CN115933733A - Fixed wing unmanned aerial vehicle longitudinal altitude speed decoupling nonlinear control method - Google Patents
Fixed wing unmanned aerial vehicle longitudinal altitude speed decoupling nonlinear control method Download PDFInfo
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Abstract
The invention discloses a fixed wing unmanned aerial vehicle longitudinal height speed decoupling nonlinear control method, which comprises the following steps: determining a desired airspeed and a desired altitude for the drone; designing an expected height change rate, and calculating to obtain an expected track inclination angle through the expected height change rate and an expected flight airspeed; calculating to obtain a corresponding trim attack angle and a corresponding expected attitude angle according to the expected flight airspeed and the expected track inclination; designing a thrust-attack angle control law based on an incremental nonlinear dynamic inverse method according to the deviation of the trim attack angle and the actual attack angle to obtain a thrust control instruction; and designing a pitch angle control law based on an incremental nonlinear dynamic inverse method according to the deviation between the expected attitude angle and the actual attitude angle to obtain an elevator yaw instruction. The invention is applied to the field of unmanned aerial vehicles, not only solves the nonlinear influence of the aerodynamic characteristics in a large attack angle state, but also has better adaptability to model errors, is suitable for flying in a large attack angle state, effectively reduces the landing speed and improves the safety in the taking-off and landing process.
Description
Technical Field
The invention relates to the technical field of unmanned aerial vehicle control, in particular to a fixed wing unmanned aerial vehicle longitudinal height speed decoupling nonlinear control method.
Background
Autonomous take-off and landing are one of key technical links in the development process of the unmanned aerial vehicle, and are particularly important for the fixed-wing unmanned aerial vehicle with higher cruising speed. Compared with the landing process, the safety of the take-off process is relatively high, if a wheel type undercarriage is adopted, the key technology lies in the sliding deviation correction, and the technical verification can be completed through the ground test. And the landing process relates to unmanned aerial vehicle body gesture stable control, and height, rate of sinking accurate control reach processes such as race deceleration control, and the complexity is high, and the risk is great. Among them, the landing speed is one of important factors affecting safety. At present, large attack angle landing becomes a technical scheme for effectively reducing landing speed, and intensive research is carried out in various countries. Furthermore, accurate control of altitude and speed prior to landing is also critical.
The key points of large angle of attack flight are attitude stabilization and angle of attack maintenance. The attitude stabilization refers to attitude stabilization control in a large attack angle state. For a conventional fixed wing aircraft, aerodynamic parameters in a small attack angle range change along with an approximately linear trend of an attack angle, and when the attack angle is increased (smaller than a stalling attack angle), a certain degree of nonlinear characteristics appear; for unconventional aircrafts such as tailless layout and the like, the nonlinearity of the aerodynamic characteristic parameters of the aircraft along with the change of an attack angle and a control rudder deflection is strong. Therefore, in designing the attitude control law in a large angle of attack range, it is necessary to design the nonlinear control law by sufficiently considering the influence of the nonlinearity of the pneumatic model. The angle of attack keeps, indicates that unmanned aerial vehicle except realizing the stable attitude promptly, still needs to keep great angle of attack state. In a conventional control method, a logic scheme that an elevator controls a pitch angle and an accelerator controls the flying speed is adopted. Although the method can play a role in indirectly controlling the attack angle by controlling the speed, the method ignores the constraint relation among the states of the control targets on one hand, and the situation that the speed cannot be reduced or the altitude is reduced too fast when the throttle is zero can occur. Meanwhile, according to the balance of the lift weight, the speed and the attack angle have the unique corresponding relation, when the angle of attack is large, the change of the lift coefficient is slowed down, the speed change amplitude is reduced, the attack angle is more sensitive to the speed change at the moment, and the influence of the measurement error and the disturbance on the control effect is intensified.
On the other hand, in the height and speed control, the traditional control mode has a strong coupling effect, namely in the process of controlling the attitude by the elevator, the flying height can be changed, so that the speed change is influenced, and in the process of controlling the accelerator, the speed change can change the attack angle, so that the flying track and attitude are changed, and the flying height is further influenced. Although the total energy control method provides a decoupling strategy for controlling the total energy of the airplane and controlling the kinetic energy/potential energy balance relationship of an elevator or a pitch angle by using an accelerator, the total energy control method still has some defects: small-angle simplification is carried out in the theoretical derivation process, and the change of a large-angle range cannot be adapted; the adaptability to the nonlinear model is poor by adopting proportional-integral control; more control parameters are needed, and the complexity of parameter optimization is increased.
In conclusion, the conventional control method for the longitudinal speed and the height of the unmanned aerial vehicle has certain defects in the application scenes of accurate track control and low speed and large attack angle, and the actual requirements are difficult to meet.
Disclosure of Invention
Aiming at the defects in the design process of the existing unmanned aerial vehicle control method, the invention provides the nonlinear control method for decoupling the longitudinal altitude speed of the fixed-wing unmanned aerial vehicle, which has strong model parameter adaptability and low instruction calculation complexity and can effectively realize the altitude speed decoupling control in a large attack angle range.
In order to achieve the purpose, the invention provides a fixed wing unmanned aerial vehicle longitudinal altitude speed decoupling nonlinear control method, which is characterized by comprising the following steps of:
step 2, designing an expected altitude change rate according to the relation between the expected flying altitude and the actual altitude, and calculating to obtain an expected track inclination angle according to the expected altitude change rate and the expected flying airspeed;
step 3, calculating to obtain a corresponding trim attack angle and a corresponding desired attitude angle by using a general trim method according to the desired flight airspeed and the desired track inclination angle;
step 4, designing a thrust-attack angle control law based on an incremental nonlinear dynamic inverse method according to the deviation between the trim attack angle and the actual attack angle to obtain a thrust control instruction and realize attack angle-speed control;
and 5, designing a pitch angle control law based on an incremental nonlinear dynamic inverse method according to the deviation between the expected attitude angle and the actual attitude angle to obtain an elevator yaw instruction, and realizing attitude-track control.
The invention has the following beneficial technical effects:
1. the invention indirectly converts the direct commands of the expected flying airspeed and the expected flying altitude into the commands of the attack angle and the attitude angle by using the state balancing constraint relation, thereby realizing decoupling control. The thrust-attack angle control law enables the unmanned aerial vehicle to be maintained in a wide-range attack angle state, and the thrust-attack angle control law is combined with the attitude angle control law, so that the unmanned aerial vehicle is stably maintained in a current attack angle state, and track control is realized. The logic that thrust is used for directly controlling the attack angle is adopted, on one hand, the problem that the attack angle is sensitive to speed change in a large attack angle area is avoided, and the influence of speed disturbance on the attack angle is weakened; on the other hand, if the thrust is adopted to control the speed, and the speed is related to the attack angle, an intermediate link of attack angle change exists, a control loop can be increased, and the control efficiency is reduced;
2. in the invention, the attack angle and the attitude control are controlled by utilizing an incremental nonlinear dynamic inverse theory and acceleration feedback, so that the influence of part of model nonlinearity on the control effect is avoided, and the adaptability and the anti-interference capability of the control law are improved;
3. on one hand, in the change process of the attack angle, the pitch angle can be considered to reach the expected value, and the variable in the control process of the attack angle is reduced, so that the change meets the increment nonlinear dynamic inverse simplification condition; on the other hand, in the track angle control process, the track angle control process can be finely adjusted by adjusting the expected pitch angle, track angle change in the attack angle control process is compensated, and track control precision is improved.
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In order to more clearly illustrate the embodiments or technical solutions of the present invention, the drawings used in the embodiments or technical solutions of the prior art will be briefly described below, it is obvious that the drawings in the following description are only some embodiments of the present invention, and for those skilled in the art, other drawings can be obtained according to the structures shown in the drawings without creative efforts.
FIG. 1 is a logic structure diagram of a fixed wing unmanned aerial vehicle longitudinal altitude speed decoupling nonlinear control method in the embodiment of the invention;
FIG. 2 is a block diagram of an angle of attack control law in an embodiment of the present invention;
FIG. 3 is a block diagram of an attitude control law according to an embodiment of the present invention.
The implementation, functional features and advantages of the objects of the present invention will be further explained with reference to the accompanying drawings.
Detailed Description
The technical solutions in the embodiments of the present invention will be clearly and completely described below with reference to the drawings in the embodiments of the present invention, and it is obvious that the described embodiments are only a part of the embodiments of the present invention, and not all of the embodiments. All other embodiments, which can be derived by a person skilled in the art from the embodiments given herein without making any creative effort, shall fall within the protection scope of the present invention.
In addition, the technical solutions in the embodiments of the present invention may be combined with each other, but it must be based on the realization of the technical solutions by those skilled in the art, and when the technical solutions are contradictory to each other or cannot be realized, such a combination of the technical solutions should not be considered to exist, and is not within the protection scope of the present invention.
Fig. 1 shows a fixed-wing drone longitudinal altitude speed decoupling nonlinear control method disclosed in this embodiment, and with reference to fig. 1, the method specifically includes the following steps 1 to 5.
Step 2, designing an expected height change rate according to the relation between the expected flying height and the actual height of the fixed-wing unmanned aerial vehicle, and calculating to obtain an expected track inclination angle according to the expected height change rate and the expected flying airspeed, wherein the specific implementation process is as follows:
step 2.1, designing the expected height change rate as a first-order link, specifically:
in the formula (I), the compound is shown in the specification,
to expect a high degree of change, K h The reciprocal of the first-order link time constant can be designed according to the expected dynamic response characteristic, for example, if the expected adjustment time is T, then K is calculated according to the first-order link response characteristic h The value is 3/T, but is limited by the control capability of the unmanned aerial vehicle, so that the expected adjusting time is not too short, and the specific size is determined according to the motion characteristics of the unmanned aerial vehicle body; h is c To the desired flying height, h t The current actual height;
step 2.2, calculating an expected track inclination angle according to the expected height change rate and the expected flying airspeed, wherein the expected track inclination angle is as follows:
in the formula, gamma c Indicating desired track inclination, V a,c Representing the desired airspeed without regard to the wind.
And 3, calculating to obtain a corresponding trim attack angle and a corresponding desired attitude angle by using a general trim method through the desired flight airspeed and the desired track inclination angle, wherein the specific implementation process of the method is as follows:
step 3.1, according to the expected flight airspeed and the expected track inclination angle, obtaining the corresponding trim attack angle under the current expected flight airspeed and expected track inclination angle states by using an unmanned aerial vehicle dynamic model and a general trim calculation method, wherein the trim attack angle is as follows:
α trim ＝trim(V a,c ,γ c )
in the formula, alpha trim Trim incidence, trim (-) is the trim calculation process;
step 3.2, calculating to obtain an expected attitude angle according to the expected track inclination angle and the trim attack angle, wherein the expected attitude angle comprises the following steps:
θ c ＝γ c +α trim
in the formula, theta c Is the desired attitude angle.
In a specific process, for the trimming process in step 3.1, the flight performance and the control capability of the aircraft body need to be considered, corresponding constraint conditions are set, the rationality of the trimming state is judged and corresponding adjustment is performed, so that the target state is ensured to be theoretically reachable. For example: the maximum flight attack angle can be determined through the minimum flight speed limit of the unmanned aerial vehicle; by elevator yaw limiting, a range of pitch angles may be determined that can maintain pitch moment balance.
Step 4, designing a thrust-attack angle control law based on an incremental nonlinear dynamic inverse method according to the trim attack angle and the actual attack angle deviation to obtain a thrust control instruction and realize attack angle-speed control, wherein the specific implementation process comprises the following steps:
the differential equation of the angle of attack dynamics is as follows:
in the formula (I), the compound is shown in the specification,
is the angle of attack rate of change, m is the unmanned aerial vehicle mass, V a The flight airspeed is, L is the lift force, gamma is the track inclination, T is the engine thrust, and alpha is the attack angle;
referring to fig. 2, according to the incremental nonlinear dynamic inverse theory, a thrust-attack angle control law is obtained, which is:
in the formula, subscript (t-1) represents the last time step state, and subscript t represents the current time step state. At the initial moment, the last time step is zero or the initial leveling value, K α Parameters are designed for the control process.
Step 5, designing a pitch angle control law based on an incremental nonlinear dynamic inverse method according to the deviation between the expected attitude angle and the actual attitude angle to obtain an elevator yaw instruction, and realizing attitude-track control, wherein the specific implementation process is as follows with reference to fig. 3:
step 5.1, according to the differential equation of the longitudinal attitude motion dynamics, the pitching moment coefficient is subjected to first-order Taylor expansion at a certain state, and the differential equation of the longitudinal attitude dynamics in the neighborhood of the state is obtained, wherein the differential equation is as follows:
wherein θ represents a pitch angle, Q represents a pitch angular velocity, Q represents a dynamic pressure, and I yy Is the moment of inertia in the pitch direction, S represents the wetted area, C is the longitudinal reference length (mean chord length), C m Representing the coefficient of the pitching moment (deviation delta from the current angle of attack alpha and elevator) e Related to);
step 5.2, according to the longitudinal attitude dynamics differential equation in the step 5.1, applying a nonlinear dynamic inverse theory to a pitch angle to pitch angle velocity loop to obtain an angle loop control law, wherein the angle loop control law is as follows:
q c ＝K θ (θ c -θ)
in the formula, q c To expect pitch angle velocity, control will be made through the angular velocity to the elevator yaw loop, K θ Determining the response speed of the angle loop for the control coefficient;
and 5.3, applying an increment nonlinear dynamic inverse theory to obtain an angular velocity loop control law according to the angular velocity in the step 5.2 to the elevator deflection loop, wherein the control law is as follows:
in the formula, delta e(t) Deviation of the elevator at the current time step, delta e(t-1) Deviation of the elevator at the previous time step, theta (t) Pitch angle, q, for the current time step (t) Pitch angular velocity, q, at the current time step (t-1) Pitch angle rate at the last time step, C mδe The partial derivative of the pitching moment coefficient to the elevator deflection is related to the model parameter and the flight state, and the incremental nonlinear dynamic inverse method can reduce the model dependence to a certain extent, K q Are control parameters.
In the specific process, in the angular velocity loop control law of the step 5.3, the pitch angle command is obtained by calculation according to the expected track angle and the actual attack angle, and the tracking speed of the pitch angle command is considered to be higher than the tracking speed of the attack angle command so as to compensate the track error caused by the slower tracking speed of the attack angle. The desired pitch angle command theta in the angle loop control law can be set c Adjusted to gamma c + α, the sum of the desired track inclination and the current actual angle of attack, to compensate for angle of attack tracking errors.
The above description is only a preferred embodiment of the present invention, and is not intended to limit the scope of the present invention, and all equivalent structural changes made by using the contents of the present specification and the drawings, or any other related technical fields, which are directly or indirectly applied to the present invention, are included in the scope of the present invention.
Claims (8)
1. A fixed wing unmanned aerial vehicle longitudinal altitude speed decoupling nonlinear control method is characterized by comprising the following steps:
step 1, determining an expected flying airspeed and an expected flying height of a fixed-wing unmanned aerial vehicle;
step 2, designing an expected altitude change rate according to the relation between the expected flying altitude and the actual altitude, and calculating to obtain an expected track inclination angle according to the expected altitude change rate and the expected flying airspeed;
step 3, calculating to obtain a corresponding trim attack angle and a corresponding desired attitude angle by using a general trim method according to the desired flight airspeed and the desired track inclination angle;
step 4, designing a thrust-attack angle control law based on an incremental nonlinear dynamic inverse method according to the deviation between the trim attack angle and the actual attack angle to obtain a thrust control instruction and realize attack angle-speed control;
and 5, designing a pitch angle control law based on an incremental nonlinear dynamic inverse method according to the deviation between the expected attitude angle and the actual attitude angle to obtain an elevator yaw instruction, and realizing attitude-track control.
2. The fixed-wing drone longitudinal altitude speed decoupling nonlinear control method according to claim 1, characterized in that in step 2, the expected altitude rate of change is designed as a first-order link:
in the formula (I), the compound is shown in the specification,
to expect a high degree of change, K h Is the reciprocal of the time constant of the first order link, h c To the desired flying height, h t Is the current actual height.
3. The fixed-wing drone longitudinal altitude speed decoupling nonlinear control method of claim 2, wherein in step 2, the desired track inclination is:
in the formula, gamma c Indicating desired track inclination, V a,c Representing the desired airspeed without regard to the wind.
4. The fixed-wing drone longitudinal altitude speed decoupling nonlinear control method of claim 1, 2 or 3, characterized in that step 3 specifically includes:
step 3.1, according to the expected flight airspeed and the expected track inclination angle, obtaining the corresponding trim incidence angle under the current expected flight airspeed and expected track inclination angle states by using an unmanned aerial vehicle dynamic model and a general trim calculation method, wherein the trim incidence angle is as follows:
α trim ＝trim(V a,c ,γ c )
in the formula, alpha trim For trim angle of attack, trim () is a process representing a trim calculation, V a,c Representing the desired airspeed, gamma, of the flight without regard to the influence of wind c Representing a desired track inclination;
step 3.2, calculating to obtain an expected attitude angle according to the expected track inclination angle and the trim incidence angle, wherein the expected attitude angle comprises the following steps:
θ c ＝γ c +α trim
in the formula, theta c Is the desired attitude angle.
5. The fixed-wing unmanned aerial vehicle longitudinal altitude speed decoupling nonlinear control method according to claim 4, characterized in that in the balancing process of step 3.1, corresponding constraint conditions are set for the flight performance and control capability of the aircraft body, and the rationality of the balancing state is judged and correspondingly adjusted to ensure that the target state is theoretically reachable.
6. The fixed-wing drone longitudinal altitude speed decoupling nonlinear control method according to claim 4, characterized in that step 4 specifically includes:
establishing an angle of attack dynamics differential equation as follows:
in the formula (I), the compound is shown in the specification,
is the angle of attack rate of change, m is the unmanned aerial vehicle mass, V a The flight airspeed is, L is the lift force, gamma is the track inclination, T is the engine thrust, and alpha is the attack angle;
according to the increment nonlinear dynamic inverse theory, the thrust-attack angle control law is obtained, and the law comprises the following steps:
in the formula, subscript (t-1) represents the last time step state, subscript t represents the current time step state, K α Parameters are designed for the control process.
7. The fixed-wing drone longitudinal altitude speed decoupling nonlinear control method according to claim 4, characterized in that step 5 specifically includes:
step 5.1, according to the differential equation of the longitudinal attitude motion dynamics, and performing first-order Taylor expansion on the pitching moment coefficient at a certain state to obtain the differential equation of the longitudinal attitude dynamics in the neighborhood of the state, wherein the differential equation is as follows:
wherein θ represents a pitch angle, Q represents a pitch angular velocity, Q represents a dynamic pressure, and I yy Is the moment of inertia in the pitch direction, S represents the wetted area, C is the longitudinal reference length, C m Representing the angle of attack alpha and the elevator deflection delta from the current e The associated pitching moment coefficient;
step 5.2, according to the longitudinal attitude dynamics differential equation in the step 5.1, applying a nonlinear dynamic inverse theory to a pitch angle to pitch angle speed loop to obtain an angle loop control law, wherein the angle loop control law is as follows:
q c ＝K θ (θ c -θ)
in the formula, q c To expect pitch angle velocity, control will be made through the angular velocity to the elevator yaw loop, K θ Determining the response speed of the angle loop for the control coefficient;
and 5.3, applying an increment nonlinear dynamic inverse theory according to the angular velocity to the elevator deflection loop in the step 5.2 to obtain an angular velocity loop control law, wherein the law comprises the following steps:
in the formula, delta e(t) Elevator yaw, delta, for the current time step e(t-1) Deviation of the elevator at the previous time step, theta (t) Pitch angle, q, for the current time step (t) Pitch angle rate, q, for the current time step (t-1) The pitch angle rate for the previous time step,
the partial derivative of the pitching moment coefficient to the elevator deviation is a value related to model parameters and flight states, and the model dependency, K, can be reduced to a certain extent by the incremental nonlinear dynamic inverse method q Are control parameters.
8. The fixed-wing unmanned aerial vehicle longitudinal altitude speed decoupling nonlinear control method according to claim 7, characterized in that in the angular speed loop control law of step 5.3, the pitch angle command tracking speed is faster than the attack angle command tracking speed to compensate for a track error caused by the slower attack angle tracking speed.
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Cited By (2)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN117389320A (en) * | 2023-12-07 | 2024-01-12 | 陕西德鑫智能科技有限公司 | Unmanned aerial vehicle cruise control method and system |
| CN117452973A (en) * | 2023-12-22 | 2024-01-26 | 中国航空工业集团公司西安飞机设计研究所 | Method and device for optimizing short-distance landing flight path of front airport of conveyor |
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Cited By (4)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN117389320A (en) * | 2023-12-07 | 2024-01-12 | 陕西德鑫智能科技有限公司 | Unmanned aerial vehicle cruise control method and system |
| CN117389320B (en) * | 2023-12-07 | 2024-03-08 | 陕西德鑫智能科技有限公司 | Unmanned aerial vehicle cruise control method and system |
| CN117452973A (en) * | 2023-12-22 | 2024-01-26 | 中国航空工业集团公司西安飞机设计研究所 | Method and device for optimizing short-distance landing flight path of front airport of conveyor |
| CN117452973B (en) * | 2023-12-22 | 2024-03-19 | 中国航空工业集团公司西安飞机设计研究所 | Method and device for optimizing short-distance landing flight path of front airport of conveyor |
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