CN106970531B - Method for determining mode conversion control strategy of tilt wing vertical take-off and landing unmanned aerial vehicle - Google Patents
Method for determining mode conversion control strategy of tilt wing vertical take-off and landing unmanned aerial vehicle Download PDFInfo
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Abstract
The invention discloses a method for determining a mode conversion control strategy of a tilt wing vertical take-off and landing unmanned aerial vehicle, which comprises the steps of establishing a nonlinear model of the tilt wing unmanned aerial vehicle, determining a mode conversion stage transition corridor by deducing a force and moment balance equation in a longitudinal symmetrical plane, carrying out balancing calculation through constraint conditions of forward conversion and reverse conversion, and determining the mode conversion control strategy according to a balancing result. The method is simple and effective in calculation and clear in physical significance, the influence of the adjustment of the modal conversion control strategy on the conversion process is obvious, and the conversion control strategy can be adjusted according to the object characteristics, so that the method has higher robustness and higher engineering use value.
Description
Technical Field
The invention belongs to the field of unmanned aerial vehicle control, and particularly relates to a method for determining a mode conversion control strategy of a tilt-wing vertical take-off and landing unmanned aerial vehicle.
Background
The wing-propeller integrated tilting unmanned aerial vehicle has the advantages of high aerodynamic efficiency of the fixed-wing unmanned aerial vehicle and capability of vertical take-off and landing of the rotor unmanned aerial vehicle, and has high application value and wide development prospect in the fields of military operations, civil disaster relief and the like. The GL-10 distributed power tilting wing vertical take-off and landing unmanned aerial vehicle similar to the NASA in America is characterized in that four motors are respectively installed on two sides of a wing, one motor is installed on each side of an empennage, the wing and the empennage tilt forwards and upwards horizontally in the mode conversion process, and the conversion of the flight modes of the tilting wing is realized through the combined action of the tension of a propeller of the wing and the empennage and aerodynamic force.
The specific gravity of pneumatic bearing and power bearing in the mode conversion process is closely related to the tilting angle and the flying speed of the wings. Under the same wing inclination angle condition, the speed ranges corresponding to different propeller pulling forces are different and are limited by conditions such as the maximum pulling force of the propeller and the wing stall, the wing inclination angle and the flight speed jointly determine the available flight envelope of the tilt-wing unmanned aerial vehicle and define the available flight envelope as a transition corridor. The forward conversion of the tilt wing is the process of turning the wing and the empennage from vertical to horizontal, and the reverse conversion is the process of turning the wing and the empennage from horizontal to vertical. The determination of the transition corridor mainly adopts a pitch angle balancing method, firstly a nonlinear simulation model of the tilt wing unmanned aerial vehicle is established, then a proper balancing constraint condition is selected according to different index requirements of forward conversion and backward conversion, and finally balancing calculation is carried out under different wing tilt angles, so that a feasible empennage deflection range, an engine body pitch angle variation range and a propeller throttle (thrust normalization) variation range are sought.
The flight process of the tilt-wing unmanned aerial vehicle can be divided into three modes of vertical take-off and landing, transition and fixed wing, and the tilt-wing unmanned aerial vehicle completely depends on power to bear in the vertical take-off and landing mode, completely depends on pneumatic bearing in the fixed wing mode, and depends on power to bear and pneumatic bearing to mutually convert in the transition mode. The larger the wing tilting angle is, the closer the wing tilting angle is to a vertical take-off and landing mode, the larger the specific gravity of power bearing is, and the higher the regulation efficiency based on the pitching attitude of the empennage propeller is; conversely, the smaller the wing tilting angle is, the closer the wing is to the cruising state, the higher the flying speed is, the larger the aerodynamic bearing specific gravity is, and the more effective the pitching attitude adjustment based on the empennage tilting is. The unmanned aerial vehicle control of the tilting wings has the difficulties that a mode conversion control strategy and a mode conversion control method are determined, and an optimized conversion track is selected in a transition corridor, so that the mode conversion of the tilting wings is realized stably and quickly.
Disclosure of Invention
Technical problem to be solved
In order to avoid the defects of the prior art, the invention provides a method for determining a mode conversion control strategy of a tilt-wing vertical take-off and landing unmanned aerial vehicle, and the stable and quick conversion between a vertical take-off and landing mode and a cruise mode is realized.
Technical scheme
A method for determining a mode conversion control strategy of an unmanned aerial vehicle with tilting wings and vertical take-off and landing is characterized by comprising the following steps:
step 1: obtaining an expression of a force and moment equation in the tilting process according to the stress analysis of the tilting wing unmanned aerial vehicle in a transition mode, and establishing a nonlinear simulation model of the tilting wing unmanned aerial vehicle;
the expression of the force and moment equation along the body shaft in the tilting process is as follows:
wherein:
Fx、Fy、Fzrespectively, component force along the axis of the body, Mx、My、MzRespectively, the moment around the body axis, theta represents the pitch angle of the body, i represents the number of the propeller, and TiRepresenting the tension of each propeller on the wing or empennage; locpropiThe position coordinates of each propeller on the wing or the empennage relative to the mass center of the machine body; thetaFIndicating wing tilt angle, thetaBRepresenting a empennage tilting angle, wherein the tilting angle is defined as an included angle between the empennage tilting angle and the x axis of the machine body; g represents body weight, L represents lift force and
d represents a resistance andρ is the atmospheric density, VaIs airspeed, SwIs the wing area, CLIs a coefficient of lift, CDAs coefficient of resistance, Mx_aero、My_aero、Mz'aero' is respectively the pneumatic moment around the body axis;
step 2: according to the wing stall boundary and the maximum available thrust limit of the propeller, the transition corridor boundary is obtained through balancing calculation based on a nonlinear simulation model, and the specific process is as follows:
obtaining the longitudinal force F of the unmanned aerial vehicle in the tilting process according to the formulas (1) and (2)xNormal force FzAnd pitching moment MyThe dynamic balance of (a) needs to satisfy the following relationship:
suppose θ in the transition processF＝θBThe tension of each propeller on the wing is equal, the tension of each propeller on the empennage is equal, and the maximum tension value T of each propeller on the wing is takenmax(ii) a The values of the wing propeller moment arm are as follows: x is the number ofprop＝xprop1The x coordinates of the thrust lines of each propeller of the tail are the same under a body coordinate system, the ratio of the pitching moment arms of each propeller of the wing to the pitching moment arms of each propeller of the tail is 1:4, and the value x of the moment arm of the propeller of the tail takes onprop＝4xprop1Solving the equation set (3) can respectively obtain the limiting conditions of the maximum tilting boundary and the minimum tilting boundary of the transition corridor:
and step 3: modifying constraint conditions of forward conversion and reverse conversion of the tilting wings, and determining propeller tension T of different state points through balancing calculation based on a nonlinear simulation modeliEmpennage dip angle thetaBAnd the feasible range of the pitch angle theta of the machine body; the method comprises the following specific steps:
the forward conversion aims at converting the wing inclination angle from vertical to horizontal as soon as possible, establishing the forward flying speed as soon as possible, and ensuring that the wing inclination angle does not fall high and is kept relatively stable in the tilting process; the force and moment constraint conditions and corresponding trim constraint conditions in the longitudinal symmetry plane are as follows:
wherein u is the velocity component along the longitudinal axis of the airframe, and h is the flying height; selecting a state point of each 10 degrees of wing inclination angle, determining a feasible empennage deflection range, an engine pitch angle variation range and a propeller tension variation range under each state point in the wing tilting process through balancing calculation by taking the wing inclination angle and the flight speed as known quantities;
the reverse conversion is opposite to the forward conversion, and in order to reduce the forward flying speed as soon as possible, a group of forward pitch angle instructions are required to be given, so that the machine body is always in a head-up state; before the inclination angle of the wing is rotated to 50 degrees, the inclination angle of the tail wing is basically maintained horizontal, and the constraint of the z-axis velocity component w of the body can be properly released, so that the tail wing has certain head raising capability in the normal direction; meanwhile, the height of the airplane needs to be kept constant in the reverse conversion process, and the speed component u of the airplane body shaft is restrained to decelerate the airplane; the force and moment constraint conditions and corresponding trim constraint conditions in the longitudinal symmetry plane are as follows:
similar to forward conversion, the feasible empennage deflection range, the variable range of the pitching angle of the engine body and the variable range of the propeller tension at each state point in the tilting process of the wings can be determined through balancing calculation;
and 4, step 4: determining control strategies of forward conversion and reverse conversion of the tilt wings according to performance index requirements of different stages of modal conversion; the method comprises the following specific steps:
the forward conversion can be divided into three phases: the aircraft can be accelerated in a downward direction within the range of the wing inclination angle of 90-50 degrees, the wing attack angle is large, the pitching moment coefficient determined by the pitching moment characteristic curve is basically maintained near a negative value, the corresponding flight speed in the stage is small, and therefore the downward moment generated pneumatically is small; the inclination angles of the wings and the empennage are large, the pitching adjustment capacity of the propeller is strong, and the selectable instruction range is large; when the wing inclination angle is between 50 and 20 degrees, the pitching adjustment capacity of the wing propeller is weakened, the aerodynamic head-lowering characteristic is obvious, and a head-raising instruction needs to be given to prevent the airplane from rapidly lowering head; the wing inclination angle is less than 20 degrees, the wing attack angle is reduced, the pneumatic head lowering moment is weakened, and the selectable instruction range is larger; in the forward conversion process, the empennage rotating shaft is far away from the center of gravity of the engine body, so that the empennage can tilt before the wings have large pitching adjustment capacity; the empennage basically turns to the horizontal position below the wing inclination angle of 30 degrees, and at the moment, the pitching adjusting capacity of the empennage propeller accelerator on the engine body is weaker, and the rotating speed of the empennage propeller can be forced to be zero;
similar to the forward conversion, the reverse conversion can also be divided into three stages, the aircraft is raised and decelerated at the inclination angle of the wing of 0-10 degrees, the incidence angle of the wing is small, and the selectable instruction range is large; a head-up command is given between 10 and 40 degrees to reduce the head-down characteristic; after 40 degrees, the airplane is kept in a head-up state, the flying speed is low, the aerodynamic characteristics are not obvious, and the selectable instruction range is large; in the process of reverse conversion, in order to raise the head of the engine body as soon as possible, the wings need to tilt before the empennage, when the inclination angle of the wings is less than 50 degrees, the rotating speed of the empennage propeller is zero, and at the moment, the pitching adjustment of the engine body is realized by the deflection of the wing propeller and the empennage rudder.
Advantageous effects
According to the method for determining the mode conversion control strategy of the tilt-wing vertical take-off and landing unmanned aerial vehicle, the tilt-wing vertical take-off and landing unmanned aerial vehicle can take off and land vertically and cruise efficiently, the application prospect is wide, the safety and the reliability of mode conversion of the tilt-wing vertical take-off and landing unmanned aerial vehicle are necessary conditions for the unmanned aerial vehicle, and forward conversion control strategies and reverse conversion control strategies need to be designed respectively to ensure that the transition process is in a transition corridor. According to the layout and the basic characteristics of the tilt-wing vertical take-off and landing unmanned aerial vehicle, a force and moment balance equation in a longitudinal symmetrical plane is established for the tilt-wing vertical take-off and landing unmanned aerial vehicle, a mode conversion stage transition corridor is determined, the link of nonlinear model linearization processing is reduced, and the calculation complexity is simplified; the transition corridor only defines the relation between the wing tilting angle and the flying speed, the influence of the pitch angle is not reflected, the feasible pitch angle trim value selection range under different inclination angles of the wing is determined according to the trim result, and the physical significance is clear; by utilizing the method for determining the mode conversion control strategy provided by the invention, according to the full-aircraft pitching moment characteristic curve, the influence on the conversion track can be realized by correcting the balancing rule, the conversion track can be effectively corrected by integrating the tension amplitude limit of the propeller and the expected pitch angle change range, and the method has good engineering use value.
Drawings
Fig. 1 is a force distribution diagram of a transition mode of a tilt-wing drone.
Fig. 2 is a schematic diagram of a mode conversion transition corridor.
Fig. 3 is a curve of the variation range of the pitch angle of the forward and reverse conversion bodies: (a) the curve of the change range of the pitch angle of the forward conversion machine body, and the curve of the change range of the pitch angle of the reverse conversion machine body.
FIG. 4 is a diagram of simulation results for forward and reverse conversion: the upper left graph is the relation between a pitch angle and a wing inclination angle, the upper right graph is the relation between a wing propeller trim accelerator and the wing inclination angle, the lower left graph is the relation between a flight speed and the wing inclination angle, and the lower right graph is the relation between a tail propeller trim accelerator and the wing inclination angle.
FIG. 5 is a diagram of simulation results of the improved forward and reverse conversion: the upper left graph is the relation between a pitch angle and a wing inclination angle, the upper right graph is the relation between a wing propeller trim accelerator and the wing inclination angle, the lower left graph is the relation between a flight speed and the wing inclination angle, and the lower right graph is the relation between a tail propeller trim accelerator and the wing inclination angle.
FIG. 6 is a flow chart of the present invention.
Detailed Description
The invention will now be further described with reference to the following examples and drawings:
the method comprises the following steps: as shown in fig. 1, according to the stress analysis of the tilt wing drone in the transition mode, an expression of a force and moment equation in the tilting process is obtained, and a nonlinear simulation model of the tilt wing drone is established.
In the vertical take-off and landing mode, the motion state of the airplane is changed by the tension and the inclination angle of the propellers on the wings and the empennage; changing the motion state of the airplane by means of the tail dip angle and the ailerons in the fixed wing mode; and the motion state of the airplane is changed by means of the tension and the inclination angle of the ailerons, the wings and the empennage propellers in the transition mode.
The expressions of the force and moment equations of the tilt-wing drone along the axis of the drone during tilting are as follows:
wherein:
Fx、Fy、Fzrespectively, component force along the axis of the body, Mx、My、MzRespectively, the moment around the body axis, theta represents the pitch angle of the body, i represents the number of the propeller, and TiRepresenting the tension of each propeller on the wing or empennage; locpropiThe position coordinates of each propeller on the wing or the empennage relative to the mass center of the machine body; thetaFIndicating wing tilt angle, thetaBRepresenting a empennage tilting angle, wherein the tilting angle is defined as an included angle between the empennage tilting angle and the x axis of the machine body; g represents body weight, L represents lift force and
d represents a resistance and
ρ is the atmospheric density, VaIs airspeed, SwIs the wing area, CLIs a coefficient of lift, CDAs coefficient of resistance, Mx_aero、My_aero、Mz"aero" is the aerodynamic moment about the body axis, respectively.
Step two: a transition corridor containing three modalities is shown in fig. 2. The transition corridor is an available flight envelope defined by two boundary curves in a longitudinal symmetry plane, the boundary curves are determined according to normal force, tangential force and pitching moment balance conditions, wherein the lower boundary is mainly calculated by the normal force balance conditions, the upper boundary is mainly calculated by the tangential force balance conditions, and the pitching angle in the transition process is basically kept near the horizontal level.
For equations (1) and (2), assuming | θ | ≦ 10 degrees in the transition process, the correlation term corresponding to sin θ is small, and the resistance is small compared to the normal force and can be ignored. The tension of each propeller on the wing is equal, and the tension of each propeller on the empennage is equal. The ratio of the pitching moment arm of each propeller of the wing to the pitching moment arm of each propeller of the empennage is 1:4, and the pitching moment arms of the propellers of the wing are the same, so that the value of the pitching moment arm of each propeller of the wing is xpropi＝xprop1The force arm of the empennage propeller takes x as valuepropi＝4xprop1. The simplified tilt wing unmanned aerial vehicle needs to satisfy the following relations in the tilt process of longitudinal force, normal force and pitching moment balance:
suppose θ in the transition processF＝θBThe tension of each propeller on the wing is equal, the tension of each propeller on the empennage is equal, and the maximum tension value T of each propeller on the wing is takenmax. Solving the equation set (3) can respectively obtain the limiting conditions of the maximum tilting boundary and the minimum tilting boundary of the transition corridor:
from the above formula, it can be seen that the upper boundary of the wing inclination angle depends on the tangential force, the lower boundary depends on the normal force balance condition, and the factors affecting the transition corridor also include the pitching moment characteristic of the unmanned aerial vehicle, the pitching angle, the propeller tension, the transition strategy, and the transition principle (whether to allow dropping, etc.), and the value of the influencing factors is modified to adjust the boundary of the transition corridor.
The designed transition corridor should have the following two characteristics: the larger the wing inclination angle is, the smaller the corresponding flying speed is, and the conversion relation between pneumatic bearing and power bearing is met; has a certain width, and ensures that the conversion process has certain disturbance resistance.
Step three: in the forward conversion process, in order to enable the aircraft to establish forward flight speed as soon as possible, the pitch angle needs to be kept relatively stable without falling high in the tilting process. The force and moment constraint conditions and corresponding trim constraint conditions in the longitudinal symmetry plane are as follows:
selecting the wing inclination angle of every 10 degrees as a state point, taking the wing inclination angle and the flight speed as known quantities, and determining the feasible variation range of the engine pitch angle at each state point corresponding to the boundary shown by two dotted lines in the figure 3(a) in the process of wing tilting through balancing calculation. Because the wing dip angle is between 50 and 20 degrees, the head lowering characteristic is obvious, the selectable range is small, the pitching moment balance constraint condition can be released to raise the head of the airplane, namely, the M is enabledy>0, the wing propeller throttle is increased to a certain extent to ensure sufficient head-up torque. The variation range of the body pitch angle after the trimming constraint is relaxed corresponds to the solid line segment in fig. 3(a), and has a larger selectable area in each stage of the conversion.
The reverse conversion is opposite to the forward conversion, and in order to reduce the forward flying speed as soon as possible, a group of forward pitch angle instructions needs to be given, so that the machine body is always in a head-up state. Before the inclination angle of the wing is rotated to 50 degrees, the inclination angle of the tail wing is basically maintained horizontal, and the constraint of the Z-axis velocity component w of the body can be properly released, so that the tail wing has certain head raising capability in the normal direction. Meanwhile, the speed component u of the X axis of the airframe needs to be kept high in the reverse conversion process, and the speed component u of the X axis of the airframe is restrained to enable the aircraft to decelerate. The force and moment constraint conditions and corresponding trim constraint conditions in the longitudinal symmetry plane are as follows:
similar to the forward conversion, the feasible empennage deflection range, the variable range of the engine pitch angle and the variable range of the propeller tension under each wing inclination angle in the wing tilting process can be determined through the balancing calculation.
Selecting a state point with the wing inclination angle of every 10 degrees as a state point, and operating the aircraftThe range of variation of the pitch angle of the body that is feasible at each state point during tilting of the wing is determined by trim calculations using the wing pitch angle and the flying speed as known quantities, corresponding to the boundaries indicated by the two dashed lines in fig. 3 (b). Similar to the corresponding phase of the forward transition, because of the significant low head characteristic and the small selectable range between 20 and 40 degrees of wing tilt angle, the aircraft can be raised by releasing the trim pitch angle rate constraint, i.e., the aircraft can be raised
The head-up moment is provided by increasing the propeller throttle of the wing. The improved variation range of the pitching angle of the machine body corresponds to a solid line segment in fig. 3(b), and the variation range has larger selectable areas at each stage of the conversion process.
Step four: and (3) rotating the given wing inclination angle instruction from 90 degrees to 0 degree at the speed of-10 degrees/second, and respectively establishing an interpolation table of an empennage inclination angle instruction, an engine body pitch angle instruction, a forward flying speed instruction, a wing propeller throttle and an empennage propeller throttle leveling value as expected input instructions of the system according to a balancing result. The simulation result of the forward conversion is shown in fig. 4(a), where the dashed line segment represents the desired input instruction. The diagram reflects that the response tracking effect of the pitch angle and the speed of the wing inclination angle between 40 degrees and 20 degrees is poor, the phenomena of head-down acceleration and head-up deceleration appear, the corresponding wing propeller throttle has the response characteristic of increasing firstly and decreasing secondly in the section, the tail wing propeller throttle basically tracks an expected input instruction, and the tail wing propeller throttle is zero below the wing inclination angle of 40 degrees.
Aiming at the pitch angle and the speed fluctuation in the forward conversion transition process, a control instruction needs to be adjusted according to the characteristics of the unmanned aerial vehicle, and the attitude and the speed fluctuation in the rotation process are reduced. A larger forward pitch angle command can be given at the wing inclination angles of 40 degrees and 50 degrees, the aircraft is raised first, the rapid head lowering characteristic of the aircraft is realized after the 40 degrees are slowed down, and the difference value between the actual pitch angle response and the expected pitch angle command is kept in a smaller range. The improved forward transition is shown in fig. 5(a), where the dashed line segment represents the desired input instruction.
The simulation result of the reverse conversion is shown in fig. 4(b), in which the dotted line segment represents the desired input instruction. The figure reflects that the wing inclination angle is between 20 degrees and 50 degrees, the pitch angle tracking deviation is large, and the response effect is poor. Similar to the forward transition, a larger forward pitch command may be given at wing pitches of 20 and 30 to increase the pitch response, and the modified reverse transition is shown in fig. 5(b), where the dashed line segment represents the desired input command.
Based on the pitching moment characteristic curve of the unmanned aerial vehicle, after the pitching angle instruction is corrected according to the simulation result, the pitching angle response change in the conversion process is stable, and the robustness and the engineering practicability of the control strategy can be further improved.
Claims (1)
1. A method for determining a mode conversion control strategy of an unmanned aerial vehicle with tilting wings and vertical take-off and landing is characterized by comprising the following steps:
step 1: obtaining an expression of a force and moment equation in the tilting process according to the stress analysis of the tilting wing unmanned aerial vehicle in a transition mode, and establishing a nonlinear simulation model of the tilting wing unmanned aerial vehicle;
the expression of the force and moment equation along the body shaft in the tilting process is as follows:
wherein:
Fx、Fy、Fzrespectively, component force along the axis of the body, Mx、My、MzRespectively, the moment around the body axis, theta represents the pitch angle of the body, i represents the number of the propeller, and TiRepresenting the tension of each propeller on the wing or empennage; locpropiFor each propeller on the wing or empennage relative to the bodyA position coordinate of the centroid; thetaFIndicating wing tilt angle, thetaBRepresenting a empennage tilting angle, wherein the tilting angle is defined as an included angle between the empennage tilting angle and the x axis of the machine body; g represents body weight, L represents lift force and
d represents a resistance and
ρ is the atmospheric density, VaIs airspeed, SwIs the wing area, CLIs a coefficient of lift, CDAs coefficient of resistance, Mx_aero、My_aero、Mz'aero' is respectively the pneumatic moment around the body axis;
step 2: according to the wing stall boundary and the maximum available thrust limit of the propeller, the transition corridor boundary is obtained through balancing calculation based on a nonlinear simulation model, and the specific process is as follows:
obtaining the longitudinal force F of the unmanned aerial vehicle in the tilting process according to the formulas (1) and (2)xNormal force FzAnd pitching moment MyThe dynamic balance of (a) needs to satisfy the following relationship:
suppose θ in the transition processF＝θBThe tension of each propeller on the wing is equal, the tension of each propeller on the empennage is equal, and the maximum tension value T of each propeller on the wing is takenmax(ii) a The values of the wing propeller moment arm are as follows: x is the number ofprop＝xprop1The x coordinates of the thrust lines of each propeller of the tail are the same under a body coordinate system, the ratio of the pitching moment arms of each propeller of the wing to the pitching moment arms of each propeller of the tail is 1:4, and the value x of the moment arm of the propeller of the tail takes onprop＝4xprop1Solving the equation set (3) can respectively obtain the limiting conditions of the maximum tilting boundary and the minimum tilting boundary of the transition corridor:
and step 3: modifying constraint conditions of forward conversion and reverse conversion of the tilting wings, and determining propeller tension T of different state points through balancing calculation based on a nonlinear simulation modeliEmpennage dip angle thetaBAnd the feasible range of the pitch angle theta of the machine body; the method comprises the following specific steps:
the forward conversion aims at converting the wing inclination angle from vertical to horizontal as soon as possible, establishing the forward flying speed as soon as possible, and ensuring that the wing inclination angle does not fall high and is kept relatively stable in the tilting process; the force and moment constraint conditions and corresponding trim constraint conditions in the longitudinal symmetry plane are as follows:
wherein u is the velocity component along the longitudinal axis of the airframe, and h is the flying height; selecting a state point of each 10 degrees of wing inclination angle, determining a feasible empennage deflection range, an engine pitch angle variation range and a propeller tension variation range under each state point in the wing tilting process through balancing calculation by taking the wing inclination angle and the flight speed as known quantities;
the reverse conversion is opposite to the forward conversion, and in order to reduce the forward flying speed as soon as possible, a group of forward pitch angle instructions are required to be given, so that the machine body is always in a head-up state; before the inclination angle of the wing is rotated to 50 degrees, the inclination angle of the tail wing is basically maintained horizontal, and the constraint of the z-axis velocity component w of the body is properly released, so that the tail wing has certain head raising capability in the normal direction; meanwhile, the height of the airplane needs to be kept constant in the reverse conversion process, and the speed component u along the longitudinal axis of the airplane body is restrained to decelerate the airplane; the force and moment constraint conditions and corresponding trim constraint conditions in the longitudinal symmetry plane are as follows:
similar to forward conversion, the feasible empennage deflection range, the variable range of the pitching angle of the engine body and the variable range of the propeller tension at each state point in the tilting process of the wings can be determined through balancing calculation;
and 4, step 4: determining control strategies of forward conversion and reverse conversion of the tilt wings according to performance index requirements of different stages of modal conversion; the method comprises the following specific steps:
the forward conversion can be divided into three phases: the aircraft can be accelerated in a downward direction within the range of the wing inclination angle of 90-50 degrees, the wing attack angle is large, the pitching moment coefficient determined by the pitching moment characteristic curve is basically maintained near a negative value, the corresponding flight speed in the stage is small, and therefore the downward moment generated pneumatically is small; the inclination angles of the wings and the empennage are large, the pitching adjustment capacity of the propeller is strong, and the selectable instruction range is large; when the wing inclination angle is between 50 and 20 degrees, the pitching adjustment capacity of the wing propeller is weakened, the aerodynamic head-lowering characteristic is obvious, and a head-raising instruction needs to be given to prevent the airplane from rapidly lowering head; the wing inclination angle is less than 20 degrees, the wing attack angle is reduced, the pneumatic head lowering moment is weakened, and the selectable instruction range is larger; in the forward conversion process, the empennage rotating shaft is far away from the center of gravity of the engine body, so that the empennage can tilt before the wings have large pitching adjustment capacity; the empennage basically turns to the horizontal position below the wing inclination angle of 30 degrees, and at the moment, the pitching adjusting capacity of the empennage propeller accelerator on the engine body is weaker, and the rotating speed of the empennage propeller can be forced to be zero;
similar to the forward conversion, the reverse conversion can also be divided into three stages, the aircraft is raised and decelerated at the inclination angle of the wing of 0-10 degrees, the incidence angle of the wing is small, and the selectable instruction range is large; a head-up command is given between 10 and 40 degrees to reduce the head-down characteristic; after 40 degrees, the airplane is kept in a head-up state, the flying speed is low, the aerodynamic characteristics are not obvious, and the selectable instruction range is large; in the process of reverse conversion, in order to raise the head of the engine body as soon as possible, the wings need to tilt before the empennage, when the inclination angle of the wings is less than 50 degrees, the rotating speed of the empennage propeller is zero, and at the moment, the pitching adjustment of the engine body is realized by the deflection of the wing propeller and the empennage rudder.
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| CN109946971B (en) * | 2019-04-04 | 2021-09-24 | 南京航空航天大学 | Smooth switching control method for transition section of tilt rotor unmanned aerial vehicle |
| CN111522356B (en) * | 2020-03-27 | 2021-06-04 | 北京航空航天大学 | Strong-robustness full-envelope integrated control method for tilt rotor unmanned aerial vehicle |
| CN112685832B (en) * | 2020-12-29 | 2022-06-17 | 清华大学 | Vertical take-off and landing fixed wing aircraft transition process corridor construction method |
| CN113859518A (en) * | 2021-10-28 | 2021-12-31 | 南京晓航机器人科技有限公司 | Multi-rotor unmanned aerial vehicle and method for improving speed and endurance |
| CN118034071A (en) * | 2024-04-15 | 2024-05-14 | 四川沃飞长空科技发展有限公司 | Aircraft control distribution method and device, aircraft, storage medium and product |
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