CN108528692B - Folding wing dual-rotor aircraft and control method thereof - Google Patents
Folding wing dual-rotor aircraft and control method thereof Download PDFInfo
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- B—PERFORMING OPERATIONS; TRANSPORTING
- B64—AIRCRAFT; AVIATION; COSMONAUTICS
- B64C—AEROPLANES; HELICOPTERS
- B64C27/00—Rotorcraft; Rotors peculiar thereto
- B64C27/04—Helicopters
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- B—PERFORMING OPERATIONS; TRANSPORTING
- B64—AIRCRAFT; AVIATION; COSMONAUTICS
- B64C—AEROPLANES; HELICOPTERS
- B64C1/00—Fuselages; Constructional features common to fuselages, wings, stabilising surfaces or the like
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- B—PERFORMING OPERATIONS; TRANSPORTING
- B64—AIRCRAFT; AVIATION; COSMONAUTICS
- B64C—AEROPLANES; HELICOPTERS
- B64C1/00—Fuselages; Constructional features common to fuselages, wings, stabilising surfaces or the like
- B64C1/30—Parts of fuselage relatively movable to reduce overall dimensions of aircraft
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- B—PERFORMING OPERATIONS; TRANSPORTING
- B64—AIRCRAFT; AVIATION; COSMONAUTICS
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Abstract
The invention discloses a folding wing double-rotor aircraft and a control method thereof, and relates to the field of aircraft. The aircraft can work in a helicopter mode and an airplane mode and share one set of control system. Compared with the existing vertical take-off and landing aircraft which uses two sets of control systems, the design of the invention is simpler and more reliable, and no redundant control mechanism exists in the vertical take-off and landing or level flight process of the aircraft, thereby improving the utilization rate of structural components of the aircraft and reducing the redundant structural weight. The invention is suitable for aerial photography and can also be used for individual soldier investigation.
Description
Technical Field
The invention relates to the field of aircrafts, in particular to a folding wing dual-rotor aircraft and a control method thereof.
Background
Although the aerodynamic efficiency of wing is higher for traditional fixed wing aircraft, no matter it is that the hand is thrown or launch and all have certain requirement to the place, and the transmission is thrown to the hand moreover and needs operating personnel to expose in spacious place, is unfavorable for operating personnel's safety to be hidden. Although the four-rotor and multi-rotor aircraft is easy to take off, land and control, the flying speed, the endurance time and the flying radius of the aircraft are very limited due to the overlarge flying resistance of the aircraft body, and the investigation task in a larger range is difficult to complete. The existing vertical take-off and landing aircraft simply overlaps the functions of the fixed-wing aircraft and the four-rotor aircraft to achieve the purpose of vertical take-off and landing by utilizing the long-endurance characteristics of the fixed-wing aircraft and the good stability of the four rotors, however, the working efficiency of the design is not optimal under different flight modes, and the endurance potential of the aircraft is difficult to be fully exerted. The larger the diameter of the rotor is, the higher the hovering efficiency is, and the hovering efficiency of the aircraft with the same takeoff weight is higher by using two rotors than that of four rotors or multiple rotors.
In conclusion, the existing vertical take-off and landing aircraft has the problems of complex structure and low aerodynamic efficiency.
Disclosure of Invention
The purpose of the invention is as follows: aiming at the defects and shortcomings of the prior art, the invention provides a dual-rotor aircraft with folding wings, which realizes the stable operation of the aircraft in a helicopter mode and an airplane mode through a dual-rotor system with an automatic tilter, does not need a conventional pneumatic control surface when flying in the airplane mode, and still uses an operation system in the helicopter mode to perform flying operation.
Another object of the present invention is to provide a control method for the above folding wing dual rotor aircraft.
The technical scheme is as follows: in order to achieve the purpose, the invention provides a folding wing dual-rotor aircraft, which comprises a fuselage and wings respectively positioned on two sides of the fuselage, wherein the two wings are hinged on the fuselage and can rotate around the longitudinal axis of the fuselage, and the end parts of the two wings far away from the fuselage are respectively provided with a rotor system used for providing power. The rotor system utilizes the periodic variable pitch generated by the motion of the built-in automatic tilter to change the aerodynamic force and aerodynamic moment of the rotor, and realizes the control of the pitching, yawing and rolling motions of the aircraft.
The two wings are variable dihedral wings, and when the aircraft flies flatly, the wings rotate around the longitudinal axis of the aircraft body to change the dihedral angles of the wings, so that the gravity center of the aircraft is reduced, and the stability of the aircraft is improved. The wings can be folded around the fuselage after the aircraft lands, so that the size of the aircraft in the unfolding direction is reduced by half, and the aircraft is convenient to store and carry.
The rotor wing system comprises a propeller hub, blades arranged on the propeller hub and a nacelle connected with the propeller hub, wherein a variable-pitch pull rod, an automatic inclinator, a steering engine and a steering engine pull rod are arranged in the nacelle, the variable-pitch pull rod is correspondingly connected with the blades of the blades, the automatic inclinator is connected with the variable-pitch pull rod, the steering engine is connected with the automatic inclinator through the steering engine pull rod, and the steering engine drives the steering engine pull rod to move up and down to drive the automatic inclinator to move up and down and tilt, so that the blades are driven to change the pitch. The automatic tilter moves up and down along the rotor shaft to change the rotor total pitch, and the automatic tilter tilts around the spherical hinge to enable the rotor to generate periodic pitch change. And a motor is also arranged in the short cabin, and the motor is connected with the rotor wing shaft and used for providing power for the propeller hub. The blades in the two rotor systems rotate in opposite directions to balance the reactive torque generated during rotation.
In order to increase course stability, a ventral fin is arranged at the tail part of the machine body, a foldable undercarriage structure is arranged at the tail part of the nacelle, and when the nacelle vertically takes off and lands, three-point supports are formed by the two undercarriages and the end part of the ventral fin.
The lifting motion of the double-rotor aircraft with the folding wings is controlled by a rotor system, the aircraft ascends when the pulling force generated by the rotation of the rotors is greater than the total weight of the aircraft, and the aircraft descends when the pulling force is less than the total weight of the aircraft. There are two ways to vary the tension of the aircraft rotor: firstly, under the condition that the rotating speed of the rotor wing is constant, the actuating direction of the steering engine is controlled to enable the automatic inclinator to move up and down along the rotor wing shaft to drive the blade pitch to increase or decrease, so that the pulling force of the rotor wing is changed; secondly, the rotary wing is driven to change the rotating speed by controlling the change of the rotating speed of the motor, so that the tension of the rotary wing is changed.
The automatic tilter tilt produces a cyclic variation in the angle of attack of the blades, i.e. a cyclic pitch of the rotor. When the aircraft is in a helicopter flight mode and the longitudinal periodic variable pitch is carried out, the aircraft moves longitudinally; when the distance is changed in a transverse period, the aircraft moves transversely; and the longitudinal periodic variable pitch differential is adopted, and the aircraft changes the course. In the airplane flight mode, the longitudinal periodic variable pitch performs pitching operation on the aircraft, the transverse periodic variable pitch performs course operation on the aircraft, and the longitudinal periodic variable pitch performs rolling operation on the aircraft in a differential mode.
Has the advantages that: compared with the prior art, the invention has the following beneficial effects:
1. the folding wing double-rotor aircraft adopts a vertical flying wing structure, adopts two rotors with automatic inclinators as power and control surfaces of a helicopter mode and an airplane mode of the aircraft, can ensure the maneuverability and the stability in the helicopter mode, can realize the control of the aircraft in the airplane mode, does not need an additional pneumatic control surface, simplifies the structure of the vertical take-off and landing aircraft, lightens the self weight of the aircraft, and increases the effective load. And the stability can be actively increased by means of flight control, and the static stability can be achieved by means of the self pneumatic design of the aircraft instead of the conventional aircraft mode.
2. The aircraft provided by the invention realizes hovering/vertical takeoff by using the double rotors, after the aircraft is in an airplane mode, the wings mainly provide lift force, and the rotors work in an axial flow mode, so that the total distance of the rotors can be increased, the rotating speed can be reduced, the profile resistance of the rotors can be effectively reduced on the premise of ensuring that the rotors can provide enough tension, the required power of the aircraft is reduced, and the capability of longer endurance time is realized.
3. The variable dihedral wing has the advantages that the variable dihedral wing is adopted by the aircraft, the center of gravity is effectively reduced, meanwhile, the aerodynamic force borne by the wing is more beneficial to the stability of the flight of the aircraft, the variable dihedral structure acts in an aircraft helicopter mode, the center of gravity of a fuselage is deviated, the transition to the aircraft mode can be completed more efficiently by matching with the periodic pitch variation of the rotor wing, and the high dependence of the tilting transition of the aircraft on the control efficiency of the rotor wing is reduced.
4. The invention adopts the hinge folding mechanism, so that the aircraft can be folded in half, the geometric overall dimension of the aircraft is greatly reduced, and the folding type aircraft is convenient for carrying by an individual soldier or carrying by a backpack for travel. Meanwhile, the variable dihedral angle structure and the folding hinge of the aircraft are designed in a fusion mode, so that the complexity of the aircraft structure is reduced, and the weight of the aircraft structure is also reduced. The variable dihedral characteristic enables the aircraft to increase flight stability and reduce sensitivity to weather conditions by increasing the dihedral under adverse flight conditions. When in a helicopter mode, the mechanism with the variable dihedral angle is used for driving, so that the gravity center of the aircraft can be deviated, and the longitudinal aerodynamic force generated by the periodic variable pitch control of the rotor wing is matched, so that the interconversion between the helicopter mode and the airplane mode can be better completed.
5. The invention adopts the ventral fin design, and is beneficial to keeping the course stability of the aircraft in the process of mutually switching between the vertical flight mode and the horizontal flight mode. The ventral fins play a role in increasing the course stability and keeping balance when the aircraft sideslips. The ventral fin layout ingeniously utilizes the structure of the ventral fin and the landing gear which extends out of the nacelle to form a three-point support, so that unnecessary structural weight is reduced.
Drawings
In order to more clearly illustrate the technical solutions in the embodiments of the present invention, the drawings needed to be used in the embodiments will be briefly described below, and it is obvious that the drawings in the following description are only some embodiments of the present invention, and it is obvious for those skilled in the art that other drawings can be obtained according to these drawings without creative efforts.
Fig. 1 is a schematic structural diagram (vertical takeoff) of a dual-rotor aircraft with folding wings according to an embodiment of the present invention;
fig. 2 is a side view of a main body structure of a folding wing dual-rotor aircraft (helicopter state) according to an embodiment of the present invention;
fig. 3 is a structural view (level flight) of a folding wing dual rotor aircraft provided in an embodiment of the present invention;
FIG. 4 is a schematic view of a nacelle system of a dual-rotor aircraft with folding wings according to an embodiment of the present invention;
FIG. 5 is a helicopter mode flight control strategy for a folded wing dual rotor aircraft provided in an embodiment of the present invention;
FIG. 6 is an illustration of a folded wing dual rotor aircraft mode flight control strategy provided in an embodiment of the present invention;
fig. 7 is an actuation schematic diagram of a steering engine of a dual-rotor aircraft with folding wings provided in an embodiment of the present invention.
Detailed Description
The present invention will be described in further detail with reference to the accompanying drawings and examples. For the purpose of illustration, only some structures relevant to the present invention are shown in the drawings, and not all structures are shown. It is further understood that the terms "upper," "lower," "inner," "aft," and the like, as used herein, refer to an orientation or positional relationship based on that shown in FIG. 1 for the purpose of describing the present invention and simplifying the description, but do not indicate or imply that the device referred to must have a particular orientation, be constructed and operated in a particular orientation, and therefore should not be considered as limiting the present invention.
As shown in fig. 1, the dual-rotor aircraft with folding wings provided by the invention comprises a fuselage 1 and wings 2 respectively located at two sides of the fuselage 1, wherein the two wings 2 are connected to the fuselage 1 through hinges and can rotate around the fuselage, the end parts of the two wings 2 far away from the fuselage are both connected with a nacelle 6, and the head part of the nacelle 6 is provided with a blade 4 and a fairing 5. Referring to fig. 2 and 3, nacelle 6 is attached to hub 3, blades 4 are mounted on hub 3, and blades 4 rotate as hub 3 rotates. The afterbody of fuselage 1 is equipped with ventral fin 9, and the afterbody of nacelle 6 is equipped with two undercarriage support arms 7, and undercarriage 8 is installed on undercarriage support arm 7, and two undercarriage support arms can rotate around the inside fulcrum of nacelle, lead the undercarriage and receive and release. During vertical take-off and landing, the two landing gears and the end of the ventral fin form a three-point support. The optoelectronic device 10 is provided on the head of the body 1 for aerial photography or reconnaissance.
The two wings 2 are variable dihedral wings, and when the aircraft flies flatly, the wings rotate around the longitudinal axis of the fuselage to change the dihedral angles of the wings, so that the gravity center of the aircraft is reduced, and the stability of the aircraft is improved. The wings can be folded around the fuselage after the aircraft lands, so that the size of the aircraft in the unfolding direction is reduced by half, and the aircraft is convenient to store and carry.
Fig. 4 shows the structure inside the nacelle 6, the nacelle 6 includes a pitch-variable pull rod 11, an automatic inclinator 12, a steering engine 13, a steering engine pull rod 14, and a motor 15 connected with a rotor shaft, the pitch-variable pull rod 11 is connected with the blades of the blades 4 in a one-to-one correspondence, the automatic inclinator 12 is connected with the pitch-variable pull rod 11, the steering engine 13 is connected with the automatic inclinator 12 through the steering engine pull rod 14, and the motor 15 is used for inputting power to the hub 3 to drive the hub to rotate. The steering engine 13 drives the steering engine pull rod 14 to actuate, so that the automatic inclinator 12 moves up and down or inclines to drive the blades 4 to change the pitch. When all steering engines in the nacelle 6 have the same actuation amount, the automatic inclinator 12 moves up and down along the rotor shaft to increase or decrease the attack angle of the blades 4, thereby changing the total pitch of the rotor. When the steering engines have different working amounts, the automatic inclinator 12 tilts around the spherical hinge, so that the attack angle of the blades 4 generates periodic change, namely the periodic variable pitch of the rotor wing is generated, and the lift force generated by the blades 4 at different rotating azimuth angles is different. The blade is generally made for elastic material, and is different when the lift that different azimuths received, and the blade is also different in the deformation of waving the direction, and rotor oar dish can produce the slope promptly, and the pulling force that the rotor produced has a component to oar dish incline direction.
The lifting motion of the aircraft is controlled by a rotor system, the aircraft rises when the pulling force generated by the rotation of the rotor is larger than the total weight of the aircraft, and the aircraft descends when the pulling force is smaller than the total weight of the aircraft. There are two ways to vary the tension of the aircraft rotor: firstly, under the condition that the rotating speed of the rotor wing is constant, the actuating direction of the steering engine is controlled to enable the automatic inclinator to move up and down along the rotor wing shaft to drive the blade pitch to increase or decrease, so that the pulling force of the rotor wing is changed; secondly, the rotary wing is driven to change the rotating speed by controlling the change of the rotating speed of the motor, so that the tension of the rotary wing is changed.
When the aircraft takes off vertically, the aircraft is lifted to a certain height by using the pulling force generated by the double rotors (namely, the two blades 4), the two landing gears 8 are respectively retracted into the corresponding nacelles 6, and the motion and the stability of the aircraft in all directions are controlled by the automatic inclinators 12 to reach the safe height. As described above, the automatic inclinator periodically changes the pitch to enable the rotor disc to generate lateral force to form tilting moment to the fuselage, the two wings 2 rotate around the fuselage 1 to enable the center of gravity of the aircraft to deviate towards the tilting side, the aircraft is accelerated in the horizontal direction by the tilting of the rotor disc, the lift force generated by the two wings is gradually increased along with the increase of the forward flying speed, the power consumption of the rotors is gradually reduced until the lift force of the aircraft is completely provided by the two wings, the rotors only generate forward flying pulling force and control the flight attitude of the aircraft to be stable, the required power of the rotors is reduced under the condition that the pulling force is not changed by improving the total pitch of the rotors and reducing the rotating speed of the rotors, and the endurance time of the aircraft is prolonged. When the aircraft needs to land, the horizontal flight attitude of the aircraft is changed into the vertical flight attitude by matching the automatic inclinators of the rotors with the rotation of the two wings, the dihedral angle of the aircraft is zero at the moment, the periodic pitch of the rotors enables the aircraft to generate a head-up moment, when the head-up attack angle of the aircraft is increased, the resistance of the aircraft is increased, the aircraft is decelerated by the aid of the pneumatic braking effect, the helicopter mode is converted at the moment, the aircraft is controlled to land accurately according to a set landing point, and a flight task is completed.
When the aircraft is in a helicopter mode and the longitudinal periodic variable pitch is carried out, the aircraft can move longitudinally; when the distance is changed in a transverse period, the aircraft moves transversely; and the longitudinal periodic variable pitch differential is adopted, and the aircraft changes the course. In the airplane mode, the longitudinal cyclic variable pitch performs pitching operation on the aircraft, the transverse cyclic variable pitch performs course operation on the aircraft, and the longitudinal cyclic variable pitch differential performs rolling operation on the aircraft.
As shown in fig. 5, in the helicopter mode, the total pitch of the rotor, the motor speed, the longitudinal cyclic pitch differential, and the transverse cyclic pitch can be controlled, and the total pitch and the motor speed need to be controlled independently. As shown in fig. 6, in the airplane mode, the collective pitch, the motor rotation speed, the longitudinal cyclic pitch differential, and the lateral cyclic pitch can be directly controlled. The control of the rotor total pitch is realized by the automatic inclinator moving up and down along the rotor shaft to increase or decrease the blade pitch, the control of the motor rotating speed is realized by the adjustable pulse width signal sent by the flight control system, and the change of the total pitch and the motor rotating speed can change the rotor tension, thereby controlling the lifting motion of the aircraft. The control of longitudinal and transverse periodic variable distances is realized by controlling the movement of an automatic inclinator, the movement of the automatic inclinator is realized by controlling the actuating direction of a steering engine, the actuating schematic diagrams of the steering engine are shown in figures 7(a) -7(h), the operating numbers in the figures represent different steering engine driving positions, wherein the steering engines 1-3 are positioned in one nacelle, the steering engines 4-6 are positioned in the other nacelle, and the arrow directions represent the driving direction of the steering engines. Fig. 7(a) -7(h) show different automatic recliner inputs for different motions of the aircraft, respectively, fig. 7(a) is a schematic diagram of total distance increasing, fig. 7(b) is a schematic diagram of total distance decreasing, fig. 7(c) is a schematic diagram of longitudinal periodic variable distance (backward), fig. 7(d) is a schematic diagram of longitudinal periodic variable distance (forward), fig. 7(e) is a schematic diagram of longitudinal periodic variable distance differential (left turn), fig. 7(f) is a schematic diagram of longitudinal periodic variable distance differential (right turn), fig. 7(g) is a schematic diagram of transverse periodic variable distance (right turn), and fig. 7(h) is a schematic diagram of transverse periodic variable distance (left turn). Taking fig. 7(c) as an example, to make the aircraft lean backward, the steering engines No. 1 and 4 are driven upward, and the steering engines No. 2, 3, 5, and 6 are driven downward. Similarly, in other cases, the driving direction of the steering engine can be obtained by referring to corresponding diagrams, and the description is omitted.
Claims (7)
1. A double-rotor aircraft with folding wings is characterized by comprising an aircraft body, a left wing and a right wing which are hinged on the aircraft body and can rotate around the longitudinal axis of the aircraft body, wherein the end parts of the two wings, far away from the aircraft body, are respectively provided with rotor systems with opposite rotating directions, the two wings are variable dihedral wings, when the aircraft flies flatly, the wings rotate around the longitudinal axis of the aircraft body to change the dihedral angles of the wings, so that the gravity center of the aircraft is reduced, and the wings can be folded around the aircraft body after the aircraft lands, so that the spanwise size of the aircraft is reduced by a half; the rotor system changes the aerodynamic force and aerodynamic moment of the rotor by utilizing the cyclic pitch generated by the motion of the built-in automatic tillers, so that the pitching, yawing and rolling motions of the aircraft are controlled, and the flying motions of the aircraft in a helicopter mode and an airplane mode are controlled by the cyclic pitch of the rotor system.
2. The dual-rotor aircraft with the folding wings as claimed in claim 1, wherein the rotor system comprises a hub, blades mounted on the hub, and a nacelle connected with the hub, a variable-pitch pull rod, an automatic tilter, a steering engine and a steering engine pull rod are arranged in the nacelle, the variable-pitch pull rod is correspondingly connected with the blades of the blades, the automatic tilter is connected with the variable-pitch pull rod, the steering engine is connected with the automatic tilter through the steering engine pull rod, and the steering engine drives the steering engine pull rod to move up and down so as to drive the automatic tilter to move up and down and tilt, so that the blades are driven to change the pitch.
3. A folded wing dual rotor aircraft according to claim 2, wherein a motor is further provided within the nacelle for powering the hub, the motor being coupled to the rotor shaft.
4. The foldable wing twin rotor aircraft according to claim 2, wherein a ventral fin is provided at the aft portion of the fuselage, and the nacelle aft portion is provided with a foldable landing gear structure, wherein the two landing gears and the end of the ventral fin form a three-point support during vertical take-off and landing.
5. A method for controlling a folding wing dual rotor aircraft according to any one of claims 1-4, wherein the lifting and lowering movements of the aircraft are controlled by a rotor system, the rotor rotating generates a pulling force which is greater than the gross weight of the aircraft, the aircraft is lifted, and the aircraft is lowered; the flight motion of the aircraft in helicopter mode and airplane mode is controlled by rotor cyclic pitch.
6. The method of controlling a folded wing dual rotor aircraft according to claim 5, wherein the method of varying the aircraft rotor tension comprises:
under the condition that the rotating speed of the rotor wing is constant, the automatic inclinator moves up and down along the rotor wing shaft by controlling the actuating direction of the steering engine to drive the blade pitch to increase or decrease, so that the pulling force of the rotor wing is changed; and
the rotating speed of the rotor wing is changed by driving the rotor wing through controlling the change of the rotating speed of the motor, so that the pulling force of the rotor wing is changed.
7. The control method for a folding wing twin rotor aircraft according to claim 5, wherein in the helicopter mode, the aircraft moves in the longitudinal direction during longitudinal cyclic pitch; when the distance is changed in a transverse period, the aircraft moves transversely; the longitudinal periodic variable pitch differential is adopted, and the course of the aircraft is changed;
in the airplane mode, the longitudinal cyclic variable pitch performs pitching operation on the aircraft, the transverse cyclic variable pitch performs course operation on the aircraft, and the longitudinal cyclic variable pitch differential performs rolling operation on the aircraft.
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| Application Number | Priority Date | Filing Date | Title |
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| CN201810364530.2A CN108528692B (en) | 2018-04-23 | 2018-04-23 | Folding wing dual-rotor aircraft and control method thereof |
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| CN201810364530.2A CN108528692B (en) | 2018-04-23 | 2018-04-23 | Folding wing dual-rotor aircraft and control method thereof |
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| CN108528692B true CN108528692B (en) | 2021-08-06 |
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Families Citing this family (7)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN111232192A (en) * | 2018-11-29 | 2020-06-05 | 戴瑾 | Double-rotor aircraft |
| CN109823511B (en) * | 2019-03-01 | 2024-05-24 | 北京航空航天大学 | Variable wing structure with transverse automatic stability augmentation function |
| CN111038693B (en) * | 2019-12-04 | 2023-04-18 | 王挺 | Mechanical control system of multi-rotor aircraft |
| CN113071668A (en) * | 2020-04-16 | 2021-07-06 | 灵遥机器人(深圳)有限责任公司 | Unmanned aerial vehicle |
| CN111678386A (en) * | 2020-07-03 | 2020-09-18 | 南京航空航天大学 | Aircraft head deflection control device |
| CN112327922B (en) * | 2020-11-18 | 2022-04-22 | 南京航空航天大学 | Autonomous take-off and landing integrated control method for flying wing unmanned aerial vehicle |
| CN114348250B (en) * | 2022-01-12 | 2023-04-07 | 广东汇天航空航天科技有限公司 | Transverse double-rotor aircraft, flight control method thereof and electronic equipment |
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| CN103287576A (en) * | 2013-05-24 | 2013-09-11 | 北京航空航天大学 | Tailless layout single tail seat type vertical take-off and landing aircraft |
| CN106828918A (en) * | 2015-12-04 | 2017-06-13 | 中航通飞研究院有限公司 | A kind of Three-wing-surface vertically taking off and landing flyer |
| CN106892094A (en) * | 2017-01-22 | 2017-06-27 | 南京航空航天大学 | A kind of individually controllable four rotor unmanned aircraft of space six degree of freedom and its control method |
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| US9174731B2 (en) * | 2013-08-14 | 2015-11-03 | Bell Helicopter Textron Inc. | Fixed engine and rotating proprotor arrangement for a tiltrotor aircraft |
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|---|---|---|---|---|
| CN103287576A (en) * | 2013-05-24 | 2013-09-11 | 北京航空航天大学 | Tailless layout single tail seat type vertical take-off and landing aircraft |
| CN106828918A (en) * | 2015-12-04 | 2017-06-13 | 中航通飞研究院有限公司 | A kind of Three-wing-surface vertically taking off and landing flyer |
| CN106892094A (en) * | 2017-01-22 | 2017-06-27 | 南京航空航天大学 | A kind of individually controllable four rotor unmanned aircraft of space six degree of freedom and its control method |
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