Distributed contra-rotating rotor hybrid tilting rotorcraft

Abstract

The invention belongs to the field of vertical take-off and landing aircrafts, and relates to a distributed contra-rotating rotor hybrid tilting rotorcraft. Comprising a front tilting electric contra-rotating rotor nacelle, a front cabin body, a wing end tilting electric contra-rotating rotor nacelle, a V-shaped empennage, a V-shaped empennage aileron, a vertical empennage, a retractable undercarriage, a green hybrid power energy system and an empennage tilting electric contra-rotating rotor nacelle, the front tilting electric contra-rotating rotor nacelle is hinged with the other end of the front cabin body; an empennage tilting electric contra-rotating rotor nacelle is arranged at the end part of the V-shaped empennage; wing end tilting electric contra-rotating rotor wing nacelles are arranged at the ends of the wings; and rotor wing units are arranged in the rotor wing nacelles, the six rotor wing units all adopt a coaxial reverse upper and lower two-paddle structure, and the rotating directions of the rotor wings in the six rotor wing units are that the adjacent rotating directions and the opposite rotating directions are opposite.

Description

Technical Field

[0001] The present invention belongs to the technical field of vertical take-off and landing aircraft development, and in particular relates to a distributed counter-rotating rotor hybrid tiltrotor aircraft. Background Art

[0002] As a special aircraft, helicopters have special capabilities such as vertical take-off and landing, hovering in the air, flying forward, backward, left and right, and flying at ultra-low altitude. These characteristics and capabilities are irreplaceable by fixed-wing aircraft and other means of transportation, and also determine that helicopters have an important strategic position in maintaining national security, national economic construction, and social development.

[0003] First of all, the biggest feature of a helicopter is that it can only provide lift and forward force through the main rotor. The main rotor control must achieve the functions of collective pitch and cyclic adjustment, and has a complex variable pitch and cyclic control system. The helicopter mode generates the highest lift efficiency, but forward flight is also the biggest shortcoming of the helicopter mode. That is, the power required by its rotor in cruising state is about 40% of the maximum take-off power, which is much greater than that of a fixed-wing aircraft with the same maximum take-off weight. As a result, the maximum range is short, generally about 600-900km, which is equivalent to a fraction of that in fixed-wing state; and its flight speed is slow, generally the maximum flight speed is about 300km / h, not exceeding 400km / h;

[0004] Secondly, typical helicopters such as the AgustaWestland AW139 and Sikorsky S-70 Black Hawk are twin-engine, single-rotor helicopters with tail rotors. The rotor system, transmission system, and control system in the main power transmission chain have no redundant design. Single-point failures can easily cause a Class I accident. The safety and reliability of helicopters are not high, about 10^ -5 ;

[0005] Third, helicopters are noisy and vibrate loudly, and their flight routes are strictly restricted. They cannot fly in urban ultra-low altitudes (300m). The main sources of noise are the rotors, tail rotors, and engines. Currently, the noise level inside a conventional helicopter cabin is 100dB, and at 50m the noise level is 120dB.

[0006] Fourth, helicopters have high manufacturing costs, high maintenance costs, and high operating costs, far exceeding those of fixed-wing aircraft of the same weight.

[0007] Products based on traditional mechanical thinking have the problem that the turboshaft engine cannot be decoupled from the transmission system and the dual-rotor mechanical connection, and cannot solve the problem of high efficiency of the turboshaft engine in vertical take-off and landing in helicopter mode, high power state during transition flight, and low power state in fixed-wing mode, and the maximum range increase is limited; its configuration power transmission chain is similar to that of traditional typical helicopter power transmission chains, and the redundancy reliability indicators are limited, which cannot solve user pain points; subject to the limitations of large rotor disc load and blade tip linear speed, the noise and vibration of the whole machine are limited; the similarity of the power transmission chain structure determines that its manufacturing cost, maintenance cost, and operating cost improvement are limited, and even in small batches, the manufacturing, maintenance and operating costs are much higher than those of helicopters.

[0008] With the development of vertical take-off and landing aircraft technology, the requirements for aviation aircraft are becoming increasingly higher, requiring long range, higher safety and reliability, low noise and low vibration, and lower manufacturing, maintenance and use costs. Summary of the Invention

[0009] Purpose of the invention: To provide a distributed counter-rotating rotor hybrid tiltrotor aircraft to improve safety and reliability.

[0010] Technical solution:

[0011] A distributed counter-rotating rotor hybrid tilt-rotor aircraft, comprising: a front tilt-electric counter-rotating rotor nacelle 1, a front cabin body 2, a wingtip tilt-electric counter-rotating rotor nacelle 3, a fuselage 4, a V-shaped tail 5, a V-shaped tail aileron 6, a vertical tail 7, a retractable landing gear 8, a wing 9, an aileron 10, a flap 11, a green hybrid energy system 12, and a tail tilt-electric counter-rotating rotor nacelle 13, wherein one end of the front cabin body 2 is fixed to the front of the wing 9, the wing 9 is provided with ailerons 10 and flaps 11, the front tilt-electric counter-rotating rotor nacelle 1 is hinged to the other end of the front cabin body 2; a V-shaped tail aileron 6 is provided on the V-shaped tail 5, a vertical tail 7 is provided below the V-shaped tail 5, a tail tilt-electric counter-rotating rotor nacelle 13 is provided at the end of the V-shaped tail 5, and a wingtip tilt- Electric counter-rotating rotor nacelle 3; a retractable landing gear 8 is provided under the fuselage 4, and a green hybrid energy system 12 is used to power the motors in the front tilt electric counter-rotating rotor nacelle 1, the wingtip tilt electric counter-rotating rotor nacelle 3, and the tail tilt electric counter-rotating rotor nacelle 13. Rotor units are arranged in the wingtip tilt electric counter-rotating rotor nacelles 3 at the wingtips of the left and right wings, and rotor units are arranged in the front tilt electric counter-rotating rotor nacelle 1 in front of the left and right wings. Rotor units are arranged in the tail tilt electric counter-rotating rotor nacelles 13 on both sides of the V-shaped tail oblique wing. The six distributed rotor units all adopt a coaxial inverted upper and lower propeller configuration. The rotation direction of the rotors in the six rotor units is: adjacent and relative rotation directions are opposite, that is, the front left, middle right and rear left are clockwise, and the front right, middle left and rear right are counterclockwise.

[0012] Furthermore, all six rotor units are tilt-rotating electric counter-rotating rotor units.

[0013] Furthermore, the green hybrid energy system 12 adopts a series hybrid system.

[0014] Furthermore, the tilting electric counter-rotating rotor unit includes: a front fairing 20, a rear pitch electric pitch-changing actuator 21, a front pitch electric pitch-changing actuator 22, a three-blade front rotor hub 23, a front rotor blade 24 at the tip of a lower reverse-swept winglet, a front blade pitch-changing mechanism 25, a front blade bearing assembly 26, a front blade pitch-changing pin disc 27, an elastic external spline shaft 28, a bearing assembly 29 connecting the front rotor hub and the rear rotor hub, a rear blade pitch-changing mechanism 30, a three-blade rear rotor hub 31, a rear rotor blade 32 at the tip of a lower reverse-swept winglet, a rear blade bearing assembly 33, a rear blade pitch-changing pin disc 34, a differential planetary reducer 35, a DC permanent magnet motor 36, and a power line control line riser assembly 37, wherein the output end of the DC permanent magnet motor 36 is positioned The boss and the mounting inner hole of the differential planetary reducer 35 are matched to ensure concentric positioning, the DC permanent magnet motor 36 is directly connected to the casing flange of the differential planetary reducer 35 by a screw pile group, and the output spline shaft of the DC permanent magnet motor 36 is connected to the elastic shaft 47 of the differential planetary reducer 35; the outer rotor shaft 42 of the differential planetary reducer 35 is concentrically positioned by matching the flange boss and the inner hole of the 3-blade rear rotor hub 31, and then the outer rotor shaft 42 flange and the 3-blade rear rotor hub 31 are directly connected by a screw pile assembly; the 3-blade front rotor hub 23 is directly connected to the 3-blade rear rotor hub 31 through the front rotor hub and the rear rotor hub connecting bearing assembly 29, which will transmit all the loads of the rear rotor and all the loads of the front rotor except the torque; the differential The inner rotor shaft 41 of the planetary reducer 35 and the inner spline on the three-blade front rotor hub 23 are connected through the outer splines at both ends of the elastic outer spline shaft 28, and only torque is transmitted; at the same time, the power line control line riser assembly 37 also passes through the rear blade pitch change mechanism 30 and the front blade pitch change mechanism 25 from the center of the DC permanent magnet motor 36 and the differential planetary reducer 35, and is connected to the rear blade pitch electric pitch change actuator 21 and the front blade pitch electric pitch change actuator 22 in the front fairing 20 through the internal collecting ring assembly of the power line control line riser assembly 37; the front rotor blades 24 with three lower anti-swept winglet tips are respectively installed on the three-blade front rotor hub 23 through three groups of front blade bearing assemblies 26 and front blade pitch change pin disks 27; the front blade pitch electric pitch change actuator 22 pulls The front propeller pitch changing mechanism 25 moves back and forth, and the front propeller pitch changing mechanism 25 drives the front propeller pitch changing pin disk 27 to rotate clockwise or counterclockwise around the central axis of the front propeller bearing assembly 26 through the internal sliding groove of the front propeller pitch changing mechanism 25, thereby realizing the total pitch change adjustment of the front rotor blades; the rear rotor blades 32 with three downward-swept winglet tips are respectively installed on the three-blade rear rotor hub 31 through three groups of rear propeller bearing assemblies 33 and rear propeller pitch changing pin disks 34, and the rear pitch electric pitch changing actuator 21 pulls the rear propeller pitch changing mechanism 30 to move back and forth, and the rear propeller pitch changing mechanism 30 drives the rear propeller pitch changing pin disk 34 to rotate clockwise or counterclockwise around the central axis of the rear propeller bearing assembly 33 through the internal sliding groove of the rear propeller pitch changing mechanism 30, thereby realizing the total pitch change adjustment of the rear rotor blades.

[0015] Furthermore, the rear propeller pitch changing mechanism 30 includes: a pull rod 71, an inner pressure plate connecting screw pile assembly 72, an inner pressure plate 73, a bidirectional tension and pressure angular contact bearing 74, an outer pressure plate connecting screw pile assembly 75, an outer pressure plate 76, an outer ring three-pronged member 77 and a pull fork plate 78, wherein the three claws of the pull fork plate 78 respectively pass through three evenly distributed grooves provided on the elastic outer spline shaft 28, the three claws of the pull fork plate 78 are connected to the pull rod 71, and the pull fork plate 78 and the outer ring three-pronged member 77 are installed with a bidirectional tension and pressure angular contact bearing 74 via the inner pressure plate connecting screw pile assembly 72, the inner pressure plate 73, the outer pressure plate connecting screw pile assembly 75, and the outer pressure plate 76. The pressure angular contact bearing 74 ensures that the pull rod 71, the pull fork plate 78, the inner pressure plate connecting screw assembly 72, and the inner pressure plate 73 rotate along with the elastic external spline shaft 28 and the three-blade front rotor hub 23; at the same time, it ensures that the outer ring trident 77, the outer pressure plate connecting screw assembly 75, and the outer pressure plate 76 rotate in the opposite direction following the three-blade rear rotor hub 31, thereby achieving that they respectively follow the front and rear rotors in opposite rotation; and at the same time, it ensures that the pull rod 71 can pull the pull fork plate 78, the two-way pull pressure angular contact bearing 74 and the outer ring trident 77 to move forward and backward when following the opposite rotation of the front and rear rotors, thereby achieving the total pitch adjustment of the rear rotor blades.

[0016] Furthermore, the rear pitch electric pitch variable actuator 21 is connected to the front pitch electric pitch variable actuator 22, and the outer frame of the front pitch electric pitch variable actuator 22 is connected to the three-blade front rotor hub 23; the inner shaft of the front pitch electric pitch variable actuator 22 is connected to the flange of the front pitch variable mechanism 25, and the output flange of the rear pitch electric pitch variable actuator 21 is connected to the flange of the rear pitch variable mechanism 30 through a bolt assembly, and the straight pipe of the rear pitch variable mechanism 30 directly passes through the straight pipe of the front pitch variable mechanism 25; the front fairing 20 wraps the rear pitch electric pitch variable actuator 21 and the front pitch electric pitch variable actuator 22, and is directly connected to the three-blade front rotor hub 23.

[0017] Furthermore, the differential planetary reducer 35 includes: an inner rotor shaft 41, an outer rotor shaft 42, a driven inner ring gear 43, an intermediate planetary gear 44, an outer planetary gear 45, a double-tooth sun gear 46 and an elastic shaft 47; wherein, the differential planetary reducer has two degrees of freedom, the torque of the inner rotor shaft 41 and the outer rotor shaft 42 is equal, the inner rotor shaft 41 is connected to the front rotor hub 23, the outer rotor shaft 42 is connected to the rear rotor hub 31, one end of the elastic shaft 47 is connected to the output shaft of the DC permanent magnet motor 36; the other end of the elastic shaft 47 is connected to the double-tooth sun gear 46; the double-tooth sun gear 46 is meshed with the outer planetary gear 45; the outer planetary gear 45 is coaxially connected to the intermediate planetary gear 44; the driven inner ring gear 43 is meshed with the intermediate planetary gear 44.

[0018] Furthermore, the overall aerodynamic layout adopts a distributed 6-rotor + upper monoplane high lift-to-drag ratio wing + V-shaped tail.

[0019] Beneficial effects:

[0020] The present invention is based on the design concept and layout of multi-system and cross-system integration of the entire aircraft, and follows the design ideas of high safety, high reliability and appropriate redundancy. It innovatively integrates a distributed new power transmission chain configuration with high safety, high reliability and appropriate redundancy, a tilt-electric counter-rotating rotor unit and aerodynamic coupling analysis and optimization of the entire aircraft, a high cruise lift-to-drag ratio (16-18) wing, a low wind resistance fuselage, a low-noise and high-efficiency counter-rotating rotor design, a differential planetary reducer, and a green hybrid energy system into the tilt-rotor aircraft, eliminating the single-point failure mode of the power transmission chain for the first time. This is the first time in my country that a new configuration scheme of a distributed counter-rotating rotor hybrid tilt-rotor aircraft has been proposed.

[0021] Compared with typical helicopters, the new configuration of the distributed counter-rotating rotor hybrid tiltrotor aircraft eliminates the rotor system, transmission system and control system with automatic tilt device, which are large in size, complex in structure and have single-point failure mode, ensuring that the aircraft has a long range (over 3000km), high speed (cruise speed 520km / h, maximum speed 650km / h), high safety and high reliability (based on the helicopter reliability 10^ -5 Improved to distributed tiltrotor aircraft 10^ -6 ), low noise and vibration (cruising noise at 500m is about 45dB) and lower manufacturing, maintenance and use costs. BRIEF DESCRIPTION OF THE DRAWINGS

[0022] Figure 1 It is a distributed counter-rotating rotor hybrid tiltrotor aircraft appearance diagram;

[0023] Figure 2 The present invention is a structural diagram of a distributed counter-rotating rotor hybrid tiltrotor aircraft;

[0024] Figure 3 Another structural diagram of a distributed counter-rotating rotor hybrid tiltrotor aircraft;

[0025] Figure 4 It is a structural diagram of a distributed tilt-rotor electric counter-rotating rotor unit;

[0026] Figure 5 It is a diagram of the rear propeller pitch-changing mechanism of a distributed tilt-rotating electric counter-rotating rotor unit;

[0027] Figure 6 It is a simplified diagram of a differential planetary reducer for a distributed tilt-rotating electric counter-rotating rotor unit;

[0028] Figure 7 This is a simplified diagram of a green hybrid energy system for a distributed counter-rotating rotor hybrid tiltrotor aircraft;

[0029] Among them, the front tilt electric counter-rotating rotor nacelle 1, the front nacelle 2, the wingtip tilt electric counter-rotating rotor nacelle 3, the fuselage 4, the V-shaped tail 5, the V-shaped tail aileron 6, the vertical tail 7, the retractable landing gear 8, the wing 9, the aileron 10, the flap 11, the green hybrid energy system 12, and the tail tilt electric counter-rotating rotor nacelle 13;

[0030] Front fairing 20, rear pitch electric pitch-changing actuator 21, front pitch electric pitch-changing actuator 22, three-blade front rotor hub 23, front rotor blades at the tip of the reverse-swept winglet 24, front blade pitch-changing mechanism 25, front blade bearing assembly 26, front blade pitch-changing pin disc 27, elastic external spline shaft 28, bearing assembly 29 connecting the front rotor hub and the rear rotor hub, rear blade pitch-changing mechanism 30, three-blade rear rotor hub 31, rear rotor blades at the tip of the reverse-swept winglet 32, rear blade bearing assembly 33, rear blade pitch-changing pin disc 34, differential planetary reducer 35, DC permanent magnet motor 36, power line and control line riser assembly 37;

[0031] Inner rotor shaft 41, outer rotor shaft 42, driven inner ring gear 43, intermediate planetary gear 44, outer planetary gear 45, double-toothed sun gear 46, elastic shaft 47;

[0032] Turboshaft engine 51, FADEC electronic regulator 52, AC generator 53, controlled rectifier 54, integrated power conversion controller 55, aircraft integrated flight control system 56, bidirectional DC-DC module 57, motor controller 58, contra-rotating rotor 59, supercapacitor 60, lithium energy storage power battery 61, battery management system (BMS) 62, green biofuel tank 63;

[0033] Pull rod 71 , inner pressure plate connecting screw pile assembly 72 , inner pressure plate 73 , bidirectional tension and pressure angular contact bearing 74 , outer pressure plate connecting screw pile assembly 75 , outer pressure plate 76 , outer ring three-pronged piece 77 , and pull fork plate 78 . DETAILED DESCRIPTION

[0034] In order to make the purpose, technical solutions and advantages of the implementation of this application clearer, the technical solutions in the implementation of this application will be described in more detail below in conjunction with the drawings in the implementation of this application. In the drawings, the same or similar numbers throughout represent the same or similar elements or elements with the same or similar functions. The described implementation is a part of the implementation of this application, not all of the implementations. The implementation described below with reference to the drawings is exemplary and is intended to be used to explain this application, and should not be understood as a limitation on this application. Based on the implementation in this application, all other implementations obtained by ordinary technicians in this field without making creative work are within the scope of protection of this application. The implementation of this application is described in detail below in conjunction with the drawings.

[0035] In the description of the present invention, it should be understood that the terms "center", "axial", "vertical", "up", "down", "upper end", "bottom end", "inside", "outside", etc., indicating the orientation or position relationship, are based on the orientation or position relationship shown in the accompanying drawings, and are only for the convenience of describing the present invention and simplifying the description, rather than indicating or implying that the device or element referred to must have a specific orientation, be constructed and operated in a specific orientation, and therefore cannot be understood as limiting the scope of protection of the present invention.

[0036] The present invention relates to a new distributed power transmission chain structure, a tilting counter-rotating rotor and a green hybrid power energy system for use on a rotorcraft; the system can adapt to the layout of the new distributed power transmission chain structure, the tilting counter-rotating rotor and the hybrid power green energy system in the rotorcraft, and can also realize deceleration and torque increase, heat dissipation and cooling of high-speed and high-power-to-weight ratio aviation electric motors, and innovatively proposes a differential planetary reducer configuration for the counter-rotating rotor path and an efficient and low-noise downward-swept winglet tip three-blade front and rear rotor configuration aerodynamic design and a long-flight green hybrid power energy system. At the same time, the front and rear counter-rotating rotor pitch change technology is adopted to realize the counter-rotating rotor speed and torque control and a low-noise counter-rotating rotor electric drive unit technology and an ultra-long-flight hybrid power energy system.

[0037] The present invention is based on the design concept and layout of multi-system and cross-system integration of the entire aircraft, and follows the design ideas of high safety, high reliability and appropriate redundancy. Through in-depth research on the overall layout of the distributed counter-rotating rotor hybrid tilt-rotor aircraft and the configuration analysis of the distributed new power transmission chain with high safety and high reliability of the entire aircraft, the aerodynamic analysis of the counter-rotating rotor of the downward-swept winglet tip and its aerodynamic coupling analysis and optimization with the aerodynamics of the entire aircraft, the implementation method path of the low-noise and high-efficiency electric rotor and the construction of a green hybrid energy system, a new configuration scheme for a distributed counter-rotating rotor hybrid tilt-rotor aircraft is proposed.

[0038] Compared with typical helicopters, the new configuration of this distributed counter-rotating rotor hybrid tiltrotor aircraft eliminates the rotor system, transmission system and control system with automatic tilt device, which are huge in size, complex in structure and have single-point failure modes; compared with the tilt-rotor twin-rotor aircraft, the power transmission chain is only a purely mechanical twin-turboshaft engine, dual-path transmission system and dual rotors and dual control system. The design ensures that it only has self-protection capability in an emergency, with low reliability and lack of OEI mission continuation capability; the new configuration of this distributed counter-rotating rotor hybrid tiltrotor aircraft can continue to perform the mission when OEI occurs in all 6 sets of electric counter-rotating rotor drive units (one set of electric counter-rotating rotor drive unit fails), and its reliability is far superior to the previous two configuration aircraft (helicopters and tilt-rotor twin-rotors).

[0039] Its new configuration scheme consists of a new distributed power transmission chain structure with high safety, high reliability and appropriate redundancy, 6 sets of low-noise and high-efficiency tilting electric counter-rotating rotor units, high cruise lift-to-drag ratio (16-18) wings, low wind resistance fuselage and high power-to-weight ratio green hybrid energy system, etc., which ensure that the aircraft has a long range (over 3000km), high speed (cruising speed 520km / h, maximum speed 650km / h), high safety and high reliability (based on the helicopter reliability 10^ -5 Improved to distributed tiltrotor aircraft 10^ -6 ), low noise and vibration (cruising noise at 500m is about 45dB) and lower manufacturing, maintenance and use costs.

[0040] The distributed new power transmission chain architecture scheme is mainly as described later. According to the target of cruising speed of 520Km / h and maximum speed of 650Km / h, the aerodynamic shape and aerodynamic drag value of the whole aircraft are designed. The preliminary scheme of the upper monoplane wing with a lift-to-drag ratio of 16-18 in cruise state and the distributed 6 sets of tilt-rotating electric counter-rotating rotor units and their relative layout positions are designed. The aerodynamic coupling analysis and optimization of the distributed 6 sets of tilt-rotating electric counter-rotating rotor units and the whole aircraft layout, the relative position of the wings and the aerodynamic shape of the whole aircraft are carried out in the helicopter mode vertical take-off and landing, transition flight and fixed-wing mode cruise flight. The six tilting electric counter-rotating rotor units are basically evenly distributed in two large circles, with a counter-rotating rotor drive unit nacelle arranged on each side of the wing and the V-shaped tail wing. The left and right wings each have a front cabin close to the fuselage, with a counter-rotating rotor drive unit arranged in front of it. When a counter-rotating rotor drive unit suffers an OEI single-engine failure, the two adjacent counter-rotating rotor drive units will increase their power and thrust to keep the aircraft in stable flight and balance so that it can continue to perform its mission. The above forms the distributed new power transmission chain architecture and the combined configuration scheme of the whole aircraft aerodynamic design.

[0041] The aircraft's new configuration adopts a design concept of high safety, high reliability, and appropriate redundancy: each motor in the distributed six tilt-rotor electric counter-rotating rotor units consists of two sets of coils, and each set of motor coils is controlled by two independent motor controllers; each motor controller is powered by two independent battery packs; together, they form a high-reliability and high-safety power transmission chain consisting of six sets of tilt-rotor electric counter-rotating rotor units and four sets of isolated independent battery packs. There is no single point of failure in the power transmission chain, and the overall reliability of the aircraft is approximately 10^ -6 The reliability index far exceeds that of typical helicopters such as the twin-engine, single-rotor, tail-rotor helicopters AgustaW139 and Sikorsky S-70 Black Hawk. The rotor system, transmission system, and control system in the power transmission chain have no redundant design. Single-point failures can easily cause a Class I accident. The safety and reliability of the helicopter are not high, with a total reliability of approximately 10^ -5 .

[0042] In helicopter mode, by changing the collective pitch and rotation speed of the six sets of tilt-electric counter-rotating rotor units, the tilt-rotor aircraft can maintain stable flight, balance and change flight direction; in fixed-wing mode, by changing the thrust of the six sets of tilt-electric counter-rotating rotor units, the different opening degrees of the flaps and ailerons of the wings and the ailerons on the tail oblique wing and their combination, the tilt-rotor aircraft can maintain stable flight and change flight direction.

[0043] Each set of distributed tilt-electric counter-rotating rotor units includes a high-speed 23,000 rpm DC permanent magnet motor 36, a herringbone helical gear differential planetary reducer 35, a set of coaxial front and rear counter-rotating rotors 59 for changing the pitch, and a tilt mechanism.

[0044] The high efficiency of the tilt-rotor electric counter-rotating rotor unit is guaranteed by the design technologies such as CFD aerodynamic optimization design of rotor blade airfoil and its combination, the tip of the downward-swept winglet and the counter-rotating rotor configuration. First, the CFD aerodynamic design of the rotor blade airfoil and its relative position to the aerodynamic shape of the whole aircraft and the wing, coupled aerodynamic analysis optimization and experimental verification can ensure the high aerodynamic efficiency of the rotor and the whole aircraft; secondly, the tip of the downward-swept winglet can greatly reduce the tip vortex and greatly improve the efficiency; then the counter-rotating rotor configuration, differential planetary reducer and counter-rotating pitch rotor together The cooperation can realize the opposite direction of the front and rear propellers, equal speed and torque. It is a combined propeller with the advantages of high efficiency and torque balance. On the rotor propulsion axis, two rotors are arranged one in front and one in the back. The diameter of the front propeller is slightly larger than that of the rear propeller, and their rotation directions are opposite. Compared with a single rotor of the same size and power, the vortex energy in its wake cannot be utilized, while in the counter-rotating twin rotors, the vortex energy generated by the front propeller that is not effectively utilized can be utilized by the rear propeller. Considering the propulsion performance and torque balance, the spacing between the front and rear propellers of the counter-rotating rotors can be taken as L=0.2D p It is appropriate. After the joint action of the front and rear rotors, the axial speed of the slipstream is greater, and the vortex speed is significantly reduced, and the flow tends to flow axially. Therefore, the aerodynamic efficiency of the counter-rotating twin rotors is improved by about 13%-20% and the thrust is about twice that of a single propeller. Its external comprehensive torque is about 5% of that of a single propeller, which is almost negligible, greatly reducing the load level transmitted to the aircraft structure and the design difficulty.

[0045] The low-noise design measures for the pitch-shifting rotors adopt technologies such as low disc load, low blade tip linear speed and high-contact herringbone helical gear planetary reducer. First, the rotor tip linear speed in cruise state adopts a low linear speed scheme, which is about 110m / s, far lower than the 210m / s level of the helicopter main rotor blade tip linear speed, which can greatly reduce the rotor noise. Secondly, the low rotor disc load is about 60Kgf / m 2 , which is much lower than the 113.5Kgf / m of the classic tiltrotor V222 level, and can also greatly reduce rotor noise; the third pair of variable-pitch rotor blades adopts a high-rigidity design, and its rigidity is much greater than that of helicopter rotor blades; finally, the differential planetary reducer adopts high-overlap herringbone helical planetary technology to control the noise level of transmission gear meshing; through the above measures, the cruise noise of the distributed electric counter-rotating rotor hybrid tilt-rotor aircraft is expected to drop to about 45dB, which is significantly reduced compared with the current conventional helicopter cabin noise of 100dB and the noise of 120dB at 50m.

[0046] In helicopter mode, vertical take-off and landing (incoming air velocity is zero, rotor thrust is maximum, thrust, torque, efficiency) and transition flight (incoming air velocity is less than 50m / s, with a certain angle to the rotor) and cruising state in fixed-wing mode, finally, three-dimensional CFD aerodynamic design, aerodynamic coupling analysis and optimization of the efficient aerodynamic shape rotor blades and their relative position to the wings and the overall aircraft shape are adopted. With a maximum takeoff weight of 3200kg, the aircraft utilizes six distributed tilting electric counter-rotating rotor units with a rotor diameter of 3.35m, operating at 1100rpm (323kW transient power) and 910rpm (200kW steady-state power). In OEI mode, two adjacent units replace the original three. Steady-state motor power = 1.5*3200 / 6 / 4 = 200kW. A high-speed, 23,000rpm DC permanent magnet motor is required, achieving a power-to-weight ratio of 10-15kW / kg. The differential planetary reducer utilizes a high-contact herringbone helical gear configuration, offering a high power-to-weight ratio and low noise, achieving a power-to-weight ratio of approximately 15-20kW / kg. The motor and planetary reducer are integrated, undergoing design analysis and optimization for electromagnetics, thermal analysis, load-bearing, lubrication and cooling, forged magnesium alloy casing, combined oil injection and centrifugal oil passages, and gear and bearing oil film thickness verification, ensuring high reliability and safety.

[0047] The tilt mechanism enables each set of distributed electric tilt pairs to switch between the rotor unit and the fuselage structure from vertically upward in helicopter mode to horizontally forward in fixed-wing mode, with a maximum angle of approximately 95°; the tilt speed is given by the flight control system based on the flight stability of the transition flight state.

[0048] The long flight time and long range of the distributed counter-rotating rotor tilt-rotor aircraft are achieved by using a high lift-to-drag ratio of the wing cruise to generate lift for the entire aircraft with less power and the high efficiency of the green hybrid system. It decouples the direct mechanical connection between the turboshaft engine and the transmission system and rotor system in the helicopter power transmission chain, and combines the turboshaft engine, generator and electric motor (EM) in the new configuration separately and efficiently together to achieve a unique propulsion architecture, using higher energy density fuel and more efficient electricity conversion to achieve more efficient energy utilization.

[0049] The green hybrid energy system is composed of a renewable biofuel turbine generator and a high-safety lithium battery pack, a power electronics integrated power conversion controller, a high-voltage distribution system and other electric drive systems, which account for 31.3% of the total weight and weigh 1000Kg; 4 isolated independent battery packs, two by two, power the motor; the current maximum energy density that commercial batteries can achieve is 300w·h / kg, and the laboratory can reach 700w·h / kg. Compared with the energy density of fuel 12700w·h / kg, comprehensive calculation and analysis give the optimal solution, using a high-safety and reliable battery pack of 100Kwh, a discharge rate of 15C, and a weight of 333.4Kg. The remaining 666.6Kg is the weight of the bio-turbine shaft generator system. The total weight of the 500Kw-class turboshaft engine + engine + fuel system is about 300Kg, which can ensure that the maximum fuel volume is more than 350Kg of renewable biofuel, ensuring green and renewable recycling.

[0050] This aircraft avoids the mechanical rigidity of the powertrains of helicopters and tiltrotors, which prevents the turboshaft engine from balancing the high-power optimization requirements of the entire system in all states, such as vertical takeoff, vertical landing, and transition flight, with the low-power requirements of cruise flight. This results in excessive energy consumption and short range for both helicopters and tiltrotors. This aircraft configuration addresses these shortcomings of helicopters by leveraging the unique size independence of electric propulsion to ensure high-efficiency operation of the entire system (turboshaft engine, generator, electric motor, differential planetary reducer, and contra-rotating rotors) in high-power and high-thrust states, such as vertical takeoff and vertical landing, inclined climb, and slow-glide transition flight, while also ensuring high-efficiency operation of the entire system in low-power cruise flight. Furthermore, a wing with a high cruise lift-to-drag ratio of 16-18 is used to provide lift in forward flight. In fixed-wing mode, the tiltrotors are solely responsible for providing forward thrust, significantly reducing energy consumption in cruise mode.

[0051] This model uses 6 sets of tilting electric counter-rotating rotor units and a green energy system to replace the huge and complex rotor systems, transmission systems and control systems of helicopters and tiltrotors. It eliminates the complex rotor cycle control mechanism of helicopters and only retains the blade collective pitch change mechanism. The reliability of the entire aircraft is about 10 times higher. At the same time, the improved reliability, reduced vibration stress level and improved energy efficiency will greatly reduce the maintenance and use costs. Due to the 6-fold scale effect, its production cost will drop rapidly. The manufacturing, maintenance and use costs of this configuration of tiltrotor aircraft will be greatly reduced.

[0052] Through the above measures, the distributed counter-rotating rotor hybrid tilt-rotor aircraft eliminates the rotor system, transmission system and control system with automatic tilt device that are huge in size, complex in structure and have single-point failure mode, and adopts a distributed new power transmission chain structure with high safety and high reliability redundant design, 6 sets of tilting electric counter-rotating rotor units, high cruise lift-to-drag ratio (16-18) wings and green high power-to-weight ratio hybrid energy system to form a configuration scheme. It successfully solves the shortcomings of helicopters such as low safety and reliability, short range, high noise and vibration, high manufacturing cost, high maintenance cost and high operating cost, and ensures that the distributed counter-rotating rotor hybrid tilt-rotor aircraft has long range (more than 3000km), high speed (cruise speed 520km / h, maximum speed 650km / h), high safety and high reliability (derived from the helicopter reliability 10^ -5 Improved to distributed tiltrotor aircraft 10^ -6 ), low noise and low vibration (cruising noise at 500m is about 45dB), and the goal of lower manufacturing, maintenance and use costs.

[0053] The distributed 6-group counter-rotating rotor hybrid tilt-rotor configuration scheme is used as an example to illustrate the composition, installation method and innovation of the distributed counter-rotating rotor hybrid tilt-rotor aircraft. Figure 1 、 Figure 2 、 Figure 3 、 Figure 4 、 Figure 5 、 Figure 6 and Figure 7 .like Figure 1 、 Figure 2 and Figure 3As shown, the distributed counter-rotating rotor hybrid tilt-rotor aircraft includes a front tilt-electric counter-rotating rotor nacelle 1, a front cabin body 2, a wingtip tilt-electric counter-rotating rotor nacelle 3, a fuselage 4, a V-shaped tail 5, a V-shaped tail aileron 6, a vertical tail 7, a retractable landing gear 8, a wing 9, ailerons 10, a flap 11, a green hybrid energy system 12, and a tail tilt-electric counter-rotating rotor nacelle 13; wherein, one end of the front cabin body 2 is fixed to the front of the wing 9, the wing 9 is provided with ailerons 10 and flaps 11, the front tilt-electric counter-rotating rotor nacelle 1 is hinged to the other end of the front cabin body 2; a V-shaped tail aileron 6 is provided on the V-shaped tail 5, a vertical tail 7 is provided below the V-shaped tail 5, a tail tilt-electric counter-rotating rotor nacelle 13 is provided at the end of the V-shaped tail 5, and a wingtip tilt- Electric counter-rotating rotor nacelle 3; a retractable landing gear 8 is provided under the fuselage 4, and a green hybrid energy system 12 is used to power the motors in the front tilt electric counter-rotating rotor nacelle 1, the wingtip tilt electric counter-rotating rotor nacelle 3, and the tail tilt electric counter-rotating rotor nacelle 13. Rotor units are arranged in the wingtip tilt electric counter-rotating rotor nacelles 3 at the wingtips of the left and right wings, and rotor units are arranged in the front tilt electric counter-rotating rotor nacelle 1 in front of the left and right wings. Rotor units are arranged in the tail tilt electric counter-rotating rotor nacelles 13 on both sides of the V-shaped tail oblique wing. The six distributed rotor units all adopt a coaxial inverted upper and lower propeller configuration. The rotation direction of the rotors in the six rotor units is: adjacent and relative rotation directions are opposite, that is, the front left, middle right and rear left are clockwise, and the front right, middle left and rear right are counterclockwise. The entire aircraft adopts an overall aerodynamic layout of distributed 6 rotors + upper monoplane high lift-to-drag ratio wing + V-shaped tail, and the retractable landing gear reduces wind resistance in cruising state.

[0054] like Figure 4 As shown, the front tilt-rotor electric counter-rotating rotor nacelle 1, the wingtip tilt-rotor electric counter-rotating rotor nacelle 3, and the tail tilt-rotor electric counter-rotating rotor nacelle 13 of the distributed counter-rotating rotor hybrid tilt-rotor aircraft are mainly typically a tilt-rotor electric counter-rotating rotor unit, which includes: a front fairing 20, a rear pitch electric variable pitch actuator 21, a front pitch electric variable pitch actuator 22, a three-blade front rotor hub 23, and a front rotor with a downward reverse swept winglet blade tip. Blades 24, front propeller pitch control mechanism 25, front propeller bearing assembly 26, front propeller pitch control pin disc 27, elastic external spline shaft 28, bearing assembly 29 connecting the front rotor hub and the rear rotor hub, rear propeller pitch control mechanism 30, three-blade rear rotor hub 31, rear rotor blades 32 with downward reverse swept winglet tips, rear propeller bearing assembly 33, rear propeller pitch control pin disc 34, differential planetary reducer 35, DC permanent magnet motor 36, power line control line riser assembly 37.

[0055] The positioning boss at the output end of the DC permanent magnet motor 36 and the mounting inner hole of the differential planetary reducer 35 cooperate to ensure concentric positioning. The DC permanent magnet motor 36 is directly connected to the casing flange of the differential planetary reducer 35 by a screw pile group, and the output spline shaft of the DC permanent magnet motor 36 is connected to the elastic shaft 47 of the differential planetary reducer 35.

[0056] The outer rotor shaft 42 of the differential planetary reducer 35 is concentrically positioned through its flange boss and the inner hole of the 3-blade rear rotor hub 31, and then the outer rotor shaft 42 flange and the 3-blade rear rotor hub 31 are directly connected by a screw assembly; the 3-blade front rotor hub 23 is connected to the 3-blade rear rotor hub 31 through the front rotor hub and the rear rotor hub connecting bearing assembly 29, which transmits all the loads of the rear rotor and all the loads of the front rotor except the torque; its inner rotor shaft 41 and the 3-blade The internal spline on the front rotor hub 23 is connected through the external splines at both ends of the elastic external spline shaft 28, and only transmits torque; at the same time, the power line control line riser assembly 37 also passes through the rear blade pitch change mechanism 30 and the front blade pitch change mechanism 25 from the center of the DC permanent magnet motor 36 and the differential planetary reducer 35, and is connected to the rear blade pitch electric pitch change actuator 21 and the front blade pitch electric pitch change actuator 22 in the front fairing through the internal collector ring assembly of the power line control line riser assembly 37.

[0057] The front rotor blades 24 with three downward-swept winglet tips are respectively installed on the three-blade front rotor hub 23 through three sets of front propeller bearing assemblies 26 and front propeller pitch change pin discs 27; the front pitch electric pitch change actuator 22 pulls the front propeller pitch change mechanism 25 to move forward and backward, and the front propeller pitch change mechanism 25 drives the front propeller pitch change pin disc 27 to rotate clockwise or counterclockwise around the central axis of the front propeller bearing assembly 26 through its internal slide groove, thereby realizing the total pitch change adjustment of the front rotor blades.

[0058] Similarly, the three rear rotor blades 32 with downward-swept winglet tips are respectively installed on the three-blade rear rotor hub 31 through three sets of rear propeller bearing assemblies 33 and rear propeller pitch change pin discs 34. The rear pitch electric pitch change actuator 21 pulls the rear propeller pitch change mechanism 30 to move forward and backward. The rear propeller pitch change mechanism 30 drives the rear propeller pitch change pin disc 34 to rotate clockwise or counterclockwise around the central axis of the rear propeller bearing assembly 33 through its internal slide groove, thereby realizing the total pitch change adjustment of the rear rotor blades.

[0059] like Figure 5As shown, the rear propeller pitch change mechanism 30 includes a pull rod 71, an inner pressure plate connecting screw pile assembly 72, an inner pressure plate 73, a bidirectional tension and pressure angular contact bearing 74, an outer pressure plate connecting screw pile assembly 75, an outer pressure plate 76, an outer ring three-pronged member 77 and a pull fork plate 78. The three claws of the pull fork plate 78 pass through three evenly distributed sliding grooves provided on the elastic outer spline shaft 28 respectively, and the three claws of the pull fork plate 78 are connected to the pull rod 71. The pull fork plate 78 and the outer ring three-pronged member 77 are installed with a bidirectional tension and pressure angular contact bearing 74 through the inner pressure plate connecting screw pile assembly 72, the inner pressure plate 73, the outer pressure plate connecting screw pile assembly 75, and the outer pressure plate 76. The bidirectional tension and pressure angular contact bearing 74 ensures that the pull rod 71, the pull fork plate 78, the inner pressure plate connecting screw pile assembly 72, and the inner pressure plate 73 follow the elastic The outer spline shaft 28 and the three-blade front rotor hub 23 rotate; at the same time, the outer ring trident 77, the outer pressure plate connecting screw assembly 75, and the outer pressure plate 76 follow the three-blade rear rotor hub 31 to rotate in the opposite direction, so that they can follow the front and rear rotors in opposite rotation respectively; at the same time, it is ensured that the pull rod 71 can pull the pull fork plate 78, the two-way pull and pressure angular contact bearing 74 and the outer ring trident 77 to move forward and backward when following the front and rear rotors in opposite rotation, so as to realize the total pitch adjustment of the rear rotor blades.

[0060] The rear pitch electric pitch variable pitch actuator 21 is connected to the front pitch electric pitch variable pitch actuator 22, and the outer frame of the front pitch electric pitch variable pitch actuator 22 is connected to the three-blade front rotor hub 23; the inner shaft of the front pitch electric pitch variable pitch actuator 22 is connected to the flange of the front pitch variable mechanism 25, and the output flange of the rear pitch electric pitch variable pitch actuator 21 is connected to the flange of the rear pitch variable mechanism 30 through a bolt assembly, and the straight pipe of the rear pitch variable mechanism 30 directly passes through the straight pipe of the front pitch variable mechanism 25; the front fairing 20 wraps the rear pitch electric pitch variable pitch actuator 21 and the front pitch electric pitch variable pitch actuator 22, and is connected to the three-blade front rotor hub 23.

[0061] Through the above structure, the functions of electrically driving the front and rear rotors to rotate in opposite directions and changing the collective pitch of the front and rear rotor blades are realized.

[0062] like Figure 6 As shown, the differential planetary reducer 35 includes: an inner rotor shaft 41, an outer rotor shaft 42, a driven inner ring gear 43 connected to the outer rotor shaft 42, an intermediate planetary gear 44, an outer planetary gear 45, a double-tooth sun gear 46 and an elastic shaft 47; the differential planetary reducer has two degrees of freedom. In order to ensure that the working state of the motor remains unchanged, the torque of the inner and outer rotor shafts needs to be equal; the inner rotor shaft 41 is connected to the front rotor hub 23, the outer rotor shaft 42 is connected to the rear rotor hub 31, and the elastic shaft 47 is connected to the output shaft of the DC permanent magnet motor 36; the front and rear rotors both have matching pitch electric pitch change mechanisms, which work together with the differential planetary reducer 35 to achieve opposite steering, equal speed, and equal torque for the front and rear rotors.

[0063] like Figure 7 As shown, the green hybrid energy system 12 adopts a series-structured hybrid system, which consists of an aircraft integrated flight control system 56, a green biofuel tank 63, a turboshaft engine 51, a FADEC electronic regulator 52, an AC generator 53, an integrated power conversion controller 55, a motor controller 58, a supercapacitor 60, a lithium-ion energy storage battery 61, a battery controller BMS 62, etc.

[0064] The aircraft's integrated flight control system 56 controls the electric boost pump in the green biofuel tank 63 and the turboshaft engine 51 to generate pressure and suction, drawing the green biorenewable fuel in the green biofuel tank 63 into the turboshaft engine 51 for spraying, atomization, and mixed combustion. The fuel is converted into mechanical kinetic energy and output by the turboshaft to drive the aircraft's high power-to-weight ratio AC generator 53. The AC power generated by the generator is converted into DC power by the controlled rectifier 54 and bidirectional DC-DC module 57 in the integrated power conversion controller 55. The DC power is then charged into four independently separated lithium-ion energy storage power batteries 61 and supercapacitors 60 or directly fed into the motor controller 58 to drive the DC permanent magnet motor 36. At the same time, the DC power in the lithium-ion energy storage power battery 61 can also be converted into suitable DC power by the bidirectional DC-DC module 57 in the integrated power conversion controller 55, and then passed through the motor controller 58 to drive the DC permanent magnet motor 36. The DC power is then reduced in speed and torque by the differential planetary reducer 35, and the output forward and reverse counter-rotating power is output, driving the forward and reverse relative rotation of the front and rear rotors of the counter-rotating rotors 59.

[0065] In helicopter mode, the maximum motor power in vertical take-off, vertical landing, transition flight and other states is greater than the maximum power of the turboshaft engine 51 and the AC generator 53, which is about 2 times; in helicopter mode, when in high-power states such as vertical take-off, vertical landing, transition flight and other states, the turboshaft engine 51 and the AC generator 53 work at full load, and together with the lithium energy storage power battery 61 and the supercapacitor 60 with high-rate rapid discharge, they meet the maximum power and extreme state usage requirements.

[0066] When the distributed counter-rotor tilt-rotor aircraft is in high-power states such as vertical takeoff, vertical landing, oblique climb, and slow-down descent during transition flight in helicopter mode, the aircraft integrated flight control system 56 provides control signals to the FADEC electronic regulator 52, battery controller BMS 62, controlled rectifier 54, bidirectional DC-DC module 57, and motor controller 58, allowing the electric energy in the lithium energy storage power battery 61 to be converted into appropriate DC power through the bidirectional DC-DC module 57. At the same time, the turboshaft engine 51 in the green turboshaft generator system drives the AC generator 53 to generate AC power, which is then rectified into DC power by the controlled rectifier 54 in the integrated power conversion controller 55. The supercapacitor 60 can discharge quickly directly to the motor controller 58 to meet transient power requirements caused by transient airflow, etc. Together, power is supplied to the six DC permanent magnet motors 36 through the six motor controllers 58 to drive the six sets of tilt-rotor electric counter-rotating rotor units, generating maximum rotor thrust.

[0067] When in fixed-wing mode, the aircraft's integrated flight control system 56 sends control signals to the FADEC electronic regulator 52, battery controller BMS 62, controlled rectifier 54, bidirectional DC-DC module 57, and motor controller 58, allowing the turboshaft engine 51 and AC generator 53 in the green turboshaft generator system to operate at full load in the high-efficiency range to generate AC current, which is then converted into DC power by the controlled rectifier 54 in the integrated power conversion controller 55. The DC current is first supplied to the DC permanent magnet motor 36 through the motor controller 58 to drive the counter-rotating rotors 59 to rotate and drive the aircraft forward. The excess DC power can be directly charged into the supercapacitor 60 and converted into DC power of appropriate voltage by the bidirectional DC-DC module 57 in the integrated power conversion controller 55, and then charged into four independently separated lithium-ion energy storage power batteries 61. This replenishes the lithium-ion energy storage power batteries 61 after they have consumed energy during vertical takeoff and transition flight. This is the energy replenishment process of the lithium-ion energy storage power batteries 61.

[0068] When the lithium energy storage power battery 61 is fully charged, the aircraft integrated flight control system 56 gives control signals to the FADEC electronic regulator 52, battery controller BMS62, controlled rectifier 54, bidirectional DC-DC module 57 and motor controller 58; the green turbine shaft generator system is turned off, and the lithium energy storage power battery 61 is converted into a suitable voltage DC power through the bidirectional DC-DC module 57 in the integrated power conversion controller 55 and together with the super capacitor 60, it is supplied to the six DC permanent magnet motors 36 through the six motor controllers 58 to drive the six sets of tilt-rotating electric counter-rotating rotors 59, and the rotors generate forward flight thrust. This is the discharge working process of the lithium energy storage power battery 61.

[0069] Until the power of the lithium energy storage power battery (SOC) drops to a certain ratio, such as 20%, the green turboshaft generator system is started again to work at full load to perform the lithium energy storage power battery 61 and the supercapacitor 60 energy replenishment process. This process is repeated so that the green turboshaft generator system is always in the most efficient working area to burn and generate electricity, realizing the mechanical connection decoupling between the rotor and the green turboshaft generator system. They work in their respective high-efficiency areas within the full flight envelope, achieving the full flight range.

[0070] The present invention, "A Distributed Counter-Rotating Rotor Hybrid Tilt-Rotor Aircraft," has the following key features:

[0071] It is based on the design concept and layout of the whole aircraft multi-system cross-system integration, and in accordance with the design ideas of high safety, high reliability and appropriate redundancy, it innovatively proposes a high safety and high reliability distributed new power transmission chain configuration, tilt-rotating electric counter-rotating rotor unit, low noise and high efficiency rotor design, differential planetary reducer, green hybrid energy system and high cruise lift-to-drag ratio 16-18 high wing monoplane wing combination whole aircraft solution. Compared with helicopters and tilt-rotor aircraft, Figure 1 and Figure 2 The new configuration of the distributed electric counter-rotating rotor hybrid tiltrotor aircraft shown in the figure eliminates the large, complex rotor system, transmission system, and control system including the automatic tilter, which are characterized by single-point failure modes. The new aircraft configuration adopts a new distributed power transmission chain architecture with high safety, high reliability, and appropriate redundancy, and a distributed set of six low-noise electric counter-rotating rotor units. This constructs a high-reliability and high-safety power transmission chain that eliminates single-point failure modes, greatly improving the safety and reliability of the entire aircraft. The overall reliability of the aircraft is approximately 10^ -6 ; far exceeding the typical helicopter full-machine reliability of about 10^ -5 , and has the ability to continue to perform missions in OEI state; it has achieved the design of the entire aircraft's aerodynamic shape and aerodynamic wind resistance value according to the target of a cruising speed of 520Km / h and a maximum speed of 650Km / h, and designed a high-wing wing airfoil combination configuration scheme with a cruise lift-to-drag ratio of 16-18, ensuring the high speed of the aircraft.

[0072] It is based on the high efficiency of the rotor. The high efficiency of the counter-rotating rotor unit is guaranteed by relying on CFD aerodynamic optimization design of the rotor blade airfoil, the blade tip of the downward-swept winglet and the counter-rotating rotor configuration and other design technologies. The aerodynamic efficiency of the counter-rotating rotor is improved by about 13%-20% and the thrust is twice that of a single propeller. Its external comprehensive torque is about 5% of that of a single propeller, which greatly reduces the load level transmitted to the aircraft structure and the design difficulty, ensuring the high efficiency of the counter-rotating rotor.

[0073] It is based on the low-noise design concept of rotor blades. The low-noise measures for rotor blades first adopt the design technology of low disc load and low blade tip linear speed, and at the same time adopt the design technology of high-rigidity blades and high-overlap herringbone helical gear differential planetary reducer. The comprehensive result shows that the cruise noise of this distributed counter-rotating rotor tilt-rotor aircraft is about 45dB, which is greatly reduced compared with the current conventional helicopter cabin noise of 100dB and the noise of 120dB at 50m.

[0074] It is based on the construction of a counter-rotating rotor with equal rotation speed and adjustable pitch. Figure 5 The herringbone helical gear differential planetary reducer shown and Figure 4 The structure, movement and load transfer of the counter-rotating variable-pitch rotor assembly shown.

[0075] Its long flight time and large range based on the tilt-rotor aircraft are achieved through a green hybrid power system and a wing with a cruise lift-to-drag ratio of 16-18; the green hybrid power system decouples the direct mechanical connection between the turboshaft engine and the transmission system and rotor system, and efficiently combines the turboshaft engine, generator and electric motor (EM) to achieve a unique propulsion architecture, using higher energy density fuel and the turboshaft engine's more efficient electrical energy conversion within the full envelope to achieve more efficient energy utilization; the design has an excellent aerodynamic shape of the entire aircraft, low wind resistance value and a wing airfoil combination configuration with a cruise lift-to-drag ratio of 16-18. In fixed-wing mode, the wing provides lift for forward flight, and the tilt-rotor is only responsible for providing forward flight thrust, greatly reducing energy consumption in cruise state; the estimated range is more than 5,000 km.

[0076] Furthermore, this model replaces the huge and complex rotor systems, transmission systems and control systems of helicopters and tiltrotors with 6 sets of tilting electric counter-rotating rotor units and a green energy system, eliminates the complex rotor cycle control mechanism of helicopters, and only retains the blade collective pitch change mechanism. The reliability of the entire aircraft is about 10 times higher. At the same time, the improved reliability, reduced vibration stress level and improved energy efficiency will greatly reduce the maintenance and use costs. Due to the 6-fold scale effect, its production cost will drop rapidly. The manufacturing, maintenance and use costs of this configuration of tiltrotor aircraft will be greatly reduced.

[0077] The combination of the above key points ensures that the aircraft has long range (over 3000 km), high speed (cruising speed 520 km / h, maximum speed 650 km / h), high safety and high reliability (based on the helicopter reliability 10^ -5 Improved to distributed tiltrotor aircraft 10^ -6 ), low noise and vibration (cruising noise at 500m is about 45dB) and lower manufacturing, maintenance and use costs.

[0078] The above description is merely a specific embodiment of the present application, but the scope of protection of the present application is not limited thereto. Any changes or substitutions that can be easily conceived by a person skilled in the art within the technical scope disclosed in this application should be included in the scope of protection of the present application. Therefore, the scope of protection of the present application should be based on the scope of protection of the claims.

Claims

1. A distributed counter-rotating rotor hybrid tiltrotor aircraft, characterized in that: include: Front tilt electric counter-rotating rotor nacelle, front cabin body, wingtip tilt electric counter-rotating rotor nacelle, fuselage, V-shaped tail, V-shaped tail aileron, vertical tail, retractable landing gear, wing, aileron, flap, green hybrid energy system, tail tilt electric counter-rotating rotor nacelle, wherein one end of the front cabin body is fixed to the front of the wing, ailerons and flaps are provided on the wing, and the front tilt electric counter-rotating rotor nacelle is hinged to the other end of the front cabin body; V-shaped tail is provided with V-shaped tail aileron, a vertical tail is provided below the V-shaped tail, tail tilt electric counter-rotating rotor nacelle is provided at the end of the V-shaped tail, wingtip tilt electric counter-rotating rotor nacelle is provided at the end of the wing; retractable landing gear is provided below the fuselage. The green hybrid energy system is used to power the motors in the front tilt-rotor electric counter-rotating rotor nacelle, wingtip tilt-rotor electric counter-rotating rotor nacelle, and tail tilt-rotor electric counter-rotating rotor nacelle. Rotor units are arranged in the wingtip tilt-rotor electric counter-rotating rotor nacelles at the wingtips of the left and right wings, and rotor units are arranged in the front tilt-rotor electric counter-rotating rotor nacelles in front of the left and right wings. Rotor units are arranged in the tail tilt-rotor electric counter-rotating rotor nacelles on both sides of the V-shaped tail oblique wing. The six distributed rotor units all adopt a coaxial inverted upper and lower propeller configuration. The rotation direction of the rotors in the six rotor units is: adjacent and opposite rotation directions are opposite, that is, the front left, center right and rear left are clockwise, and the front right, center left and rear right are counterclockwise.

2. The distributed counter-rotating rotor hybrid tiltrotor aircraft according to claim 1, characterized in that: All six rotor units are tilt-rotating electric counter-rotating rotor units.

3. The distributed counter-rotating rotor hybrid tiltrotor aircraft according to claim 2, characterized in that: The green hybrid energy system adopts a series hybrid system.

4. The distributed counter-rotating rotor hybrid tiltrotor aircraft according to claim 3, characterized in that: The tilt-rotor electric counter-rotating rotor unit includes: a front fairing, a rear pitch electric pitch change actuator, a front pitch electric pitch change actuator, a three-blade front rotor hub, front rotor blades with lower reverse swept winglet tips, a front blade pitch change mechanism, a front blade bearing assembly, a front blade pitch change pin disc, an elastic external spline shaft, a bearing assembly connecting the front rotor hub and the rear rotor hub, a rear blade pitch change mechanism, a three-blade rear rotor hub, rear rotor blades with lower reverse swept winglet tips, a rear blade bearing assembly, a rear blade pitch change pin disc, a differential planetary reducer, a DC permanent magnet motor, and a power line control line riser assembly.

5. The distributed counter-rotating rotor hybrid tiltrotor aircraft according to claim 4, characterized in that: The positioning boss at the output end of the DC permanent magnet motor and the inner hole of the differential planetary reducer are matched to ensure concentric positioning. The DC permanent magnet motor and the casing flange of the differential planetary reducer are directly connected by a screw pile group, and the output spline shaft of the DC permanent magnet motor is connected to the elastic shaft of the differential planetary reducer; the outer rotor shaft of the differential planetary reducer is concentrically positioned by matching its flange boss and the inner hole of the 3-blade rear rotor hub, and then the outer rotor shaft flange and the 3-blade rear rotor hub are directly connected by a screw pile assembly; the 3-blade front rotor hub is connected to the 3-blade rear rotor hub through the front rotor hub and the rear rotor hub connecting bearing assembly, which will transmit all the loads of the rear rotor and all the loads of the front rotor except torque; the inner rotor shaft of the differential planetary reducer and the inner spline on the 3-blade front rotor hub are connected through the outer splines at both ends of the elastic outer spline shaft, and only torque is transmitted; at the same time, the power line control line riser assembly also passes through the rear blade pitch change mechanism and the front blade pitch change mechanism from the center of the DC permanent magnet motor and the differential planetary reducer, and passes through The internal collecting ring assembly of the power line control line riser assembly connects the rear pitch electric pitch change actuator and the front pitch electric pitch change actuator in the front fairing; the front rotor blades with three downward-swept winglet tips are respectively installed on the three-blade front rotor hub through three sets of front propeller bearing assemblies and front propeller pitch change pin discs; the front propeller pitch electric pitch change actuator pulls the front propeller pitch change mechanism to move forward and backward, and the front propeller pitch change mechanism drives the front propeller pitch change pin disc clockwise around the central axis of the front propeller bearing assembly through the internal slide groove of the front propeller pitch change mechanism. The rear rotor blades are mounted on the 3-blade rear rotor hub through 3 sets of rear propeller bearing assemblies and rear propeller pitch change pins respectively. The rear propeller pitch electric variable pitch actuator pulls the rear propeller pitch change mechanism to move forward and backward. The rear propeller pitch change mechanism drives the rear propeller pitch change pin to rotate clockwise or counterclockwise around the central axis of the rear propeller bearing assembly through the internal slide groove of the rear propeller pitch change mechanism to achieve total pitch adjustment of the rear rotor blades.

6. The distributed counter-rotating rotor hybrid tiltrotor aircraft according to claim 5, characterized in that: The rear propeller pitch changing mechanism includes: a pull rod, an inner pressure plate connecting screw pile assembly, an inner pressure plate, a bidirectional tension and pressure angular contact bearing, an outer pressure plate connecting screw pile assembly, an outer pressure plate, an outer ring trident and a pull fork plate, wherein the three claws of the pull fork plate respectively pass through three evenly distributed grooves provided on the elastic outer spline shaft, the three claws of the pull fork plate are connected to the pull rod, the pull fork plate and the outer ring trident are connected through the inner pressure plate connecting screw pile assembly, the inner pressure plate, the outer pressure plate connecting screw pile assembly, and the outer pressure plate installing a bidirectional tension and pressure angular contact bearing, the bidirectional tension and pressure angular contact bearing is both Ensure that the pull rod, pull fork plate, inner pressure plate connecting screw pile assembly, and inner pressure plate follow the rotation of the elastic external spline shaft and the three-blade front rotor hub; at the same time, ensure that the outer ring trident, outer pressure plate connecting screw pile assembly, and outer pressure plate follow the three-blade rear rotor hub to rotate in the opposite direction, so that they follow the opposite rotation of the front and rear rotors respectively; at the same time, ensure that the pull rod can pull the pull fork plate, bidirectional pull and pressure angular contact bearing, and outer ring trident to move forward and backward when following the opposite rotation of the front and rear rotors, so as to realize the total pitch adjustment of the rear rotor blades.

7. The distributed counter-rotating rotor hybrid tiltrotor aircraft according to claim 6, characterized in that: The rear pitch electric pitch variable actuator is connected to the front pitch electric pitch variable actuator, and the outer frame of the front pitch electric pitch variable actuator is connected to the three-blade front rotor hub; the inner shaft of the front pitch electric pitch variable actuator is connected to the flange of the front pitch variable mechanism, and the output flange of the rear pitch electric pitch variable actuator is connected to the flange of the rear pitch variable mechanism through a bolt assembly, and the straight pipe of the rear pitch variable mechanism directly passes through the straight pipe of the front pitch variable mechanism; the front fairing wraps the rear pitch electric pitch variable actuator and the front pitch electric pitch variable actuator, and is directly connected to the three-blade front rotor hub.

8. The distributed counter-rotating rotor hybrid tiltrotor aircraft according to claim 7, characterized in that: The differential planetary reducer includes: an inner rotor shaft, an outer rotor shaft, a driven inner ring gear, an intermediate planetary gear, an outer planetary gear, a double-tooth sun gear and an elastic shaft; wherein, the differential planetary reducer has two degrees of freedom, the torque of the inner rotor shaft and the outer rotor shaft are equal, the inner rotor shaft is connected to the front rotor hub, the outer rotor shaft is connected to the rear rotor hub, one end of the elastic shaft is connected to the output shaft of the DC permanent magnet motor; the other end of the elastic shaft is connected to the double-tooth sun gear; the double-tooth sun gear is meshed with the outer planetary gear; the outer planetary gear is coaxially connected to the intermediate planetary gear; the driven inner ring gear is meshed with the intermediate planetary gear.

9. The distributed counter-rotating rotor hybrid tiltrotor aircraft according to claim 8, characterized in that: The overall aerodynamic layout adopts a distributed 6-rotor + upper monoplane high lift-to-drag ratio wing + V-shaped tail.

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