CN116853487A - Tilt rotor aircraft and transition conversion control device thereof - Google Patents
Tilt rotor aircraft and transition conversion control device thereof Download PDFInfo
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- CN116853487A CN116853487A CN202310755162.5A CN202310755162A CN116853487A CN 116853487 A CN116853487 A CN 116853487A CN 202310755162 A CN202310755162 A CN 202310755162A CN 116853487 A CN116853487 A CN 116853487A
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- B—PERFORMING OPERATIONS; TRANSPORTING
- B64—AIRCRAFT; AVIATION; COSMONAUTICS
- B64C—AEROPLANES; HELICOPTERS
- B64C27/00—Rotorcraft; Rotors peculiar thereto
- B64C27/22—Compound rotorcraft, i.e. aircraft using in flight the features of both aeroplane and rotorcraft
- B64C27/28—Compound rotorcraft, i.e. aircraft using in flight the features of both aeroplane and rotorcraft with forward-propulsion propellers pivotable to act as lifting rotors
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- B—PERFORMING OPERATIONS; TRANSPORTING
- B64—AIRCRAFT; AVIATION; COSMONAUTICS
- B64C—AEROPLANES; HELICOPTERS
- B64C19/00—Aircraft control not otherwise provided for
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- B—PERFORMING OPERATIONS; TRANSPORTING
- B64—AIRCRAFT; AVIATION; COSMONAUTICS
- B64C—AEROPLANES; HELICOPTERS
- B64C27/00—Rotorcraft; Rotors peculiar thereto
- B64C27/52—Tilting of rotor bodily relative to fuselage
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Abstract
The invention discloses a tiltrotor aircraft and a transition conversion control device thereof. The device comprises: the instruction filtering module is used for receiving a host control instruction, and combining sensor data and aircraft state processing to obtain a control expectation for the aircraft motion state; the control module is used for receiving control expectations, comprehensively processing the real-time motion state and the position data of the aircraft according to the pneumatic and quality characteristics of the aircraft and the real-time motion state and the position data of the aircraft input by the sensor to obtain forces and moments in all directions which the aircraft should provide, calculating the difference between the real-time parameters of the aircraft and the control expectations, and realizing the conversion and output of the error value to the forces or the moments based on a control algorithm; the control distribution module is used for receiving the force and the moment output by the control module and calculating the control instructions of different actuators according to the current motion state of the aircraft. By adopting the technical scheme of the invention, the problems in the transition and conversion process of the vertical take-off and landing aircraft are solved, and the purpose of assisting a pilot in controlling the aircraft is realized.
Description
Technical Field
The invention relates to the technical field of aircrafts, in particular to a tilting rotor wing aircraft and a transition conversion control device thereof.
Background
At present, the saturation of road vehicles is continuously increased, so that the traffic in cities is increasingly crowded, and the daily travel time of people is gradually increased. With the vigorous development of navigation products in recent years, the low-altitude airspace is gradually released, and the problem that people travel in short distance is gradually increased by using a relatively idle low-altitude airspace is solved. The solution of using low-altitude airspace traveling is mainly an electric vertical take-off and landing aircraft, and the main electric vertical take-off and landing aircraft is mainly divided into three types of tilting, compound wing and multi-rotor wing. Compared with the traditional aircraft, one of the biggest characteristics and development emphasis of the vertical take-off and landing aircraft is how to realize the air transition control after vertical take-off or before vertical landing.
The vertical take-off and landing aircrafts of the tilting type, the composite wing type and the multi-rotor type have great difference in cruising ability, flying speed and the like, wherein the aircraft is the most clean in cruising state in the configuration of the tilting type and has the lowest resistance coefficient. Therefore, the aircraft of the "tilting class" is highest in cruising efficiency, the "compound wing class" is next to the "multi-rotor class" is the next to the "multi-rotor class" in cruising efficiency; in the flying speed, the multi-rotor wing type is the lowest, and compared with the tilting type, the compound wing type has larger zero lift resistance, so that the cruising lift coefficient can be higher on the premise of ensuring the cruising lift-drag ratio, the cruising speed is lower, and the tilting type cruising speed is the fastest; in flight control, both the "tilting type" and the "composite wing type" face the problem of conversion balance between aerodynamic force and power, and the "tilting type" aircraft introduces a control dimension of a power angle in control, which can certainly cause the increase of control difficulty. Such control problems are not faced by conventional navigable and even civil aviation aircraft. In addition, the conversion from the power system to wing aerodynamic provides the primary lift, and the driver is expected to achieve accurate control of the lift direction and thrust direction power respectively is almost impossible or requires a great deal of effort, during which the driver is required to decide if other emergency situations occur, and the driver may enter a state in which the drivers are out of each other, which seriously affects the flight safety. There is a need for an auxiliary flight control or automatic flight control method that can slow pilot-operated aircraft pressures for use with this type of aircraft.
Disclosure of Invention
The present invention provides a tiltrotor aircraft comprising: the aircraft comprises a wing (1), a fuselage (2), a horizontal tail (3), a vertical tail (4), a landing gear (5) and a stay bar (6); the wing (1) comprises a main wing (11) and winglets (12) at two sides, the horizontal tail (3) comprises a horizontal stabilizer (31) and nacelle (32) at two sides, and left and right stay bars (6) are respectively arranged at the front sides of the left and right wings of the main wing (11); the power systems are arranged on the end parts of the stay bars (6), the winglets (12) on the two sides and the nacelle (32) on the two sides; the power system arranged on the winglets (12) on the two sides and the nacelle (32) on the two sides changes the thrust direction along with the tilting of the winglets (12) on the two sides and the nacelle (32) on the two sides; the power system arranged at the end parts of the left and right stay bars (6) inclines along with the rotation axis in the direction away from the machine body.
In the tilt rotor aircraft, the first rotor (71) mounted on the right side stay (6) rotates anticlockwise, the second rotor (72) mounted on the right side winglet (12) rotates anticlockwise, the third rotor (73) mounted on the right side nacelle (32) rotates clockwise, the fourth rotor (74) mounted on the left side nacelle (32) rotates anticlockwise, the fifth rotor (75) mounted on the left side winglet (12) rotates clockwise, and the sixth rotor (76) mounted on the left side stay (6) rotates clockwise, wherein the rotation direction is defined as a overlooking direction when the winglet (12) and the nacelle (32) are in a vertical state.
In the tilting rotor aircraft, the rotation directions of the first rotor (71) and the sixth rotor (76) ensure that the aircraft respectively bear aerodynamic moments in the clockwise direction and the anticlockwise direction, and when the heading is controlled by changing the rotation speed of a certain rotor in a low-speed state, the contribution of component force generated by camber to the yaw moment of the whole aircraft is consistent with the contribution of aerodynamic resistance moment caused by the rotation direction to the yaw moment.
A tiltrotor aircraft as described above, wherein the third rotor (73) is counter-rotating to the first rotor (71), the fourth rotor (74) is counter-rotating to the sixth rotor (76), and a typical four-rotor condition is created when the symmetric rotor control balance is closed when either the second rotor (72) or the fifth rotor (75) fails.
The invention also provides a transition control device for a tiltrotor aircraft, the control device being applied to any one of the tiltrotor aircraft, and comprising:
in the tilting transition process, the angles of the winglets at two sides or the nacelle at two sides are changed synchronously or asynchronously, and the power output is changed;
controlling the winglet and the nacelle to a large angle in a deceleration transition stage;
the asynchronous tilting winglet and the nacelle realize decoupling of pitch control and acceleration control;
and the transition flight is realized by synchronously tilting the winglet and the nacelle.
The invention also provides a transition conversion control device of the tiltrotor aircraft, which comprises the following components: the control device is used for controlling any one of the tilting rotor aircraft, and specifically comprises an instruction filtering module, a control module and a control distribution module;
the instruction filtering module is used for receiving a host control instruction, and combining sensor data and aircraft state processing to obtain a control expectation for the aircraft motion state;
the control module is used for receiving the control expectation of the motion state of the aircraft output by the command filtering module, comprehensively processing the real-time motion state and the position data of the aircraft input by the sensor according to the pneumatic and quality characteristics of the aircraft, and obtaining the force and the moment of each direction which the aircraft should provide, calculating the difference between the real-time parameters of the aircraft and the control expectation, and realizing the conversion and the output from the error value to the force or the moment based on a control algorithm;
the control distribution module is used for receiving the force and the moment output by the control module and calculating the control instructions of different actuators according to the current motion state of the aircraft.
A tiltrotor aircraft transition control apparatus as described above wherein the control of the aircraft motion state desirably includes directional velocity, acceleration, attitude angle, angular velocity.
The transition conversion control device of the tiltrotor aircraft comprises a command filtering module, a control module and a control module, wherein the command filtering module comprises a limit envelope construction strategy and an optimization principle strategy; limiting the envelope, namely comprehensively calculating a tilting angle-airspeed curve according to the multi-dimensions of the aircraft such as self-quality characteristics, aerodynamic characteristics, dynamic characteristics, control safety, structural strength limitation and the like under various aircraft states; the optimization principle is a set of a series of optimization judgment conditions, and aims to screen out and obtain an optimal control expectation in the current state in a limiting envelope according to the actual flight state of the aircraft.
The transition conversion control device of the tiltrotor aircraft specifically comprises the following steps of:
traversing the tilting angle, and obtaining the power ranges of the lift force and the thrust direction by using a trigonometric function based on the known power characteristics of the aircraft;
traversing the attack angle and airspeed of the aircraft, and obtaining the lift force and the resistance by using an aerodynamic calculation method based on the known aerodynamic characteristics of the aircraft;
assuming that the tilting transition Cheng Fei altitude is kept unchanged, obtaining the available airspeed range of the aircraft in each attack angle and tilting angle state by using a vertical direction dynamic balance equation based on the known aircraft mass characteristics;
based on the calculated airspeed range, the aircraft horizontal acceleration is calculated using a horizontal dynamic equilibrium equation.
A tiltrotor aircraft transition control device as described above, wherein the optimization principle strategy comprises: the principle of the fastest transition, the principle of energy saving and the principle of passenger comfort, and the weight duty ratio of each principle is automatically adjusted under different flight tasks or aircraft states.
The beneficial effects achieved by the invention are as follows: the technical scheme of the invention solves the difficult problem in the transition and conversion process of the vertical take-off and landing aircraft, achieves the aim of assisting the pilot in controlling the aircraft, greatly reduces the workload of the pilot in the period, and can ensure that the pilot keeps enough energy to make a decision on the safe operation of the aircraft.
Drawings
In order to more clearly illustrate the embodiments of the present invention or the technical solutions in the prior art, the drawings used in the embodiments or the description of the prior art will be briefly described below, and it is obvious that the drawings in the following description are only some embodiments described in the present invention, and other drawings may be obtained according to these drawings for a person having ordinary skill in the art.
FIG. 1 is a schematic view of a tiltrotor aircraft according to one embodiment of the present invention;
FIG. 2 is a schematic diagram of a rotor layout on a tiltrotor aircraft;
fig. 3 is a schematic diagram of a transition control device for a tiltrotor aircraft according to a second embodiment of the present invention;
FIG. 4 is a flowchart of a limit envelope construction strategy;
FIG. 5 is a graph of aircraft pitch angle versus airspeed.
Detailed Description
The following description of the embodiments of the present invention will be made clearly and fully with reference to the accompanying drawings, in which it is evident that the embodiments described are some, but not all embodiments of the invention. All other embodiments, which can be made by those skilled in the art based on the embodiments of the invention without making any inventive effort, are intended to be within the scope of the invention.
Example 1
As shown in fig. 1, a first embodiment of the present invention provides a tiltrotor aircraft, including: wing 1, fuselage 2, horizontal tail 3, vertical tail 4, undercarriage 5 and stay bar 6.
The wing 1 comprises a main wing 11 and two winglets 12, the winglets 12 being connected to the main wing 11 by a first tilting axis about which the winglets 12 are rotatable.
The horizontal tail 3 comprises a horizontal stabilizer 31 and two side short cabins 32, the short cabins 32 are connected with the horizontal stabilizer 31 through a second tilting shaft, and the short cabins 32 can rotate around the second tilting shaft.
Left and right stay bars 6 are respectively mounted on the front sides of the left and right wings of the main wing 11. The power system is mounted on the end of the strut 6, on both winglets 12 and on both nacelle 32. The power systems mounted on the winglets 12 and nacelle 32 change thrust direction as the winglets 12 and nacelle 32 tilt; the power system mounted at the ends of the left and right stay bars 6 is inclined (i.e., camber) with the rotation axis in a direction away from the fuselage: on the one hand, the power system can provide lateral component force through the camber arrangement mode, and then heading moment is generated. The yaw moment can be generated by changing the output quantity of the stay bar power system in a differentiated way to enhance the heading control capability; on the other hand, the camber of the power system can enable the rotation plane of the rotor to not pass through the passenger area of the cabin, so that the possibility of injury to passengers caused by rotor blasting is effectively reduced.
Specifically, referring to fig. 2, the first rotor 71 mounted on the right strut 6 is rotated counterclockwise, the second rotor 72 mounted on the right winglet 12 is rotated counterclockwise, the third rotor 73 mounted on the right nacelle 32 is rotated clockwise, the fourth rotor 74 mounted on the left nacelle 32 is rotated counterclockwise, the fifth rotor 75 mounted on the left winglet 12 is rotated clockwise, and the sixth rotor 76 mounted on the left strut 6 is rotated clockwise, the rotation being defined as a top view when the winglet 12 and nacelle 32 are in a vertical position.
The rotation directions of the first rotor wing 71 and the sixth rotor wing 76 ensure that the aircraft bears aerodynamic moments in the clockwise direction and the anticlockwise direction respectively, so that when the course is controlled by changing the rotation speed of a certain rotor wing in a low-speed state, the contribution of component force generated by camber to the yaw moment of the whole aircraft is consistent with the contribution of aerodynamic resistance moment caused by the rotation direction to the yaw moment, and the course control efficiency is improved; the third rotor 73 and the first rotor 71 have opposite rotation directions, and the fourth rotor 74 and the sixth rotor 76 have opposite rotation directions, so that a typical four-rotor state can be formed when the symmetrical rotor control balance is closed when the second rotor 72 or the fifth rotor 75 fails, and the control is facilitated; the second rotor 72 and the fifth rotor 75 are designed in the direction of rotation to increase the aerodynamic efficiency of the machine.
Example two
Referring to fig. 3, a second embodiment of the present invention provides a transition control device for a tiltrotor aircraft, including a main control system and an optional control system.
The first main control system comprises an instruction filtering module, a control module and a control distribution module. Wherein:
(1) The instruction filtering module is used for receiving a host control instruction, and combining sensor data and aircraft state processing to obtain a control expectation for the aircraft motion state;
the host control instruction can be a control instruction input by a driver or a control instruction automatically planned by the system. Control desires for the state of motion of the aircraft include, but are not limited to, directional speeds, accelerations, attitude angles, angular velocities, and the like.
The instruction filtering module comprises a constraint envelope construction strategy and an optimization principle strategy. The limiting envelope is a tilting angle-airspeed curve comprehensively calculated according to the multi-dimensions of the aircraft such as the mass characteristic, aerodynamic characteristic, dynamic characteristic, control safety, structural strength limit and the like under various aircraft states. The optimization principle is a set of a series of optimization judgment conditions, and aims to screen out and obtain an optimal control expectation in the current state in a limiting envelope according to the actual flight state of the aircraft.
(1) Referring to fig. 4, the constraint envelope construction strategy specifically includes:
step 410, traversing the tilting angle, and obtaining the power ranges of the lift force and the thrust direction by using a trigonometric function based on the known power characteristics of the aircraft;
step 420: traversing the attack angle and airspeed of the aircraft, and obtaining the lift force and the resistance by using an aerodynamic calculation method based on the known aerodynamic characteristics of the aircraft;
step 430: assuming that the tilting transition Cheng Fei altitude is kept unchanged, obtaining the available airspeed range of the aircraft in each attack angle and tilting angle state by using a vertical direction dynamic balance equation based on the known aircraft mass characteristics;
step 440: based on the airspeed range determined in step 430, the aircraft horizontal acceleration is determined using a horizontal dynamic equilibrium equation.
And (3) calculating to obtain a relation curve of the inclination angle and the airspeed of the aircraft under each attack angle after the steps are completed, wherein the finally obtained limiting envelope is shown in fig. 5, and the area surrounded by the two curves and the coordinate axis is the usable range of the aircraft. For the situation that partial system failure affects the power output, different limiting envelopes are obtained according to all possible failure states of the aircraft and combinations thereof.
The limiting envelope is an inherent attribute of the aircraft, the calculation flow can be finished in the design stage in advance, and the limiting envelope is injected into the control module, so that the calculation resources are saved.
(2) The optimization principle strategy comprises the following steps: the principle of the fastest transition, the principle of energy saving, the principle of passenger comfort and the like, and the weight ratio of each principle is automatically adjusted under different flight tasks or aircraft states: for example, in an emergency state, tasks with strong timeliness such as emergency forced landing or transporting wounded medicines are needed, the weight of the principle of fastest transition can be adjusted to be 90% at the highest, the consideration weight of energy conservation is completely sacrificed to be 0%, and the comfort weight of passengers is partially sacrificed to be 10%.
The optimization principle can adjust the weight of each principle according to the real-time motion state, the flight task and the like of the aircraft, is a real-time strategy and is realized by means of the hardware data processing capacity of the control device.
(2) The control module is used for receiving the control expectation of the motion state of the aircraft output by the command filtering module, comprehensively processing the real-time motion state and the position data of the aircraft input according to the pneumatic and quality characteristics of the aircraft, and the sensor to obtain the force and the moment of each direction which the aircraft should provide, calculating the difference between the real-time parameters of the aircraft and the control expectation, and realizing the conversion and the output from the error value to the force or the moment by using the principle of a dynamic balance equation based on a PID (proportion integration differentiation) control algorithm to form closed-loop control.
(3) The control distribution module is used for receiving the force and the moment output by the control module, and calculating control instructions of different actuators according to the current motion state of the aircraft, wherein the control instructions comprise the rotating speed or the torque of a power motor, the rotating and tilting angle of a tilting rudder, the steering engine deflection angle of a control surface and the like.
(II) the optional control system has a high correlation with the aircraft layout, and based on the tiltrotor aircraft layout described above, the optional control system scheme includes: the vector control yaw, the large tilting angle negative tension and the asynchronous tilting realize decoupling of pitching and acceleration control, and the synchronous tilting winglet and the nacelle realize transitional flight. The method specifically comprises the following steps:
(1) yaw is controlled by a powertrain vector: in the tilting transition process, the angle and power output of the winglets 12 or the shortages 32 on the two sides can be synchronously or asynchronously changed, so that the effect of generating unbalanced yaw moment is achieved, and yaw control is realized;
(2) negative tension of large tilting angle: the winglet and the nacelle can be controlled to a large angle in the deceleration transition conversion stage, the horizontal component of the power can be changed into resistance while the lift force is generated, the transition deceleration is increased, the transition conversion time is reduced, and the air distance of transition conversion is shortened;
(3) decoupling of pitch control and acceleration control is achieved by the asynchronous tilting winglet and the nacelle: in the acceleration tilting transition stage, the flying speed of the aircraft is low, and the pitching attitude control is realized by matching with a power system. It is therefore possible to consider the winglet tilting in advance, so that the forward acceleration is mainly provided by the winglet power, the nacelle power system then forming with the strut power system a four-rotor control of the pitch attitude of the aircraft. Therefore, the pitching attitude control and the forward acceleration control of the aircraft with complex aerodynamic characteristics in the tilting transition stage can be decoupled, and the control complexity is reduced;
(4) the synchronous tilting winglet and the nacelle realize transitional flight, and decoupling of pitching control and acceleration control is realized by a control distribution module.
The foregoing embodiments have been provided for the purpose of illustrating the general principles of the present invention in further detail, and are not to be construed as limiting the scope of the invention, but are merely intended to cover any modifications, equivalents, improvements, etc. based on the teachings of the invention.
Claims (10)
1. A tiltrotor aircraft, comprising: the aircraft comprises a wing (1), a fuselage (2), a horizontal tail (3), a vertical tail (4), a landing gear (5) and a stay bar (6); the wing (1) comprises a main wing (11) and winglets (12) at two sides, the horizontal tail (3) comprises a horizontal stabilizer (31) and nacelle (32) at two sides, and left and right stay bars (6) are respectively arranged at the front sides of the left and right wings of the main wing (11); the power systems are arranged on the end parts of the stay bars (6), the winglets (12) on the two sides and the nacelle (32) on the two sides; the power system arranged on the winglets (12) on the two sides and the nacelle (32) on the two sides changes the thrust direction along with the tilting of the winglets (12) on the two sides and the nacelle (32) on the two sides; the power system arranged at the end parts of the left and right stay bars (6) inclines along with the rotation axis in the direction away from the machine body.
2. A tiltrotor aircraft according to claim 1, wherein the first rotor (71) mounted on the right strut (6) is rotated counter-clockwise, the second rotor (72) mounted on the right winglet (12) is rotated counter-clockwise, the third rotor (73) mounted on the right nacelle (32) is rotated clockwise, the fourth rotor (74) mounted on the left nacelle (32) is rotated counter-clockwise, the fifth rotor (75) mounted on the left winglet (12) is rotated clockwise, the sixth rotor (76) mounted on the left strut (6) is rotated clockwise, and the rotation is defined as a top view when the winglet (12) and nacelle (32) are in a vertical position.
3. A tiltrotor aircraft according to claim 2, wherein the direction of rotation of the first rotor (71) and the sixth rotor (76) ensures that the aircraft is subjected to aerodynamic moments in the clockwise and anticlockwise directions respectively, and wherein the contribution of the component of force generated by camber to the yaw moment of the machine is consistent with the contribution of aerodynamic drag moment caused by the direction of rotation when the heading is controlled by varying the rotational speed of a rotor at low speed.
4. A tiltrotor aircraft according to claim 3, wherein the third rotor (73) is counter-rotated to the first rotor (71), and the fourth rotor (74) is counter-rotated to the sixth rotor (76), and wherein a typical quadrotor condition is established when the symmetric rotor control is balanced when either the second rotor (72) or the fifth rotor (75) fails.
5. A tiltrotor aircraft transition control device, comprising: the control device is used for controlling the tiltrotor aircraft according to any one of claims 1-4, and specifically comprises a command filtering module, a control module and a control distribution module;
the instruction filtering module is used for receiving a host control instruction, and combining sensor data and aircraft state processing to obtain a control expectation for the aircraft motion state;
the control module is used for receiving the control expectation of the motion state of the aircraft output by the command filtering module, comprehensively processing the real-time motion state and the position data of the aircraft input by the sensor according to the pneumatic and quality characteristics of the aircraft, and obtaining the force and the moment of each direction which the aircraft should provide, calculating the difference between the real-time parameters of the aircraft and the control expectation, and realizing the conversion and the output from the error value to the force or the moment based on a control algorithm;
the control distribution module is used for receiving the force and the moment output by the control module and calculating the control instructions of different actuators according to the current motion state of the aircraft.
6. A tiltrotor aircraft transition control device according to claim 5, wherein the control of aircraft motion state is expected to include directional velocity, acceleration, attitude angle, angular velocity.
7. The tiltrotor aircraft transition control device according to claim 5, wherein the command filter module comprises a constraint envelope construction strategy and an optimization principle strategy; limiting the envelope, namely comprehensively calculating a tilting angle-airspeed curve according to the multi-dimensions of the aircraft such as self-quality characteristics, aerodynamic characteristics, dynamic characteristics, control safety, structural strength limitation and the like under various aircraft states; the optimization principle is a set of a series of optimization judgment conditions, and aims to screen out and obtain an optimal control expectation in the current state in a limiting envelope according to the actual flight state of the aircraft.
8. The tiltrotor aircraft transition control device according to claim 7, wherein the constraint envelope construction strategy comprises:
traversing the tilting angle, and obtaining the power ranges of the lift force and the thrust direction by using a trigonometric function based on the known power characteristics of the aircraft;
traversing the attack angle and airspeed of the aircraft, and obtaining the lift force and the resistance by using an aerodynamic calculation method based on the known aerodynamic characteristics of the aircraft;
assuming that the tilting transition Cheng Fei altitude is kept unchanged, obtaining the available airspeed range of the aircraft in each attack angle and tilting angle state by using a vertical direction dynamic balance equation based on the known aircraft mass characteristics;
based on the calculated airspeed range, the aircraft horizontal acceleration is calculated using a horizontal dynamic equilibrium equation.
9. The tiltrotor aircraft transition control device according to claim 7, wherein the optimization principle strategy comprises: the principle of the fastest transition, the principle of energy saving and the principle of passenger comfort, and the weight duty ratio of each principle is automatically adjusted under different flight tasks or aircraft states.
10. A tiltrotor aircraft transition control device, the control device being applied to a tiltrotor aircraft according to any of claims 1-4, comprising:
in the tilting transition process, the angles of the winglets at two sides or the nacelle at two sides are changed synchronously or asynchronously, and the power output is changed;
controlling the winglet and the nacelle to a large angle in a deceleration transition stage;
the asynchronous tilting winglet and the nacelle realize decoupling of pitch control and acceleration control;
and the transition flight is realized by synchronously tilting the winglet and the nacelle.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202310755162.5A CN116853487A (en) | 2023-06-25 | 2023-06-25 | Tilt rotor aircraft and transition conversion control device thereof |
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| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202310755162.5A CN116853487A (en) | 2023-06-25 | 2023-06-25 | Tilt rotor aircraft and transition conversion control device thereof |
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| CN116853487A true CN116853487A (en) | 2023-10-10 |
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| CN202310755162.5A Pending CN116853487A (en) | 2023-06-25 | 2023-06-25 | Tilt rotor aircraft and transition conversion control device thereof |
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