WO2018209911A1 - 基于双涵道风扇动力系统的可折叠式固定翼垂直起降无人飞行器 - Google Patents

基于双涵道风扇动力系统的可折叠式固定翼垂直起降无人飞行器 Download PDF

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Publication number
WO2018209911A1
WO2018209911A1 PCT/CN2017/111341 CN2017111341W WO2018209911A1 WO 2018209911 A1 WO2018209911 A1 WO 2018209911A1 CN 2017111341 W CN2017111341 W CN 2017111341W WO 2018209911 A1 WO2018209911 A1 WO 2018209911A1
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Prior art keywords
wing
power system
aerial vehicle
foldable
ducted fan
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PCT/CN2017/111341
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English (en)
French (fr)
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裴海龙
程子欢
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华南理工大学
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Priority to SG11201910744WA priority Critical patent/SG11201910744WA/en
Priority to US16/613,930 priority patent/US11634222B2/en
Publication of WO2018209911A1 publication Critical patent/WO2018209911A1/zh

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    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64C—AEROPLANES; HELICOPTERS
    • B64C39/00—Aircraft not otherwise provided for
    • B64C39/02—Aircraft not otherwise provided for characterised by special use
    • B64C39/024—Aircraft not otherwise provided for characterised by special use of the remote controlled vehicle type, i.e. RPV
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64C—AEROPLANES; HELICOPTERS
    • B64C11/00—Propellers, e.g. of ducted type; Features common to propellers and rotors for rotorcraft
    • B64C11/001—Shrouded propellers
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64C—AEROPLANES; HELICOPTERS
    • B64C29/00—Aircraft capable of landing or taking-off vertically, e.g. vertical take-off and landing [VTOL] aircraft
    • B64C29/0008—Aircraft capable of landing or taking-off vertically, e.g. vertical take-off and landing [VTOL] aircraft having its flight directional axis horizontal when grounded
    • B64C29/0016—Aircraft capable of landing or taking-off vertically, e.g. vertical take-off and landing [VTOL] aircraft having its flight directional axis horizontal when grounded the lift during taking-off being created by free or ducted propellers or by blowers
    • B64C29/0025—Aircraft capable of landing or taking-off vertically, e.g. vertical take-off and landing [VTOL] aircraft having its flight directional axis horizontal when grounded the lift during taking-off being created by free or ducted propellers or by blowers the propellers being fixed relative to the fuselage
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64C—AEROPLANES; HELICOPTERS
    • B64C29/00—Aircraft capable of landing or taking-off vertically, e.g. vertical take-off and landing [VTOL] aircraft
    • B64C29/02—Aircraft capable of landing or taking-off vertically, e.g. vertical take-off and landing [VTOL] aircraft having its flight directional axis vertical when grounded
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64C—AEROPLANES; HELICOPTERS
    • B64C39/00—Aircraft not otherwise provided for
    • B64C39/06—Aircraft not otherwise provided for having disc- or ring-shaped wings
    • B64C39/068—Aircraft not otherwise provided for having disc- or ring-shaped wings having multiple wings joined at the tips
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64U—UNMANNED AERIAL VEHICLES [UAV]; EQUIPMENT THEREFOR
    • B64U10/00—Type of UAV
    • B64U10/20—Vertical take-off and landing [VTOL] aircraft
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64U—UNMANNED AERIAL VEHICLES [UAV]; EQUIPMENT THEREFOR
    • B64U30/00—Means for producing lift; Empennages; Arrangements thereof
    • B64U30/20—Rotors; Rotor supports
    • B64U30/26—Ducted or shrouded rotors
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64U—UNMANNED AERIAL VEHICLES [UAV]; EQUIPMENT THEREFOR
    • B64U60/00—Undercarriages
    • B64U60/40—Undercarriages foldable or retractable
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64U—UNMANNED AERIAL VEHICLES [UAV]; EQUIPMENT THEREFOR
    • B64U10/00—Type of UAV
    • B64U10/25—Fixed-wing aircraft
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64U—UNMANNED AERIAL VEHICLES [UAV]; EQUIPMENT THEREFOR
    • B64U30/00—Means for producing lift; Empennages; Arrangements thereof
    • B64U30/10—Wings
    • B64U30/12—Variable or detachable wings, e.g. wings with adjustable sweep
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64U—UNMANNED AERIAL VEHICLES [UAV]; EQUIPMENT THEREFOR
    • B64U30/00—Means for producing lift; Empennages; Arrangements thereof
    • B64U30/20—Rotors; Rotor supports
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64U—UNMANNED AERIAL VEHICLES [UAV]; EQUIPMENT THEREFOR
    • B64U50/00—Propulsion; Power supply
    • B64U50/10—Propulsion
    • B64U50/13—Propulsion using external fans or propellers
    • B64U50/14—Propulsion using external fans or propellers ducted or shrouded
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64U—UNMANNED AERIAL VEHICLES [UAV]; EQUIPMENT THEREFOR
    • B64U50/00—Propulsion; Power supply
    • B64U50/10—Propulsion
    • B64U50/19—Propulsion using electrically powered motors
    • Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
    • Y02—TECHNOLOGIES OR APPLICATIONS FOR MITIGATION OR ADAPTATION AGAINST CLIMATE CHANGE
    • Y02T—CLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION
    • Y02T50/00—Aeronautics or air transport
    • Y02T50/10—Drag reduction

Definitions

  • the invention relates to the technical field of aircrafts, and in particular to a foldable fixed-wing vertical take-off and landing unmanned aerial vehicle based on a double-ducted fan power system.
  • the tilting rotor is currently the most typical vertical take-off and landing scheme. By rotating the rotor shaft, it has the functions of a helicopter rotor and a fixed-wing propeller.
  • the most typical aircraft is the American V-22 "Osprey".
  • the tilting mechanism requires a complicated rotating mechanism design, and its structure is often subjected to damage caused by unfavorable factors such as the gyro alternating torque generated by the power rotating paddle/wing, and it is difficult to achieve stable aerodynamics in the tilting airfoil during low speed and transition. The effect is that there are a large number of accidents in the practical process and even crashes frequently (the V-22 Osprey crash reported frequently), while the utilization rate of the rotating mechanism is low, which becomes an unnecessary load during cruise flight, affecting the flight of the whole aircraft. performance.
  • auxiliary vertical lift system which is equipped with a vertical rotor or a power fan on a conventional fixed-wing aircraft, and directly generates a rotor pull force in the vertical direction to vertically take off and take off and control the attitude.
  • Fixed-wing and quad-rotor composite aircraft have become a hot topic in recent years due to their simple structure and ease of design and control.
  • Latitude Engineering LLC. of the United States is currently the world leader in composite four-rotor technology. Its HQ-20
  • the compound drone has a total weight of 11kg and can be loaded with a load of 0.9kg.
  • the maximum cruising speed is about 74km/h and the maximum cruising time is 15 hours.
  • the two sets of power systems are individually activated in the vertical takeoff and landing and horizontal flight modes, resulting in low efficiency of the whole power plant and affecting the flight time / distance / mobility of the aircraft.
  • the exposed rotor also has a large aerodynamic drag during level flight, making it difficult to achieve high-speed cruising flight.
  • the tailstock type vertical take-off and landing aircraft is another vertical take-off and landing scheme that is different from the tilting rotor. Unlike the tilting rotor, the tailstock type aircraft rotor is generally not tiltable, but directly uses the propeller power to vertically take off and land. In the air, relying on the flight control system to change the flight attitude to achieve the conversion of vertical takeoff and landing and cruise flight.
  • the tail-type vertical take-off and landing aircraft can effectively overcome the performance degradation caused by redundant power/tilting systems, and can maximize the use of airborne systems in both vertical take-off and horizontal flight. The disadvantage is that the flight state transition process is difficult to control and is disturbed by the wind.
  • the power system needs to achieve a total thrust-to-weight ratio of at least 1 or more, the blade has a large radius, a high rotational speed, and aerodynamic drag during level flight, and it is difficult to achieve a high cruising speed.
  • the flying wing structure has a large windward area during vertical take-off and landing and transition, and is greatly disturbed by gusts.
  • the object of the present invention is to solve the above-mentioned drawbacks of the prior art and to provide a foldable fixed-wing vertical take-off and landing unmanned aerial vehicle based on a double-ducted fan power system.
  • the aircraft includes a fuselage, a foldable wing 3, a ducted fan power system 7, and a retractable landing gear 9, the fuselage being divided into a nose 1, a front fuselage 2, a middle fuselage 5, and
  • the rear fuselage 6, the ducted fan power system 7 is symmetrically distributed on both sides of the rear fuselage 6 in a horizontal arrangement, the foldable wing 3 adopts an upper single wing layout, and passes through the wing folding shaft 4
  • Fixed to the front of the middle fuselage 5, the retractable landing gear 9 is placed at the front of the rear fuselage 6, the aircraft adopts a tailless layout, and the center of gravity of the aircraft is located at the rear of the front fuselage 2, the middle fuselage Before 5, the combination of the specific positional relationship between the duct and the wing achieved combination optimization.
  • the handpiece 1 is an electronic cabin for incorporating various sensors and optoelectronic devices;
  • the front fuselage 2 is a main load compartment for carrying main energy and load; and
  • the middle fuselage 5 is a secondary load.
  • the rear fuselage 6 is provided with a retractable landing gear at the front, the middle portion
  • the ducted fan power system 7 is symmetrically disposed on both sides, and the rear portion is a tapered rectifier.
  • the foldable wing 3 adopts a foldable configuration, and the wing is a two-stage folding wing, which can be folded 36°-180° to the belly along the longitudinal axis, and the trailing edge of the wing is placed near the wing tip.
  • the double-ducted fan power system 7 is symmetrically distributed on both sides of the rear fuselage 6 in a horizontal and tail-push layout, the number of which is 2, and the rotating shaft is located below the lower surface of the wing.
  • the ducted fan power system 7 includes: a duct body 10, a power fan 11, a fan driving mechanism 12, a control rudder surface 13, and a control rudder surface driving mechanism 14; wherein the power fan 11 is located in the culvert In the track body 10, the fan drive mechanism 12 is connected to the duct body 10; the control rudder surface 13 is located at the exit of the duct, and the number is 4, which is a "cross" type surrounding duct rotation axis; The rotation axis of the control rudder surface 13 is perpendicular to the ridge rotation axis, one end of which is connected to the duct body, and one end is connected to the control rudder surface driving mechanism 14 disposed in the duct body.
  • the aircraft adopts a tailless layout, and the whole machine has no conventional horizontal tail, vertical tail, elevator and rudder.
  • the number of the retractable landing gears 9 is four, and the length of the single landing gear can be adjusted in real time.
  • control rudder surface 13 is movable, and the attitude control torque is provided by deflecting the control rudder surface 13 to achieve stability and control of the flight attitude.
  • the present invention uses a ducted fan as a power system, and has the advantages that the outer profile of the duct optimizes the aerodynamic performance of the fan, the ducted wall blocks the formation of the fan tip vortex, and reduces the power loss of the fan tip, and simultaneously
  • the duct itself can generate lift under the action of fan suction. Therefore, compared with the isolated propeller, the ducted fan of the same radius has a higher lift-to-weight ratio (generally about 27% higher) at the same power consumption. At the same time, the duct can generate partial lift (about 10% of the lift of the wing) when the aircraft is flying horizontally.
  • ducted fans as the power system of the vertical take-off and landing aircraft can improve the aircraft's take-off, landing and flight efficiency, and effectively reduce energy consumption; single ducted fan power source meets multi-modal lift/thrust requirements, and high efficiency is suitable.
  • the ducted attitude control rudder surface is placed in the stable high-speed slipstream in the duct, reducing the external pneumatic operating surface (tail, etc.), avoiding the conventional design of low-speed (off-and-drop) aerodynamic disturbance failure instability factors and large
  • the gust disturbance of the exposed propeller; the ducted dynamic configuration does not require a tilting mechanism to increase system reliability, while the shape is simple and effective to reduce the radar reflection area.
  • the tail push type power layout ensures that the machine head is not interfered by the propeller, and it is convenient to install a variety of sensors and communication equipment; the tail push type ducted fan can obtain the maximum attitude (pitch, roll) control arm, which can maximize Improve its anti-disturbance ability.
  • the present invention uses a specific ducted-wing relative position to achieve combination optimization.
  • the ducted fan power system is located below the trailing edge of the wing.
  • the Kangda effect is generated in the vicinity of the wing, which effectively improves the back pressure gradient on the upper surface of the wing and reduces the airflow separation of the wing boundary layer.
  • the aerodynamic performance of the wing is greatly enhanced.
  • the lift coefficient of the wing is increased by 25%
  • the stall angle of attack is increased to 40°
  • the lift-to-resistance ratio of the whole machine is increased by 15%.
  • This technology improves the stability and security of the handover process. Therefore, the aircraft of the present invention can further reduce the power loss of the take-off and landing process and prolong the flight time, and can perform short-distance running and take-off and landing in a state where the extreme is not suitable for vertical take-off and landing.
  • the folding wing is folded up during vertical take-off and landing, which can reduce the windward windward area of the wing and enhance the wind resistance of the aircraft. When the plane is deployed, the wing can be deployed to obtain greater lift.
  • the ducted fan When the ducted fan is affected by the crosswind during operation, it will generate a momentum resistance at the entrance plane of the duct, which is the main resistance when the aircraft is flying at low speed. Since the center of gravity of the aircraft is relatively high (high position when placed vertically), the ducted fan and its control rudder surface are relatively rearward (lower position when placed vertically), and the whole machine is subjected to lateral wind resistance to gravity during vertical takeoff and landing and hovering. Form a low head torque. Therefore, in a strong crosswind environment, the aircraft can automatically reach the steady state with a small inclination angle. The mechanism of action enables the aircraft of the present invention to greatly increase its anti-disturbance capability during vertical take-off and landing. At the same time, the retractable landing gear can adjust the landing angle of the aircraft to match the aircraft to achieve anti-wind take-off and landing, further improving its stability.
  • Figure 1 is a three-dimensional schematic view of an aircraft of the present invention
  • Figure 2 (a) is a front view of the aircraft of the present invention (the landing gear is stowed);
  • Figure 2 (b) is a left side view of the aircraft of the present invention (the landing gear is stowed);
  • Figure 2 (c) is a side view of the aircraft of the present invention (the landing gear is stowed);
  • FIG. 3 is a schematic structural view of the ducted fan power system of the aircraft of the present invention.
  • Figure 4 (a) is a three-dimensional schematic view of the aircraft horizontal flying wing of the present invention when deployed;
  • Figure 4 (b) is a three-dimensional schematic view of the wing of the present invention when the aircraft is vertically taken up and down;
  • Figure 5 is a schematic diagram showing the multi-modal flight (vertical take-off and landing, horizontal flight) and transition process of the aircraft according to the present invention
  • Figure 6 (a) is a schematic diagram 1 of the optimized feature size of the ducted-wing combination of the present invention.
  • Figure 6 (b) is a schematic diagram 2 of the optimized feature size of the ducted-wing combination of the present invention.
  • Figure 7 is a schematic diagram of a system in accordance with an embodiment of the present invention.
  • FIG. 8(a) and Fig. 8(b) are comparison diagrams of the flow field CFD simulation of the ducted-wing combination optimization of the present invention and the conventional fixed wing at 40° angle of attack, wherein:
  • Figure 8 (a) is a flow field CFD simulation diagram of a conventional single wing stall
  • Fig. 8(b) is a CFD simulation diagram of flow field without stalling under the optimization of ducted-wing combination
  • Figure 9 is a schematic view showing the principle of the anti-side wind of the aircraft of the present invention.
  • a three-dimensional schematic diagram of each main component of the aircraft of the present embodiment includes: a fuselage, a foldable wing 3 , a ducted fan power system 7 , a retractable landing gear 9 , and the fuselage is divided into a nose.
  • the front fuselage 2, the middle fuselage 5 and the rear fuselage 6, the ducted fan power system 7 is symmetrically distributed on both sides of the rear fuselage 6 in a horizontal arrangement, and the foldable wing 3 is adopted.
  • the upper wing layout is placed at the front of the middle fuselage 5
  • the retractable landing gear 9 is placed at the front of the rear fuselage 6
  • the aircraft adopts a tailless layout, and the center of gravity of the aircraft is located at the front fuselage 2 In the rear part, the combination of the positional relationship between the duct and the wing is optimized.
  • the aerodynamic layout adopted by the aircraft of the present embodiment is: a foldable rectangular upper single-wing, a double-tailed ducted fan power system, and a retractable landing gear layout, as shown in Fig. 2(a)(b)(c).
  • the main dimensions of the whole machine include:
  • Total weight of the whole machine 20kg (including 5Kg payload)
  • the aircraft uses electric power, uses a motor as a power source, and a lithium battery serves as an energy source.
  • the head 1 is equipped with an electronic cabin and a load compartment 1 , and various sensors are installed, including an airspeed tube, a radar, a visible light/infrared camera, and an electronic compass.
  • the load compartment 2 In the front fuselage 2 of the embodiment, the load compartment 2, the main power battery and the airborne avionics system (including the sensor, the main control computer, the navigation flight control module, the communication module, the energy management module) are placed, and the main task load is mounted on the belly. This part is also where the weight of the whole machine is concentrated.
  • the fuselage 5 is placed with a sub-power battery, a folding mechanism, an operating mechanism of the landing gear, and a driving motor, and a secondary load is mounted on the belly, and the portion is a secondary weight concentration of the whole machine.
  • the foldable wing 3 of this embodiment is laid out with an upper single wing, a rectangular flat wing, and a Clark-Y airfoil to improve the medium speed performance (low speed performance is ensured by the wing/duct system design), the foldable type
  • the wing 3 is placed in the front of the middle fuselage 5 in an upper single-wing configuration in a foldable configuration.
  • the foldable wing 3 is a two-stage folding wing that can be folded 36° to 180° to the belly along the longitudinal axis.
  • Ailerons 8 are placed at the trailing edge of the wing near the wing tips.
  • the double-ducted fan power system 7 is symmetrically distributed on both sides of the rear fuselage 6 in a row-type and tail-push layout, and the number thereof is 2, and the rotating shaft is located below the lower surface of the wing.
  • the double-ducted fan power system 7 includes: a ducted body 10, a power fan 11, a fan drive mechanism 12 (this embodiment is a motor), a control rudder surface 13, and a control rudder surface drive mechanism 14 (this embodiment is an electric rudder) ,As shown in Figure 3.
  • the double-ducted fan power system 7 has a power fan 11 located in the duct, and uses a variable-width 4-blade fan to be connected to the duct body 10 through the fan drive mechanism 12; the control rudder surface 13 is located at the exit of the duct. 4, in a "cross" type around the ducted axis of rotation.
  • the rotating surface of the control rudder surface 13 is perpendicular to the rotating shaft of the duct. One end thereof is connected to the duct body, and one end is connected to the control rudder surface driving mechanism 14 disposed in the duct body.
  • the section of the duct body 10 adopts a specific streamline design, and the structure layout can improve the vertical take-off and landing performance, improve the hovering efficiency and the anti-disturbance capability, and the duct can also generate partial lift when leveling;
  • the structure of the single wing is affected by the vortex of the trailing edge of the wing and the position of the wing, even At 0° angle of attack, the duct can still generate partial lift (about 10% of the lift of the wing), thus improving the efficiency of the whole machine.
  • the movable control rudder surface 13 is placed at the exit of the duct, and the attitude control of the aircraft is realized by the tilting of the rudder surface.
  • the position of the center of gravity is placed in the area between the front fuselage and the leading edge of the wing according to the conventional fixed wing layout, and the control rudder surface 13 can generate a large control torque to the center of gravity, thereby obtaining excellent control performance of the aircraft.
  • This embodiment uses a specific ducted-wing relative position to achieve combined optimization.
  • the ducted suction can produce the Coanda effect at the trailing edge of the wing, slow the separation of the airflow in the boundary layer, increase the airfoil stall angle of attack, and generate a low pressure zone in the upper part of the wing to increase the lift coefficient of the wing.
  • the flow field CFD simulation of the aircraft of the present embodiment and the conventional fixed-wing aircraft flying at a speed of 30 m/s at an angle of attack of 40° is shown in Fig. 8, and the simulation results show that the configuration is compared with the conventional fixed-wing structure.
  • the lift coefficient is increased by 25%, the stall angle of attack is increased to 40°, and the lift-to-resistance ratio of the whole machine is increased by 15%.
  • the aircraft of this embodiment adopts a tailless layout. There are no conventional horizontal tails, vertical tails, elevators and rudders.
  • the attitude control torque is provided by the deflection control rudder surface 13 to achieve stability and control of the flight attitude.
  • the retractable landing gear 9 is placed at the front of the rear fuselage 6 in a number of four, and the length of the single landing gear can be adjusted in real time.
  • the foldable wing 3 is retracted during vertical take-off and landing, which can reduce the windward windward area of the wing and enhance the wind resistance of the aircraft; as shown in Fig. 4(a), When the plane is flying, the wing can be deployed to obtain a larger lift.
  • the aircraft of the present invention adopts a vertical attitude when the ground stops, and the landing gear is lowered.
  • the generated lift is vertically upward, thereby achieving vertical takeoff and landing of the aircraft.
  • the foldable wing 3 is stowed to reduce the windward area of the side wind of the aircraft during take-off and landing, and to improve its anti-disturbance capability.
  • the attitude of the aircraft is controlled by controlling the deflection of the rudder surface 13 to generate a control torque.
  • the lift generated by the ducted fan power system 7 will produce a component in the horizontal direction, thereby achieving horizontal flight of the aircraft.
  • the foldable wing 3 When the flight attitude and speed reach a certain range, the foldable wing 3 is deployed, and the fuselage enters a horizontal flight state. In this state, the characteristics of the aircraft are similar to those of the conventional fixed-wing aircraft, and the speed can be higher and lower. Energy consumption for cruising flight.
  • G gravity
  • T ducted lift
  • Fd momentum resistance
  • F control surface control force
  • ⁇ wind resistance balance inclination angle.
  • the anti-disturbance mechanism of the aircraft of the present invention in crosswind environment is :
  • the ducted power fan 11 can generate a momentum resistance Fd at the vicinity of the inlet of the duct body 10 in the crosswind, which is the main resistance under the lifting and lowering conditions, and the present invention adopts a high (front) center of gravity position design, Momentum resistance produces a head moment on the center of gravity, and the duct can automatically tilt the angle ⁇ to the wind and reach a state of self-stability;
  • the control rudder surface 13 is placed in the high-speed slipstream of the duct, and is less interfered by the flight condition, and can generate a stable control force F, which has a large torque on the center of gravity of the aircraft.
  • F a stable control force
  • the duct can reach a steady state under the crosswind at a small dip angle and keep the attitude within a certain range.
  • the maximum equilibrium angle of the aircraft of the present invention is 14.8.
  • the landing gear is extended and contracted with the angle of inclination of the aircraft, so that the landing plane of the landing gear is always parallel with the landing platform.
  • the maximum adjustable angle of the landing gear plane is 25°
  • the maximum required wind resistance angle of the aircraft is 15 °, so the aircraft can achieve a sloping landing.

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Abstract

一种基于双涵道风扇动力系统的可折叠式固定翼垂直起降无人飞行器，采用置于机身尾部、横列式、尾推布局的双涵道风扇动力系统（7），为飞行器提供垂直起降的升力和水平飞行的推力；通过偏转置于涵道出口处的控制舵面提供矢量推力，实现快速姿态转换；机翼采用折叠翼构型（3），飞行器垂直起降／低速飞行时机翼折叠以减小侧风迎风面积，水平飞行时机翼展开以获得较大升力；涵道一机翼组合优化，机翼置于特定涵道气流区内，涵道抽吸在机翼后缘产生康达效应，以提高机翼的性能。该飞行器可垂直起降与高速巡航等多模态飞行作业；垂直起降飞行器悬停／低速飞行时的气动效率高；起降／悬停抗扰动能力强；能耗低、噪声小、安全可靠性高。

Description

基于双涵道风扇动力系统的可折叠式固定翼垂直起降无人飞行器 技术领域
本发明涉及飞行器技术领域，具体涉及一种基于双涵道风扇动力系统的可折叠式固定翼垂直起降无人飞行器。
背景技术
垂直起降飞行器一直都是航空工程研究的热点，近年来，由于材料、能源、动力和控制技术的进步，许多先进的方案被应用于垂直起降飞行器上。目前主要的几类可实现舰载垂直起降飞行器构型设计主要有：倾转动力机构(含倾转翼等)、辅助垂直升力和尾座式结构。
(1)倾转旋翼是目前最为典型的一种垂直起降方案，通过旋转旋翼轴使其兼具直升机旋翼和固定翼螺旋桨的功能，最典型的飞机是美国的V-22“鱼鹰”。然而，倾转机构需要复杂的转动机构设计，其结构往往要承受动力旋转桨/翼产生的陀螺交变力矩等不利因素产生损伤，而且在低速和过渡过程中倾转翼面难以达到稳定的气动效果，目前实用过程中出现了大量的事故甚至经常坠毁(V-22鱼鹰坠毁的报道时有发生)，同时转动机构利用率低，在巡航飞行时成为了不必要的负载，影响了整机飞行性能。
(2)另一种可行的方案是采用辅助垂直升力系统，在传统固定翼飞机上加装垂直方向上的旋翼或动力风扇，依靠垂直方向上直接产生旋翼拉力来垂直起降和控制姿态。固定翼与四旋翼复合型飞机由于其结构简单，易于设计和控制，近年来成为了一热点。美国Latitude Engineering LLC.公司是目前世界复合四旋翼技术领域领先者。其HQ-20 复合无人机全机重量11kg，可以装载0.9kg载荷。最大巡航速度达到约74km/h，最大续航时间15小时。辅助垂直升力系统固定翼飞行器，其两套动力系统在垂直起降和水平飞行模态中各自单独启用，造成了整机动力装置效率低下从而影响飞行器的飞行时间/距离/机动性等作战性能，同时暴露的旋翼也在平飞时有较大的气动阻力，难以实现高速巡航飞行。
(3)尾座式垂直起降飞行器是区别于倾转旋翼的另一种垂直起降方案，与倾转旋翼不同，尾座式飞行器旋翼一般不可倾转，而是直接使用螺旋桨动力垂直起降，在空中依靠飞行控制系统变化飞行姿态达到垂直起降与巡航飞行的转换。采用尾座式的垂直起降飞行器能有效克服冗余动力/倾转系统带来的性能下降，在垂直起降和水平飞行时均能最大限度地利用机载系统。其缺点在于飞行状态过渡过程控制困难，受风扰动较大。另外，其动力系统由于需要达到至少1以上的全机推重比，桨盘半径较大，转速较高，平飞时气动阻力大，难以达到较高的巡航速度。同时飞翼结构在垂直起降和过渡过程中有较大的迎风面积，受阵风干扰大。
现有的垂直起降飞行器构型大都面临着起降过程中操作翼面气动作用小因而易受不稳定气流影响，暴露在外的大尺寸的旋转翼同样也会受到阵风的作用(同时机头安装往往影响电子舱通讯与传感设备)；倾转动力机构复杂易损控制稳定性差，辅助冗余动力效率低不适合大载荷长航程设计。
发明内容
本发明的目的是为了解决现有技术中的上述缺陷，提供一种基于双涵道风扇动力系统的可折叠式固定翼垂直起降无人飞行器。
本发明的目的可以通过采取如下技术方案达到：
一种基于双涵道风扇动力系统的可折叠式固定翼垂直起降无人飞行 器，所述飞行器包括机身、可折叠式机翼3、涵道风扇动力系统7、可伸缩式起落架9，所述机身分为机头1、前机身2、中机身5与后机身6，所述涵道风扇动力系统7采用横列式布局对称分布于所述后机身6两侧，所述可折叠式机翼3采用上单翼布局，并通过机翼折叠轴4固定于所述中机身5前部，所述可伸缩式起落架9置于所述后机身6前部，飞行器采用无尾式布局，飞机重心位于前机身2后部、中机身5之前，涵道与机翼之间采用特定位置关系实现组合优化。
进一步地，所述机头1为电子舱，用于内置多种传感器和光电设备；所述前机身2为主负载舱，用于搭载主要能源和载荷；所述中机身5为次负载舱，用于搭载航电系统、次要能源、机翼折叠轴4的驱动机构、可伸缩式起落架9的驱动机构；所述后机身6的前部置有可伸缩式起落架，中部两侧对称安置所述涵道风扇动力系统7，后部为锥形整流体。
进一步地，所述可折叠式机翼3采用可折叠式构型，机翼为二段式折叠翼，可沿纵向轴线向机腹折叠36°～180°，机翼后缘靠近翼梢处安置有副翼8。
进一步地，所述双涵道风扇动力系统7采用横列式、尾推布局对称分布于所述后机身6两侧，其数量为2，其旋转轴位于机翼下表面以下。
进一步地，所述涵道风扇动力系统7包括：涵道体10、动力风扇11、风扇驱动机构12、控制舵面13、控制舵面驱动机构14；其中，所述动力风扇11位于所述涵道体10中，通过所述风扇驱动机构12与所述涵道体10相连接；所述控制舵面13位于涵道出口处，数量为4，呈“十字”型环绕涵道旋转轴；所述控制舵面13的旋转轴与涵道旋转轴垂直，其一端与涵道体相连，一端连接置于涵道体中的所述控制舵面驱动机构14。
进一步地，所述涵道与机翼之间实现组合优化的特定相对位置关系满足：
所述可折叠式机翼3的后缘距涵道入口平面距离l1与涵道入口直径d关系为：
0.35d≤l1≤0.45d；
所述可折叠式机翼3的弦线平面距涵道中心轴线距离l2与涵道入口直径d关系为：
0.25d≤l2≤0.4d。
进一步地，所述飞行器采用无尾式布局，全机无常规水平尾翼、垂直尾翼、升降舵与方向舵。
进一步地，所述可伸缩式起落架9的数量为4，单个起落架长度可实时调节。
进一步地，所述控制舵面13为可活动的，通过偏转所述控制舵面13提供姿态控制力矩，实现飞行姿态的稳定与控制。
本发明相对于现有技术具有如下的优点及效果：
1、本发明使用涵道风扇作为动力系统，其优势在于：涵道的外廓作用优化了风扇的气动性能、涵道壁阻挡了风扇桨尖涡的形成、降低了风扇桨尖动力损失、同时涵道自身能在风扇抽吸作用下产生升力。因此，与孤立螺旋桨相比，相同半径的涵道风扇在相同能耗功率的情况下其升力重量比更高(一般高出27％左右)。同时，涵道在飞行器水平飞行时能产生部分升力(占机翼升力10％左右)。因此，采用涵道风扇作为垂直起降飞行器的动力系统能提高飞行器的起降、悬停和飞行效率，有效降低能耗；单一涵道风扇动力源满足多模态升力/推力要求，效率高适合垂直起降/远距离飞行；同时涵道风扇系统气动噪声小，安全性，可靠性高。
2、涵道姿态控制舵面置于涵道内稳定高速滑流中，减少外部气动操作面(尾翼等)，避免常规设计低速(起降)气动扰动失效不稳定因素以及大 尺寸外露螺旋桨的阵风扰动；涵道动力构型无需倾转机构增加系统可靠性，同时外形简洁有效减少雷达反射面积。
3、尾推式动力布局保证了机头不受螺旋桨干扰，可方便安装多种传感器及通讯设备；尾推式涵道风扇能够获得最大的姿态(俯仰、滚转)控制力臂，能够最大限度地提升其抗扰动能力。
4、本发明采用特定的涵道-机翼相对位置，实现组合优化。涵道风扇动力系统位于机翼后缘下方，通过涵道风扇的抽吸作用，在机翼附近产生康达效应，有效改善机翼上表面的逆压梯度，减缓机翼附面层气流分离，使得机翼气动性能大幅增强。与传统固定翼结构相比，机翼的升力系数提高25％，失速迎角提高至40°，全机升阻比提高15％。该技术提高了切换过程的稳定安全性。因此本发明飞行器可进一步减少起降过程的动力损耗延长飞行时间，在极端恶劣不适合垂直起降的状态下可进行短距滑跑起降。
5、折叠翼在垂直起降时收起，能减小机翼的侧风迎风面积，增强飞行器抗风能力,在平飞时机翼展开，能够获得较大的升力。
6、涵道风扇在运转时受侧风影响会在涵道入口平面处产生一动量阻力，该阻力在飞行器低速飞行时为主要阻力。由于飞行器重心位置较为靠前(垂直放置时为高位)，涵道风扇及其控制舵面较为靠后(垂直放置时为低位)，全机在垂直起降和悬停时受侧风阻力对重力形成一低头力矩。因此，在较强侧风环境下，飞行器能够以小倾角自动迎风达到稳定状态。该作用机理使得本发明飞行器能在垂直起降时大大提高其抗扰动能力。同时可伸缩式起落架可调整飞行器着陆角度，以配合飞行器实现倾斜抗风起降，进一步提高其稳定性。
附图说明
附图用来提供对本发明的进一步理解，并且构成说明书的一部分，与本发明的实施例共同用于解释本发明，并不构成对本发明的限制。在附图中：
图1是本发明飞行器的三维示意图；
图2(a)是本发明飞行器的主视图(起落架收起)；
图2(b)是本发明飞行器的左视图(起落架收起)；
图2(c)是本发明飞行器的侧视图(起落架收起)；
图3是本发明飞行器之所述涵道风扇动力系统的结构示意图；
图4(a)是本发明飞行器水平飞行机翼展开时的三维示意图；
图4(b)是本发明飞行器垂直起降时机翼折叠时的三维示意图；
图5是本发明所述飞行器多模态飞行(垂直起降、水平飞行)及其过渡过程转换示意图；
图6(a)是本发明所述涵道—机翼组合优化特征尺寸示意图1；
图6(b)是本发明所述涵道—机翼组合优化特征尺寸示意图2；
图7是本发明实施例的系统示意图；
图8(a)和图8(b)是本发明所述涵道—机翼组合优化与传统固定翼在40°迎角飞行时的流场CFD仿真比较图，其中：
图8(a)为传统单独机翼发生失速的流场CFD仿真图；
图8(b)为涵道—机翼组合优化下无失速的流场CFD仿真图；
图9是本发明飞行器抗侧风原理示意图。
具体实施方式
为使本发明实施例的目的、技术方案和优点更加清楚，下面将结合本发明实施例中的附图，对本发明实施例中的技术方案进行清楚、完整地描述，显然，所描述的实施例是本发明一部分实施例，而不是全部的实施例。 基于本发明中的实施例，本领域普通技术人员在没有做出创造性劳动前提下所获得的所有其他实施例，都属于本发明保护的范围。
实施例
如图1所示，本实施例飞行器各主要部件三维示意图，包括：机身、可折叠式机翼3、涵道风扇动力系统7、可伸缩式起落架9，所述机身分为机头1、前机身2、中机身5与后机身6，所述涵道风扇动力系统7采用横列式布局对称分布于所述后机身6两侧，所述可折叠式机翼3采用上单翼布局置于所述中机身5前部，所述可伸缩式起落架9置于所述后机身6前部，飞行器采用无尾式布局，飞机重心位于所述前机身2后部，涵道与机翼之间采用特定位置关系实现组合优化。
本实施例飞行器采用的气动布局为：可折叠式矩形上单翼、双尾推涵道风扇动力系统、可伸缩式起落架布局，如图2(a)(b)(c)所示。全机主要尺寸包括：
机翼展长：1.5m
展弦比：7.5
机身尺寸：0.16m×0.16m×1.2m
涵道外径：0.33m
风扇半径：0.116m采用可变距4叶桨风扇
全机总重：20kg(含5Kg有效载荷)
飞行时间：1h
全机主要尺寸和系统分布如图7所示。
本实施例飞行器采用电动力，使用电机作为动力源，锂电池作为能量源。
本实施例机头1部分搭载电子舱及载荷舱1，安置多种传感器，包括空速管，雷达，可见光/红外摄像头及电子罗盘。
本实施例前机身2放置载荷舱2、主动力电池和机载航电系统(包括传感器、主控计算机、导航飞控模块、通讯模块、能源管理模块)，机腹挂载主要任务载荷,该部分也是全机重量集中之处。
本实施例中机身5放置副动力电池,折叠机构、起落架的作动机构和驱动电机，机腹挂载次要载荷，该部分是全机次要重量集中之处。
本实施例可折叠式机翼3布局采用上单翼、矩形平直翼、Clark-Y翼型以提高其中速性能(低速性能利用机翼/涵道组合系统设计保证)，所述可折叠式机翼3采用上单翼布局置于中机身5前部，采用可折叠式构型。可折叠式机翼3为二段式折叠翼，可沿纵向轴线向机腹折叠36°～180°。机翼后缘靠近翼梢处安置有副翼8。
本实施例双涵道风扇动力系统7采用横列式、尾推布局对称分布于后机身6两侧，其数量为2，其旋转轴位于机翼下表面以下。双涵道风扇动力系统7包括：涵道体10、动力风扇11、风扇驱动机构12(本实施例为电机)、控制舵面13、控制舵面驱动机构14(本实施例为电动舵机)，如图3所示。
所述双涵道风扇动力系统7，其动力风扇11位于涵道中，采用可变距4叶风扇，通过风扇驱动机构12与涵道体10相连接；控制舵面13位于涵道出口处，数量为4，呈“十字”型环绕涵道旋转轴。控制舵面13旋转轴与涵道旋转轴垂直，其一端与涵道体相连，一端连接置于涵道体中的控制舵面驱动机构14。
所述涵道体10剖面采用特定流线形设计，采用该结构布局能提高垂直起降的性能，提高悬停效率和抗扰动能力，同时涵道在平飞时也能产生部分升力；由于上单翼的结构，受机翼后缘脱体涡及机翼位置影响，即使 在0°迎角的情况下涵道仍能产生部分升力(占机翼升力10％左右)，从而提高整机效率。可活动的控制舵面13置于涵道出口处，通过舵面倾转实现飞行器的姿态控制。重心位置按常规固定翼布局置于前机身与机翼前缘之间的区域，控制舵面13能对重心产生较大控制力矩，从而使飞行器获得优秀的控制性能。
本实施例采用特定涵道—机翼相对位置实现组合优化。如图6(a)、(b)所示，本实施例机翼与涵道相对位置采用l1＝0.4d、l2＝0.35d。涵道抽吸能够在机翼后缘产生康达效应，减缓附面层气流分离，增大翼型失速迎角，同时在机翼上部产生低压区，提高机翼升力系数。本实施例飞行器与传统固定翼飞行器在40°迎角以30m/s速度飞行时的流场CFD仿真如,8所示，仿真结果表明，该构型与传统固定翼结构相比，机翼的升力系数提高25％，失速迎角提高至40°，全机升阻比提高15％。
本实施例飞行器采用无尾式布局。全机无常规水平尾翼、垂直尾翼、升降舵与方向舵。通过偏转控制舵面13提供姿态控制力矩，实现飞行姿态的稳定与控制。
所述可伸缩式起落架9置于后机身6前部，数量为4，单个起落架长度可实时调节。
本发明的工作原理及过程：
如图4(b)所示，可折叠式机翼3在垂直起降时收起，能减小机翼的侧风迎风面积，增强飞行器抗风能力；如图4(a)所示，在平飞时机翼展开，能够获得较大的升力。
如图5所示，本发明飞行器在地面停靠时采用竖直姿态，起落架放下。当涵道风扇运转时，产生的升力竖直向上，从而实现飞行器的垂直起降。 同时可折叠式机翼3收起，减小起降时飞行器对侧风的迎风面积，提高其抗扰动能力。通过控制舵面13的偏转产生控制力矩控制飞行器的姿态。当飞行器脱离竖直姿态时，涵道风扇动力系统7产生的升力将在水平方向上产生分量，从而实现飞行器的水平飞行。当飞行姿态与速度达到一定范围后，可折叠式机翼3展开，机身进入水平飞行状态，在该状态下，飞行器各项特性与传统固定翼飞机类似，可以较高的速度和较低的能耗进行巡航飞行。
如图9所示，图中G为重力，T为涵道升力，Fd为动量阻力，F为舵面控制力，α为抗风平衡倾角，本发明飞行器在侧风环境下的抗扰动机制为：
a.涵道风扇对侧风的稳定作用。涵道的动力风扇11在侧风中能够在涵道体10近入口上方处产生一动量阻力Fd，该阻力在起降工况下为主要阻力，本发明采用高(前)重心位置设计，使得动量阻力对重心产生低头力矩，涵道能够自动迎风倾斜α角度，达到自稳状态；
b.控制舵面13置于涵道高速滑流中，受飞行状况干扰较小，能够产生稳定控制力F，该力对飞行器重心有较大的力矩。通过舵面控制，使得涵道能够在侧风下以小倾角迎风达到稳定状态，并保持该姿态在一定范围内。对于不大于16m/s的侧风，本发明飞行器最大平衡角度为14.8°。
c.起落架配合飞行器倾斜角度进行伸缩，使得起落架着陆平面始终与起降平台保持平行，本实施例中，起落架平面可调节的最大角度为25°，飞行器最大需要的抗风角度为15°，因此飞行器可以实现倾斜降落。
上述实施例为本发明较佳的实施方式，但本发明的实施方式并不受上 述实施例的限制，其他的任何未背离本发明的精神实质与原理下所作的改变、修饰、替代、组合、简化，均应为等效的置换方式，都包含在本发明的保护范围之内。

Claims (9)

  1. 一种基于双涵道风扇动力系统的可折叠式固定翼垂直起降无人飞行器，所述飞行器包括机身、可折叠式机翼(3)、涵道风扇动力系统(7)、可伸缩式起落架(9)，所述机身分为机头(1)、前机身(2)、中机身(5)与后机身(6)，其特征在于，
    所述涵道风扇动力系统(7)采用横列式布局对称分布于所述后机身(6)两侧，所述可折叠式机翼(3)采用上单翼布局，并通过机翼折叠轴(4)固定于所述中机身(5)前部，所述可伸缩式起落架(9)置于所述后机身(6)前部，飞行器采用无尾式布局，飞机重心位于前机身(2)后部、中机身(5)之前，涵道与机翼之间采用特定位置关系实现组合优化。
  2. 根据权利要求1所述的基于双涵道风扇动力系统的可折叠式固定翼垂直起降无人飞行器，其特征在于，
    所述机头(1)为电子舱，用于内置多种传感器和光电设备；所述前机身(2)为主负载舱，用于搭载主要能源和载荷；所述中机身(5)为次负载舱，用于搭载航电系统、次要能源、机翼折叠轴(4)的驱动机构、可伸缩式起落架(9)的驱动机构；所述后机身(6)的前部置有可伸缩式起落架，中部两侧对称安置所述涵道风扇动力系统(7)，后部为锥形整流体。
  3. 根据权利要求1所述的基于双涵道风扇动力系统的可折叠式固定翼垂直起降无人飞行器，其特征在于，
    所述可折叠式机翼(3)采用可折叠式构型，机翼为二段式折叠翼，可沿纵向轴线向机腹折叠36°～180°，机翼后缘靠近翼梢处安置有副翼(8)。
  4. 根据权利要求1所述的基于双涵道风扇动力系统的可折叠式固定翼垂直起降无人飞行器，其特征在于，
    所述双涵道风扇动力系统(7)采用横列式、尾推布局对称分布于所述后机身(6)两侧，其数量为2，其旋转轴位于机翼下表面以下。
  5. 根据权利要求4所述的基于双涵道风扇动力系统的可折叠式固定翼垂直起降无人飞行器，其特征在于，
    所述涵道风扇动力系统(7)包括：涵道体(10)、动力风扇(11)、风扇驱动机构(12)、控制舵面(13)、控制舵面驱动机构(14)；其中，所述动力风扇(11)位于所述涵道体(10)中，通过所述风扇驱动机构(12)与所述涵道体(10)相连接；所述控制舵面(13)位于涵道出口处，数量为4，呈“十字”型环绕涵道旋转轴；所述控制舵面(13)的旋转轴与涵道旋转轴垂直，其一端与涵道体相连，一端连接置于涵道体中的所述控制舵面驱动机构(14)。
  6. 根据权利要求1所述的基于双涵道风扇动力系统的可折叠式固定翼垂直起降无人飞行器，其特征在于，
    所述涵道与机翼之间实现组合优化的特定相对位置关系满足：
    所述可折叠式机翼(3)的后缘距涵道入口平面距离l1与涵道入口直径d关系为：
    0.35d≤l1≤0.45d；
    所述可折叠式机翼(3)的弦线平面距涵道中心轴线距离l2与涵道入口直径d关系为：
    0.25d≤l2≤0.4d。
  7. 根据权利要求1所述的基于双涵道风扇动力系统的可折叠式固定翼垂直起降无人飞行器，其特征在于，
    所述飞行器采用无尾式布局，全机无常规水平尾翼、垂直尾翼、升降舵与方向舵。
  8. 根据权利要求1所述的基于双涵道风扇动力系统的可折叠式固定 翼垂直起降无人飞行器，其特征在于，
    所述可伸缩式起落架(9)的数量为4，单个起落架长度可实时调节。
  9. 根据权利要求5所述的基于双涵道风扇动力系统的可折叠式固定翼垂直起降无人飞行器，其特征在于，
    所述控制舵面(13)为可活动的，通过偏转所述控制舵面(13)提供姿态控制力矩，实现飞行姿态的稳定与控制。
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