's angle of attack and flight speed are changed by extending and retracting the main rotor and tilt propellers. The flight process and mode conversion process are designed. Under normal circumstances, a complete flight process can be regarded as the following steps:

[0145] Step 1: The aircraft takes off vertically in helicopter mode and accelerates forward horizontally at low altitude. During this process, the aircraft maintains a zero angle of attack and the main rotor begins to tilt forward, providing the thrust required for the helicopter to move forward. At the same time, the lift it can provide decreases. During this process, the wing lift increases with the increase in flight speed, thereby gradually sharing the lift on the wing.

[0146] Step 2: When the aircraft reaches the fixed-wing mode's minimum stall speed of 39.3 m / s, stop horizontal acceleration and maintain this speed. The aircraft then begins to increase its angle of attack, reducing the total main rotor torque, further distributing lift to the wing. When the angle of attack reaches 13°, the wing and canards can provide full lift. The aircraft continues to accelerate to 1.2 Vmin, or 47.16 m / s. At this point, the main rotor stops and retracts into the fuselage, and the propellers tilt, completing the mode transition.

[0147] Step 3: The aircraft then begins to accelerate horizontally to 64.525 m / s, gradually reducing its angle of attack, and eventually flies forward horizontally at an angle of attack of 3.5°. The aircraft then steadily climbs to an altitude of 3000 m, still flying forward horizontally at a 3.5° angle of attack, at a speed of 74.987 m / s, entering the cruise process.

[0148] Step 4: After the 3000m cruise, the aircraft descends steadily to a low altitude, flying horizontally at a 3.5° angle of attack and a speed of 64.525 m / s. It then begins horizontal forward flight, gradually increasing the angle of attack to 13°.

[0149] Step 5: The aircraft continues until its speed drops to 39.3 m / s. The main rotors extend from the fuselage and begin rotating. The propellers tilt again, returning to helicopter mode. The aircraft maintains a speed of 39.3 m / s, and the angle of attack gradually decreases from 13° to 0°. During this process, the main rotors increase their total torque to maintain lift. Once the angle of attack reaches 0°, the aircraft resumes horizontal forward flight, decelerating. The main rotors begin to reduce their forward tilt, gradually distributing the entire aircraft's lift to the main rotors. The aircraft then completes a vertical landing in helicopter mode, concluding the flight.

[0150] In this example, Xflr5 was used for shape design and aerodynamic analysis. The principles and key formulas involved in the calculations were documented. Key design points throughout the flight process were selected for rigorous and detailed aerodynamic calculations. Based on this, the flight process and mode transitions were designed to refine the mission profile. By switching between helicopter and fixed-wing modes, the aircraft's speed bottleneck was overcome, achieving the goal of a hybrid high-speed aircraft.

[0151] The various embodiments in this specification are described in a progressive manner, with each embodiment focusing on the differences from the other embodiments. Reference can be made to the description of the similarities between the various embodiments. The aircraft control methods disclosed in the embodiments are described briefly because they correspond to the aircraft disclosed in the embodiments. For relevant details, refer to the description of the aircraft.

[0152] This patent is not limited to the above-mentioned optimal implementation method. Anyone can derive various other forms of a composite high-speed aircraft and its control method under the inspiration of this patent. All equal changes and modifications made according to the scope of the patent application of this invention should be covered by this patent.

Claims

1. A composite high-speed aircraft, characterized by: The two wings (30) are respectively arranged on both sides of the fuselage (10) and close to the tail end of the fuselage; The two canards (40) are respectively arranged on both sides of the fuselage (10) and close to the front end of the fuselage; a propeller is respectively installed at the wingtip of the two canards; the two propellers (50) are connected by a coordination shaft, and have the same rotation speed and opposite rotation directions; the coordination shaft passes through the two canards (40) transversely; The main rotor automatic retractable mechanism (20) is installed on the top of the fuselage and is located between the wing (30) and the canard (40) in the longitudinal direction of the fuselage; when the flight mode is switched, the main rotor of the main rotor automatic retractable mechanism (20) is folded or unfolded; The main rotor automatic retractable mechanism (20) comprises: a main rotor, a main rotor box and a hydraulic press; The main rotor is connected to the main rotor box through a rotating main shaft; the main rotor box is fixedly connected to the hydraulic press through ears on both sides of the hydraulic press; the hydraulic press is fixedly installed inside the fuselage, and drives the main rotor box and the main rotor to extend out of the fuselage or retract into the fuselage through telescopic movement.

2. A composite high-speed aircraft according to claim 1, characterized in that: The flight modes of the aircraft include: helicopter mode and fixed-wing mode; In the helicopter mode, the main rotor in the main rotor automatic retractable mechanism (20) is in an extended state, and during vertical take-off and landing, hovering, and low-speed forward movement, the propeller of the main rotor is adjusted to be parallel to the fuselage by tilting to provide partial lift, and the pitch angle of one of the two propellers (50) is positive and the other is negative, so as to obtain a yaw moment to balance the anti-torque of the main rotor; In the fixed-wing mode, the main rotor in the main rotor automatic retractable mechanism is in a folded state, and in a high-speed forward state, the propeller (50) is tilted to a direction perpendicular to the fuselage to serve as a thrust propeller.

3. A composite high-speed aircraft according to claim 1, characterized in that: The installation angle of the wing (30) is 3.5°.

4. A composite high-speed aircraft according to claim 2, characterized in that: When the aircraft switches between helicopter mode and fixed-wing mode, the aircraft's angle of attack changes between 0° and 13°.

5. The composite high-speed aircraft according to claim 1, characterized in that: Ailerons are respectively installed at the trailing edges of the two wings (30), and vertical tails (31) are respectively installed at the wing tips.

6. A composite high-speed aircraft according to claim 5, characterized in that: The aileron area is 5% to 7% of the wing area; The chord length of the aileron is 20% to 25% of the chord length of the wing.

7. The composite high-speed aircraft according to claim 1, characterized in that: The inclination angle of the two propellers during vertical flight is denoted as α, the total weight of the aircraft is W, and the total thrust of the two propellers is , the length of the force arm relative to the center of gravity is , the total lift of the main rotor is , the length of the force arm relative to the center of gravity is , then the balance of the longitudinal pitching moment needs to satisfy × cos α×= × , the vertical force balance needs to satisfy × cos α+ = W .

8. A method for controlling a composite high-speed aircraft according to any one of claims 1 to 7, characterized in that: The aircraft takes off vertically in helicopter mode and reaches the minimum stall speed V in fixed-wing mode in forward flight. min The angle of attack is maintained at 0 before; during the acceleration phase, the main rotor provides the thrust required for the helicopter to move forward; In the aircraft from V min Accelerate to 1.2V min During the flight, the angle of attack is first adjusted from 0 to 13°, and then the main rotor is retracted and the propeller is tilted to convert the mode to fixed-wing mode.

9. A method for controlling a composite high-speed aircraft according to any one of claims 1 to 7, characterized in that: The aircraft matches the optimal speed and cruise state with an angle of attack of 3.5°, and adjusts the angle of attack to 13° when deceleration is required. min When the mode is switched to helicopter mode, the speed is kept at V min to adjust the angle of attack to 0 for further deceleration and / or landing in helicopter mode.

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