CN113184166A - Stability augmentation control method for flying wing layout unmanned aerial vehicle - Google Patents

Stability augmentation control method for flying wing layout unmanned aerial vehicle Download PDF

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CN113184166A
CN113184166A CN202110353313.5A CN202110353313A CN113184166A CN 113184166 A CN113184166 A CN 113184166A CN 202110353313 A CN202110353313 A CN 202110353313A CN 113184166 A CN113184166 A CN 113184166A
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rudder
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邹阳
张瞿辉
肖奔
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Chengdu Aircraft Industrial Group Co Ltd
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    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
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Abstract

The invention relates to the technical field of unmanned aerial vehicle flying wing layout, in particular to a method for controlling stability augmentation of an unmanned aerial vehicle with flying wing layout, which comprises the steps of arranging eight control surfaces on two wings of the unmanned aerial vehicle, arranging a control surface deflection virtual line on each wing, and enabling all the control surfaces to deflect upwards and downwards along the corresponding control surface deflection virtual lines; and designing a control law of the unmanned aerial vehicle controller on the basis of a transverse course linear track tracking control mode. The technical scheme is designed for decoupling course stability augmentation control of the flying wing unmanned aerial vehicle, can solve the problem of course divergence of the flying wing unmanned aerial vehicle, and is fast, stable and practical compared with the existing control method.

Description

Stability augmentation control method for flying wing layout unmanned aerial vehicle
Technical Field
The invention relates to the technical field of unmanned aerial vehicle flying wing layout, in particular to a method for controlling stability augmentation of an unmanned aerial vehicle with flying wing layout.
Background
The aerodynamic layout of the airplane refers to the configuration of the appearance of the airplane and the arrangement form of large components, including the positions and the shapes of components directly related to aerodynamic force, such as wings, a fuselage, an air inlet channel and the like. The aircraft aerodynamic layout can be divided into a conventional layout, a canard layout, a tailless layout, a three-wing layout and a flying wing layout according to wing surfaces, wherein the flying wing layout is the best choice for the aerodynamic layout integrated design, and has the advantages of high aerodynamic efficiency, large lift-drag ratio, small radar reflection sectional area, good stealth performance and the like. Flying wing overall arrangement unmanned aerial vehicle, the design that the wing body fuses into an organic whole makes its infiltration area reduce 33% than the aircraft of conventional overall arrangement, and the aircraft wholly becomes a lift face, has obtained bigger lift-drag ratio, has reduced the oil consumption, has increased the time of flight, in addition, because it has adopted high wing body to fuse the technique, has good radar stealth performance, consequently, flying wing overall arrangement unmanned aerial vehicle relies on self characteristics to become the hot door of unmanned aerial vehicle research and development. The current flying wing layout has a series of problems, the most obvious of which are: because of no vertical tail and the rudder generally adopts a split resistance rudder, obvious coupling between rolling channels and yawing channels is easily caused, so that an airplane control system needs to be controlled before a single channel is designed, and the control is a great problem.
Specifically, the flying wing layout is a tailless aerodynamic layout form, and previous research experiences show that the flying wing layout unmanned aerial vehicle has the characteristics of insufficient stability, inconvenient arrangement of control surfaces, and unstable longitudinal direction and course or insufficient static stability. Although adopt the all-wing aircraft overall arrangement can promote unmanned aerial vehicle's aerodynamic characteristic by a wide margin, its vertical manipulation arm of force is short nevertheless for saturation phenomenon appears easily in the elevator, and the static poor stability of yaw direction if appear comparatively obvious sideslipping in the horizontal sideslip, makes the coupling between each passageway more serious, has further increased flight control's the degree of difficulty.
At present, a plurality of methods, such as an adaptive control method, a nonlinear dynamic inverse control and the like, are developed for the control research of a flying wing layout unmanned aerial vehicle, so that a good control effect is obtained, and control indexes can reach design indexes, but in the engineering application of actual flight, some modern advanced control methods like the above are difficult to apply to the actual flight, and only stay in theoretical research, and lack of practice for the engineering application proves that there are two main reasons: firstly, the modern control method cannot carry out engineering verification, and the physical relation of flight parameters of the unmanned aerial vehicle cannot be proved; secondly, the cost of the unmanned aerial vehicle with the flying wing layout is too high, the cost is too high, and the risk is too high when the modern control method is applied to the unmanned aerial vehicle with the flying wing layout.
Disclosure of Invention
The invention provides a stability augmentation control method for a flying wing layout unmanned aerial vehicle aiming at some technical difficulties, aiming at the characteristic of static and unstable transverse course of the flying wing layout, the course of the unmanned aerial vehicle is augmented and controlled, and the actual flight trial verification is carried out by using the flying wing layout unmanned aerial vehicle, so that the correctness of the control method is obtained.
The purpose of the invention is realized by the following technical scheme:
a stability augmentation control method for a flying wing layout unmanned aerial vehicle is characterized by comprising the following steps: the method comprises the steps of controlling control surface distribution and setting a control rate;
the control surface allocation comprises the following steps: eight control surfaces are arranged on two wings of the unmanned aerial vehicle, a control surface deflection virtual line is arranged on each wing, the deflection virtual line is not a solid line, and only one virtual straight line for determining the installation position of the control surface is arranged, namely, the eight control surfaces are respectively the control surface deflection virtual line on the left wing, the left elevators L1, the left ailerons L2, the left inner side differential type rudder L3 and the left outer side differential type rudder L4 which are arranged at intervals from inside to outside in sequence, and the right elevators R1, the right ailerons R2, the right inner side differential type rudder R3 and the right outer side differential type rudder R4 which are arranged at intervals from inside to outside in sequence along the control surface deflection virtual line on the right wing; all the control surfaces can be deflected upwards and downwards along the corresponding control surface deflection virtual line. In the technical scheme, the flying wing layout unmanned aerial vehicle is not provided with a vertical tail, all control surfaces can only be placed on wings, the control on the course needs to generate resistance through deflection of the control surfaces, and the resistance is multiplied by a moment arm to form yawing moment to change the course of the unmanned aerial vehicle.
The setting control law: to realize unmanned aerial vehicle autonomous control, a controller needs to be arranged for the unmanned aerial vehicle, and a control law of the unmanned aerial vehicle controller is designed on the basis of a horizontal-course linear trajectory tracking control mode. The linear track following control is generally used in an air cruise section, an airplane flies along a specified air route (route point), the lateral offset distance between the airplane and the air route is eliminated in the flying process, the lateral offset between the airplane and the air route is calculated in real time by a navigation algorithm, the transverse distance between the airplane is changed by controlling a rolling angle, and the track course of the airplane is corrected at the same time, and a control principle structure block diagram is shown in fig. 2. For a classical control method, an aircraft with a statically unstable transverse heading generally adopts two stability augmentation control methods, one is to introduce lateral overload for stability augmentation, and the other is to introduce a lateral slip angle for stability augmentation, in practical application, the signal noise of a lateral overload sensor is too large, and the signal oscillation frequency is too fast, so that the design of a stability augmentation controller is carried out by adopting the lateral slip angle, and the control rate of the controller is as follows:
Figure BDA0003001036440000031
Figure BDA0003001036440000032
Figure BDA0003001036440000033
Figure BDA0003001036440000034
Figure BDA0003001036440000035
Figure BDA0003001036440000036
Figure BDA0003001036440000037
wherein, deltaaThe control deflection of the aileron control surface, i.e. the deflection of the aileron control surface carried in the aileron control signal sent by the controller, can be delta according to the aileron control signalaSkewness of (d);
Figure BDA0003001036440000038
for the roll angle proportional coefficient, a mathematic model of the airplane can be built by utilizing matlab software, parameters can be adjusted by utilizing tools in the mathematic model, and the numerical value of the roll angle proportional coefficient can be obtained by fine adjustment according to actual conditions; phi is the roll angle, namely the actual roll angle of the airplane; phi is agSetting target value for the roll angle, namely the target value of the roll angle through internal operation of a controller, and calculating the deflection delta of the control surface of the ailerona；
Figure BDA0003001036440000039
For the roll angular velocity ratio coefficient, a mathematic model of the airplane can be built by utilizing matlab software, parameters can be adjusted by utilizing tools in the mathematic model, and the numerical value of the roll angular velocity ratio coefficient can be obtained by fine adjustment according to actual conditions; p is the roll rate.
Figure BDA00030010364400000310
For booking flight path angle proportionality coefficient, mathematic model of airplane can be built by matlab software, and parameter adjustment is carried out by using tools in mathematic model, and then fine adjustment is carried out according to actual conditions, so that numerical value of booking flight path angle proportionality coefficient can be obtained; psigThe target value given for the predetermined course angle, i.e. the control parameter input to the controller in accordance with the predetermined course, is used to calculate a given target value phi for the roll anglegIs heavyA parameter is required; psi is the track angle, i.e. the actual track angle of the aircraft;
Figure BDA00030010364400000311
for the lateral deviation speed proportionality coefficient, mathematic models of airplanes can be built by utilizing matlab software, parameters are adjusted by utilizing tools in the mathematic models, and the numerical value of the lateral deviation speed proportionality coefficient can be obtained by fine adjustment according to actual conditions;
Figure BDA00030010364400000312
the target value given for the lateral deviation speed, i.e. the target value of the lateral deviation speed given by the internal operation of the controller, is used for calculating the given target value phi of the roll anglegThe important parameters of (a);
Figure BDA00030010364400000313
is the side offset speed, i.e. the actual side offset speed of the aircraft;
Figure BDA00030010364400000314
for the side deviation speed integral coefficient, a mathematic model of the airplane can be built by utilizing matlab software, the reference is adjusted by utilizing tools in the mathematic model, and the numerical value of the side deviation speed integral coefficient can be obtained by fine adjustment according to the actual condition; dt is the integration time.
Figure BDA0003001036440000041
For the lateral offset proportionality coefficient, mathematic models of airplanes can be built by using matlab software, parameters can be adjusted by using tools in the mathematic models, and the numerical value of the lateral offset proportionality coefficient can be obtained by fine adjustment according to actual conditions; y isgThe target value given for the lateral offset is calculated by inputting into the controller a control parameter dependent on the distance between the actual course and the predetermined course of the aircraft
Figure BDA0003001036440000042
The important parameters of (a); and Y is the offset distance, namely the actual offset distance of the airplane.
δrThe deviation degree of the rudder is the deviation degree of the rudder control surface carried in the rudder control signal sent by the controller, and the deviation degree of the rudder control surface can be generated to be delta according to the rudder control signalrSkewness of (d);
Figure BDA0003001036440000043
for the yaw rate proportionality coefficient, mathematic models of airplanes can be built by using matlab software, parameters can be adjusted by using tools in the mathematic models, and the numerical value of the yaw rate proportionality coefficient can be obtained by fine adjustment according to actual conditions; r is the yaw rate, i.e. the actual yaw rate of the aircraft;
Figure BDA0003001036440000044
for the sideslip angle proportionality coefficient, a mathematic model of the airplane can be built by utilizing matlab software, parameters are adjusted by utilizing tools in the mathematic model, and the numerical value of the sideslip angle proportionality coefficient can be obtained by fine adjustment according to actual conditions; beta is the sideslip angle, i.e., the actual sideslip angle of the aircraft.
δeFor elevator control skewness, i.e. the degree of deflection of the elevator control surface carried in the elevator control signal from the controller, the elevator control surface can be brought to a magnitude delta in response to the rudder control signaleSkewness of (d);
Figure BDA0003001036440000045
for the pitch angle proportionality coefficient, mathematic models of the airplanes can be built by matlab software, the reference is adjusted by tools in the mathematic models, and the numerical value of the pitch angle proportionality coefficient can be obtained by fine adjustment according to actual conditions; theta is a pitch angle, namely the actual pitch angle of the airplane; thetagThe target value given for pitch angle, i.e. the target value for pitch angle given by the internal operation of the controller, is used to calculate the elevator control yaw δeThe important parameters of (a);
Figure BDA0003001036440000046
for the pitch angle damping coefficient, a mathematic model of the airplane can be built by utilizing matlab software, and tools in the mathematic model are utilized to carry outAdjusting parameters, and performing fine adjustment according to actual conditions to obtain a numerical value of a pitch angle damping coefficient; k is a radical ofqFor the pitch angle rate proportionality coefficient, mathematic models of the airplanes can be built by matlab software, the reference is adjusted by tools in the mathematic models, and the numerical value of the pitch angle rate proportionality coefficient can be obtained by fine adjustment according to actual conditions; q is the pitch angle rate, i.e. the actual pitch angle rate of the aircraft.
Figure BDA0003001036440000051
For the lifting speed proportionality coefficient, mathematic models of the airplanes can be built by utilizing matlab software, the parameters can be adjusted by utilizing tools in the mathematic models, and the numerical value of the lifting speed proportionality coefficient can be obtained by fine adjustment according to actual conditions;
Figure BDA0003001036440000052
the target value for the lifting speed is calculated by the controller internal operationgThe important parameters of (a);
Figure BDA0003001036440000053
is the lifting speed, i.e. the actual lifting speed of the aircraft;
Figure BDA0003001036440000054
for the lifting speed integral coefficient, a mathematic model of the airplane can be built by utilizing matlab software, the reference is adjusted by utilizing tools in the mathematic model, and the numerical value of the lifting speed integral coefficient can be obtained by fine adjustment according to the actual condition.
Figure BDA0003001036440000055
For the height proportionality coefficient, mathematic models of the airplanes can be built by utilizing matlab software, parameters can be adjusted by utilizing tools in the mathematic models, and the numerical value of the height proportionality coefficient can be obtained by fine adjustment according to actual conditions; hgGiving a target value for altitude, i.e. input control of altitude of aircraft flight according to actual demandThe control parameter of the controller is to calculate a given target value of the lifting speed
Figure BDA0003001036440000056
The important parameters of (a); h is the altitude, i.e. the actual flying altitude of the aircraft.
The above equations represent the control of the state of the unmanned aerial vehicle, and the three equations at the top represent the control of the lateral offset of the unmanned aerial vehicle through the ailerons, so that the unmanned aerial vehicle is ensured to fly along the designated air route; the third equation is given by the lateral offset YgCalculating the given lateral deviation speed according to the actual lateral deviation distance Y of the airplane
Figure BDA0003001036440000057
The second equation is given by the lateral offset velocity
Figure BDA0003001036440000058
And actual lateral offset velocity of the aircraft
Figure BDA0003001036440000059
Calculating a given roll angle phi by combining the actual flight path angle of the airplaneg(ii) a The first equation is through a given roll angle phigCalculating the deflection delta of the ailerons required by the airplane by combining the actual rolling angle phi of the airplane and the temperature rise of the rolling angular velocity paTo control the drone to fly along the planned route. The fourth equation represents how many rudders are needed by the airplane to realize the course control of the unmanned aerial vehicle through the yaw rate and the sideslip angle. The fourth equation represents how much rudder deflection delta is needed by the airplane to be calculated according to the yaw rate r and the sideslip angle betarThe course control of the unmanned aerial vehicle is realized, wherein the rudder in the technical scheme comprises a left outer side differential rudder L4, a left inner side differential rudder L3, a right inner side differential rudder R3 and a right outer side differential rudder R4.
The last three equations represent the control of the aircraft altitude via the elevators, the first to last equation being the control of the aircraft altitude H givengCalculating a given lifting speed from the actual flying height H of the aircraft
Figure BDA0003001036440000061
The second to last equation is given by the lifting speed
Figure BDA0003001036440000062
Calculating a given pitch angle theta from the actual acceleration and deceleration of the aircraftg(ii) a The first to last equation is given by the pitch angle θgCalculating the deflection delta of the elevator which is needed by the airplane by combining the actual pitch angle theta of the airplane and the pitch angle speed qeTo control the drone to fly along a particular altitude.
Further, in the process of controlling the allocation of the control surfaces, the left elevator L1 and the right elevator R1 are deflected downwards at the same time to be positive; the left aileron L2 is deflected upwards, and the right aileron R2 is deflected downwards to be positive; the left outboard differential rudder L4 is biased up, the left inboard differential rudder L3 is biased down, and the right inboard differential rudder R3 and the right outboard differential rudder R4 are positive when held in the neutral position; the right outboard differential rudder R4 is biased upward, the right inboard differential rudder R3 is biased downward, and the left inboard differential rudder L3 and the left outboard differential rudder L4 are negative when held in the neutral position. Wherein "positive" and "negative" specifically refer to the yaw direction of the corresponding control surface.
The correlation equation of the control rate calculates the control surface required by the unmanned aerial vehicle to realize the control of the unmanned aerial vehicle through the difference between the actual state and the target state of the unmanned aerial vehicle, and calculates the rudder output (including delta)e、δa、δr) Then, the rudder output is required to be distributed to each control surface. Specifically, all control surfaces are enabled to be positive when deflected downwards, and the control surfaces are controlled to be distributed as follows: control plane deflection delta of elevatore＝δL1＝δR1It means that the left elevator L1 and the right elevator R1 have the same yaw direction and both have the same yaw degree equal to the elevator control yaw degree deltaeAileron control plane deflection deltaa＝δL2＝δR2It is shown that the yaw directions of the left aileron L2 and the right aileron R2 are the same, and the degree of deviation of both is equal to the control degree of deviation delta of the control surface of the aileronsa(ii) a Deflection of rudder surface
Figure BDA0003001036440000063
Indicating the control deflection delta of the rudderrIs positive (i.e. delta)r> 0), the left outer differential rudder L4 and the left inner differential rudder L3 are controlled to deflect, wherein the left outer differential rudder L4 and the left inner differential rudder L3 are controlled in opposite directions, and the left outer differential rudder L4 and the left inner differential rudder L3 are both controlled to be equal to the rudder control deflection δr(ii) a When the control deflection delta of the rudderrIs negative (i.e. delta)r< 0), the deflection of the right outer differential rudder R4 and the right inner differential rudder R3 is controlled, wherein the deflection of the right outer differential rudder R4 and the deflection of the right inner differential rudder R3 are opposite, and the deflection of the right outer differential rudder R4 and the deflection of the right inner differential rudder R3 are both equal to the control deflection δ of the ruddersr. Wherein, deltaL1For left-hand elevator deflection, deltaR1Right elevator deflection; deltaL2Left aileron deflection, deltaR2Right aileron skewness; deltaL3For left-inner differential rudder deflection, -deltaL4Is a left outer differential rudder deflection of-deltaR3Is a right-inboard differential rudder deflection, deltaR4The rudder deflection is a right-outer differential rudder deflection. Wherein the negative sign represents the direction.
When the left outer differential rudder L4 is deflected upward and the left inner differential rudder L3 is deflected downward, the aerodynamic forces of the right inner differential rudder R3 and the right outer differential rudder R4 deflected within ± 10 ° are dead sections, and therefore, further, the minimum deflection angle of the left outer differential rudder L4, the left inner differential rudder L3, the right outer differential rudder R4, and the right inner differential rudder R3 is 10 ° for stability enhancement.
The beneficial effect that this technical scheme brought:
1) according to the technical scheme, 8 control surface structures which can correspond to the upward deviation and the downward deviation of the deflection virtual line of the control surface are arranged on the wing, the course of the unmanned aerial vehicle can be effectively controlled by controlling the deflection angle of the corresponding control surface, the structure is simple, the realization is easy, repeated practice can be supported, and the unmanned aerial vehicle has the advantages of low cost and low risk;
2) this technical scheme is for flying wing overall arrangement unmanned aerial vehicle's course increases steady control decoupling zero design, can solve the problem that flying wing unmanned aerial vehicle course diverges, and for current control method, this technical scheme is quick, stable, practical, and has passed through flight test and verified, and whole flight process control is respond well, and flight control process is mild, and the sideslip angle is controlled throughout within a safe small-angle scope.
Drawings
FIG. 1 is a schematic illustration of an aircraft control surface layout;
FIG. 2 is a control flow block diagram of the controller;
FIG. 3 is a frequency domain bode plot obtained experimentally;
FIG. 4 is a plot of experimentally obtained root traces;
FIG. 5 is a graph of time domain response obtained experimentally;
fig. 6 is a flight curve of sideslip angle of the unmanned aerial vehicle obtained through experiments.
Wherein:
1. and deflecting the virtual line.
Detailed Description
The invention is further described in the following with reference to the drawings and examples, but it should not be understood that the invention is limited to the examples below, and variations and modifications in the field of the invention are intended to be included within the scope of the appended claims without departing from the spirit of the invention.
Example 1
The embodiment discloses a stability augmentation control method for an unmanned aerial vehicle with flying wing layout, which is used as a basic implementation scheme of the invention and comprises the steps of controlling control surface distribution and setting a control rate;
controlling the allocation of control surfaces: as shown in fig. 1, eight control surfaces are arranged on two wings of an unmanned aerial vehicle, and a control surface deflection virtual line is arranged on each wing; the eight control surfaces are respectively a virtual deflection line along the control surface on the left wing, a left elevator L1, a left aileron L2, a left inner differential rudder L3 and a left outer differential rudder L4 which are sequentially arranged from inside to outside at intervals, a virtual deflection line along the control surface on the right wing, and a right elevator R1, a right aileron R2, a right inner differential rudder R3 and a right outer differential rudder R4 which are sequentially arranged from inside to outside at intervals; all the control surfaces can deflect upwards and downwards along the corresponding control surface deflection virtual lines;
setting a control law: designing a control law of the unmanned aerial vehicle controller based on a transverse course linear track tracking control mode, and then
Figure BDA0003001036440000081
Figure BDA0003001036440000082
Figure BDA0003001036440000083
Figure BDA0003001036440000084
Figure BDA0003001036440000085
Figure BDA0003001036440000086
Figure BDA0003001036440000087
δaIn order to control the deflection of the control surface of the aileron,
Figure BDA0003001036440000088
is a roll angle proportional coefficient, phi is a roll angle, phigFor a given target value of the roll angle,
Figure BDA0003001036440000089
the roll angular velocity ratio coefficient is shown, and p is the roll angular velocity;
Figure BDA00030010364400000810
for booking the track angle scaling factor, #gA target value is given for a predetermined course path angle, ψ is the path angle,
Figure BDA00030010364400000811
for the side-shift speed scaling factor,
Figure BDA00030010364400000812
a target value is given for the side shift speed,
Figure BDA00030010364400000813
in order to set the lateral shift speed to the lateral shift speed,
Figure BDA00030010364400000814
is the side offset velocity integral coefficient, dt is the integral time;
Figure BDA00030010364400000815
is the coefficient of lateral offset proportionality, YgSetting a target value for the lateral offset, wherein Y is the lateral offset;
δris the control deflection of the rudder,
Figure BDA00030010364400000816
is the yaw rate proportionality coefficient, r is the yaw rate,
Figure BDA00030010364400000817
is the sideslip angle proportionality coefficient, beta is the sideslip angle;
δein order to control the deflection of the elevator,
Figure BDA0003001036440000091
is a pitch angle proportionality coefficient, theta isAngle of pitch, thetagA target value is given for the pitch angle,
Figure BDA0003001036440000092
for damping coefficient of pitch angle, kqIs a pitch angle rate proportionality coefficient, q is the pitch angle rate;
Figure BDA0003001036440000093
in order to obtain the proportional coefficient of the lifting speed,
Figure BDA0003001036440000094
a target value is given for the lifting speed,
Figure BDA0003001036440000095
in order to increase or decrease the lifting speed,
Figure BDA0003001036440000096
is the lifting speed integral coefficient;
Figure BDA0003001036440000097
to a high scale factor, HgGiven a target value for the height, H is the height.
This technical scheme has set up 8 rudder face structures that can correspond rudder face deflection virtual line top and bottom partially on the wing, through the deflection angle of control correspondence rudder face, can realize the effective control to unmanned aerial vehicle course, and its simple structure, easily realizes, and can support practice repeatedly, has low cost, low risk's advantage, and its concrete control method is: the lateral offset of the unmanned aerial vehicle is controlled through the ailerons, so that the unmanned aerial vehicle is ensured to fly along a designated air route; calculating a given lateral offset speed according to the given lateral offset and the actual lateral offset of the airplane; calculating a given roll angle by combining the given lateral offset speed and the actual lateral offset speed of the airplane with the actual heading angle of the airplane; calculating the deviation of the ailerons required by the airplane to control the unmanned plane to fly along the planned route by the given roll angle and the actual roll angle of the airplane and combining with the roll angle rate temperature rise; and calculating the number of rudder required by the airplane to realize the course control of the unmanned aerial vehicle through the yaw angular speed and the sideslip angle. And then, according to the calculation structure, the corresponding control surface is controlled to deflect, so that the stable control of the course of the unmanned aerial vehicle can be realized, the method is simple and reliable, compared with the existing control method, the method has the advantages of rapidness, stability and practicability, and the experimental results shown in figures 3-6 prove that the whole flight process of the unmanned aerial vehicle has a good control effect, the flight control process is mild, and the sideslip angle is always controlled within a safe small-angle range.
Example 2
The embodiment discloses a stability augmentation control method for an unmanned aerial vehicle with flying wing layout, which is a preferable implementation scheme of the invention, namely in the embodiment 1, in the process of controlling the allocation of control surfaces, a left elevator L1 and a right elevator R1 are deflected downwards to be positive at the same time; the left aileron L2 is deflected upwards, and the right aileron R2 is deflected downwards to be positive; the left outboard differential rudder L4 is biased up, the left inboard differential rudder L3 is biased down, and the right inboard differential rudder R3 and the right outboard differential rudder R4 are positive when held in the neutral position; the right outboard differential rudder R4 is biased upward, the right inboard differential rudder R3 is biased downward, and the left inboard differential rudder L3 and the left outboard differential rudder L4 are negative when held in the neutral position.
Further, all the control surfaces are enabled to be positive when deflected downwards, and the control surfaces are controlled to be distributed as follows: control plane deflection delta of elevatore＝δL1＝δR1(ii) a Aileron control surface deflection deltaa＝δL2＝δR2(ii) a Deflection of rudder surface
Figure BDA0003001036440000101
Wherein, the deviation degree delta of the control surface of the elevatoreI.e. the elevator control deflection, deltaL1For left-hand elevator deflection, deltaR1Right elevator deflection; aileron control surface deflection deltaaI.e. to control the deflection, delta, of the aileron control surfaceL2Left aileron deflection, deltaR2Right aileron skewness; rudder surface deflection delta of rudderrI.e. the control deflection, delta, of the rudderL3For left-inner differential rudder deflection, -deltaL4Is a left outer differential rudder deflection of-deltaR3Is a right-inboard differential rudder deflection, deltaR4The rudder deflection is a right-outer differential rudder deflection. Further, the minimum deflection angle of the left outer differential rudder L4, the left inner differential rudder L3, the right outer differential rudder R4, and the right inner differential rudder R3 is 10 °.

Claims (4)

1. A stability augmentation control method for a flying wing layout unmanned aerial vehicle is characterized by comprising the following steps: the method comprises the steps of controlling control surface distribution and setting a control rate;
the control surface allocation comprises the following steps: eight control surfaces are arranged on two wings of the unmanned aerial vehicle, and a control surface deflection virtual line is arranged on each wing; the eight control surfaces are respectively a virtual deflection line along the control surface on the left wing, a left elevator L1, a left aileron L2, a left inner differential rudder L3 and a left outer differential rudder L4 which are sequentially arranged from inside to outside at intervals, a virtual deflection line along the control surface on the right wing, and a right elevator R1, a right aileron R2, a right inner differential rudder R3 and a right outer differential rudder R4 which are sequentially arranged from inside to outside at intervals; all the control surfaces can deflect upwards and downwards along the corresponding control surface deflection virtual lines;
the setting control law: designing a control law of the unmanned aerial vehicle controller based on a transverse course linear track tracking control mode, and then
Figure FDA0003001036430000011
Figure FDA0003001036430000012
Figure FDA0003001036430000013
Figure FDA0003001036430000014
Figure FDA0003001036430000015
Figure FDA0003001036430000016
Figure FDA0003001036430000017
δaIn order to control the deflection of the control surface of the aileron,
Figure FDA0003001036430000018
is a roll angle proportional coefficient, phi is a roll angle, phigFor a given target value of the roll angle,
Figure FDA0003001036430000019
the roll angular velocity ratio coefficient is shown, and p is the roll angular velocity;
Figure FDA00030010364300000110
for booking the track angle scaling factor, #gA target value is given for a predetermined course path angle, ψ is the path angle,
Figure FDA00030010364300000111
for the side-shift speed scaling factor,
Figure FDA00030010364300000112
a target value is given for the side shift speed,
Figure FDA00030010364300000113
in order to set the lateral shift speed to the lateral shift speed,
Figure FDA00030010364300000114
is the side offset velocity integral coefficient, dt is the integral time;
Figure FDA00030010364300000115
is the coefficient of lateral offset proportionality, YgSetting a target value for the lateral offset, wherein Y is the lateral offset;
δris the control deflection of the rudder,
Figure FDA00030010364300000116
is the yaw rate proportionality coefficient, r is the yaw rate,
Figure FDA00030010364300000117
is the sideslip angle proportionality coefficient, beta is the sideslip angle;
δein order to control the deflection of the elevator,
Figure FDA00030010364300000118
is the pitch angle proportionality coefficient, theta is the pitch angle, thetagA target value is given for the pitch angle,
Figure FDA0003001036430000021
for damping coefficient of pitch angle, kqIs a pitch angle rate proportionality coefficient, q is the pitch angle rate;
Figure FDA0003001036430000022
in order to obtain the proportional coefficient of the lifting speed,
Figure FDA0003001036430000023
a target value is given for the lifting speed,
Figure FDA0003001036430000024
in order to increase or decrease the lifting speed,
Figure FDA0003001036430000025
is the lifting speed integral coefficient;
Figure FDA0003001036430000026
to a high scale factor, HgGiven a target value for the height, H is the height.
2. The stability augmentation control method for the flying wing layout unmanned aerial vehicle as claimed in claim 1, wherein: in the process of controlling the allocation of the control surfaces, the left elevator L1 and the right elevator R1 are deflected downwards to be positive at the same time; the left aileron L2 is deflected upwards, and the right aileron R2 is deflected downwards to be positive; the left outboard differential rudder L4 is biased up, the left inboard differential rudder L3 is biased down, and the right inboard differential rudder R3 and the right outboard differential rudder R4 are positive when held in the neutral position; the right outboard differential rudder R4 is biased upward, the right inboard differential rudder R3 is biased downward, and the left inboard differential rudder L3 and the left outboard differential rudder L4 are negative when held in the neutral position.
3. The stability augmentation control method for the flying wing layout unmanned aerial vehicle as claimed in claim 2, wherein: and making all the control surfaces to be positive when deflected downwards, and controlling the control surfaces to be distributed as follows: control plane deflection delta of elevatore＝δL1＝δR1(ii) a Aileron control surface deflection deltaa＝δL2＝δR2(ii) a Deflection of rudder surface
Figure FDA0003001036430000027
Wherein, deltaL1For left-hand elevator deflection, deltaR1Right elevator deflection; deltaL2Left aileron deflection, deltaR2Right aileron skewness; deltaL3For left-inner differential rudder deflection, -deltaL4Is a left outer differential rudder deflection of-deltaR3Is a right-inboard differential rudder deflection, deltaR4The rudder deflection is a right-outer differential rudder deflection.
4. The stability augmentation control method for the flying wing layout unmanned aerial vehicle as claimed in claim 2, wherein: the minimum deflection angle of the left outboard differential rudder L4, the left inboard differential rudder L3, the right outboard differential rudder R4, and the right inboard differential rudder R3 is 10 °.
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