Electric tilt rotor aircraft and control system thereof

Abstract

An electric tilt rotor aircraft and a control system thereof. The aircraft comprises a fuselage (1), wings (2), horizontal stabilizers (3) and a landing gear system (4), wherein the fuselage (1) comprises a fuselage body and a rudder (10) for controlling and adjusting a flight direction of the aircraft; the wings (2) each comprise an integral wing box section (21), an aileron (22), a wing left / right nacelle (23) and a supporting rod left / right nacelle (24); left and right supporting rods are integrated on front sides of the integral wing box sections (21); the wing left / right nacelle (23) is connected to the integral wing box section (21) by means of a tilting shaft (25), and the wing left / right nacelle (23) rotates around the tilting shaft (25); the horizontal stabilizers (3) each comprise a horizontal stabilizer box section (31) and a horizontal stabilizer left / right nacelle (32); the horizontal stabilizer left / right nacelle (32) is connected to the horizontal stabilizer box section (31) by means of the tilting shaft (25), and the horizontal stabilizer left / right nacelle (32) rotates around the tilting shaft (25); the landing gear system (4) uses a tricycle landing gear system, and comprises a nose landing gear (41) and two main landing gears (42); brake systems are integrated in the main landing gears (42); and the nose landing gear (41) plays a role in balancing and controlling direction. By using the electric tilt rotor aircraft, urban air traffic can be realized.

Description

An electric tilt-rotor aircraft and its control system

[0001] CROSS-REFERENCE TO RELATED APPLICATIONS

[0002] This application is based on the Chinese patent application with application number 2023107155089 and application date June 15, 2023, and claims the priority of the Chinese patent application. The entire content of the Chinese patent application is hereby introduced into this application as a reference. Technical Field

[0003] The present invention relates to the technical field of aircraft, and in particular to an electric tilt-rotor aircraft and system. Background Art

[0004] With the increasing density of vehicles on urban roads, daily commuting times are increasing. Capitalizing on the untapped availability of low- and medium-altitude airspace in cities, the concept of Urban Air Mobility (UAM) has been proposed. Electric vertical take-off and landing (EVTOL) aircraft, with their environmentally friendly nature and minimal reliance on infrastructure, have become the leading UAM solution. Currently, two common layouts exist in the industry. One is the multi-rotor layout: This type of aircraft relies on multiple lift rotors for vertical take-off and landing, as well as forward flight propulsion. Because they lack the complex mechanical structure of helicopter rotors, their flight speed is limited by rotor loads, and their endurance is poor due to lower flight efficiency. The other common layout is the "lift + thrust" combination: This type of aircraft combines a traditional fixed-rotor layout with multiple lift rotors to achieve vertical take-off and landing. During cruise, the lift rotors are deactivated, relying solely on propulsion for forward flight. The drag generated by these lift rotors during cruise significantly reduces flight efficiency.

[0005] To this end, the present invention designs an electric tilt-rotor aircraft, and simultaneously designs an operation control method and an electrical control method of the electric tilt-rotor aircraft.

[0006] Summary of the Invention

[0007] The present invention provides an electric tilt-rotor aircraft, comprising: a fuselage, wings, a horizontal tail, and a landing gear;

[0008] The fuselage consists of a main body and a rudder. The main body integrates the cabin door, windshield, lower window and rear window. The rudder is installed at the rear of the main body and is used to control and adjust the flight direction of the aircraft. The cockpit is set in the main body and includes a display console, pilot seat, front passenger seat, rear passenger seat, and the rear of the cockpit is the luggage compartment.

[0009] The wing-body structure is a separate structure, including an integral wing box section, ailerons, left / right wing nacelles, and left / right strut nacelles. The left / right struts are integrated into the front side of the integral wing box section. The left / right wing nacelles are connected to the integral wing box section through a tilt axis, and the left / right wing nacelles rotate around the tilt axis. The left / right wing nacelles are equipped with motors and controllers, and the rotors are installed on the motor output shafts. The wings adopt an upper single gull-wing layout to increase the ground clearance of the wings and rotors.

[0010] The horizontal tail includes a horizontal tail box section and a left / right horizontal tail nacelle. The left / right horizontal tail nacelle is connected to the horizontal tail box section via a tilt axis, and the left / right horizontal tail rotates around the tilt axis.

[0011] Rotors are installed on the left / right strut nacelles, the left / right wing nacelles, and the left / right horizontal tail nacelles. The rotors on the left / right wing nacelles and the left / right horizontal tail nacelles are rotors that rotate around a tilt axis; the rotors on the left / right strut nacelles are foldable fixed rotors that are foldable but fixed in position and cannot be tilted; the left / right wing nacelles and the left / right horizontal tail nacelles each include an integrally assembled nacelle rib and a servo mounting frame, the servo mounting frame fixing the servo bracket, the servo bracket installing the servo via a threaded pin, and the servo rotates around the threaded pin axis; one end of the servo is fixed to the tilt axis via a tilt servo arm; the servo drives the nacelle portion to tilt by driving a retractable push rod of the servo;

[0012] The landing gear adopts a three-point landing gear, including a front landing gear and two main landing gears. The main landing gear is integrated with a braking system, which acts as a brake when the aircraft lands, and the front landing gear plays a role of balance and direction control.

[0013] As described above, an electric tilt-rotor aircraft has a cabin layout of 1+2+2, that is, the first row is a pilot seat, which is deployed in the center, the second row is two front passenger seats, and the third row is two rear passenger seats; after opening the left / right cabin door, the passengers in the third row can reach their seats through the space between the first and second row seats; after the pilot boards the aircraft through the left / right cabin door, he or she can reach the pilot seat through the aisles on both sides.

[0014] In the electric tilt-rotor aircraft as described above, an instrument panel is located in front of the pilot's seat, with the front instrument panel slightly offset to the right to ensure that the pilot's eyes are in the center of the PFD. A throttle panel and a side stick are respectively deployed on the left and right sides of the seat. The pilot controls the throttle panel and the side stick to accelerate / decelerate, ascend / descend, pitch, roll, and turn the aircraft.

[0015] As described above, an electric tilt-rotor aircraft is provided, wherein the rotor is a five-blade rotor, which is connected to a pitch-changing mechanism. The pitch-changing mechanism is a torque transmission structure used to amplify the control input displacement, amplify the control input force, or convert the arc motion control input into a strict linear motion output, and is used to change the rotor blade angle. The pitch-changing mechanism includes a torque input point structure, a torque transmission structure, and a torque output point structure.

[0016] As described above, an electric tilt-rotor aircraft, wherein the torque input point structure includes an integrally formed joystick and an input end block, the input end block is fixed to one end of the joystick, the input end block is connected to the servo arm / push rod of the servo, and the joystick is driven to rotate by the movement of the servo arm / push rod; the other end of the joystick is connected to the torque transmission structure, the top of the torque transmission structure is connected to multiple torque output point structures, the rotation of the joystick drives the torque transmission structure to move, the torque transmission structure transfers the motion to the torque output point structure, driving the torque output point structure to move; the upper end of each torque output point structure is directly connected to the blade hub pitch hinge corresponding to the rotor, and the movement of the torque output point structure drives the blade hub pitch hinge to move, thereby realizing the change of the blade angle.

[0017] In the electric tilt-rotor aircraft as described above, the fixed rotor is configured as a four-blade foldable rotor.

[0018] As described above, the electric tilt-rotor aircraft, wherein the servo specifically includes a servo rod, a retractable push rod and a tilt servo arm, the threaded pin on the servo bracket fixes the servo rod, the retractable push rod is inserted into the servo rod and is retractably connected to the servo rod, a servo motor is arranged in the servo rod, the servo motor drives the retractable push rod to retract and retract along the servo rod, the other end of the retractable push rod is fixed to one end of the tilt servo arm by a fastener, and the other end of the tilt servo arm is sleeved on the tilt shaft and fixed to the tilt shaft.

[0019] An electric tilt-rotor aircraft as described above, wherein the aircraft's configuration modes include a multi-rotor mode, a transition mode, and a fixed-rotor mode; the aircraft uses the multi-rotor mode for takeoff and landing, and dynamically switches between the multi-rotor mode, transition mode, and fixed-rotor mode in the air.

[0020] The beneficial effects achieved by the present invention are as follows: The present invention provides an electric tilt-rotor aircraft, which introduces the overall construction mode of the aircraft from the perspectives of structure, power system, and control system, respectively, to solve the difficulties faced in the transition conversion process of vertical take-off and landing aircraft. The power system provides the entire aircraft with a variety of high-voltage architectural power and can effectively prevent safety issues caused by single-point failures. The control system realizes precise vertical take-off and landing control, tilt transition control, and fixed-rotor cruise flight control, thereby improving the flight accuracy and safety of the aircraft. BRIEF DESCRIPTION OF THE DRAWINGS

[0021] In order to more clearly illustrate the embodiments of the present invention or the technical solutions in the prior art, the following briefly introduces the drawings required for use in the embodiments or the description of the prior art. Obviously, the drawings described below are only some embodiments described in the present invention. For ordinary technicians in this field, other drawings can also be obtained based on these drawings.

[0022] 1 and 2 are schematic diagrams of the external structure of an electric tilt-rotor aircraft provided in Embodiment 1 of the present invention;

[0023] FIG3 is a view of the aircraft with doors and windows;

[0024] FIG4 is a view of the interior of the cabin without doors and windows;

[0025] Figures 5 and 6 are disassembled views of the aircraft;

[0026] Figures 7 and 8 are schematic diagrams of the left and right nacelles of the wing;

[0027] Figure 9 is a schematic diagram of the arrangement of the tilt axis and the steering gear;

[0028] Figure 10 is a schematic diagram of the nacelle tilting motion;

[0029] Figure 11 is a schematic diagram of rotor rotation;

[0030] Figures 12 and 13 are schematic diagrams of the integrated structure of the reducer;

[0031] Figures 14-22 are schematic diagrams of various situations of the aircraft high-pressure architecture;

[0032] Figure 23 is a schematic diagram of a low voltage control architecture;

[0033] Figure 24 is a schematic diagram of a low voltage power distribution architecture;

[0034] Figure 25 is a schematic diagram of the internal avionics and flight control architecture of the aircraft;

[0035] Figure 26 is a schematic diagram of the flight control system;

[0036] Figure 27 is a schematic diagram of motor tilt control;

[0037] Figures 28 and 29 are schematic diagrams of multi-rotor control at low speed;

[0038] FIG30 is a schematic diagram of a transitional mode flight control mode;

[0039] Figure 31 is a schematic diagram of the fixed rotor flight phase;

[0040] FIG32 is a schematic diagram of the functional module architecture of the control and distribution system;

[0041] Figure 33 is a schematic diagram of the flight management module control;

[0042] Figure 34 is a schematic diagram of attitude angle control;

[0043] FIG35 is a schematic diagram of the desired angular velocity calculation;

[0044] FIG36 is a schematic diagram of an attitude angular velocity control loop;

[0045] Figure 37 is a schematic diagram of the actuator allocation module architecture. DETAILED DESCRIPTION

[0046] The following is a clear and complete description of the technical solutions in the embodiments of the present invention, in conjunction with the accompanying drawings. Obviously, the described embodiments are only part of the embodiments of the present invention, not all of them. Based on the embodiments of the present invention, all other embodiments obtained by those skilled in the art without creative work are within the scope of protection of the present invention.

[0047] Example 1

[0048] As shown in Figures 1 and 2, the first embodiment of the present invention provides an electric tilt-rotor aircraft including: a fuselage 1, wings 2, a horizontal tail 3, and a landing gear 4.

[0049] The fuselage 1 includes a main body and a rudder 10. The main body integrates the cabin door, windshield, lower window, and rear window. The rudder is mounted at the rear of the main body and is used to control and adjust the aircraft's flight direction. The external configuration of the fuselage 1 is similar to that of a standard aircraft and will not be described in detail here.

[0050] As shown in Figures 3 and 4, Figure 3 shows the aircraft with doors and windows, while Figure 4 shows the interior of the cabin without doors and windows. The exterior of the fuselage 1 includes a windshield 11, a lower window 12, a door 13, a rear window 14, and a luggage door 15. The cabin within the fuselage 1 includes a display console 16, a pilot's seat 17, front passenger seats 18, and rear passenger seats 19. The rear of the cabin is the luggage compartment, which has a luggage door 15. The preferred cabin layout of the present invention is a 1+2+2 arrangement, with a centrally located pilot seat 17 in the first row, two front passenger seats 18 in the second row, and two rear passenger seats 19 in the third row. The width between the two seats in the second row is no less than 305 mm, meeting regulatory requirements. After opening the cabin door, passengers in the third row can reach their seats by passing through the space between the first and second rows. The cabin door can be configured as a left or right door, or as a cabin door on one side with an emergency exit on the other, also meeting regulatory requirements. The aisle formed by the first row of pilot seats 17 and the structures on both sides is more than 305 mm wide. After boarding the aircraft from the left / right cabin door, the pilot must pass through the aisles on both sides to reach the pilot seat 17.

[0051] The instrument panel is located in front of the pilot's seat 17. The front instrument panel is slightly offset to the right to ensure the pilot's eye position is centered over the PFD. Throttle and sidestick controls are located on either side of the seat. The pilot controls the throttle and sidestick to accelerate / decelerate, ascend / descend, pitch, roll, and turn the aircraft. The front instrument panel is equipped with two integrated display units (DUs), a backup instrument (ISIS), an audio control panel, a flight control panel, a power control panel, a data recorder, and a cockpit voice recorder.

[0052] Directly in front of the pilot is a continuous windshield, with two lower left and right windows located to the lower sides, both mounted on the main cockpit structure. This continuous windshield provides the necessary forward visibility for aircraft control, while the lower left and right windows provide downward visibility during vertical landings, facilitating alignment with the landing surface. A display console 16, located in front of the pilot's seat 17 and mounted on the main cockpit structure, provides the pilot with the control interface for aircraft operation.

[0053] Left door and right door windows are designed above the left and right doors respectively, and left and right porthole windows are designed between the left / right door windows and the rear sides. The left / right doors and left / right porthole windows are all installed on the main cabin structure, so that every passenger in the second and third rows can directly obtain a view outside the aircraft through the doors and windows or portholes, effectively ensuring the riding experience of every passenger.

[0054] Behind the third row of seats, an independent space is designed to be separated by the main cabin structure and can accommodate four 20-inch suitcases. Luggage can be put in or taken out through the 15-door luggage compartment fixed to the main cabin structure.

[0055] As shown in the disassembled views of Figures 5 and 6 , the wing 2 comprises an integrated wing box 21, ailerons 22, left / right wing nacelles 23, and left / right strut nacelles 24. The wing-body structure is separate, with left and right struts integrated into the front of the integrated wing box 21. The left / right wing nacelles 23 are connected to the integrated wing box 21 via a tilt axis 25, which rotates around the tilt axis. Figure 7 shows a schematic diagram of the left / right wing nacelles, which include winglets 231 and outer ailerons 232. The wiring harnesses of the left / right wing nacelles 23 are routed through the tilt axis to prevent interference with the structure during tilting (as shown in Figure 8 ). The left / right nacelles 23 of the wing are equipped with motors and controllers, and the rotors are installed on the motor output shafts; the wings adopt an upper gull-wing layout and are swept back, which increases the height of the rotor rotation plane on the struts and increases the distance from the cabin door, which can reduce the risk of injury to passengers when boarding and leaving the aircraft, and can increase the minimum ground clearance of the wing tilt nacelle, thereby ensuring the maximum roll angle.

[0056] The tailplane 3 consists of a tail box section 31 and left and right tailplane nacelles 32. Elevators 33 are mounted on the tail of the tail box section 31. The left and right tailplane nacelles 32 are connected to the tail box section 31 via a tilt axis, which rotates around the tilt axis. The wings and tailplane share a common tilt mechanism, reducing technical difficulty and design complexity.

[0057] Rotors are installed on the left / right strut nacelles 24, the left / right wing nacelles 23, and the left / right horizontal tail. The rotors on the left / right wing nacelles 23 and the left / right horizontal tail nacelles are rotors that can rotate around the tilt axis to achieve a tilt of -10° to 110°. The rotors on the left / right strut nacelles 24 are foldable fixed rotors that can be folded but are fixed in position and cannot be tilted. In addition, it is preferred to set the rotor to a five-blade rotor. The five-blade rotor has the function of variable collective pitch, taking into account both hovering efficiency and cruising efficiency. The fixed rotor is set to a four-blade foldable rotor. The four-blade rotor has a reliable folding function, ensuring that the blades follow the heading airflow when the fixed rotor is cruising, reducing aerodynamic drag and balancing the unbalanced loads generated by the forward and backward blades.

[0058] The five-bladed rotor is connected to a pitch-changing mechanism. The pitch-changing mechanism is a torque-transmitting structure used to amplify the displacement or force input from the control input, or to convert arc-shaped control input into a strictly linear motion output, thereby changing the rotor blade angle. The pitch-changing mechanism comprises a torque input point structure, a torque transmission structure, and a torque output point structure. The torque input point structure comprises an integrally formed joystick and an input end block. The input end block is fixed to one end of the joystick and connected to the servo arm / push rod of the servo. The movement of the servo arm / push rod drives the joystick to rotate. The other end of the joystick is connected to the torque transmission structure. The top of the torque transmission structure is connected to multiple torque output point structures. The rotation of the joystick drives the torque transmission structure to move, which in turn transfers the motion to the torque output point structures, driving the movement of the torque output point structures. The top end of each torque output point structure is directly connected to the corresponding rotor blade hub pitch hinge. The movement of the torque output point structure drives the blade hub pitch hinge to change the blade angle. The variable pitch control structure of the present invention is a torque transmission structure used to amplify the control input displacement, amplify the control input force, or convert the arc motion control input into a strict linear motion output.

[0059] The left / right wing nacelles and the left / right horizontal tail nacelles each consist of an integrally assembled nacelle rib and a servo mounting frame. The servo mounting frame secures the servo bracket, which is then mounted with a threaded pin, allowing the servo to rotate about the pin's axis. One end of the servo is secured to the tilt axis via a tilt servo arm. The servo drives the servo's retractable push rod, which partially tilts the nacelle. Figure 9 shows a schematic diagram of the tilt axis and servo arrangement. The servo specifically comprises a servo rod, a retractable push rod, and a tilt servo arm. The threaded pin on the servo bracket secures the servo rod. The retractable push rod is inserted into and retractably connected to the servo rod. A servo motor is located within the servo rod, driving the retractable push rod to extend and retract along the servo rod. The other end of the retractable push rod is secured to one end of the tilt servo arm via a fastener. The other end of the tilt servo arm is sleeved onto the tilt axis and secured to the tilt axis. The nacelle's tilting motion is illustrated in Figure 10.

[0060] When the aircraft is on the ground or in level flight, the nacelle is not tilting, and the servo's retractable push rod is extended. When the aircraft is in a vertical takeoff or landing position and needs to tilt, the servo directly drives the retractable push rod inward. The push rod's extension and retraction speed are uniform and controllable. Under the reaction force, the nacelle rotates at a constant speed around the tilt axis. The servo has a power-off lock function to lock the nacelle's position before and after tilting until the nacelle reaches the vertical position, completing the tilting process. During this entire rotation process, the box section and the tilt axis remain stationary, as does the tilt servo arm attached to the tilt axis. The servo's retractable push rod retracts and retracts, driving the nacelle's rotation. Simultaneously, the servo makes additional rotation around the threaded pin to prevent it from getting stuck.

[0061] As used herein, "rotor" may refer to any suitable rotary aerodynamic actuator, commonly referred to as a rotor, rotary wing, rotary airfoil, or the like. A rotor may refer to a rotary aerodynamic actuator utilizing an articulated or semi-rigid hub (e.g., wherein the connection between the blades and the hub may be articulated, flexible, rigid, and / or otherwise connected), a rotor may refer to a rotary aerodynamic actuator utilizing a rigid hub (e.g., wherein the connection between the blades and the hub may be articulated, flexible, rigid, and / or otherwise connected), and the use of a rotor may refer to any configuration and any other possible configuration of articulated or rigid blades, and / or any other possible configuration of blades connected to a central member or hub. The rotor utilizes a multi-blade, large-diameter, low-speed rotor design that can reduce power requirements in multi-rotor flight mode, reduce takeoff and landing noise levels, and reduce vibration of the multi-rotor and airframe structure during conversion.

[0062] The configuration modes of the aircraft include multi-rotor mode, transition mode and fixed-rotor mode; the aircraft uses the multi-rotor mode for takeoff and landing, and dynamically switches between the multi-rotor mode, transition mode and fixed-wing mode in the air. During vertical takeoff and landing, all tilt-rotors of the aircraft are directed vertically upward, and all rotors provide lift for the aircraft. In the multi-rotor mode, multiple rotors are directed upward, and the specific rotation angle of the rotors is calculated by the controller in the aircraft, which is specifically implemented in the control method of Example 2. In the fixed-wing mode, multiple rotors are directed forward along the direction of the fuselage and arranged in a forward state. The foldable fixed rotor is folded into a two-blade state and feathered to reduce resistance, thereby minimizing the additional resistance generated by it and improving endurance performance.

[0063] Furthermore, the aircraft rotors have different rotation modes, and multiple rotors are used to optimize hover mode performance and fixed-wing mode performance. When switching between the hover configuration and the cruise configuration, at least some of the rotors are shifted along an axis parallel to at least one of the longitudinal axis and the lateral axis of the aircraft, so that the position of each rotor is shifted.

[0064] Specifically, as shown in Figure 11, the first rotor 101 installed on the right strut rotates counterclockwise, the second rotor 102 installed on the right winglet rotates counterclockwise, the third rotor 103 installed on the right nacelle rotates clockwise, the fourth rotor 104 installed on the left nacelle rotates counterclockwise, the fifth rotor 75 installed on the left winglet rotates clockwise, and the sixth rotor 106 installed on the left strut rotates clockwise. The rotation direction is defined as the downward direction when the winglet and nacelle are in a vertical state.

[0065] The rotation directions of the first rotor 101 and the sixth rotor 106 ensure that the aircraft is subjected to aerodynamic moments in the clockwise and counterclockwise directions respectively. In this way, when the heading is controlled by changing the rotation speed of a certain rotor in the low-speed state, the contribution of the component force generated by the camber to the yaw moment of the entire aircraft is consistent with the contribution of the aerodynamic drag moment caused by the rotation direction, thereby increasing the heading control efficiency; the rotation direction of the third rotor 103 is opposite to that of the first rotor 101, and the rotation direction of the fourth rotor 104 is opposite to that of the sixth rotor 106. In this way, when the second rotor 102 or the fifth rotor 75 fails, the symmetrical rotor control balance can be closed to form a typical four-rotor state, which is conducive to control; the rotation direction of the second rotor 102 and the fifth rotor 75 is designed to increase the aerodynamic efficiency of the entire aircraft.

[0066] When used in urban air mobility operations, the aircraft initially assumes a multi-rotor configuration. The left and right wing tilt nacelles and the left and right horizontal tail tilt nacelles rotate upward to 90 degrees, and the foldable rotors on the nacelles deploy as four-bladed propellers. Once passengers are on board, six propulsion systems provide lift for a vertical takeoff to a safe altitude. Subsequently, the left and right wing tilt nacelles and the left and right horizontal tail tilt nacelles are slowly tilted forward to provide forward thrust. The aircraft's forward speed increases until a certain speed is reached, at which point the foldable rotors on the left and right nacelles fold, feather, and fully close. The left and right wing tilt nacelles and the left and right horizontal tail tilt nacelles rotate to 0 degrees, and the propeller pitch gradually increases. The rotors on the left and right wing tilt nacelles and the left and right horizontal tail tilt nacelles then provide thrust to maintain the aircraft's cruise phase. During the approach phase, the left and right foldable rotors activate to provide lift. The left and right wing tilt nacelles and the left and right horizontal tail tilt nacelles gradually rotate from 0° to 90°, gradually reducing the aircraft's forward speed until it reaches a hovering state. During this process, the propeller pitch gradually decreases. Six power systems then provide lift to vertically descend to the destination, completing the mission.

[0067] The landing gear 4 is a tricycle type, consisting of a nose landing gear 41 and two main landing gears 42. The main landing gear 42 has an integrated braking system, which acts as a brake during landing, while the nose landing gear 41 provides balance and directional control. The tricycle landing gear supports autonomous ground mobility and low-speed rolling capabilities, and can be used to support emergency rolling landings.

[0068] Example 2

[0069] A second embodiment of the present invention provides an internal power system for an aircraft, the power system including six power units and an energy system.

[0070] A single power unit consists of a motor (including motor controller), a speed reducer, a rotor, and associated lubrication and thermal management components. The motor, speed reducer, and rotor in each power unit are structurally integrated. Power units are typically constructed of metal and are suitably designed using finite element analysis methods such as topology optimization to optimize force transmission and minimize structural weight. The rotor base and speed reducer cover are one structural component, while the speed reducer bottom cover and motor cover are also one structural component. The reducer integrated structure is shown in Figures 12 and 13, including an engine mounting frame 111, a reinforced corner box 112, a reducer connecting frame 113, a reducer connecting leg 114 and a reducer mounting adjustment ring 115. The engine mounting frame 111, the reducer connecting frame 113 and the reducer mounting adjustment ring 115 are all annular sheets, one in number each, and are arranged parallel to each other; the engine mounting frame 111 is used to mount and fix the electric engine 119, and the reducer mounting adjustment ring 115 is used to mount and fix the reducer 1110; both ends of the multiple reinforced corner boxes 112 are detachably connected to the engine mounting frame 111 and the reducer connecting frame 111 by fasteners 86. 13; and both ends of the multiple reducer connecting legs 114 are also removably connected between the reducer connecting frame 113 and the reducer mounting adjustment ring 115 via fasteners 86. Multiple reinforced corner boxes 112 are evenly distributed around the centerline of the engine mounting frame 111 and the reducer connecting frame 113, and multiple reducer connecting legs 114 are also evenly distributed around the centerline of the reducer connecting frame 113 and the reducer mounting adjustment ring 115. This ensures that when the reducer 1110 transmits various loads and bending moments from the rotor to the fuselage 1 structure, the load is more evenly transferred, reducing the possibility of stress concentration and thus extending the service life of the equipment. Both the motor and reducer are designed with cooling channels in the housing and share a water cooling system. Each motor has two redundant power and control channels, with a total of twelve channels for the six motors. The nacelle that houses the power unit and tilt-steering gear is designed as a semi-monocoque structure, meeting strength requirements while leaving space for the layout and installation of equipment, wiring harnesses, and air ducts. Among them, in addition to push rod type servos, the tilting servos can also be set to worm gear type, harmonic gear type, cycloid pinwheel type, etc.

[0071] The energy system includes multiple battery modules, multiple battery management and energy distribution units, and related thermal management accessories. The battery modules consist of series-connected cells, cooling plates, and monitoring sensors. The module housing is constructed of high-temperature, flame-retardant materials, providing structural support and an exhaust duct for thermal runaway. Externally, the battery modules feature high-voltage connections, communication ports, cooling water ports, exhaust ducts, and mechanical mounting ports.

[0072] A single battery management and energy distribution unit includes power on and off functions, pre-charging function, charging function, pressure relief function, battery management function, DC-DC conversion function and energy distribution function, etc. The aircraft has four or six sets of battery modules, and one set of battery modules consists of six to eight battery modules. According to different installation positions, the battery modules have different installation arrangements and methods with the wing struts, the inner section of the wing or the outer section of the wing. At the same time, insulating and high-temperature resistant materials are used in the wing structure where the battery modules are installed. The present invention arranges the power battery packs on both sides of the wing and on the struts on both sides, making full use of the wing / strut space, optimizing the fuselage space layout, reducing the fuselage size, avoiding major battery failures affecting the cabin, and streamlining the cabin design / protection; the batteries are physically isolated and will not affect each other, reducing the flight load of the wing. In addition, the power battery is not set in the fuselage section, which can achieve the following effects: 1. The power battery is located higher than the occupants, so if thermal runaway occurs, it is not easy to burn or suffocate the occupants; 2. The space of the aircraft's wing struts, inner and outer sections is fully utilized to integrate the parts into a whole, and does not occupy valuable space such as the fuselage section (cabin, luggage compartment); 3. The power battery is very heavy, so the force transmission path can be shortened to increase the structural rigidity and reduce the structural weight.

[0073] The high-voltage interfaces of each module in a set of battery modules are connected in series and ultimately connected to the corresponding battery management and energy distribution unit; the communication interfaces of the battery modules are also connected to the corresponding battery management and energy distribution unit; the water cooling interface of the module is connected in parallel to the thermal management system of the battery module; the exhaust channel of the module is connected to the exhaust port on the wing structure.

[0074] The battery module provides an operating voltage platform (high-voltage platform) of approximately 500V-800V. Each set of battery modules is designed to be combined with a battery management and energy distribution unit. When four sets of battery modules are installed, each set of battery modules provides energy to three channels; when six sets of battery modules are installed, each set of battery modules provides energy to two channels. This design ensures that when a single set of battery modules fails, the motor can still obtain sufficient energy to ensure the safe flight and landing of the aircraft. A set of battery modules can be installed on the wing struts, the inner section of the wing, the outer section of the wing, etc., and the corresponding battery management and energy distribution units are also installed in adjacent positions.

[0075] A single BMEDU integrates the functions of multiple traditional devices, such as a battery management system, DC-DC converter, and energy distribution unit. This reduces the number of devices, connectors, and space requirements, while optimizing wiring harness routing and layout. The BMEDU's DC-DC conversion function converts the 500V-800V operating voltage to a medium-voltage platform (MV) and a 28V LV platform (LV) suitable for civil aviation. Together with the 28V LV battery, it provides redundant power to all low-voltage devices on the aircraft. The BMEDU's connectors are positioned according to the wiring harness's routing, significantly simplifying wiring harness layout.

[0076] Example 3

[0077] Different voltages are required within an aircraft to power various devices. Generally, both high-voltage and low-voltage architectures are required to meet these requirements. In Example 1, the present invention incorporates multiple batteries within the struts and wings of the aircraft structure. All batteries support both recharging and battery replacement. These multiple batteries combine to generate different voltages to power various devices.

[0078] A third embodiment of the present invention provides an aircraft high-voltage architecture, specifically including:

[0079] As shown in Figure 14, the aircraft has six electric motors (respectively, motor 1#, motor 2#, motor 3#, motor 4#, motor 5#, and motor 6#. Motor 1# and motor 6# are used to drive the rotation of the foldable rotors on the left and right struts, respectively. Motor 2# and motor 5# are used to drive the rotation of the rotors on the left / right nacelles of wing 2, respectively. Motor 3# and motor 4# are used to drive the rotation of the rotors on the left / right horizontal tail 3 nacelles of wing 2). The aircraft has four power batteries (respectively, power battery 1#, power battery 2#, power battery 3#, and power battery 4#). The four power batteries are evenly and symmetrically connected to the motor channels, making the aircraft energy and power architecture completely axially symmetrical or centrally symmetrical, completing the safety redundant power distribution function. In the figure, power battery 1#, power battery 2#, and motor 3# are on the same side, and motor 4#, motor 5#, and motor 6# are on the same side.

[0080] Each motor has two redundant channels. If one channel fails, the other channel can still ensure that the motor can operate at half its rated power. Each power battery can output three high-voltage distribution channels, which are connected to the inverters of different motors. For example, in Figure 14, the first high voltage of power battery 1# is connected to inverter 1B of motor 1#, the second high voltage of power battery 1# is connected to inverter 3B of motor 3#, and the third high voltage of motor 1# is connected to inverter 5B of motor 5#; the first high voltage of power battery 2# is connected to inverter 2B of motor 2#, the second high voltage of power battery 1# is connected to inverter 4B of motor 4#, and the third high voltage of motor 2# is connected to inverter 6B of motor 6#; the first high voltage of power battery 3# is connected to inverter 1A of motor 1#, the second high voltage of power battery 3# is connected to inverter 3A of motor 3#, and the third high voltage of motor 3# is connected to inverter 5A of motor 5#; the first high voltage of power battery 4# is connected to inverter 2A of motor 2#, the second high voltage of power battery 4# is connected to inverter 4A of motor 4#, and the third high voltage of motor 4# is connected to inverter 6A of motor 6#.

[0081] When a single inverter fails, the motor torque drops to 70% of its rated torque. Stable operation can be achieved by reducing the power of the opposite motor. For example, as shown in Figure 15, when the connection between power battery 4# and inverter 6A of motor 6# fails, the torque of motor 6# decreases. This reduces the power of opposite motor 3# and inverter 3A, thereby achieving balanced stability on both sides.

[0082] When a single power battery fails, the power of the three motors connected to the failed power battery is reduced. For example, as shown in FIG16 , the three motors are just in phase, so the torque is still balanced to ensure regional safety.

[0083] When two adjacent power batteries fail, all motors are in a single inverter state. For example, as shown in Figure 17, both power battery 3# and power battery 4# fail, and power battery 3# and power battery 4# are connected to one of the inverters of all motors (1# to 6#). At this time, all motors are in a single inverter state. Although this situation is an extremely low-probability risk, all motors are in a single inverter state at this time, and balance is still maintained. The single-channel power can still be increased. However, in order to reduce the risk, a controlled forced landing is required to reduce losses.

[0084] When in fixed rotor mode, motor 1# and motor 6# are shut down, as shown in Figure 18. At this time, power battery 1# and power battery 3# do not supply power to motor 1#, and power battery 2# and power battery 4# do not supply power to motor 2#. However, the discharge conditions of the four power batteries are still consistent. Even if motor 3# and motor 4# are operated in feathering mode, the discharge conditions of the power batteries are still consistent.

[0085] When a single motor fails, stable operation can be achieved by reducing the power of the opposite motor. For example, as shown in Figure 19, when left motor 6# fails, the power supply from motor 6# to power battery 2# and power battery 4# is disconnected. In order to achieve stable left and right control of the aircraft, the power of the right motor can be reduced. Since one of the power supply lines from power battery 2# and power battery 4# is disconnected, the power of the motor connected to power battery 1# or power battery 3# can be reduced. Preferably, since motor 3# is connected to power battery 1# and power battery 3#, the power of inverter 3A and inverter 3B on the right motor 3# is reduced.

[0086] When two symmetrical power batteries fail, for example, as shown in Figure 20, power battery 2# and power battery 4# on the opposite side fail, motor 2#, motor 4# and motor 6# will all fail to work. At this time, motor 1#, motor 3# and motor 5# can still work, and the aircraft can maintain normal flight. Although this situation is an extremely low-probability event, in order to ensure the safety of the aircraft's flight, a controlled forced landing is still required to reduce losses.

[0087] As shown in FIG21 , in order to prevent an emergency situation where multiple power batteries become ineffective, the present invention can further add a fifth power battery 5# to connect to the power distribution circuit of all power batteries (1# to 4#). In an emergency, power battery 5# replenishes electric energy for all power batteries, enabling each power battery to replenish electric energy to the corresponding motor.

[0088] In the aircraft high-voltage architecture shown in another embodiment, as shown in Figure 22, the aircraft has four power batteries (respectively power battery 1#, power battery 2#, power battery 3#, power battery 4#), two distribution boxes (respectively distribution box 1# and distribution box 2#) and six motors (respectively motor 1#, motor 2#, motor 3#, motor 4#, motor 5#, motor 6#). The four power batteries are respectively connected to the two distribution boxes. In the figure, power battery 1# and power battery 2# are connected to distribution box 1#, and power battery 3# and power battery 4# are connected to distribution box 2#. Each motor still has two redundant channels, but each is connected to two distribution boxes. In the figure, redundant channel 1# of motor 1# is connected to distribution box 2#, redundant channel 2# of motor 1# is connected to distribution box 1#, redundant channel 1# of motor 2# is connected to distribution box 2#, redundant channel 2# of motor 2# is connected to distribution box 1#, redundant channel 1# of motor 3# is connected to distribution box 2#, redundant channel 2# of motor 3# is connected to distribution box 1#, redundant channel 1# of motor 4# is connected to distribution box 2#, redundant channel 2# of motor 4# is connected to distribution box 1#, redundant channel 1# of motor 5# is connected to distribution box 2#, redundant channel 2# of motor 5# is connected to distribution box 1#, redundant channel 1# of motor 6# is connected to distribution box 2#, and redundant channel 2# of motor 6# is connected to distribution box 1#. If one channel fails, the other channel can still ensure that the motor operates at half its rated power.

[0089] In addition, six power batteries and six electric motors can be used, with each electric motor connected to the inverters of two power batteries, which can also achieve the above technical effects.

[0090] Example 4

[0091] A fourth embodiment of the present invention provides a low-voltage architecture for an aircraft. The low-voltage architecture for the aircraft is divided into a low-voltage control architecture ( FIG. 23 ) and a low-voltage power distribution architecture ( FIG. 24 ).

[0092] As shown in Figure 23, in the aircraft's low-voltage control architecture, the aircraft's internal management computer is connected to the low-voltage power distribution switch box via communication line CAN1 and to the DC-DC converter via communication line CAN2. The DC-DC converter provides two 28V power lines, A and B, to the low-voltage power distribution switch box. Together with the 28V low-voltage battery, the low-voltage power distribution switch box forms a redundant low-voltage power supply, powering low-voltage equipment. The DC-DC converter has two redundant inputs, capable of simultaneously drawing high-voltage power from power battery 1 and power battery 2. Power battery 1 and power battery 2 are connected to the flight management computer via communication line CAN3, and are controlled and monitored by the flight management computer.

[0093] Figure 24 shows the aircraft's low-voltage power distribution architecture, primarily the 28V low-voltage architecture within the power distribution switch box. Low-voltage power distribution architectures generally employ two types of multi-input safety redundancy designs. In the first form, the DC-DC converter is connected in parallel with the 28V low-voltage battery, forming a relatively safe 28V bus. A diode is added to the 28V low-voltage battery branch to prevent backflow from the bus to the battery. In the second form, two redundant DC-DC converter channels, ChA and ChB, and a 28V low-voltage battery are connected to the power distribution switch box. The 28V battery bus is the most reliable single power source, providing power to the DC-DC converter buses, ChA and ChB, and the 28V emergency bus via three diodes. Therefore, even if the DC-DC converter buses ChA and ChB fail, they can obtain power from the 28V battery bus to maintain a powered state; and the 28V emergency bus can operate stably and with the highest safety as long as one of the DC-DC buses ChA, ChB and the 28V emergency bus is working.

[0094] Example 5

[0095] As shown in FIG25 , the second embodiment of the present invention provides an internal avionics flight control architecture for an aircraft, including a flight control system, an indication and recording system, a navigation sensor, and a communication system, while also providing avionics optional functions for different customer groups.

[0096] The flight control system includes the flight control computer, cockpit control system, servos, and motor controllers. There are four flight control computers, including three independent, redundant ones and one backup. Developed according to DAL A, they feature heterogeneous monitoring channels to ensure data integrity and meet common-mode failure analysis requirements. The cockpit control system includes a sidestick and a throttle lever, providing the pilot with a human-machine interface for flight control and manual aircraft control. The sidestick provides three-axis control: longitudinal, lateral, and axial rotation; the throttle lever provides longitudinal control. The sidestick and throttle lever control the aircraft's flight path, altitude, speed, heading, and ground taxiing. The control variables can be attitude, speed, or altitude, determined by the control law. The sidestick and throttle lever also support functions such as air-to-ground communication control, automatic flight disconnect, and display cursor control. They feature automatic return to center when the stick is released and manual feel. Multiple redundant sensors are used to convert pilot control commands into electrical signals, which are then transmitted to the flight control computer. The servos include four aileron servos, four elevator servos, two rudder servos, four tilt servos, and four variable-pitch servos. These receive control commands from the flight control computer and operate accordingly. The four ailerons, four elevators, and two rudders are each independently deflected by a separate servo. This facilitates redundant reconstruction in the event of a single control surface failure, ensuring safe flight and reducing control surface servo performance requirements. Furthermore, the inboard ailerons and elevators can be deflected downward at large angles, reducing power loss in hover. Six motor controllers receive control commands from the flight control computer and drive the motors. The cockpit joysticks include a sidestick and a throttle.

[0097] The indication and recording system includes two integrated display units (DUs), a backup instrument (ISIS), a data recorder (FDR), and a cockpit voice recorder (CVR). Designed specifically for single-pilot operation, each DU can be conveniently switched at the touch of a button, providing comprehensive information. The ISIS provides critical flight information such as airspeed, altitude, and attitude, serving as a backup instrument in the event of primary display failure. The FDR and CVR record flight data and cabin voice, respectively.

[0098] Navigation sensors include three Air Data Systems (ADSs), three Attitude Reference Systems (AHRSs), three magnetometers, two GPS receivers, and two Radio Altimeters (RAs). These sensors provide navigation information, including airspeed, air pressure, flight attitude, heading, position, and altitude, to the flight control system and the indication and recording system.

[0099] The communication system includes two VHF radios, an audio controller, a Mode S transponder, and an ADS-B radio, along with an intercom port and headset. It enables intercom communication between the pilot and ground air traffic control, as well as between the pilot and passengers.

[0100] To meet the needs of mid-range and high-end customers, the aircraft provided by this invention also features optional avionics features, including a flight management system, a detection and collision avoidance system, automatic flight, and centralized maintenance. Data links, optical detection, and radar equipment can also be added for qualified customers. Regarding bus communications, the avionics flight control system takes into account low cost, reliability, and mature civil aviation technology. The flight control computer uses RS485 and CAN buses as field control buses to control the servos and motors. The ARINC429 bus is used between the flight control system and the avionics system to receive navigation sensor data and interact with the display system.

[0101] The integrated display exchanges data with external sensors via an RS422 interface and interacts with the flight control computer via the ARINC429 bus. The audio control board is connected to the integrated display via a high-speed bus, and the pilot configures it through the integrated display. The audio control board is also connected to two VHF units and an intercom interface unit for audio transmission. The flight control panel is connected to the flight control computer via discrete signals to switch flight modes. The power control panel is used to control the E20 aircraft's high-voltage and low-voltage power on and off, and also transmits motor enable signals. A data recorder is used to record flight data, and a cockpit voice recorder is used to record cabin audio information.

[0102] The integrated display system is specially designed for single-pilot aircraft. The two integrated displays can switch between PFD, ND, FMS, SYS and other screens through shortcut keys. The default screen is the PFD on the left and the power page in the SYS screen on the right. When one display fails, the other display can still provide the pilot with all the information to complete the flight mission. PFD display information includes main flight information such as pressure altitude, indicated airspeed, ground speed, attitude, climb rate, heading, radio altitude, VHF frequency, flight mode, etc.; supports horizontal status indication and course deviation indication; provides CAS warning information display. FMS display information includes flight plan editing and management, flight guidance management (including horizontal navigation and vertical navigation), performance-based calculation and management (including optimal battery power estimation, position prediction, range estimation, etc.), vertical information, etc. ND display information includes digital map display, track display, position information display, heading display, weather information, traffic information, etc. YS display information includes flight control system status information, such as aileron 22 deflection, aileron 22 status, rudder deflection, rudder status, elevator deflection, elevator status, and motor status: motor torque, motor speed, rotor pitch angle, motor temperature, gearbox, nacelle tilt angle, battery status: remaining power, voltage, current, temperature, power system status: including high-voltage power supply and low-voltage power supply, door status, brake status, etc.

[0103] Example 6

[0104] The human-machine interaction of the flight control system should support the concept of consistent control, based on the control of fixed-rotor aircraft, integrating the requirements of rotor control, simplifying hovering and vertical take-off and landing. Through envelope protection functions, status monitoring and warnings, and advanced control laws, the pilot can achieve worry-free control within the full flight envelope and reduce the pilot's burden.

[0105] The flight control system communicates via a bus with the Air Data System (ADS), Attitude and Heading Reference System (AHRS), three magnetometers, GPS, and Radio Altimeter (RA). It collects flight data and position information, executes flight control algorithms, and performs aircraft control functions, including vertical takeoff and landing control (in speed and attitude modes), tilt transition control, and fixed-rotor cruise flight control. It also includes track control, heading control, altitude control, and ground taxiing. It also implements envelope protection functions such as stall protection, overspeed protection, and bank angle protection, and monitors the status of the entire flight control system. It also works in conjunction with the Flight Management System, Detection and Collision Avoidance System, and Autopilot to achieve performance optimization, automatic navigation, automatic flight, and automatic obstacle avoidance, reducing the pilot's workload and improving flight safety.

[0106] Specifically, as shown in FIG26 , the flight control system includes a cockpit control room and a flight control law.

[0107] In the cockpit, the pilot inputs control instructions to perform automated flight missions, and the control console in the cockpit sends flight control commands to the flight control law. These commands include throttle, yaw, pitch, and roll, which are used to control flight modes and phases.

[0108] The flight control law includes the controller, control distribution system and battery balancing control system.

[0109] The controller includes a parameter management module, a flight management module, a control instruction preprocessing module, a position control module, a speed control module, an attitude control module and an angular velocity control module;

[0110] The controller inputs forces and torques into the control distribution system, including the force Fbx desired to be applied on the X-axis of the body, the force Fbz desired to be applied on the X-axis of the body, the torque L desired to be applied on the X-axis of the body, the torque M desired to be applied on the Y-axis of the body, and the torque N desired to be applied on the Z-axis of the body.

[0111] The following requirements are imposed on the control performance of the aircraft during flight: 1. The vertical touchdown speed shall not exceed 1.8 m / s; the vertical touchdown pitch attitude shall be less than 3°, and the roll attitude shall be less than 3° (positive crosswind is less than 5 m / s); the aircraft shall be able to safely perform roll maneuvers (roll angle between -30° and 30°) during climb, level flight and descent; when the flight state changes, the state shall be able to transition smoothly (normal overload ≤ 2.5 g); for the control mode conversion process, transient control or compensation amounts that meet the requirements shall be set.

[0112] As shown in Error! Reference source not found. 27, the flight control system can control the following actuations: motor torque / speed / pitch control, which indirectly controls the thrust of the six motors by controlling the torque / speed / pitch of the six motors; tilt nacelle angle control, in which the nacelles at motor positions 2, 3, 4, and 5 of the aircraft have a tilt function, and tilt control of the nacelles can change the thrust direction of positions 2, 3, 4, and 5; and control surface control, in which the aircraft's control surfaces are used to control the pitch, yaw, and roll torques when the rotor configuration is fixed. The control quantities include pitch control, roll control, and yaw control.

[0113] The rotor modes of the aircraft include low-speed multi-rotor control mode, transition mode flight mode and fixed rotor stage control mode.

[0114] As shown in Figures 28 and 29, during the low-speed multirotor control mode (ground speed 0-5 m / s), the aircraft's roll, pitch, yaw, and vertical movements are controlled by controlling the thrust of the six motors. The low-speed multirotor control mode includes roll, pitch, yaw, and vertical control. In low-speed multirotor mode, the aircraft's forward lateral speed is achieved by varying the pitch and roll angles. Motors 1 and 6 are tilted inward or forward to enhance yaw control.

[0115] Under roll control, the desired lateral velocity of the aircraft is generated through the velocity control loop to form the desired roll attitude angle. This desired roll attitude angle is then generated through the attitude control loop to form the desired roll angular velocity. This desired roll angular velocity is then generated through the angular velocity control loop to form the desired roll moment. Roll moment control is achieved by controlling the force difference between motors 1, 2, and 3 and motors 4, 5, and 6. Under pitch control, the desired forward velocity of the aircraft is generated through the velocity control loop to form the desired pitch attitude angle. This desired pitch attitude angle is then generated through the attitude control loop to form the desired pitch angular velocity. This desired pitch angular velocity is then generated through the angular velocity loop to form the desired pitch moment. Pitch moment control is achieved by controlling the force difference between motors 1 and 6 and motors 3 and 4. In yaw control, the aircraft's desired yaw angle is generated through the angle control loop to generate the desired yaw velocity. This desired yaw velocity is then used to generate the desired yaw moment through the velocity control loop. Yaw moment control is achieved by controlling the differential force between motors 1, 3, and 5 and motors 2, 4, and 5, as well as the tilt angle differential between motors 2, 3, 4, and 5. In vertical control, the aircraft's desired vertical velocity is generated through the velocity control loop to generate the desired vertical acceleration. Vertical acceleration control is achieved by simultaneously controlling the force of motors 1, 2, 3, 4, 5, and 6.

[0116] As shown in Figure 30, in the transition mode flight control mode, the forward speed of the aircraft is achieved by changing the inclination angles of nacelles 2, 3, 4, and 5 to generate forward acceleration, and the lateral movement is achieved by controlling the roll angle.

[0117] During transition flight control, the desired forward velocity, desired roll angle, desired yaw angle, and desired altitude are calculated through the control loop to yield the desired forward force (Fxb), desired vertical force (Fzb), desired rolling moment (L), desired pitching moment (M), and desired yaw moment (N) in the aircraft coordinate system. During transition flight, Wing 2 generates aerodynamic forces as airspeed increases, and the control surfaces generate aerodynamic torques. At this point, the torque control weights for the rotor and fixed rotor are distributed based on the "fusion speed," ensuring a smooth transfer of torque control authority during the transition from the rotor to the fixed rotor state. The "fusion speed" is the fusion speed of airspeed and ground speed.

[0118] As shown in FIG31 , during the fixed rotor flight phase, the roll moment is controlled by deflecting the aileron 22 control surface, the pitch moment is controlled by deflecting the elevator control surface, and the pitch moment is controlled by deflecting the yaw control surface.

[0119] Figure 32 shows the functional module architecture of the control distribution system. Airframe 1 navigation equipment collects real-time aircraft status information and feeds it back to the corresponding control module, forming a closed control loop. The flight control system's complete process, from pilot input to flight control output, includes the following main functional modules: flight management module, control module, and actuator distribution module. The control module is divided into the following loops (ordered from inner to outer): angular velocity control loop, attitude angle control loop, acceleration control (tilt control), velocity control loop, and position control loop.

[0120] Specifically, as shown in Figure 33, in the flight management module, the control methods are divided into autonomous flight and manual control. Autonomous flight is generally used for control in the fixed rotor configuration and tilt configuration stages, while manual control is mostly used for control in the rotor configuration stage. Among them, during autonomous flight, the flight control calculates the error between the current position of the aircraft and the preset route position in real time based on the pre-set route, generates guidance instructions, and makes the aircraft fly along the preset route. The preset route feature waypoints in the flight control are shown in Table 1 below:

[0121] Table 1 Description of near-field waypoints

[0122] Manual control, or pilot control, is based on the Unified Flight Control principle. Aircraft control input devices are divided into the left-side throttle stick and the right-side stick. The corresponding control variable mappings are shown in Table 2 below (actual application is not limited to this mapping; the main purpose is to reduce pilot control difficulty):

[0123] Table 2 Manipulation quantity mapping relationship

[0124] According to the different mapping of the driver's control quantity, it can be divided into inner loop control and outer loop control, as shown in Table 3:

[0125] Table 3 Rotor state control quantity mapping table

[0126] Among them, in the control module: the input and output of the position control loop are as follows:

[0127] Input: desired longitude and latitude, desired altitude, aircraft's current longitude and latitude, current altitude;

[0128] Output: expected speed;

[0129] The position control loop calculates the current desired speed of the aircraft based on the north error, east error, and vertical error between the current and desired positions of the aircraft, as follows:

[0130] Among them, Vnexp 、Ve exp and Vd exp are the expected north velocity, east velocity and vertical velocity respectively; K pe , K pn and K pd They are position control gain, Lat dxp and Lat is the desired latitude and the current latitude of the aircraft, Lon exp and Lon is the desired longitude and the current longitude of the aircraft, H exp and H are the desired altitude and the current altitude of the aircraft.

[0131] The input and output of the speed control loop are as follows:

[0132] Inputs: desired forward speed, desired lateral speed, desired vertical speed; forward speed, lateral speed and vertical speed;

[0133] Output: expected forward acceleration, expected lateral acceleration and expected vertical acceleration;

[0134] The speed control loop uses the PID algorithm to calculate the current desired forward acceleration, lateral acceleration, and vertical acceleration of the aircraft based on the current forward speed error, lateral speed error, and vertical speed error of the aircraft.

[0135] In the low-speed multi-rotor stage, when the aircraft's tilt angle is 90°, the aircraft's height, front-back, and left-right directions are controlled by controlling the magnitude and direction of the desired acceleration and combining the roll angle, pitch angle, and yaw angle controls.

[0136] The input and output quantities of acceleration control are as follows:

[0137] Inputs: desired forward acceleration, desired lateral acceleration, desired vertical acceleration, and desired yaw angle;

[0138] Outputs: tilt angle, desired pitch angle, desired roll angle, and desired yaw angle;

[0139] In the multirotor phase, acceleration control calculates the desired total thrust vector based on the desired forward, lateral, and vertical accelerations as follows:

[0140] Among them, ψ is the current heading angle of the aircraft, a frt 、a lat and a dn are the forward, lateral and vertical accelerations respectively, a n 、a e and a d are the north, east and earth accelerations respectively, and g is the gravitational constant;

[0141] After obtaining the total thrust vector, combined with the current yaw angle, the desired pitch and roll angles under this vector are calculated as follows:

[0142] Among them, θ exp and φ exp are the desired pitch angle and the desired roll angle, and A n 、a e and a d The normalized variable, ψ, is the aircraft's current heading angle. The roll angle represents the proportion of the aircraft's forward acceleration vector and vertical acceleration vector in the direction of the total thrust vector. A smaller roll angle results in a greater forward acceleration.

[0143] The input and output of the attitude angular velocity control loop are as follows:

[0144] Input: desired roll angle, desired pitch angle, desired yaw angle, aircraft's current roll angle, pitch angle, and yaw angle;

[0145] Output: desired roll angular velocity, desired pitch angular velocity, and desired yaw angular velocity;

[0146] As shown in Figure 34, attitude angle control mainly realizes the calculation of the desired angular velocity, thereby providing the desired input quantity for the attitude angular velocity control loop, which is the superposition of the desired quantity and the compensation quantity.

[0147] As shown in Figure 35, the expected amount calculates the attitude angular rate error based on the error between the expected attitude angle and the current attitude angle. The expected attitude and the current attitude are expressed in quaternions in the north-east coordinate system. The attitude error, tilt error, and torsion error are also expressed in quaternions in the body coordinate system. After the difference between the expected attitude and the current attitude is taken, the attitude error in the body coordinate system is obtained. In the multi-rotor stage, since the tilt angle control (x-axis and y-axis direction of the body) responds faster than the torsion angle control (z-axis direction of the body), under limited control capabilities, the tilt control is satisfied first, and then the torsion control (i.e., tilt separation) is satisfied. Therefore, when calculating the expected angular velocity, the attitude error is split into the following two parts:

[0148] Tilt error: In the body coordinate system, the difference between the desired posture z-axis and the current posture z-axis is the tilt error (or the tilt error angle and its unit vector. Because it does not include torsion, the third component of the unit vector is 0).

[0149] Torsion error: In the body coordinate system, the torsion error is obtained by taking the difference between the tilt error and the attitude error vector (or the torsion error angle and its unit vector. Since the tilt is not included, the first and second components of the unit vector are 0).

[0150] Finally, the roll and pitch angular rates are extracted from the tilt error and limited according to the attitude constraints to obtain the desired roll and pitch angular rates. The yaw angular rate is extracted from the torsion error and diluted according to the error size to output the desired yaw angular rate.

[0151] Compensation: When the aircraft has an attitude angle, compensation correction is required when calculating the expected angular velocity. The compensation correction amount varies according to the attitude angle. When the pitch and roll attitude angles are 0°, the roll angular velocity and pitch angular velocity compensation amounts are 0. The compensation amount is calculated as follows:

[0152] Among them, p Δ ,q Δ and r Δ are the pitch, roll and yaw angular velocity compensation respectively.

[0153] As shown in Figure 36, the attitude angular velocity control loop is divided into three control channels: roll angular velocity control, pitch angular velocity control, and yaw angular velocity control. Its input and output quantities are as follows:

[0154] Inputs: desired roll rate, desired pitch rate, and desired yaw rate; aircraft's current roll rate, pitch rate, and yaw rate; aircraft's roll moment of inertia, pitch moment of inertia, and yaw moment of inertia;

[0155] Output: roll control, pitch control and yaw control;

[0156] The angular velocity control loop uses the PID algorithm to calculate the current aircraft's roll, pitch, and yaw channel control variables (or control torques) based on the error between the aircraft's desired angular velocity and its current angular velocity.

[0157] The architecture of the actuator allocation module is shown in Figure 37. In the actuator allocation link, in order to improve the response speed of the actuator, advance compensation is set for the actuator control channel with slower response. The compensation amount and the allocation amount are superimposed and output to each actuator channel. The rudder mapping module superimposes the total amount of the rudder front feed amount and the rudder allocation amount, and maps it to the control amount within the range of the rudder servo stroke according to a certain mapping relationship, thereby controlling the stroke of the rudder servo. The motor mapping module is similar. The superimposed total amount is mapped to the torque control amount between 0 and 400 N.m, and the motor is controlled after output. The corresponding mapping relationship is as follows:

[0158] T is the total expected pulling force after the feedforward amount and the distribution amount are superimposed, M cmd is the motor torque control quantity.

[0159] The lead compensation amount is set according to the motor characteristics and the compensation ratio of the corresponding control input amount.

[0160] Control distribution distributes the desired force generated by the acceleration control loop and the desired torque generated by the angular velocity loop to the six motors. To meet the requirements of independent tilt and twist control, the control distribution is divided into force distribution, roll and pitch distribution, and yaw distribution. The distribution relationship is as follows:

[0161] in, and They are the force distribution, pitch and roll distribution, and yaw distribution, respectively. Fz, N, L, and M are the force control, yaw control, roll control, and pitch control, respectively.

[0162] The total control amount of the motor is the superposition of the above distribution results. To prevent a single motor from reaching the maximum power state, a correction factor is added during the superposition process. The correction priority from high to low is: pitch and roll distribution > force distribution > yaw distribution. The correction expression is as follows:

[0163] in, Total motor control quantity, K f , K n They are the force distribution ratio and the yaw control ratio, and the calculation relationship is as follows:

[0164] in, and The above and

[0165] Furthermore, the flight control law also includes a PID module, and the control law PID module is as follows:

[0166] Where Δ=In exp -In; In exp is the control input, In is the feedback, is a first-order low-pass filter of Δ (with a time constant of τ), K p , K I and K d They are PID control coefficients, Out, Out max and Out min are the output value, the maximum output value, and the minimum output value respectively.

[0167] The battery balancing control system provides an aircraft high-voltage architecture and an aircraft low-voltage architecture. The aircraft high-voltage architecture and the aircraft low-voltage architecture are respectively described in detail in the above-mentioned third embodiment and fourth embodiment.

[0168] Redundant navigation sensors are installed inside the aircraft, including three sets of ADS + two sets of AHRS + one set of IRS + two sets of GPS + two sets of RA. The data of redundant navigation sensors are integrated to obtain accurate navigation data and improve the operation accuracy of the aircraft.

[0169] evtol integrated display: A unique data display transmission path collects data from various aircraft systems (motor, power supply, BMS, mechanical, and environmental control) through the flight control computer (FCC), packages it into A429 data, and sends it to the DU. Dedicated display content includes integrated displays for the motor, battery, BMS, and evtol flight control system.

[0170] The specific implementation methods described above further illustrate the objectives, technical solutions and beneficial effects of the present invention in detail. It should be understood that the above description is only a specific implementation method of the present invention and is not intended to limit the scope of protection of the present invention. Any modifications, equivalent replacements, improvements, etc. made on the basis of the technical solution of the present invention should be included in the scope of protection of the present invention.

Claims

1. An electric tilt-rotor aircraft, characterized in that: include: Fuselage, wings, horizontal tail, landing gear; The fuselage includes a fuselage body and a rudder. The fuselage body integrates a cabin door, a windshield, a lower window and a rear window. The rudder is installed at the rear of the fuselage body and is used to control and adjust the flight direction of the aircraft. The fuselage body is provided with a cockpit including a display console, a pilot seat, a front passenger seat, and a rear passenger seat. The rear of the cockpit is a luggage compartment. The wing-body structure is a separate structure, including an integral wing box section, ailerons, left / right wing nacelles and left / right strut nacelles. The left / right struts are integrated on the front side of the integral wing box section. The left / right wing nacelles are connected to the integral wing box section through a tilt axis, and the left / right wing nacelles rotate around the tilt axis. A motor and a controller are arranged inside the left / right wing nacelles, and the rotor is installed on the motor output shaft. The wing adopts an upper single gull wing layout to increase the ground clearance of the wing and rotor. The horizontal tail includes a horizontal tail box section and a horizontal tail left / right nacelle. The horizontal tail left / right nacelle is connected to the horizontal tail box section through a tilt axis, and the left / right horizontal tail rotates around the tilt axis. Rotors are installed on the left / right nacelles of the strut, the left / right nacelles of the wing, and the left / right nacelles of the horizontal tail. The rotors on the left / right nacelles of the wing and the left / right horizontal tail nacelles are rotors that rotate around the tilt axis; the rotors on the left / right nacelles of the strut are foldable fixed rotors that can be folded but are fixed in position and cannot be tilted; the left / right nacelles of the wing and the left / right horizontal tail nacelles each include a nacelle rib and a steering gear mounting frame that are assembled into one body, the steering gear mounting frame fixes the steering gear bracket, the steering gear bracket is installed with a threaded pin, and the steering gear rotates around the threaded pin axis; one end of the steering gear is fixed to the tilt axis through a tilting steering gear arm; the steering gear drives the nacelle part to tilt by driving the retractable push rod of the steering gear; The landing gear adopts a three-point landing gear, including a front landing gear and two main landing gears. The main landing gear is integrated with a braking system, which acts as a brake when the aircraft lands, and the front landing gear plays a role in balance and direction control.

2. An electric tilt-rotor aircraft according to claim 1, characterized in that: The cockpit layout adopts 1+2+2, that is, the first row is a pilot seat with a centrally deployed position, the second row is two front passenger seats, and the third row is two rear passenger seats; after opening the left / right cabin doors, passengers in the third row can reach their seats through the space between the first and second row seats; after the pilot boards the plane from the left / right cabin doors, he reaches the pilot seat through the aisles on both sides.

3. An electric tilt-rotor aircraft according to claim 2, characterized in that: In front of the pilot's seat is the instrument panel. The front instrument panel is slightly tilted to the right to ensure that the pilot's eyes are in the middle of the PFD. The throttle panel and side stick are deployed on the left and right sides of the seat respectively. The pilot controls the throttle panel and side stick to achieve acceleration / deceleration, ascent / descent, pitch, roll, and turn of the aircraft.

4. The electric tilt-rotor aircraft according to claim 1, characterized in that: The rotor is set to be a five-blade rotor, and the five-blade rotor is connected to a pitch-changing mechanism. The pitch-changing mechanism is a torque transmission structure used to amplify the control input displacement, amplify the control input force, or convert the arc motion control input into a strict linear motion output, and is used to change the blade angle of the rotor. The pitch-changing mechanism includes a torque input point structure, a torque transmission structure and a torque output point structure.

5. An electric tilt-rotor aircraft according to claim 4, characterized in that: The torque input point structure includes an integrally formed joystick and an input end block. The input end block is fixed at one end of the joystick, and the input end block is connected to the servo arm / push rod of the servo, and the joystick is driven to rotate through the movement of the servo arm / push rod; the other end of the joystick is connected to the torque transmission structure, and the top of the torque transmission structure is connected to multiple torque output point structures. The rotation of the joystick drives the torque transmission structure to move, and the torque transmission structure transfers the movement to the torque output point structure, driving the torque output point structure to move; the upper end of each torque output point structure is directly connected to the blade hub pitch hinge corresponding to the rotor, and the movement of the torque output point structure drives the blade hub pitch hinge to move, thereby realizing the change of the blade angle.

6. The electric tilt-rotor aircraft according to claim 1, characterized in that: The fixed rotor is set as a four-blade foldable rotor.

7. The electric tilt-rotor aircraft according to claim 1, characterized in that: The servo specifically comprises a servo rod, a retractable push rod and a tilt servo arm. The threaded pin on the servo bracket fixes the servo rod. The retractable push rod is inserted into the servo rod and is telescopically connected to the servo rod. A servo motor is arranged in the servo rod. The servo motor drives the retractable push rod to retract and retract along the servo rod inward and outward. The other end of the retractable push rod is fixed to one end of the tilt servo arm through a fastener. The other end of the tilt servo arm is sleeved on the tilt shaft and fixed to the tilt shaft.

8. The electric tilt-rotor aircraft according to claim 1, characterized in that: The configuration modes of the aircraft include multi-rotor mode, transition mode and fixed-rotor mode; the aircraft uses the multi-rotor mode for takeoff and landing, and dynamically switches between the multi-rotor mode, transition mode and fixed-wing mode in the air.

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