CN116520704A - Gain self-adaptive tailstock aircraft control system and control method based on distributed dynamic pressure detection - Google Patents
Gain self-adaptive tailstock aircraft control system and control method based on distributed dynamic pressure detection Download PDFInfo
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Abstract
The invention discloses a gain self-adaptive tailstock aircraft control system and a control method based on distributed dynamic pressure detection. The invention better solves the problem of irregular change of the model of the tailstock type aircraft in the control of vertical transition conversion, and improves the safety and smoothness of the transition conversion control of the tailstock type aircraft.
Description
Technical Field
The invention relates to a gain self-adaptive tailstock aircraft control system based on distributed dynamic pressure detection, which belongs to the field of flight control, and on one hand, the method better solves the problem of irregular model change in transition control of a tailstock aircraft with control surface control, improves the safety and smoothness of the transition conversion process, ensures that the tailstock type vertical take-off and landing aircraft does not need mode switching and gain scheduling between flat flight cruising and vertical take-off and landing, realizes single-mode continuous control in the transition process, and meanwhile, the method can be suitable for larger tailstock type unmanned aircraft and has better platform universality.
Background
The vertical take-off and landing fixed wing is an aircraft which integrates the characteristics of multiple rotors and fixed wings, has the advantages of multiple rotors and vertical take-off and landing of a helicopter, is convenient to take off and land, has low requirements on deployment environment facilities, has the same high-speed flat flight capacity as the fixed wing, and is long in voyage and air time. The tailstock type aircraft is the simplest type of aircraft, has no complex tilting mechanism like tilting and lifting, and has no redundant power and wind resistance like compound and lifting, and is an important research direction in the field of vertical take-off and landing fixed wings. Has great application value in the fields of aviation mapping, military reconnaissance, logistics transportation and the like. The tailstock type aircraft is generally an unmanned aerial vehicle, the flight state of the tailstock type aircraft is generally divided into a flat flight cruise, a vertical hover and a transition flight between the flat flight cruise and the vertical hover, and the horizontal flight and the vertical hover can be respectively controlled as a fixed wing and a multi-rotor aircraft, and the control method is mature and stable, and the control of the transition stage, especially the transition process from high-speed flat flight to hover take-off and landing, has severe model change and the highest flight control difficulty.
In the conventional transition control, particularly the transition control from flat fly to hover, in order to avoid the control problem under a large attack angle, a flying mode of first vertical pulling and then slow descending is generally selected, so that the speed is slow and the energy consumption is high. In order to realize vertical-horizontal conversion control of tailstock type vertical-up, a general control method can be divided into a multi-mode switching method and a parameter dynamic estimation method to cope with the change of model structure parameters, the multi-mode switching method is discontinuous in control and is easy to cause problems in switching, and the parameter dynamic estimation method needs an accurate model, so that the two methods have no good general robustness.
Disclosure of Invention
Aiming at the defects of the prior art, the invention provides a gain self-adaptive tailstock aircraft control system and a control method based on distributed dynamic pressure detection.
In order to achieve the above object, the technical scheme of the present invention is summarized as follows:
the gain self-adaptive tailstock aircraft control system based on distributed dynamic pressure detection is characterized by mainly comprising a distributed dynamic pressure detection subsystem and an angular velocity control algorithm, wherein the distributed dynamic pressure detection subsystem outputs average control surface dynamic pressureThe angular speed control algorithm is based on the average control surface dynamic pressure +.>Adjusting the control quantity gain; the distributed dynamic pressure detection subsystem consists of a hardware detection part and a software processing part; the hardware detection part of the distributed dynamic pressure detection subsystem consists of a dynamic pressure probe array and a plurality of corresponding differential pressure airspeed sensors, wherein the dynamic pressure probe array is arranged on the surface of the wing and in front of a control surface rotating shaft; the dynamic pressure probe array is connected with the differential pressure airspeed sensors through an air duct; the software processing part of the distributed dynamic pressure detection subsystem has input quantity of dynamic pressure distribution data p acquired by the hardware detection part of the distributed dynamic pressure detection subsystem 1 ,p 2 ,…,p n And the corresponding control surface area S 1 ,S 2 ,…,S n The output quantity is the average control surface dynamic pressureThe specific calculation method is as follows:
wherein the method comprises the steps ofFor average control surface dynamic pressure, p i The control surface dynamic pressure obtained by the detection of the ith dynamic pressure probe is S i Is the ith dynamic pressureThe control surface area of the area corresponding to the probe is S, the total control surface area is +.>
The input quantity of the angular velocity control algorithm is the measured angular velocity omega obtained by a gyroscope b Given target angular velocity ω r The output quantity is uncompensated control quantity u 1k Where k is one of the body axis designations pitch, roll, and yaw. The angular speed control algorithm allows parameters to change within a certain range, and if the distributed dynamic pressure detection subsystem fails or is inaccurate in measurement, the final transition control effect is not greatly influenced as long as the distributed dynamic pressure detection subsystem can output approximate measurement parameters along with the change of the flight state.
Further, the angular velocity control algorithm may be specifically an active disturbance rejection control algorithm, where the active disturbance rejection control algorithm includes an extended state observer, a tracking differentiator, a state feedback and a total disturbance compensation; the active disturbance rejection control algorithm comprises the following steps:
step1: updating the extended state observer: angular velocity omega measured from gyroscopes b With the control quantity u output by the controller at the previous moment k The average control surface dynamic pressureSubstituting the state quantity z into the dilation observer to update the state quantity z jk Wherein j is the order of the corresponding state quantity of each machine body axis k, and k is one of machine body axis marks pitch, roll and yaw;
step2: arranging a transition process: for a given angular velocity ω of an input using a tracking differentiator r Preprocessing, planning a uniform acceleration and uniform deceleration process for the step change of the given angular velocity, and outputting the preprocessed given angular velocity and angular acceleration v 1k ,v 2k Wherein k is one of the body shaft labels pitch, roll, and yaw;
step3: calculating state feedback: the extended state observer outputs a state quantity z according to Step1 1k ,z 2k A given angular velocity and a given angular acceleration v after the pretreatment described in Step2 1k ,v 2k Substituting the state feedback rate to obtain an ideal control quantity u designed for the pure integral model 0k Wherein k is one of the body shaft labels pitch, roll, and yaw;
step4: total disturbance compensation: an estimate z of the equivalent total disturbance output by the extended state observer according to Step1 3k Disturbance compensation offset is carried out on the ideal control quantity described in Step3, so that the actual dynamic state of the aircraft approaches to the ideal pure integral model u 1k Wherein k is one of the body shaft labels pitch, roll, and yaw;
further, the control method of the gain self-adaptive tailstock aircraft control system for distributed dynamic pressure detection in the invention is characterized by comprising the following steps:
step1: obtaining control surface dynamic pressure distribution data p through a hardware part of the distributed dynamic pressure detection subsystem 1 ,p 2 ,…,p n ；
Step2: according to the dynamic pressure distribution data p of the control surface obtained in the step1 1 ,p 2 ,…,p n And the corresponding control surface area S 1 ,S 2 ,…,S n Calculating average control surface dynamic pressure through software part of the distributed dynamic pressure detection subsystem
Step3: measured angular velocity omega obtained from gyroscopes b And a given angular velocity ω of the input r The uncompensated control quantity u is output through the angular speed control algorithm 1k Wherein k is one of the body shaft labels pitch, roll, and yaw;
step4: average control surface dynamic pressure obtained according to step2Dynamic pressure gain compensation is carried out on the output quantity obtained in the step3, and a proper control quantity u is obtained k Realizes the accurate control of the aircraft,wherein k is one of the machine axis marks pitch, roll, and yaw;
further, the dynamic pressure compensation gain mode described in step4 is specifically as follows:
wherein u is pitch ,u roll ,u yaw The control amounts of final pitch and roll after compensation, u 1pitch ,u 1roll ,u 1yaw The control amounts are respectively uncompensated and,is the average control surface dynamic pressure.
Drawings
The accompanying drawings are included to provide a further understanding of the invention and are incorporated in and constitute a part of this specification, illustrate the invention and together with the embodiments of the invention, serve to explain the invention. In the drawings:
FIG. 1 is a block diagram of a gain adaptive control method of distributed dynamic pressure detection compensation;
FIG. 2 is a schematic view of the aircraft body axes;
FIG. 3 is a schematic view of the aircraft body axes;
FIG. 4 is a simplified two-region airfoil flow field profile;
FIG. 5 is a profile of the simplest two-zone dual-dynamic probe on the wing;
FIG. 6 is a schematic side mounting of a single dynamic pressure probe;
FIG. 7 is a block diagram of a one-dimensional second order active disturbance rejection control algorithm;
FIG. 8 is a graph showing the effect of the simulation control of the angular velocity with the dynamic pressure change and the measured value thereof unchanged;
FIG. 9 is a graph of simulated control effects of slower angular velocity for dynamic pressure follow-up measurements;
FIG. 10 is a graph of simulated control effects of angular velocity with accurate dynamic pressure follow-up measurements;
reference numerals: 1. a left motor; 2. a right motor; 3. a left steering engine; 4. a right steering engine; 5. a left propeller; 6. a right propeller; 7. a left wing; 8. a right wing; 9. a left control surface; 10. a right control surface; 11. a body; 101. dynamic pressure probes outside the slip flow area of the propeller; 102. dynamic pressure probes in the propeller slip flow area; a. a distributed dynamic pressure detection subsystem; a1, a distributed dynamic pressure detection subsystem hardware detection part; a2, a software processing part of the distributed dynamic pressure detection subsystem; b. an angular velocity control algorithm; c. controlling dynamic pressure gain adjustment; d. a gyroscope.
Detailed Description
In order that the above-recited objects, features and advantages of the present invention will become more apparent, a more particular description of the invention will be rendered by reference to specific embodiments thereof.
In the following description, numerous specific details are set forth in order to provide a thorough understanding of the present invention, but the present invention may be practiced in other ways other than those described herein, and persons skilled in the art will readily appreciate that the present invention is not limited to the specific embodiments disclosed below.
Further, reference herein to "one embodiment" or "an embodiment" means that a particular feature, structure, or characteristic can be included in at least one implementation of the invention. The appearances of the phrase "in one embodiment" in various places in the specification are not necessarily all referring to the same embodiment, nor are separate or alternative embodiments mutually exclusive of other embodiments.
The invention provides a gain self-adaptive tailstock aircraft control system and a control method based on distributed dynamic pressure detection, and the aiming tailstock aircraft mainly comprises a1 left motor, a2 right motor, a 3 left steering engine, a 4 right steering engine, a 5 left propeller, a 6 right propeller, a 7 left wing, an 8 right wing, a 9 left control surface, a 10 right control surface and an 11 fuselage.
As shown in fig. 1, distributed dynamic pressure detection sensors are arranged according to wing flow field distribution to obtain dynamic pressure distribution data of a control surface; calculating average control surface dynamic pressure according to the control surface dynamic pressure distribution data obtained by the distributed dynamic pressure detection sensor; outputting a control quantity through a control algorithm according to the measured angular velocity obtained by the gyroscope and the input given angular velocity; according to the average control surface dynamic pressure, dynamic pressure compensation is carried out on the output quantity, so that proper rudder deflection is obtained, and accurate control of the aircraft is realized;
for the convenience of the subsequent algorithm description, the coordinate axes of the machine body are defined according to the requirement, and the definition mode is shown in fig. 2 and 3.
The following is a specific implementation procedure.
1) Obtaining control surface dynamic pressure distribution data through a hardware part of the distributed dynamic pressure detection subsystem
As shown in FIG. 4, the wing flow field is divided into uncovered areas of the propeller slipstream, and the control surface coverage area is s 1 As shown in a 101 position in fig. 5, a dynamic pressure probe is arranged in the area to obtain the dynamic pressure p of the control surface in the area 1 And a propeller slipstream coverage area with a control surface coverage area s2, and a dynamic pressure probe is arranged in the area as shown at 202 in fig. 5 to obtain the dynamic pressure p of the control surface in the area 2 The dynamic pressure probe sides of the two areas are arranged in front of the control surface rotating shaft as shown in fig. 6.
2) Calculating average control surface dynamic pressure through software part of distributed dynamic pressure detection subsystem
According to the control surface dynamic pressure distribution condition, calculating average control surface dynamic pressure, wherein the average control surface dynamic pressure calculating method comprises the following steps:
wherein the method comprises the steps ofFor average control surface dynamic pressure, p 1 ,p 2 The dynamic pressure of the control surface outside the propeller slipstream coverage area and the dynamic pressure of the control surface inside the propeller slipstream coverage area are respectively S 1 ,S 2 The control surface area outside the propeller slipstream coverage area and the control surface area inside the propeller slipstream coverage area are respectively.
3) An active disturbance rejection control algorithm is adopted to obtain uncompensated control quantity
Taking a one-dimensional second-order active disturbance rejection control algorithm block diagram as shown in fig. 7 as an example, a specific angular velocity control algorithm is designed. Wherein v is d To reference the input signal v 1 For reference signals processed by tracking differentiators, e is the state error, u 0 For ideal control quantity, u is the actual control quantity after disturbance compensation, b 0 The gain coefficient estimator for the control quantity, w is the external disturbance, y is the controlled quantity,for the measurement filter value of the controlled quantity, the default value is obtained by a measurement filter system, z 1 ,z 2 Z is the observed quantity corresponding to the state quantity 3 For extended state observance, for tracking and estimating other internal and external disturbance forces at the input in addition to the control force.
The three body axis expansion state observers are designed according to the model, and the fal functions used before and after the expansion state observers are described:
fal is a nonlinear function, and the specific expression is:
a is a nonlinear factor and delta is a filtering factor. The fal function is formed by sectioning a proportional function and an exponential function, wherein the exponent function is generally cut off by a straight line of 0 to 1, x=delta, a part of x > delta is reserved, and a function image is symmetrical about an origin to form the fal function, wherein the linear interval width range is +delta.
Then, the process of calculating the control quantity output by the active disturbance rejection control algorithm is as follows:
step1: firstly, constructing a triaxial extended state observer, and updating the extended state observer: substituting the angular velocity obtained by the gyroscope measurement, the control quantity output by the controller at the last moment and the average control surface dynamic pressure into an extended observer model to update the state;
the x-axis is:
wherein e x Observer error of angular velocity for x-axis, z 1x ,z 2x Angular velocity omega of x-axis respectively x And its rate of changeCorresponding observed quantity, z 3x Is the observed quantity of the equivalent total disturbance of the x-axis, beta 01x ,β 02x ,β 03x Gain factor for x-axis observer, f 1x Second order term coefficients for the x-axis angular velocity observer,/>For average control surface dynamic pressure, b 1x Is the pitch control surface coefficient.
The y-axis is:
wherein e y Observer error of angular velocity in y-axis, z 1y ,z 2y Angular velocity omega of y-axis respectively y And its rate of changeCorresponding observed quantity, z 3y Is the observed quantity of the equivalent total disturbance of the y axis, beta 01y ,β 02y ,β 03y Gain factor for y-axis observer, f 1y For the second order term coefficient of the y-axis, b y Gain for differential control amount.
The z axis is:
wherein e z Observer error of angular velocity as z-axis, z 1z ,z 2z Respectively the z-axisAngular velocity omega z And its rate of changeCorresponding observed quantity, z 3z Is the observed quantity of the equivalent total disturbance of the z axis, beta 01z ,β 02z ,β 03z Gain factor for z-axis observer, f 1z Second order term coefficients for the z-axis angular velocity observer,/->For average control surface dynamic pressure, b 1z Is the roll control surface coefficient.
Step2: arranging a transition process: preprocessing the input given angular velocity by using a tracking differentiator, planning a process of uniform acceleration and uniform deceleration for the step change of the given angular velocity, and outputting the preprocessed given angular velocity and angular acceleration;
according to the maneuvering performance of the aircraft, a tracking differentiator on three engine body shafts is designed, a differential tracker constructed by a fhan function is used as a feedforward link, an input signal is preprocessed, and the processed differential signal is extracted, so that a second-order state feedback link is designed, and better input response dynamics is obtained.
Wherein omega rx ,ω ry ,ω rz For a given input triaxial angular velocity, v 1x ,v 1y ,v 1z And v 2x ,v 2y ,v 2z For a target angular velocity omega rx ,ω ry ,ω rz The instruction signal after being pre-programmed and the change rate of the instruction signal, r tdx ,r tdy ,r tdz And h tdx ,h tdy ,h tdz To track the parameters of the fastest synthesis function fhan in the differentiator. Wherein the fastest synthesis function fhan (x 1 ,x 2 R, h) defining a two-input nonlinear function, the operation of which is:
wherein d, d 0 ,y,a 0 A is a temporary variable in the calculation process, r, h is a function parameter, fh is a function output value, and x 1 ,x 2 Is an input value. sign (x) is a sign function, expressed as:
step3: calculating state feedback: substituting the angular velocity and the angular acceleration output by the extended state observer in the Step (1) and the given angular velocity and the given angular acceleration after the pretreatment in the Step (2) into a state feedback rate to obtain an ideal control quantity designed for a pure integral model;
nonlinear feedback of three engine body axes is designed by using fhan function, and ideal control quantity u of the three axes is obtained 0x ,u 0y ,u 0z ：
Wherein c x ,c y ,c z Damping coefficient r for state feedback fbx ,r fby ,r fbz And h fbx ,h fby ,h fbz Is the parameter of the fastest synthesis function fhan in the state feedback.
Step4: total disturbance compensation: according to the estimated quantity of the equivalent total disturbance output by the extended state observer in Step1, disturbance compensation offset is carried out on the ideal control quantity in Step3, so that the actual dynamic state of the aircraft approaches to an ideal pure integral model;
compensating the total disturbance of the three shafts to obtain an actual y-axis differential control quantity u y And control surface control amount u of x-axis and z-axis without dynamic pressure compensation 1x And u 1z ：
Wherein z is 2x Rate of change of angular velocity for the x-axisCorresponding observed quantity, z 3x Is the observed quantity of the equivalent total disturbance of the x-axis, f 1x Second order term coefficient for x-axis angular velocity observer, b 1x For pitch control surface coefficient, z 2y For the rate of change of the angular velocity of the y-axis +.>Corresponding observed quantity, z 3y Is the observed quantity of the equivalent total disturbance of the y axis, f 1y Second order term coefficient for y-axis angular velocity observer, b y For differential control quantity gain, z 2z For the rate of change of the angular velocity of the z-axis +.>Corresponding observed quantity, z 3z Is the observed quantity of the equivalent total disturbance of the z axis, f 1z Second order term coefficient for z-axis angular velocity observer, b 1z Is the roll control surface coefficient.
4) Dynamic pressure compensation and control quantity distribution are carried out on the control quantity
The dynamic pressure compensation method is as follows:
wherein u is x ,u z The control amounts of final pitch and roll after compensation, u 1x ,u 1z The uncompensated pitch and roll control amounts respectively,is the average control surface dynamic pressure.
Then the obtained control quantity is distributed to obtain a control signal Motor of the left Motor 1 L And a right motor 2Control signal monitor R And a control signal servo of the left steering engine 3 L And control signal servo of right steering engine 4 R ：
Wherein thr is the total throttle signal, which can be manually input by a throttle lever or automatically adjusted by a height/speed controller, and is a servo trimL ,servo trimR The neutral point zero position values of the left steering engine and the right steering engine are respectively obtained by on-site adjustment during the assembly of the aircraft.
FIG. 8 shows the dynamic pressure and attack angle to simulate the actual transition flight variation, under the dynamic pressure constant value estimation condition, namely the fixed control quantity gain, the pitching control effect of the aircraft, the total disturbance estimation of the full-automatic anti-disturbance controller is used for coping with the large variation of the model, and the pitch angle speed response has larger oscillation and delay. Fig. 9 shows the influence of dynamic pressure dynamic estimation hysteresis, and even if the average dynamic pressure measurement estimation has small error, the distributed dynamic pressure detection can reflect the real-time variation mode of the average dynamic pressure, the control effect is hardly weakened with the aid of the wide parameter characteristic of the active disturbance rejection controller, and the distributed dynamic pressure detection system can have good effect in the whole-flow flight as long as the distributed dynamic pressure detection system gives approximately accurate dynamic control quantity gain. The dynamic pressure change in the transition conversion process is simulated in fig. 10, the dynamic estimation of the distributed dynamic pressure detection system is more accurate, and the ideal control effect is achieved under the condition of low delay.
The foregoing description is only a preferred embodiment of the present invention, and the present invention is not limited thereto, but it is to be understood that modifications and equivalents of some of the technical features described in the foregoing embodiments may be made by those skilled in the art, although the present invention has been described in detail with reference to the foregoing embodiments. Any modification, equivalent replacement, improvement, etc. made within the spirit and principle of the present invention should be included in the protection scope of the present invention.
Claims (4)
1. The gain self-adaptive tailstock aircraft control system based on distributed dynamic pressure detection is characterized by mainly comprising a distributed dynamic pressure detection subsystem and an angular velocity control algorithm, wherein the distributed dynamic pressure detection subsystem outputs average control surface dynamic pressure p, and the angular velocity control algorithm is used for controlling the aircraft according to the average control surface dynamic pressureAdjusting the control quantity gain; the distributed dynamic pressure detection subsystem consists of a hardware detection part and a software processing part; the hardware detection part of the distributed dynamic pressure detection subsystem consists of a dynamic pressure probe array and a plurality of corresponding differential pressure airspeed sensors, wherein the dynamic pressure probe array is arranged on the surface of the wing and in front of a control surface rotating shaft; the dynamic pressure probe array is connected with the differential pressure airspeed sensors through an air duct; the software processing part of the distributed dynamic pressure detection subsystem has input quantity of dynamic pressure distribution data p acquired by the hardware detection part of the distributed dynamic pressure detection subsystem 1 ,p 2 ,…,p n And the corresponding control surface area S 1 ,S 2 ,…,S n The output quantity is the average control surface dynamic pressure +.>The specific calculation method is as follows:
wherein the method comprises the steps ofFor average control surface dynamic pressure, p i The control surface dynamic pressure obtained by the detection of the ith dynamic pressure probe is S i The control surface area of the area corresponding to the ith dynamic pressure probe is S, the total control surface area is +.>
The angular velocity controlThe input quantity of the control algorithm is the measured angular velocity omega obtained by the gyroscope b Given target angular velocity ω r The output quantity is uncompensated control quantity u 1k Where k is one of the body axis designations pitch, roll, and yaw. The angular speed control algorithm allows parameters to change within a certain range, and if the distributed dynamic pressure detection subsystem fails or is inaccurate in measurement, the final transition control effect is not greatly influenced as long as the distributed dynamic pressure detection subsystem can output approximate measurement parameters along with the change of the flight state.
2. The gain-adaptive tailstock aircraft control system based on distributed dynamic pressure detection according to claim 1, wherein the angular velocity control algorithm is specifically an active disturbance rejection control algorithm, and the active disturbance rejection control algorithm comprises an extended state observer, a tracking differentiator, state feedback and total disturbance compensation; the active disturbance rejection control algorithm comprises the following steps:
step1: updating the extended state observer: angular velocity omega measured from gyroscopes b With the control quantity u output by the controller at the previous moment k The average control surface dynamic pressureSubstituting the state quantity z into the dilation observer to update the state quantity z jk Wherein j is the order of the corresponding state quantity of each machine body axis k, and k is one of machine body axis marks pitch, roll and yaw;
step2: arranging a transition process: for a given angular velocity ω of an input using a tracking differentiator r Preprocessing, planning a uniform acceleration and uniform deceleration process for the step change of the given angular velocity, and outputting the preprocessed given angular velocity and angular acceleration v 1k ,v 2k Wherein k is one of the body shaft labels pitch, roll, and yaw;
step3: calculating state feedback: the extended state observer outputs a state quantity z according to Step1 1k ,z 2k A given angular velocity and a given angular acceleration v after the pretreatment described in Step2 1k ,v 2k Substituting the state feedback rate to obtain an ideal control quantity u designed for the pure integral model 0k Wherein k is one of the body shaft labels pitch, roll, and yaw;
step4: total disturbance compensation: an estimate z of the equivalent total disturbance output by the extended state observer according to Step1 3k Disturbance compensation offset is carried out on the ideal control quantity described in Step3, so that the actual dynamic state of the aircraft approaches to the ideal pure integral model u 1k Where k is one of the body axis designations pitch, roll, and yaw.
3. A control method of a gain-adaptive tailstock aircraft control system based on distributed dynamic pressure detection according to any one of claims 1-2, characterized by the steps of:
step1: obtaining control surface dynamic pressure distribution data p through a hardware part of the distributed dynamic pressure detection subsystem 1 ,p 2 ,…,p n ；
Step2: according to the dynamic pressure distribution data p of the control surface obtained in the step1 1 ,p 2 ,…,p n And the corresponding control surface area S 1 ,S 2 ,…,S n Calculating an average control surface dynamic pressure p through a software part of the distributed dynamic pressure detection subsystem;
step3: measured angular velocity omega obtained from gyroscopes b And a given angular velocity ω of the input r The uncompensated control quantity u is output through the angular speed control algorithm 1k Wherein k is one of the body shaft labels pitch, roll, and yaw;
step4: average control surface dynamic pressure obtained according to step2Dynamic pressure gain compensation is carried out on the output quantity obtained in the step3, and a proper control quantity u is obtained k The precise control of the aircraft is realized, wherein k is one of the machine body axis marks pitch, roll and yaw.
4. The method for controlling a gain-adaptive tailstock aircraft control system based on distributed dynamic pressure detection according to claim 3, wherein the dynamic pressure compensation gain mode in step4 is specifically as follows:
wherein u is pitch ,u roll ,u yaw The control amounts of final pitch and roll after compensation, u 1pitch ,u 1roll ,u 1yaw The control amounts are respectively uncompensated and,is the average control surface dynamic pressure.
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| CN202310600869.9A Pending CN116520704A (en) | 2023-05-25 | 2023-05-25 | Gain self-adaptive tailstock aircraft control system and control method based on distributed dynamic pressure detection |
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