CN110576965A - Unmanned aerial vehicle layout with least control surface configuration and control method thereof - Google Patents
Unmanned aerial vehicle layout with least control surface configuration and control method thereof Download PDFInfo
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- B—PERFORMING OPERATIONS; TRANSPORTING
- B64—AIRCRAFT; AVIATION; COSMONAUTICS
- B64C—AEROPLANES; HELICOPTERS
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- B—PERFORMING OPERATIONS; TRANSPORTING
- B64—AIRCRAFT; AVIATION; COSMONAUTICS
- B64C—AEROPLANES; HELICOPTERS
- B64C13/00—Control systems or transmitting systems for actuating flying-control surfaces, lift-increasing flaps, air brakes, or spoilers
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Abstract
The invention relates to an unmanned aerial vehicle layout with least control surface configuration, belonging to the technical field of aircraft design and flight control; the unmanned aerial vehicle consists of a fuselage, wings on two sides, engines on two sides and a full-motion horizontal tail; the bilateral engines are respectively fixed below the wings on the two sides, and the full-motion horizontal tail is fixed at the tail part of the fuselage; the lateral direction control of the airplane is realized by using engines on two sides of the airplane body, and the longitudinal control of the airplane is realized by full-motion horizontal tail deflection. The control pneumatic control surface of the airplane only has one full-motion horizontal tail, so that the resistance generation surface during airplane design is greatly reduced and the flight resistance is reduced under the condition that an additional control executing mechanism is not required to be added; the control of the least control surface is adopted, the number of control actuating mechanisms is reduced to the minimum, and the energy consumption generated by control can be reduced to a certain extent; the full-motion horizontal tail is adopted to deflect integrally, so that the steering effect of control is increased as much as possible on the premise of reducing the area of the horizontal tail, and the airplane is controlled more quickly and effectively.
Description
Technical Field
the invention belongs to the technical field of aircraft design and flight control, and particularly relates to an unmanned aerial vehicle layout with minimum control surface configuration and a control method thereof.
Background
The existing traditional fixed wing unmanned aerial vehicle mostly adopts a mode of combining an aileron, an elevator and a rudder to carry out the flight control and control of the unmanned aerial vehicle. This mode is through installing pivoting's rudder face respectively on unmanned aerial vehicle wing, vertical fin and horizontal tail, forms aileron, rudder and elevator, and when the aircraft was flown, three rudder face pivoted changes the air pressure distribution of wing and fin to produce and make unmanned aerial vehicle around the fuselage axle pivoted aerodynamic moment of three differences, realize the change of unmanned aerial vehicle gesture, and then control states such as unmanned aerial vehicle's gesture and orbit.
however, this aircraft control approach has its own drawbacks that are difficult to avoid: 1. the mounting planes (vertical tail and horizontal tail) for connecting and controlling the pneumatic control surfaces increase the flight resistance of the airplane, and generate additional energy consumption; 2. too much control surface manipulation can bring too much energy consumption, which is not beneficial to the lifting of the unmanned aerial vehicle during the voyage and the voyage; 3. the control force of the pneumatic control surface is derived from aerodynamic force, indirect force control is performed, and the control and control efficiency of the airplane is obviously deteriorated under complex aerodynamic conditions of large attack angle, large sideslip and the like; 4. for some aircraft that do not allow for the structural integrity of the wing to be compromised (e.g., solar powered aircraft, etc.), the use of ailerons is not recommended; for some light-weight airplanes with large aspect ratio, the use of the ailerons can intensify the deformation of the structure, and meanwhile, the control rudder effect of the ailerons has the condition of strong nonlinearity and even reverse effect, so that the actual use effect is difficult to meet the requirement; 5. the arrangement of the pneumatic control surface is added on the basis of the original clean pneumatic appearance of the airplane, so that the structural weight of the airplane can be increased.
The above problems are inherent in the conventional unmanned aerial vehicle layout and the control method adopted, and are difficult to avoid. Therefore, to improve these problems, new fixed-wing drone layouts and control schemes need to be reconsidered and designed. In the process, most of new unmanned aerial vehicle layout and control methods still cannot get rid of the limitation of large-scale use of the pneumatic control surface, and the problems of large flight resistance, high energy consumption, poor control effect on certain aircrafts, additional structural weight increase and the like are still not well solved.
The invention discloses a control method for an unmanned aerial vehicle with a flying wing layout to apply a resistance rudder, which is introduced in patent publication No. CN104691742B and publication No. 2017.07.25. However, the drag rudder itself belongs to the aerodynamic control surface, which has the problems of the aerodynamic control surface itself, and if it is arranged on the fuselage, there is a problem of too small control effect, if it is arranged on the wing, the structural integrity of the wing is damaged, and it is difficult to apply the drag rudder to the light wing with large aspect ratio due to the structural flexibility problem. According to the control method, under the mixed operation of manual operation and flight control, the control signals calculated by the stability augmentation control law are superposed into the manual operation according to the control authority to generate the course channel control signals, so that the problem of large manual operation burden cannot be solved essentially, and on the contrary, even the instructions of manual operation and flight control operation are offset or intensified.
Disclosure of Invention
The technical problem to be solved is as follows:
In order to avoid the defects of the prior art, the invention provides an unmanned aerial vehicle layout with least control surface configuration and a control method thereof, which mainly solve the following problems in the existing unmanned aerial vehicle layout and control mode:
(1) For the problems of control redundancy and large energy consumption caused by more control execution mechanisms, the number of the control execution mechanisms is reduced to the minimum so as to reduce the control consumption;
(2) For the problems of increased aerodynamic resistance and increased structural weight caused by the installation surface required by the installation of the aerodynamic control surface, the invention cancels the arrangement of all additional installation surfaces;
(3) for the problem of the increase of the airplane resistance caused by the deflection of the aerodynamic control surfaces, the number of the aerodynamic control surfaces is reduced to one, and the area of the only aerodynamic control surface is reduced, so that the minimization of the aerodynamic resistance is realized.
(4) For the problem that the control force of the original control mode is the aerodynamic force of indirect force, the invention fully utilizes the power on two sides of the airplane body on the premise of not changing the layout of the existing unmanned plane, exerts the control design advantage of the direct force to the maximum and has strong applicability.
(5) In order to solve the problems of structural integrity damage and poor control efficiency of the ailerons, the invention cancels the arrangement of the ailerons and controls the rolling of the airplane in other modes.
(6) On the premise of achieving the advantages, the unmanned aerial vehicle control system and the control method thereof realize the function of the unmanned aerial vehicle control system, namely effective control of the unmanned aerial vehicle. In order to better realize the function, the invention additionally provides a 'stable flight' type flight control mode and a cascade attitude control law fused into an Extended State Observer (ESO), so that on one hand, the control burden of a flying hand in the control of the airplane in the control mode is reduced, the flight experience and the safety of the airplane are greatly improved, and on the other hand, the robustness, the rapidity and the anti-interference capability of the flight control are enhanced.
The technical scheme of the invention is as follows: an unmanned aerial vehicle layout with minimum control surface configuration is characterized in that the unmanned aerial vehicle consists of a fuselage, bilateral wings, bilateral engines and a full-motion horizontal tail; the bilateral engines are respectively fixed below the wings on the two sides, and the full-motion horizontal tail is fixed at the tail part of the fuselage; the lateral direction control of the airplane is realized by using engines on two sides of the airplane body, and the longitudinal control of the airplane is realized by full-motion horizontal tail deflection.
The further technical scheme of the invention is as follows:
A method for controlling a stable flight of an unmanned aerial vehicle layout with a minimum of control surface configuration: the method is characterized by comprising the following specific steps:
The method comprises the following steps: acquiring the position of a remote controller pushed by an operating hand, wherein the remote controller consists of a throttle lever, a pitching operating lever and a rolling operating lever;
step two: the aircraft control system acquires information of a remote controller control lever position, and then calculates an attitude angle instruction corresponding to the remote controller lever position instruction, wherein the calculation formula is as follows:
Roll angle phisetpointAttitude angle command:
φsetpoint＝manual_y×φmax
If phisetpoint＞φmax，φsetpoint＝φmax；
If phisetpoint<φmax，φsetpoint＝-φmax。
where, the input value of manual _ y represents the amount of the roll lever, and full to the right is1, fully beating to the left to be-1, and setting the neutral lever position to be 0; phi is amaxThe set maximum value of the roll angle instruction is obtained;
Pitch angle thetasetpointAttitude angle command:
θsetpoint＝manual_x×θmax
If thetasetpoint＞θmax，θsetpoint＝θmax；
If thetasetpoint<θmax，θsetpoint＝-θmax。
Wherein, the manual _ x is the input value to indicate the rod amount of the pitching joystick, the forward full is 1, the backward full is-1, and the neutral rod position is 0; thetamaxsetting a maximum pitch angle instruction value;
step three: and calculating according to each attitude angle command obtained in the step two by adopting the following control law, thereby obtaining the final control mechanism control quantity Dactuator:
errorattitude＝attitudesetpoint-attitudeactual
bodyratesetpoint＝Kp×errorattitude
e＝z1-bodyrateactual
u＝Dactuator
z1＝z1prev+(z2-2×ω×e+B×u)×dt
z2＝z2prev-(ω×ω×e)×dt
wherein, attributesetpointThe attitude angle command obtained in the step two is phi obtained by resolving in the step two when the roll angle attitude control is carried outsetpointWhen pitch angle attitude control is performed, the angle is theta calculated in the step twosetpoint；attitudeactualIs the actual real-time attitude angle of the airplane, and is the actual real-time rolling of the airplane when the rolling angle attitude control is carried outangle of rotation phiactualWhen pitch attitude control is performed, it is the actual real-time pitch angle θ of the aircraftactual；errorattitudeCalculating the obtained attitude angle error; kpThe outer ring proportionality coefficient is the attitude control law; dynamic ratesetpointIn order to calculate the angular velocity command of the airplane body axis system, when the roll angle attitude control is carried out, the angular velocity command is the roll angle velocity command p of the airplanesetpointWhen pitch angle attitude control is performed, it is a pitch angle velocity command q of the aircraftsetpoint；bodyrateactualFor the measured actual angular velocity of the aircraft, the roll angle attitude control is the real-time roll angular velocity p of the aircraftactualwhen pitch attitude control is performed, it is the actual real-time pitch angular velocity q of the aircraftactual；e，u，z1，z2All intermediate variables are iteratively calculated along with a time interval dt, and omega is a state estimation parameter; kdThe inner ring proportionality coefficient of the attitude control law; b is a disturbance compensation factor; z is a radical of1prevZ calculated for the last time point1Value z2prevZ calculated for the last time point2A value;
Dactuator is the actuation quantity of the aircraft actuating mechanism, and is the total throttle difference of the power on two sides when the roll angle attitude control is carried out; when pitch angle attitude control is carried out, the pitch angle attitude control is the deflection angle of a full-motion horizontal tail;
Step four: and deflecting and differentiating the full-motion horizontal tail and the power on the two sides according to the real-time control quantity obtained by calculating the control law in the three steps.
The further technical scheme of the invention is as follows: phi in the second stepmax≤45；θmax≤45°。
Advantageous effects
The invention has the beneficial effects that: the layout form that adopts full-motion horizontal tail and both sides power to combine together compares with the overall arrangement of present vast majority of aircraft, mainly has following advantage: 1. the control pneumatic control surface of the airplane only has one full-motion horizontal tail, so that the resistance generation surface during airplane design is greatly reduced and the flight resistance is reduced under the condition that an additional control executing mechanism is not required to be added; 2. the control of the least control surface is adopted, the number of control actuating mechanisms is reduced to the minimum, and the energy consumption generated by control can be reduced to a certain extent; 3. the full-motion horizontal tail is adopted to deflect integrally, so that the steering effect of control is increased as much as possible on the premise of reducing the area of the horizontal tail, and the airplane is controlled more quickly and effectively. 4. For a partially light weight high aspect ratio aircraft, the structural rigidity does not allow the placement of ailerons on the wing in actual flight, otherwise the deformation of the aircraft wing structure is aggravated and even the adverse effects of control are aggravated, leading to failure of the overall solution of the aircraft. 5. The horizontal heading state of the airplane is controlled by adopting a power differential mode on two sides of the airplane body, the control function of the original power device of the airplane can be fully exerted, the applicability is strong, in addition, the power belongs to direct force control, the dependence on the aerodynamic conditions of the airplane during flying is not needed, the available control quantity is large, and the control quantity can be kept effective under any flying condition. In a word, this overall arrangement form has splendid application effect to the unmanned aerial vehicle of big aspect ratio, low energy consumption, light.
the control method provided by the invention is matched with the unmanned aerial vehicle layout to achieve the preset target, and comprises an improved remote control mode and an improved control law. The effect is embodied in two aspects:
(1) the remote controller rod position instruction is converted into an attitude angle instruction from an original corresponding actuating mechanism action momentum instruction, so that the problem of offset or coupling of the transverse instruction and the course instruction which is possibly caused by adopting the dynamic differential motion of two sides of the airplane to carry out transverse and course control is solved, and for the airplane layout with smaller longitudinal stability margin and insufficient operating torque and the control mode with longer response time during transverse course control, the control burden of an operator is relieved, and the flight experience of the fixed-wing airplane is improved. Actual flight tests for hundreds of times show that the remote control mode realizes one innovation of controlling the airplane to fly by a traditional remote controller manually, can greatly improve the taking-off and landing performances of the airplane under the airplane control mode, can completely avoid the situation of failed taking-off caused by overlong control and adjustment time in the taking-off process, and has the success rate of 100 percent. In addition, the remote control mode makes it possible for a common fixed-wing flight fan without long-time training to quickly get on hand to control and fly the fixed-wing unmanned aerial vehicle.
(2) An Extended State Observer (ESO) is used in an inner ring of a cascade attitude control law, and angular acceleration contributions z from the inside and outside of the system except for a given angular acceleration acting quantity B x u2Estimating, and calculating z based on the initial control amount obtained by feedback calculation2To compensate in real time for possible disturbances to which the system is exposed. Compared with the traditional cascade PID attitude control law, the method greatly enhances the anti-interference capability of the control system, enables the aircraft to take off more stably in the taking-off stage, improves the performances of rapidity and the like of the control system on the premise of not increasing the jitter of the control mechanism, and has better tracking on the attitude in the air flight. The control law can not only function in a 'stable flight' type flight control mode, but also track and control attitude instructions given in other modes (such as giving a track instruction).
Drawings
Figure 1 is a schematic view of a typical drone layout in a minimum control plane configuration;
FIG. 2 is a schematic diagram of the overall flight control system;
FIG. 3 is a schematic diagram of an improved attitude control law of a control method combining a full-motion horizontal tail and two-side power;
FIG. 4 is a graph of a simulation result of a flying wing-like layout aircraft employing a flight control mode combining full-motion horizontal tail and two-side power according to the present invention;
FIG. 5 is a graph of the change in pitch attitude of an aircraft under both conventional attitude control laws and under the control laws of the present invention.
Detailed Description
The embodiments described below with reference to the drawings are illustrative and intended to be illustrative of the invention and are not to be construed as limiting the invention.
In the description of the present invention, it is to be understood that the terms "center", "longitudinal", "lateral", "length", "width", "thickness", "upper", "lower", "front", "rear", "left", "right", "vertical", "horizontal", "top", "bottom", "inner", "outer", "clockwise", "counterclockwise", and the like, indicate orientations and positional relationships based on those shown in the drawings, and are used only for convenience of description and simplicity of description, and do not indicate or imply that the device or element being referred to must have a particular orientation, be constructed and operated in a particular orientation, and thus, should not be considered as limiting the present invention.
the following is a typical application scenario of the unmanned aerial vehicle layout with minimum control surface configuration and the control method thereof, wherein the minimum control surface configuration is formed by combining a full-motion horizontal tail and power on two sides of an airplane:
For a conventional fixed-wing aircraft with lateral stability, effective control of its attitude and trajectory is achieved, requiring at least two aspects: 1. the control of the longitudinal flight state is relatively independent in the normal flight state; 2. coupling severe lateral flight state control in normal flight conditions. This requires that at least the steering actuators be selected during the selection and design of the control mechanisms and control methods that are effective in controlling both the longitudinal and lateral heading of the aircraft, respectively. Therefore, the invention provides that a mode of combining the full-motion horizontal tail and power on two sides of the airplane is adopted to realize effective control on the attitude of the airplane, specifically, the whole unmanned aerial vehicle is provided with only one pneumatic rudder, control force rejection provided by forward power accelerator differential motion on two sides of the airplane body is fully exerted, longitudinal control on the airplane is realized through full-motion horizontal tail deflection, and transverse heading control on the airplane is realized by using a mode of power differential motion on two sides of the airplane body.
The unmanned aerial vehicle layout is characterized in that the unmanned aerial vehicle consists of a body, wings on two sides, engines on two sides and a full-motion horizontal tail; the bilateral engines are respectively fixed below the wings on the two sides, and the full-motion horizontal tail is fixed at the tail part of the fuselage; the lateral direction control of the airplane is realized by using engines on two sides of the airplane body, and the longitudinal control of the airplane is realized by full-motion horizontal tail deflection.
Second, control method for stable flight
In the above unmanned aerial vehicle layout and the control method thereof, in the concrete implementation process, the following control method needs to be supplemented.
Firstly, the conversion of the control instruction corresponding to the lever quantity output by the remote controller is carried out; secondly, it is an improvement to the conventional cascade attitude control law.
When the full-motion horizontal tail is used for controlling the longitudinal motion of the airplane, the full-motion horizontal tail has small area and is very close to the airplane body in consideration of resistance reduction and weight reduction, so that the tail capacity is small, the steering effect is not large, and the control regulation time is slightly long. When the lateral motion control of the airplane is realized by adopting the dynamic differential motion of the two sides of the airplane body, the lateral motion coupling is serious, and the dynamic difference of the two sides of the airplane is used as a control quantity, so that the situation that the control effects are mutually offset or intensified is avoided by simultaneously outputting lateral and heading control instructions on a remote controller, and meanwhile, the control of the roll angle is realized by adopting the dynamic differential motion, and for most airplanes with stability, the given step instruction is slightly longer than the adjustment time of the system, which is between 2 and 4 seconds.
Based on the situation, when the airplane is controlled by adopting the layout of the full-motion horizontal tail and the power on two sides of the airplane body, the control process has unique characteristics, and even though an experienced flyer uses a remote controller to operate the airplane, the airplane still has great difficulty. Therefore, the invention further provides a 'stable flight' type control method, a control system is integrated into the manual operation process of the airplane to be matched with the realization of the control mode of the least control surface, and the specific flow is as follows:
1. Acquiring the position of a remote controller pushed by an operating hand, wherein the remote controller consists of a throttle lever, a pitching operating lever and a rolling operating lever;
2. The aircraft control system acquires information of a remote controller control lever position, and then calculates an attitude angle instruction corresponding to the remote controller lever position instruction, wherein the calculation formula is as follows:
Roll angle phisetpointAttitude set value:
φsetpoint＝manual_y×φmax
if phisetpoint＞φmax，φsetpoint＝φmax；
If phisetpoint<φmax，φsetpoint＝-φmax。
Wherein, the manual _ y is the input value and represents the rod amount of the rolling control rod, the right full is 1, the left full is-1, the neutral rod position is 0, and the linear change is along with the real rod position. Phi is amaxFor a set maximum roll angle command, the value may be reset for different aircraft configurations, typically not exceeding 45 °.
pitch angle attitude set value:
θsetpoint＝manual_x×θmax
if thetasetpoint＞θmax，θsetpoint＝θmax；
If thetasetpoint<θmax，θsetpoint＝-θmax。
Similarly, where the manual _ x is an input value representing the amount of the pitch joystick, it is filled with 1 forward, is filled with-1 backward, and has a neutral lever position of 0, which varies linearly with the actual lever position. ThetamaxFor a set maximum pitch command value, the value can be reset for different aircraft configurations, generally not exceeding 45 °.
accelerator set value:
Tsetpoint＝manual_z
The manual _ z is an input value representing the lever amount of the throttle lever, namely, the throttle set value is not additionally processed.
It can be seen that the command and steering of the yaw joystick is cancelled; the rod amount of the yaw control rod is not processed, and the yaw angle instruction of the airplane is not obtained through calculation, namely, only the rolling motion and the pitching motion of the airplane are controlled, and the yaw of the airplane is brought by the rolling motion, so that the situation that when the rolling instruction and the yaw instruction are given at the same time, the control effect is offset or intensified possibly caused by the superposition of the dynamic differential quantities on the two sides of the airplane which is the final executing mechanism is avoided.
The stick quantity of the remote controller corresponds to the set value of the airplane attitude by the algorithm, and the specific actuating mechanism action quantity (the deflection angle of the full-motion horizontal tail and the total throttle difference value of the power on two sides) is calculated by the control law in the flight control and is obtained and executed, so that the control burden of an operating hand is greatly reduced, the flight experience is improved, and the safety of the airplane flight is improved.
3. based on each attitude angle command obtained in step 2, the following control law is used to calculate a final control mechanism controlled variable datactor:
errorattitude＝attitudesetpoint-attitudeactual
bodyratesetpoint＝Kp×errorattitude
e＝z1-bodyrateactual
u＝Dactuator
z1＝z1prev+(z2-2×ω×e+B×u)×dt
z2＝z2prev-(ω×ω×e)×dt
wherein, attributesetpointThe attitude angle command obtained in the step two is phi obtained by resolving in the step two when the roll angle attitude control is carried outsetpointWhen pitch angle attitude control is performed, the angle is theta calculated in the step twosetpoint。attitudeactualis the actual real-time attitude angle of the aircraft, and is the actual real-time roll angle phi of the aircraft when roll angle attitude control is carried outactualWhen pitch attitude control is performed, it is the actual real-time pitch angle θ of the aircraftactual。errorattitudeCalculating the obtained attitude angle error; kpthe outer ring proportionality coefficient is the attitude control law; dynamic ratesetpointIn order to calculate the angular velocity command of the airplane body axis system, when the roll angle attitude control is carried out, the angular velocity command is the roll angle velocity command p of the airplanesetpointWhen pitch angle attitude control is performed, it is a pitch angle velocity command q of the aircraftsetpoint，bodyrateactualTo measureIs the real-time roll angular velocity p of the aircraft when roll angular attitude control is performedactualWhen pitch attitude control is performed, it is the actual real-time pitch angular velocity q of the aircraftactual，e，u，z1，z2Are all intermediate variables, z, iteratively calculated over time intervals dt1，z2All of which are 0, omega is a state estimation parameter, KdInner ring proportionality coefficient of attitude control law, B disturbance compensation factor, z1prevZ calculated for the last time point1Value z2prevZ calculated for the last time point2the value Dactuator is the actuation of the aircraft actuator, for a full-motion horizontal tail, substituted into the bodyratesetpointand a bodyrateactualthe pitch angle speed instruction and the real-time pitch angle speed are the deflection angles of the full-motion horizontal tail; substituted bodyrates for aircraft fuselage bilateral dynamicssetpointAnd a bodyrateactualthe rolling angular speed instruction and the real-time rolling angular speed are the total throttle difference of the power on the two sides; namely, when the roll angle attitude is controlled, the total throttle difference of the power on the two sides is obtained; when pitch angle attitude control is performed, it is the yaw angle of the full-motion horizontal tail.
Compared with the traditional cascade PID attitude control law, the calculation process of the control law adds an Extended State Observer (ESO) in the inner ring of the traditional cascade PID attitude control law, so that the residual actual angular acceleration real-time acting quantity z except the given angular acceleration acting quantity Bxu can be observed2. With this estimated value z2Z can be compensated based on the preliminarily calculated feedback control amount2The system is changed into an integral series control system, and the control effect of the system is greatly improved. The invention provides a more accurate and stable control law method for the realization of the airplane control combining the full-motion horizontal tail and the two-side dynamic differential motion.
4. and deflecting and differentiating the full-motion horizontal tail and the power on the two sides according to the real-time control quantity obtained by calculating the control law in the three steps. For a full motion horizontal tail, the controlled variable is its deviationThe angle of rotation; for the power on two sides, when two sides respectively have a power which is parallel to the axial line of the machine body and is forward, the control quantity is the accelerator difference D of the power on two sidesδtWhen two sides of the engine respectively have more than one power which is parallel to the axis of the engine body and moves forwards, the controlled variable is the throttle difference D of the two powers which are farthest away from the axis of the engine bodyδtfarthestThis is because the farther from the fuselage axis, the greater the yaw moment that can be produced given the throttle differential, the more efficient the distribution of the throttle differential momentum. In addition, differential momentum distribution needs to be carried out on the basis of the required total forward throttle, and the situation that the total forward power of the airplane is insufficient or too large due to the power differential distribution of the two sides of the airplane body is avoided.
Referring to fig. 2, fig. 2 shows a flight control mode of a 'stable flight' type, a cascade attitude control law of an Extended State Observer (ESO) is integrated, and a layout of an unmanned aerial vehicle with a minimum control surface configuration, in which a full-motion horizontal tail and two-side power are combined, which are in a position of the whole flight control system to play a role. Referring to fig. 3, fig. 3 is a schematic diagram of an improved attitude control law of a control method combining a full-motion horizontal tail and two-side power, and a specific state estimation process of the extended state observer is shown in a formula in the invention content; the real-time action of the airplane actuating mechanism refers to the deflection angle of the full-motion horizontal tail when the pitch angle is controlled; for the power on two sides of the airplane during roll angle control, the power on two sides is the throttle difference of respective thrust.
When the airplane flies, the operation amount of the airplane operation actuating mechanism is directly or indirectly generated according to a task instruction which is transmitted to the automatic pilot by a flying hand operation remote controller or a ground station. The process of generating the control effect on the final action amount respectively given to the full-motion horizontal tail and the power systems on two sides of the airplane is as follows:
In the normal flying process of the airplane, when the full-motion horizontal tail deflects downwards, an upward aerodynamic force is generated due to the change of the upper and lower air pressure of the full-motion horizontal tail, and a negative pitching moment is further generated, so that the airplane lowers head; when the full-motion horizontal tail deflects upwards, a downward aerodynamic force is generated, so that a positive pitching moment for raising the head of the airplane is further generated, and the head of the airplane is raised. Therefore, the pitch angle attitude of the airplane is controlled, and the longitudinal flight states of the airplane, such as the climbing angle and the like, are further controlled due to the mutual influence of the longitudinal flight states of the airplane.
In the normal flying process of the airplane, when the power on the two sides of the airplane is different, so that a yaw moment for enabling the airplane to turn right is generated, the airplane can yaw right, and meanwhile, the airplane has transverse static stability, so that under the action of a negative sideslip angle, the airplane can simultaneously generate a rolling motion of rolling right; similarly, when the power on the two sides of the airplane generates a yaw moment for enabling the airplane to turn left, the airplane can generate a negative rolling motion rolling left while yawing left. Therefore, the control of the roll angle and the yaw angle of the airplane can be realized in a mode of dynamic differential motion of two sides of the airplane, and other horizontal direction motion states of the airplane can be further controlled.
As shown in fig. 1, the solar unmanned aerial vehicle has a large aspect ratio and is provided with solar cells on the upper surface of the wing. Because the airplane has a large display ratio and a light structure, the solar cell pieces are paved on the upper surface of the wing, the structural integrity is not allowed to be damaged, and therefore, the aileron cannot be arranged, and the phenomena of nonlinear control aerodynamic force and even adverse effect caused by structural deformation are avoided. In addition, in order to facilitate the long-time cruising of the airplane, the plane generating aerodynamic drag is reduced as much as possible, and the drag of the airplane is reduced, so that the airplane is only provided with a full-motion horizontal tail with a small area to control the longitudinal state, and the horizontal direction is controlled by adopting a mode of differential rotation of the propellers at two sides. And a control signal of the full-motion horizontal tail is output to the steering engine through flight control, so that the rotation of the steering engine is controlled, and the rotation of the full-motion horizontal tail is further controlled. The two side power systems are respectively composed of a brushless motor and propellers, and the flight control also gives respective control signals of the two propellers so that the propellers rotate at corresponding rotating speeds to generate forward pulling power and yawing moment.
In a 'stable flight' type flight control mode, when the airplane takes off, the airplane keeps horizontal on a runway, a hand is operated to push an accelerator lever, two forward-pulling propellers of the airplane start to rotate quickly, forward-pulling thrust is generated, and the airplane starts to accelerate. And at the moment, the roll control lever, the yaw control lever and the pitch control lever are kept at neutral positions, and the full-motion horizontal tail and the propellers at two sides of the airplane can stabilize the roll angle, the yaw angle and the pitch angle of the airplane to be near 0 degree under the action of a control law algorithm.
and then the speed of the airplane is continuously increased along with the time, the designed takeoff speed of the airplane is reached, the pitching control lever is pushed up, and a positive pitch angle instruction is given, so that the airplane lifts under the action of the control law algorithm, the attack angle is increased, the lift force of the airplane is also increased at the same time, the airplane takes off from the ground, and the airplane climbs at the given pitch angle corresponding to the lever position of the remote controller. In the process, the roll control lever is always positioned at a neutral position, so that the roll attitude angle of the airplane is always kept near 0 degree under the action of the control system, and the takeoff process of the airplane is kept stable. The throttle lever should provide the takeoff throttle for the airplane to meet the normal takeoff power requirement of the airplane.
In the takeoff process of the airplane, if the structure deforms under the action of aerodynamic force or is disturbed by wind, so that the action quantity of attitude angular acceleration changes (namely extra disturbance moment exists) and the airplane attitude changes, the disturbance sum of the airplane and the outside can be extracted from input and output information in real time through an Extended State Observer (ESO), and is solved and estimated to obtain real-time z2and then real-time compensation is carried out on the basis of the control quantity obtained through the primary feedback calculation, so that the influence of disturbance from the outside or self is eliminated, and effective anti-interference is realized. In the specific control law calculation, ω 20, for the pitch angle control system: b-30, depending on the particular aircraft, KpGenerally between 0.5 and 2.5, the larger the full-motion horizontal tail area, the farther the distance from the center of gravity of the airplane and the larger the rudder effect are, the smaller the value is, and K isdThe value is also taken according to different specific airplanes, but the value is not too large so as to avoid oscillation of a control plane. For a roll angle control system: b is 30, KpAnd KdThe value is taken according to different specific airplanes, and a good control effect of power differential motion on two sides can be obtained generally between 0 and 2.
Simulation results show that compared with the traditional PID control law, the cascade attitude control law integrated with the ESO has improved control rapidity, robustness, stability and anti-interference capability.
Referring to fig. 4, it is a specific simulation result of a flying wing-like layout airplane adopting a flight control mode combining full-motion horizontal tail and two-side power. For a given 10-degree pitch angle command value, the control results of the traditional cascade PID control law and the cascade attitude control law fused with ESO are compared as follows:
The adjusting time t of the traditional cascade PID control law and the airplane pitch angle control system is adopteds1.4s, and the ESO-integrated cascade attitude control law is adopted, and the regulation time is t for a given pitch angle commandsThe control process has no overshoot and no static error, and the rapidity is improved by 12.5 percent when the time is 1.255 s.
giving a control command that the pitch angle is kept at 0 degrees, and applying a vertical gust to the airplane from the 1 st second to the 4 th second to make the airplane subjected to gust disturbance with the instantaneous attack angle from 0 degrees to 17.69 degrees at the 1 st second; referring to fig. 5, observing and comparing pitch angle attitude change curves of the aircraft under the traditional attitude control law and the control law of the invention, it can be seen that under the traditional cascade PID control law, the pitch angle attitude disturbance peak value is as high as-5.791 degrees, and the attitude adjustment is finished in about 6.19 seconds; and the cascade control law fused with ESO can restrain the attitude disturbance peak value of the pitch angle to-3.541 degrees, and the attitude adjustment is finished in about 5.3 seconds. The result shows that compared with the traditional PID control law, the stability and the anti-interference capability of the cascade attitude control law blended with the ESO are obviously improved.
The results of hundreds of flight tests show that the safety, stability and control precision of the airplane adopting the flight control mode combining full-motion horizontal tail and two-side power can be greatly improved by adding the flight control mode of the 'stable flight' type and the cascade attitude control law fused with ESO.
The above is the takeoff process of the airplane in the "stable flight" type flight control mode. It can be seen that in the mode, the control system and the operating hand are coordinated and matched with each other, the operation process of the operating hand is greatly reduced compared with the original operation process, the stable takeoff of the airplane can be realized, and the trouble that the flying hand brings extra operation burden to the possible accidental attitude angle disturbance in the takeoff process is avoided, so that the airplane is crashed.
Although embodiments of the present invention have been shown and described above, it is understood that the above embodiments are exemplary and should not be construed as limiting the present invention, and that variations, modifications, substitutions and alterations can be made in the above embodiments by those of ordinary skill in the art without departing from the principle and spirit of the present invention.
Claims (3)
1. An unmanned aerial vehicle layout with minimum control surface configuration is characterized in that the unmanned aerial vehicle consists of a fuselage, bilateral wings, bilateral engines and a full-motion horizontal tail; the bilateral engines are respectively fixed below the wings on the two sides, and the full-motion horizontal tail is fixed at the tail part of the fuselage; the lateral direction control of the airplane is realized by using engines on two sides of the airplane body, and the longitudinal control of the airplane is realized by full-motion horizontal tail deflection.
2. a method of stable flight control using the unmanned aerial vehicle layout of minimum control surface configuration of claim 1: the method is characterized by comprising the following specific steps:
The method comprises the following steps: acquiring the position of a remote controller pushed by an operating hand, wherein the remote controller consists of a throttle lever, a pitching operating lever and a rolling operating lever;
Step two: the aircraft control system acquires information of a remote controller control lever position, and then calculates an attitude angle instruction corresponding to the remote controller lever position instruction, wherein the calculation formula is as follows:
Roll angle phisetpointAttitude angle command:
φsetpoint＝manual_y×φmax
If phisetpoint＞φmax，φsetpoint＝φmax；
If phisetpoint<φmax，φsetpoint＝-φmax。
Where the manual _ y represents roll as an input valueThe lever amount of the operating lever is fully filled to the right to be 1, fully filled to the left to be-1, and the neutral lever position is 0; phi is amaxThe set maximum value of the roll angle instruction is obtained;
pitch angle thetasetpointAttitude angle command:
θsetpoint＝manual_x×θmax
if thetasetpoint＞θmax，θsetpoint＝θmax；
If thetasetpoint<θmax，θsetpoint＝-θmax。
Wherein, the manual _ x is the input value to indicate the rod amount of the pitching joystick, the forward full is 1, the backward full is-1, and the neutral rod position is 0; thetamaxSetting a maximum pitch angle instruction value;
Step three: and calculating according to each attitude angle command obtained in the step two by adopting the following control law, thereby obtaining the final control mechanism control quantity Dactuator:
errorattitude＝attitudesetpoint-attitudeactual
bodyratesetpoint＝Kp×errorattitude
e＝z1-bodyrateactual
u＝Dactuator
z1＝z1prev+(z2-2×ω×e+B×u)×dt
z2＝z2prev-(ω×ω×e)×dt
Wherein, attributesetpointThe attitude angle command obtained in the step two is phi obtained by resolving in the step two when the roll angle attitude control is carried outsetpointWhen pitch angle attitude control is performed, the angle is theta calculated in the step twosetpoint；ttitudeactualIs the actual real-time attitude angle of the airplane, and is the airplane when the attitude control of the roll angle is carried outActual real-time roll angle phiactualwhen pitch attitude control is performed, it is the actual real-time pitch angle θ of the aircraftactual；errorattitudeCalculating the obtained attitude angle error; kpThe outer ring proportionality coefficient is the attitude control law; dynamic ratesetpointin order to calculate the angular velocity command of the airplane body axis system, when the roll angle attitude control is carried out, the angular velocity command is the roll angle velocity command p of the airplanesetpointWhen pitch angle attitude control is performed, it is a pitch angle velocity command q of the aircraftsetpoint；bodyrateactualfor the measured actual angular velocity of the aircraft, the roll angle attitude control is the real-time roll angular velocity p of the aircraftactualWhen pitch attitude control is performed, it is the actual real-time pitch angular velocity q of the aircraftactual；e，u，z1，z2All intermediate variables are iteratively calculated along with a time interval dt, and omega is a state estimation parameter; kdthe inner ring proportionality coefficient of the attitude control law; b is a disturbance compensation factor; z is a radical of1prevZ calculated for the last time point1value z2prevZ calculated for the last time point2a value;
Dactuator is the actuation quantity of the aircraft actuating mechanism, and is the total throttle difference of the power on two sides when the roll angle attitude control is carried out; when pitch angle attitude control is carried out, the pitch angle attitude control is the deflection angle of a full-motion horizontal tail;
Step four: and deflecting and differentiating the full-motion horizontal tail and the power on the two sides according to the real-time control quantity obtained by calculating the control law in the three steps.
3. the stable-flight control method of claim 2: phi in the second stepmax≤45；θmax≤45°。
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