CN117284473A - Electric tilting rotor craft and control system thereof - Google Patents

Abstract

The invention discloses an electric tilt rotor aircraft and a control system thereof. The aircraft comprises: the fuselage comprises a fuselage body and a rudder, and controls and adjusts the flight direction of the aircraft; the wing body structure of the aircraft wing is in a separation mode and comprises an integral wing box section, ailerons, wing left/right nacelle and a stay bar left/right nacelle, wherein the left/right stay bar is integrated on the front side of the integral wing box section, the wing left/right nacelle is connected with the integral wing box section through a tilting shaft, and the wing left/right nacelle rotates around the tilting shaft; the horizontal tail comprises a horizontal tail box section and a horizontal tail left/right nacelle, the horizontal tail left/right nacelle is connected with the horizontal tail box section through a tilting shaft, and the left/right horizontal tail rotates around the tilting shaft; the landing gear adopts a three-point landing gear, a brake system is integrated in the main landing gear, and the front landing gear plays roles of balancing and direction control. The electric tilting rotor craft provided by the invention can realize urban air traffic.

Description

An electric tilt-rotor aircraft and its control system

Technical Field

The present invention relates to the technical field of aircraft, and in particular to an electric tilt-rotor aircraft and a system thereof.

Background Art

With the increasing density of vehicles on urban roads, people's daily commuting time is gradually increasing. With the favorable condition that there are medium and low airspaces in cities that have not been fully developed and utilized, the concept of urban air mobility (UAM) has been proposed. Electric vertical take-off and landing aircraft have become the most important urban air mobility solution with the important advantages of green environmental protection and low dependence on infrastructure. At present, there are two common layout forms in the industry. One layout is the multi-rotor layout form: this type of aircraft relies on multiple sets of lift rotors to provide vertical take-off and landing and forward flight power for the aircraft. Because it does not have the same complex mechanical structure as the helicopter rotor, the flight speed is limited by the rotor load and is low. At the same time, the endurance is poor due to the low flight efficiency; the other common layout is the "lift + push" combination layout form: this type of aircraft combines the traditional fixed rotor layout with multiple sets of lift rotors to achieve the function of vertical take-off and landing. When cruising, the lift rotor is closed and only relies on propulsion power to fly forward. The waste resistance generated by the lift rotor during the cruising stage will greatly reduce the flight efficiency.

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.

Summary of the invention

The present invention provides an electric tilt-rotor aircraft, comprising: a fuselage, wings, a horizontal tail, and a 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.

An electric tilt-rotor aircraft as described above, wherein the cockpit layout adopts a 1+2+2 type, that is, the first row is a pilot seat, the seat deployment position is centered, 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 third row passengers can reach their seats through the space between the first and second row seats; after the pilot boards the aircraft from the left/right cabin door, he reaches the pilot seat through the aisles on both sides.

An electric tilt-rotor aircraft as described above, wherein an instrument panel is located in front of the pilot's seat, and the front instrument panel is slightly offset to the right to ensure that the pilot's eyes are in the middle of the PFD. A throttle panel and a side stick are respectively arranged on the left and right sides of the seat; the pilot controls the throttle panel and the side stick to achieve acceleration/deceleration, ascent/descent, pitch, roll, and turn of the aircraft.

An electric tilt-rotor aircraft as described above, 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 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.

An electric tilt-rotor aircraft as described above, wherein 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, 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, 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 transmits 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.

An electric tilt-rotor aircraft as described above, wherein the fixed rotor is configured as a four-blade foldable rotor.

An electric tilt-rotor aircraft as described above, wherein 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 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 along the inside and outside of the servo rod, the other end of the retractable push rod is fixed to one end of the tilt servo arm through a fastener, and the other end of the tilt servo arm is sleeved on the tilt shaft and fixed to the tilt shaft.

An electric tilt-rotor aircraft as described above, wherein the configuration modes of the aircraft 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, the transition mode and the fixed-wing mode in the air.

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 problems faced in the transition process of vertical take-off and landing aircraft. The power system provides a variety of high-voltage architecture power for the entire machine 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

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

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

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

FIG4 is an interior view of the cabin without doors and windows;

Figures 5 and 6 are disassembled views of the aircraft;

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

FIG9 is a schematic diagram of the arrangement of the tilt axis and the steering gear;

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

Figure 11 is a schematic diagram of rotor rotation;

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

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

FIG23 is a schematic diagram of a low voltage control architecture;

FIG24 is a schematic diagram of a low voltage power distribution architecture;

FIG25 is a schematic diagram of the internal avionics flight control architecture of an aircraft;

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

Fig. 27 is a schematic diagram of motor tilt control;

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

FIG30 is a schematic diagram of a transitional mode flight control mode;

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

FIG32 is a schematic diagram of the functional module architecture of the control distribution system;

Figure 33 is a schematic diagram of flight management module control;

Fig. 34 is a schematic diagram of attitude angle control;

FIG35 is a schematic diagram of the expected angular velocity calculation;

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

FIG37 is a schematic diagram of the actuator allocation module architecture.

DETAILED DESCRIPTION

The following is a clear and complete description of the technical solutions in the embodiments of the present invention in conjunction with the drawings in the embodiments of the present invention. Obviously, the described embodiments are part of the embodiments of the present invention, not all of the embodiments. 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.

Embodiment 1

As shown in FIGS. 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 .

The fuselage 1 includes a fuselage body and a rudder 10. The fuselage 1 body integrates a cabin door, a windshield, a lower window and a rear window. The rudder is installed at the rear of the fuselage 1 body to control and adjust the flight direction of the aircraft. The appearance of the fuselage 1 is set as a normal aircraft setting, which will not be described in detail here.

As shown in Figures 3 and 4, Figure 3 is a view of an aircraft with doors and windows, and Figure 4 is a view of 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 in the fuselage 1 includes a display console 16, a pilot seat 17, a front passenger seat 18, and a rear passenger seat 19. The rear of the cabin is a luggage compartment, and the fuselage has a luggage door 15. Preferably, the cabin layout of the present invention adopts a 1+2+2 type, that is, the first row is a pilot seat 17, the seat deployment position is centered, the second row is two front passenger seats 18, and the third row is two rear passenger seats 19. The width between the two seats in the second row is not less than 305mm, which meets the regulatory requirements. After opening the cabin door, the third row passengers can reach their seats through the space between the first and second row seats, wherein the cabin door can be set as a left and right cabin door, or one side can be set as a cabin door and the other side can be an emergency exit, which also meets the regulatory requirements. The width of the aisle formed by the first row of pilot seats 17 and the structures on both sides exceeds 305 mm. After the pilot boards the aircraft from the left/right cabin door, he or she reaches the pilot seat 17 through the aisles on both sides.

In front of the pilot seat 17 is the instrument panel. The front instrument panel is slightly to the right to ensure that the pilot's eyes are in the middle of the PFD. The throttle panel and side stick are respectively deployed on the left and right sides of the seat. The pilot controls the throttle panel and side stick to achieve the acceleration/deceleration, ascent/descent, pitch, roll, turn and other operations of the aircraft. The front instrument panel is equipped with 2 integrated displays DU, 1 backup instrument ISIS, 1 audio control panel, 1 flight control panel, 1 power control panel, 1 data recorder and 1 cabin voice recorder.

A continuous front windshield is designed in front of the pilot, and a left lower window and a right lower window are designed on the lower side, both of which are installed on the main structure of the cockpit. A continuous front windshield can provide the pilot with the necessary front view for controlling the aircraft, and the left/right lower windows can provide the pilot with a downward view when the aircraft lands vertically, making it easier for the pilot to align with the landing site. The display console 16 is designed in front of the pilot seat 17 and is installed on the main structure of the cockpit, providing a control interface for the pilot to operate the aircraft.

The left door and the right door are designed with a left window and a right window on the upper part respectively, and the left/right door and the window and the rear side are designed with a left/right porthole window. The left/right door and the left/right porthole window are installed on the main structure of the cockpit. This ensures that every passenger in the second and third rows can directly obtain a view of the outside of the aircraft through the doors and windows or portholes, effectively ensuring the riding experience of every passenger.

An independent space is designed behind the third row of seats, separated by the main structure of the cockpit, which can accommodate four 20-inch suitcases. Luggage can be put in or taken out through the 15-door luggage compartment fixed to the main structure of the cockpit.

As shown in the disassembled diagrams of Figures 5 and 6, the wing 2 includes an integral wing box section 21, an aileron 22, a left/right wing nacelle 23, and a left/right strut nacelle 24. The wing-body structure is separated, and the left and right struts are integrated on the front side of the integral wing box section 21. The left/right wing nacelle 23 is connected to the integral wing box section 21 through a tilt axis 25, and the left/right wing nacelle 23 rotates around the tilt axis. Figure 7 is a schematic diagram of the left/right wing nacelle. The left/right wing nacelle 23 includes a winglet 231 and an outer aileron 232. The wiring harness of the left/right wing nacelle 23 passes through the tilt axis to avoid interference between the wiring harness and the structure during the tilting process (as shown in Figure 8). A motor and a controller are arranged inside the left/right nacelle 23 of the wing, and the rotor is installed on the output shaft of the motor; the wing adopts an upper single gull-wing layout and is swept back, which increases the height of the rotor rotation plane on the strut 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.

The horizontal tail 3 includes a horizontal tail box section 31 and a horizontal tail left/right nacelle 32. The tail box section 31 is provided with an elevator 33 at its tail end. The horizontal tail left/right nacelle 32 is connected to the horizontal tail box section 31 through a tilt axis. The horizontal tail left/right nacelle 32 rotates around the tilt axis. The wing and tail share a set of tilt mechanism design, which reduces technical difficulty and design complexity.

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

The five-blade rotor is connected to a pitch-changing mechanism. The pitch-changing mechanism is a torque transmission structure used to amplify the displacement of the control input, amplify the control input force, or convert the arc motion control input into a strict linear motion output, which 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. 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 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. 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.

The left/right nacelles of the wing and the left/right horizontal tail nacelles all include a nacelle rib and a steering gear mounting frame 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 can rotate around the threaded pin axis. One end of the steering gear is fixed to the tilt shaft through a tilt steering gear arm. The steering gear drives the nacelle to tilt partially by driving the retractable push rod of the steering gear. FIG9 is a schematic diagram of the arrangement of the tilt shaft and the steering gear. The steering gear specifically includes a steering gear rod, a retractable push rod and a tilt steering gear arm. The threaded pin on the steering gear bracket fixes the steering gear rod. The retractable push rod is inserted into the steering gear rod and is retractably connected to the steering gear rod. A steering gear motor is arranged in the steering gear rod. The steering gear motor drives the retractable push rod to retract along the inside and outside of the steering gear rod. The other end of the retractable push rod is fixed to one end of the tilt steering gear arm through a fastener. The other end of the tilt steering gear arm is sleeved on the tilt shaft and fixed to the tilt shaft. The tilting movement of the nacelle is shown in FIG10.

When the aircraft is on the ground or in level flight, the nacelle part is not tilted, and the retractable push rod of the servo is in the extended state. When the aircraft is in a vertical takeoff or landing state and needs to perform a tilting action, the servo directly drives the retractable push rod to retract inward, and the propulsion and retraction speed of the retractable push rod is uniform and controllable. Under the reaction force, the nacelle part rotates at a uniform speed around the axis of the tilt axis, and the servo has a power-off locking function, which is used to lock the position of the nacelle before and after the tilting, until the nacelle part reaches the vertical state, and the tilting process ends. During the entire rotation process, the box section and the tilt axis are fixed, and the tilt servo arm fixed on the tilt axis is also fixed. The retractable push rod of the servo retracts and retracts, driving the nacelle part to rotate, and the servo makes additional rotation around the threaded pin to avoid jamming.

As used herein, "rotor" may refer to any suitable rotary aerodynamic actuator, commonly referred to as a rotor, rotary wing, rotary airfoil, etc. 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 adopts a multi-blade, large-diameter, low-speed rotor design, which can reduce the power demand in the multi-rotor flight mode, reduce the take-off and landing noise level, and reduce the vibration of the multi-rotor and the body structure during the conversion process.

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-rotor directions of the aircraft are vertically upward, and all rotors provide lift for the aircraft. In 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 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.

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 arrangement and the cruise arrangement, at least part 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.

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 the nacelle are in a vertical state.

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 whole 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 whole aircraft.

When the aircraft is used for urban air traffic operations, the aircraft is initially in a multi-rotor state, in which the left and right wing tilt nacelles and the left and right horizontal tail tilt nacelles are rotated to 90° upward, and the foldable rotors on the nacelle components are unfolded into a four-blade propeller state; after the passengers have boarded, the six power systems provide lift for vertical takeoff and fly to a safe altitude. Subsequently, the left and right wing tilt nacelles and the left and right horizontal tail tilt nacelles are controlled to slowly tilt forward to provide forward thrust for the aircraft. The forward flight speed of the aircraft continues to increase until a certain speed, and the foldable rotors on the left and right nacelle components are folded, feathered and completely closed; the left and right wing tilt nacelles and the left and right horizontal tail tilt nacelles rotate to 0°, and the pitch gradually increases. Subsequently, the rotors on the left and right wing tilt nacelles and the left and right horizontal tail tilt nacelles provide thrust to ensure the cruising stage of the aircraft. During the approach phase, the left and right foldable rotors start to provide lift, and the left and right wing tilt nacelles and the left and right horizontal tail tilt nacelles gradually rotate from 0° to 90°, so that the forward flight speed of the aircraft gradually decreases until it hovers. During this process, the pitch of the propellers gradually decreases. Subsequently, six power systems provide lift for vertical landing to the destination, completing the flight mission.

The landing gear 4 adopts a three-point landing gear, including a front landing gear 41 and two main landing gears 42. The main landing gear 42 is integrated with a brake system to play a braking role when the aircraft lands, and the front landing gear 41 plays a role of balance and direction control. The three-point landing gear supports ground autonomous mobility, provides low-speed taxiing capability support, and can be used for emergency taxiing landing support.

Embodiment 2

Embodiment 2 of the present invention provides an internal power system of an aircraft, and the power system includes six power units and an energy system.

A single power unit includes a motor (including motor controller), a reducer, a rotor, and related lubrication and thermal management accessories. The motor, reducer and rotor in each power unit are structurally integrated. The power unit is generally composed of a metal structure and is suitable for design using finite element analysis methods such as topology optimization to achieve the best force transmission path and the lightest structural weight. The rotor base and reducer top cover are one structural part, and the reducer bottom cover and motor top cover are one structural part. 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 install and fix the electric engine 119, and the reducer mounting adjustment ring 115 is used to install and fix the reducer 1110; both ends of the plurality of reinforced corner boxes 112 are detachably connected to the engine mounting frame 111 and the reducer connecting frame 114 by fasteners 86. 13; and both ends of the multiple reducer connecting legs 114 are also detachably connected between the reducer connecting frame 113 and the reducer installation adjustment ring 115 through fasteners 86; multiple reinforced corner boxes 112 are evenly distributed around the center line of the engine installation frame 111 and the reducer connecting frame 113, and multiple reducer connecting legs 114 are also evenly distributed around the center line of the reducer connecting frame 113 and the reducer installation adjustment ring 115, so that in the process of the reducer 1110 transferring various load bending moments from the rotor to the fuselage 1 structure, the load transfer is more uniform, reducing the possibility of stress concentration, thereby extending the service life of the equipment. The motor and reducer are both designed with cooling channels on the shell and share a water cooling system. Each motor has two redundant energy and control channels, and the six motors have a total of twelve channels. The nacelle for installing the power unit and the tilting servo is designed as a semi-monocoque structure, which not only meets the strength requirements, but also leaves space for the layout and installation of equipment, wiring harnesses, and air ducts. Among them, the tilting servo has not only the push rod type servo, but also can be set to the turbine worm type, harmonic gear type, cycloid pinwheel type, etc.

The energy system includes multiple battery modules, multiple battery management and energy distribution units, and related thermal management accessories. The battery module consists of battery cells in series, water-cooled plates for cooling, and sensors for monitoring. The module shell uses high-temperature resistant flame-retardant materials, provides structural support, and has a smoke exhaust channel for use in thermal runaway. The battery module has high-voltage interfaces, communication interfaces, water-cooling interfaces, smoke exhaust channels, and mechanical installation interfaces.

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 is composed 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 will be 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, which can optimize the fuselage space layout, reduce the fuselage size, avoid major battery failures affecting the cabin, and simplify the cabin design/protection; the batteries are physically isolated, will not affect each other, and reduce 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 precious space such as the fuselage section (cabin, luggage compartment) is not occupied; 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.

The high-voltage interfaces of each module in a set of battery modules are connected in series and ultimately connected to the matching battery management and energy distribution unit; the communication interfaces of the battery modules are also connected to the matching 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.

The battery module provides an operating voltage platform (high-voltage platform) of approximately 500V-800V. Each battery module is designed to be matched with a battery management and energy distribution unit. When four battery modules are installed, each battery module provides energy to three channels; when six battery modules are installed, each battery module provides energy to two channels. This design ensures that when a single battery module 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 matching battery management and energy distribution units are also installed in adjacent positions.

A single battery management and energy distribution unit integrates the functions of multiple traditional devices, such as battery management system, DC-DC converter, energy distribution unit, etc., reducing the number of devices, the number of connectors, the space occupied, and optimizing the wiring harness direction and layout. The DC-DC conversion function of the battery management and energy distribution unit will convert the 500V-800V working voltage into a working voltage platform (medium voltage platform) and a 28V working voltage platform (low voltage platform) available for civil aviation, and together with the 28V low-voltage battery, it will provide redundant power supply for the low-voltage equipment of the entire aircraft. The battery management and energy distribution unit designs the position of the connector according to the wiring harness direction, greatly simplifying the difficulty of wiring harness layout.

Embodiment 3

Different voltages need to be provided inside the aircraft to power different devices. Generally, high voltage and low voltage architectures are required to meet the voltage requirements of different devices. In the present invention, multiple batteries are installed inside the struts and wings of the aircraft structure of Example 1, and all batteries support charging and battery replacement. Different voltages are obtained by combining multiple batteries to power different devices.

Embodiment 3 of the present invention provides an aircraft high-pressure architecture, specifically including:

As shown in Figure 14, the aircraft has six motors (respectively motor 1#, motor 2#, motor 3#, motor 4#, motor 5#, motor 6#, motor 1# and motor 6# are used to drive the rotation of the foldable rotors on the left and right struts, motor 2# and motor 5# are used to drive the rotation of the rotors on the left/right nacelles of wing 2, 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#, power battery 4#), and the four power batteries are evenly and symmetrically connected to the motor channel, so that the energy power architecture of the aircraft has complete axisymmetry or central symmetry, completing the safe 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.

Each motor has two redundant channels. When one channel fails, the other channel can still ensure that the motor runs at half the rated power. Each power battery can output three high-voltage distribution channels, which are connected to the inverters of different motors. For example, in FIG. 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#.

When a single inverter fails, the motor torque drops to 70% of the rated torque, and 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, then the power of opposite motor 3# is reduced, and the power of inverter 3A is reduced to achieve balanced stability on both sides.

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, and the torque is still balanced to ensure regional safety.

When two adjacent power batteries fail, all motors are in a single inverter state. For example, as shown in Figure 17, power battery 3# and power battery 4# both 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 the balance is still maintained. The single-channel power can still be increased, but in order to reduce the risk, a controlled forced landing is required to reduce losses.

When in fixed rotor mode, motor 1# and motor 6# are shut down, as shown in FIG18 . 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.

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 the left motor 6# fails, the power supply of motor 6# connected to power battery 2# and power battery 4# is disconnected. At this time, in order to achieve stable left and right operation of the aircraft, the power of the right motor can be reduced. Since one of the power supplies of 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. It is preferred that since motor 3# is connected to power battery 1# and power battery 3#, the power of inverter 3A and inverter 3B of right motor 3# is reduced.

When two symmetrical power batteries fail, for example, as shown in FIG20 , the power batteries 2# and 4# on the opposite sides fail, then motors 2#, 4# and 6# cannot work. At this time, motors 1#, 3# and 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 flight safety of the aircraft, a controlled forced landing is still required to reduce losses.

As shown in FIG. 21 , in order to prevent an emergency situation where multiple power batteries become ineffective, the present invention can also add a fifth power battery 5# to connect the distribution circuit of all power batteries (1# to 4#). In an emergency situation, power battery 5# replenishes electric energy for all power batteries, so that each power battery replenishes electric energy to the corresponding motor.

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 motor is connected to two distribution boxes. In the figure, the redundant channel 1# of motor 1# is connected to distribution box 2#, the redundant channel 2# of motor 1# is connected to distribution box 1#, the redundant channel 1# of motor 2# is connected to distribution box 2#, the redundant channel 2# of motor 2# is connected to distribution box 1#, the redundant channel 1# of motor 3# is connected to distribution box 2#, the redundant channel 2# of motor 3# is connected to distribution box 1#, the redundant channel 1# of motor 4# is connected to distribution box 2#, the redundant channel 2# of motor 4# is connected to distribution box 1#, the redundant channel 1# of motor 5# is connected to distribution box 2#, the redundant channel 2# of motor 5# is connected to distribution box 1#, the redundant channel 1# of motor 6# is connected to distribution box 2#, and the redundant channel 2# of motor 6# is connected to distribution box 1#. When one channel fails, the other channel can still ensure that the motor runs at half the rated power.

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

Embodiment 4

Embodiment 4 of the present invention provides a low-voltage architecture of an aircraft, and the low-voltage architecture of the aircraft is divided into a low-voltage control architecture ( FIG. 23 ) and a low-voltage power distribution architecture ( FIG. 24 ).

As shown in Figure 23, in the aircraft low-voltage control architecture, the management computer inside the aircraft is connected to the low-voltage power distribution switch box through the communication line CAN1, and is connected to the DC-DC converter through the communication line CAN2. The DC-DC converter can provide two 28V power supplies, A and B, to the low-voltage power distribution switch box. The low-voltage power distribution switch box and the 28V low-voltage battery together constitute a low-voltage redundant power supply and power the low-voltage equipment. The DC-DC converter has two redundant inputs and can obtain high-voltage power from power battery 1 and power battery 2 at the same time. Power battery 1 and power battery 2 are connected to the flight management computer through the communication line CAN3 and are controlled and monitored by the flight management computer.

As shown in Figure 24, the low-voltage power distribution architecture of the aircraft mainly describes the 28V low-voltage architecture inside the power distribution switch box. The low-voltage power distribution architecture generally adopts two forms of multi-input safety redundant design. Form one, the DC-DC converter is connected in parallel with the 28V low-voltage battery to form a relatively safe 28V bus. Adding a diode to the 28V low-voltage battery branch will prevent the bus from backflowing power to the battery. Form two, two redundant DC-DC converter channels ChA and ChB, and one 28V low-voltage battery are connected to the power distribution switch box, among which the 28V battery bus is the most reliable single power source, which can power the DC-DC converter bus ChA, ChB, and 28V emergency bus through three diodes. Therefore, even if the DC converter buses ChA and ChB fail, they can obtain power from the 28V battery bus to remain powered; and the 28V emergency bus can work stably and has the highest safety as long as one of the DC buses ChA, ChB and the 28V emergency bus can work.

Embodiment 5

As shown in FIG. 25 , the second embodiment of the present invention provides an internal avionics flight control architecture of an aircraft, including a flight control system, an indication and recording system, a navigation sensor, and a communication system, and provides avionics optional functions for different customer groups.

The flight control system includes a flight control computer, a cockpit control system, a servo and a motor controller. There are 4 flight control computers, including 3 independent flight control computers with mutual redundancy and 1 backup flight control computer. They are developed according to DAL A, have heterogeneous monitoring channels to ensure data integrity, and meet the requirements of common mode fault analysis. The cockpit control system includes a side stick and a throttle lever, which provide the pilot with a human-machine interface for flight control and are used to manually control the aircraft. The side stick has three-axis control of longitudinal, lateral and axial rotation; the throttle lever has longitudinal control. Through the side stick and throttle lever, the control of the aircraft's flight track control/altitude/speed/heading and ground taxiing is achieved. The control amount of the joystick can be attitude, speed, and altitude, which is specifically determined by the control law function. At the same time, the side stick and throttle lever support air-to-ground communication control function, automatic flight disconnection control, display cursor control and other functions. The side stick and throttle lever have the functions of automatic return to center and artificial sense force when the stick is released. The side stick and throttle lever use multiple redundant sensors to convert the driver's control instructions into electrical signals and send them to the flight control computer. The servos include 4 aileron servos, 4 elevator servos and 2 rudder servos, as well as 4 tilt servos and 4 variable pitch servos, which receive control commands from the flight control computer to implement actuation. There are 4 ailerons, 4 elevators and 2 rudders, which are driven by independent servos to deflect. When a single control surface fails, it is easy to reconstruct the redundancy to ensure safe flight, and can reduce the performance requirements of the control surface servos. The inner ailerons and elevators can be deflected downward at a large angle to reduce power loss in the hovering state. There are 6 motor controllers in total, which receive control commands from the flight control computer and drive the motors to operate. The cockpit joystick includes 1 side stick and 1 throttle.

The indication and recording system includes two integrated displays DU, one backup instrument ISIS, one data recorder FDR, and one cockpit voice recorder CVR. The display system is specially designed for a single pilot. Each DU can conveniently display all information by key switching. ISIS is an emergency instrument when the main display fails, providing important flight information such as airspeed, altitude, and attitude. FDR and CVR are used to record flight data and cabin voice respectively.

Navigation sensors include 3 sets of atmospheric data systems ADS, 3 sets of heading reference systems AHRS, 3 sets of magnetometers, 2 sets of satellite navigation receivers GPS, and 2 sets of radio altimeters RA. Navigation sensors provide navigation information for the flight control system and indication and recording system, including airspeed, air pressure, flight attitude, heading, position, altitude, etc.

The communication system includes two VHF radios, an audio controller, an S-mode transponder and an ADS-B, and is equipped with an intercom interface and earphones. The communication system is used to realize intercom communication between the pilot and ground air traffic control, and between the pilot and passengers.

In order to meet the needs of mid- to high-end customers, the aircraft provided by the present invention also has avionics optional functions, including flight management system, detection and collision avoidance system, automatic flight, central maintenance, etc.; data link, optical detection and radar equipment, etc. can also be installed for customers who meet the conditions for use. In terms of bus communication, the avionics flight control comprehensively considers low cost, reliability and mature civil aviation technology. The flight control computer uses RS485 and CAN bus as the field control bus to realize the control of the servo and motor; ARINC429 bus is used between the flight control system and the avionics system to receive and obtain navigation sensor data and interact with the display system.

The integrated display exchanges data with external sensors through the RS422 interface and exchanges information with the flight control computer through the ARINC429 bus; the audio control board is connected to the integrated display through a high-speed bus, and the pilot configures it through the integrated display. The audio control board is also connected to two VHFs and one intercom interface unit to achieve audio transmission; the flight control panel is connected to the flight control computer through discrete quantities to achieve the flight mode switching function; the power control panel is used to control the high-voltage up/down and low-voltage up/down of the E20 aircraft, and at the same time completes the sending of motor enable signals; a data recorder is used to record flight data; and a cabin voice recorder is used to record audio information in the cabin.

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 displays 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 display of CAS warning information. 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 prediction, position prediction, range prediction, 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, reduction 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.

Embodiment 6

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 burden on the pilot.

The flight control system communicates with the atmospheric data system ADS, the attitude reference system AHRS, three sets of magnetometers, the satellite navigation receiver GPS, and the radio altimeter RA through the bus to collect the flight data and position information of the aircraft, execute the flight control law algorithm, and complete the control functions of the aircraft, including vertical take-off and landing control (speed mode, attitude mode), tilt transition control and fixed rotor cruise flight control, as well as track control, heading control, altitude control and ground taxiing. At the same time, it performs envelope protection functions such as stall protection, overspeed protection, and tilt angle protection, and monitors the status of the entire flight control system. It also works in coordination with the flight management system, detection and collision avoidance system, and automatic flight to achieve performance optimization, automatic navigation, automatic flight and automatic obstacle avoidance functions, reduce the burden on pilots, and improve flight safety.

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

The pilot inputs control instructions in the cockpit to perform automatic flight missions, and the control console in the cockpit sends flight control instructions to the flight control law. Among them, the flight control instructions include throttle control instructions, yaw control instructions, pitch control instructions, and roll control instructions, which are used to control the flight mode and flight phase.

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

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;

The controller inputs forces and torques to 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.

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, a transient control or compensation amount that meets the requirements shall be set.

As shown in Figure 27, the actuation quantities that can be controlled by the flight control system include: 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, the nacelles at motor positions 2, 3, 4 and 5 of the aircraft have a tilt function, and the tilt control of the nacelle can change the thrust direction of positions 2, 3, 4 and 5; rudder control, when the aircraft rudder is used to fix the rotor configuration, the torque control in the pitch, yaw and roll directions is performed, and the control quantity includes pitch control quantity, roll control quantity and yaw control quantity.

The rotor modes of the aircraft include a low-speed multi-rotor control mode, a transitional mode flight mode, and a fixed rotor stage control mode.

As shown in Figures 28 and 29, in the low-speed multi-rotor control mode, the multi-rotor flight stage (ground speed 0-5m/s) controls the aircraft's roll, pitch, yaw and vertical operation by controlling the pulling force of the six motors. The low-speed multi-rotor control mode includes roll control, pitch control, yaw control and vertical control. In the low-speed multi-rotor mode, the aircraft's forward lateral speed is achieved by changing the pitch and roll angles. Motors 1 and 6 are tilted inward or forward to improve yaw control capabilities.

Among them, the desired lateral speed of the aircraft under roll control generates the desired roll attitude angle through the speed control loop, the desired roll attitude angle generates the desired roll angular velocity through the attitude control loop, the desired roll angular velocity generates the desired roll moment through the angular velocity control loop, and the roll moment control is performed by controlling the tension difference between motors 1, 2, 3 and motors 4, 5, 6. The desired forward speed of the aircraft under pitch control generates the desired pitch attitude angle through the speed control loop, the desired pitch attitude angle generates the desired pitch angular velocity through the attitude control loop, the desired pitch angular velocity generates the desired pitch moment through the angular velocity loop, and the pitch moment control is performed by controlling the tension difference between motors 1, 6 and motors 3, 4. Under yaw control, the desired yaw angle of the aircraft passes through the angle control loop to generate the desired yaw angular velocity, and the desired yaw angular velocity passes through the angular velocity control loop to generate the desired yaw moment, and the yaw moment control is performed by controlling the tension difference between motors 1, 3, 5 and motors 2, 4, 5 and the tilt angle differential of motors 2, 3, 4, 5. Under vertical control, the desired vertical velocity of the aircraft passes through the velocity control loop to generate the desired vertical acceleration, and the vertical acceleration control is performed by simultaneously controlling the tension of motors 1, 2, 3, 4, 5, and 6.

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.

During transition flight control, the expected forward speed, expected roll angle, expected yaw angle, and expected altitude will be calculated through the control loop to obtain the expected forward force (Fxb), expected vertical force (Fzb), expected rolling moment (L), expected pitch moment (M), and expected yaw moment (N) in the body coordinate system. During transition flight, the aircraft wing 2 will generate aerodynamic force as the airspeed increases, and the control surface will generate aerodynamic torque. At this time, the torque control weights of the rotor and the fixed rotor will be allocated according to the "fusion speed", so that the "torque control right" can be smoothly handed over when the rotor transitions to the fixed rotor state. The "fusion speed" is the fusion speed of the airspeed and the ground speed.

As shown in FIG31 , during the fixed rotor flight phase, the rolling 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.

The functional module architecture of the control distribution system is shown in Figure 32. The navigation device of fuselage 1 collects the real-time status information of the aircraft body and feeds it back to the corresponding control module to form a control closed loop. The complete process from receiving the pilot's input to the flight control output includes the following main functional modules: flight management module, control module, actuator distribution module. Among them, the control module is divided into the following loops (sorted from inside to outside): angular velocity control loop, attitude angle control loop, acceleration control (tilt control), speed control loop and position control loop.

Specifically, as shown in Figure 33, in the flight management module, according to different control methods, it is divided into: autonomous flight and manual control. Autonomous flight is generally used for the control of fixed rotor configuration and tilt configuration stages, and manual control is mostly used for the control of rotor configuration stages. 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 according to the pre-set route, generates guidance instructions, and makes the aircraft fly according to the preset route. The preset route feature waypoints in the flight control are shown in Table 1 below:

Table 1 Description of near-field waypoints

Manual control is the operation of the pilot. The aircraft control input devices are divided into the left throttle stick and the right side stick according to the Unified Flight Control principle. The corresponding control quantity mapping relationship is shown in Table 2 below (the actual application is not limited to this mapping relationship, the main purpose is to reduce the difficulty of pilot control):

Table 2 Manipulation quantity mapping relationship

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:

Table 3 Rotor state control quantity mapping table

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

Input: expected longitude and latitude, expected altitude, current longitude and latitude of the aircraft, current altitude;

Output: expected speed;

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 position of the aircraft and the desired position, as follows:

Among them, Vn _exp , Ve _exp and Vd _exp are the desired north velocity, east velocity and vertical velocity respectively; K _pe , K _pn and K _pd are the position control gains respectively, Lat _exp and Lat are the desired latitude and the current latitude of the aircraft, Lon _exp and Lon are 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.

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

Inputs: expected forward speed, expected lateral speed, expected vertical speed; forward speed, lateral speed and vertical speed;

Output: expected forward acceleration, expected lateral acceleration and expected vertical acceleration;

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

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

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

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

Output: tilt angle, desired pitch angle, desired roll angle and desired yaw angle;

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

Where, 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 gravity constant;

After the total thrust vector is obtained, the expected pitch and roll angles under this vector are calculated in combination with the current yaw angle, as follows:

Where, θ _exp and φ _exp are the desired pitch angle and the desired roll angle, and are the unitized variables of a _n , a _e and a _d respectively, and ψ is the current heading angle of the aircraft. The tilt angle represents the proportion of the aircraft's forward acceleration vector and vertical acceleration vector in the direction of the total thrust vector. The smaller the tilt angle, the greater the forward acceleration.

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

Input: expected roll angle, expected pitch angle, expected yaw angle, current roll angle, pitch angle and yaw angle of the aircraft;

Output: expected roll angular velocity, expected pitch angular velocity and expected yaw angular velocity;

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.

As shown in Figure 35, the expected quantity 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 calculated, 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 and torsion separation) is satisfied. Therefore, when calculating the expected angular velocity, the attitude error is split into the following two parts:

Tilt error: In the body coordinate system, the difference between the z-axis of the desired posture and the z-axis of the current posture is used to obtain the tilt error (or the tilt error angle and its unit vector. Since it does not contain torsion, the third component of the unit vector is 0).

Torsion error: In the body coordinate system, the torsion error is obtained by subtracting 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).

Finally, the roll angular velocity and pitch angular velocity are extracted from the tilt error, and after limiting processing according to the attitude limit, the desired roll and pitch angular velocity are obtained; the yaw angular velocity is extracted from the torsion error, and after fading processing according to the error size, the desired yaw angular velocity is output.

Compensation: When the aircraft has an attitude angle, it needs to be compensated when calculating the expected angular velocity. The compensation varies according to the attitude angle. When the pitch and roll attitude angles are 0°, the roll angular velocity and pitch angular velocity compensation are 0. The compensation is calculated as follows:

Among them, p _Δ , q _Δ and r _Δ are the pitch, roll and yaw angular velocity compensations respectively.

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:

Input: expected roll rate, expected pitch rate and expected 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;

Output: roll control, pitch control and yaw control;

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

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, an 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 and the rudder allocation amount, and maps it to the control amount within the travel range of the rudder servo according to a certain mapping relationship, thereby controlling the travel of the rudder servo. The motor mapping module is similar. The superimposed total amount is mapped to a torque control amount between 0 and 400 N.m, and the motor is controlled after output. The corresponding mapping relationship is as follows:

T is the total desired pulling force after the feedforward amount and the distribution amount are superimposed, and _Mcmd is the motor torque control amount.

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

The 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. In order to meet the independent control of tilt and twist, the control distribution is divided into force distribution, roll and pitch distribution, and yaw distribution. The distribution relationship is as follows:

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

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

in, The total motor control quantity, K _f and K _n are the force distribution ratio and yaw control ratio respectively, and the calculation relationship is as follows:

in, and The above and

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

Among them, 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 , _KI and K _d are PID control coefficients respectively, Out, Out _max and Out _min are output value, maximum output value and minimum output value respectively.

The battery balancing control system provides an aircraft high-voltage architecture and an aircraft low-voltage architecture, and the aircraft high-voltage architecture and the aircraft low-voltage architecture are respectively described in detail in the above-mentioned embodiment 3 and embodiment 4.

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.

evtol integrated display: unique data display transmission path, the data of each aircraft system (motor, power supply, BMS, mechanical, environmental control) is collected by the flight control computer FCC, and then packaged into A429 data and sent to DU. Dedicated display content, motor, battery BMS, evtol flight control integrated display.

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 substitutions, 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 (9)

1. An electric tiltrotor aircraft, characterized in that the voltage architecture of the aircraft comprises: a high voltage architecture and a low voltage architecture;

the high-voltage architecture includes: six motors and four power batteries; the six motors are used for driving the rotation of the foldable rotor wings on the left and right stay bars, the rotation of the rotor wings on the left and right nacelle and the rotation of the rotor wings of the left and right horizontal tail nacelle respectively; four power batteries are connected to the motor channel evenly and symmetrically; each motor is provided with two paths of redundant channels, and when one path of the redundant channels fails, the other path of the redundant channels ensures that the motor operates at half rated power; each power battery outputs three paths of high-voltage power distribution and is respectively connected with different motor inverters;

The low pressure architecture includes: a low voltage control architecture and a low voltage power distribution architecture; the management computer in the aircraft in the low-voltage control architecture is connected with a low-voltage distribution switch box through a communication line CAN1, is connected with a direct-current/direct-current converter through a communication line CAN2, and provides AB two-way 28V power supply for the low-voltage distribution switch box, and the low-voltage distribution switch box and a 28V battery together form low-voltage redundant power supply and power supply for low-voltage equipment; the low voltage power distribution architecture includes: the direct current-direct current converter is connected with the 28V low-voltage battery in parallel to form a 28V bus bar; or two redundant DC-DC converter channels ChA and ChB and one 28V low-voltage battery are connected into the distribution switch box.

2. An electric tiltrotor aircraft according to claim 1, wherein the six motors are motor 1#, motor 2#, motor 3#, motor 4#, motor 5#, motor 6#, motor 1# and motor 6# are used to rotate the foldable rotor on the left and right struts, motor 2# and motor 5# are used to rotate the rotor on the left/right nacelle of the wing, respectively, and motor 3# and motor 4# are used to rotate the rotor on the left/right nacelle of the wing, respectively.

3. An electric tiltrotor aircraft according to claim 2, wherein the four power cells are power cell 1#, power cell 2#, power cell 3#, power cell 4#, power cell 1# being coupled to motor 1#, motor 3# and motor 5#, respectively; the power battery 2# is connected with the motor 2#, the motor 4# and the motor 6#; the power battery 3# is connected with the motor 1#, the motor 3# and the motor 5#; the power battery 4# is connected with the motor 2# and the motor 4# and the motor 6# respectively.

4. An electric tiltrotor aircraft according to claim 3, wherein 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 path of the power battery 2# is connected with the inverter 2B of the motor 2# at high voltage, the second path of the power battery 1# is connected with the inverter 4B of the motor 4# at high voltage, and the third path of the power battery 2# is connected with the inverter 6B of the motor 6# at high voltage; the first path of the power battery 3# is connected with the inverter 1A of the motor 1#, the second path of the power battery 3# is connected with the inverter 3A of the motor 3#, and the third path of the power battery 3# is connected with the inverter 5A of the motor 5#; the first path of the power battery 4# is connected with the inverter 2A of the motor 2# in a high voltage mode, the second path of the power battery 4# is connected with the inverter 4A of the motor 4# in a high voltage mode, and the third path of the motor 4# is connected with the inverter 6A of the motor 6# in a high voltage mode.

5. The electric tiltrotor aircraft according to claim 4, wherein when the single inverter fails, motor torque drops, stable operation is achieved by reducing the contralateral motor power.

6. An electric tiltrotor aircraft according to claim 4, wherein when a single power battery fails, the three motors controlling the connection of the failed power battery are powered down and the three motors are just spaced apart and the torque is balanced.

7. An electric tiltrotor aircraft according to claim 4, wherein all motors are controlled to be in a single inverter state when two adjacent power cells fail.

8. An electric tiltrotor aircraft according to claim 4, wherein a fifth power cell 5# is added to connect the power distribution circuits of all the power cells, and the power cell 5# supplements all the power cells with electric energy in an emergency, so that each power cell supplements the corresponding motor with electric energy.

9. An electric tiltrotor aircraft according to claim 4, wherein six power cells and six motors are employed, each motor being connected to an inverter of two power cells.