CN114995480B - Three-axis static unstable flying wing unmanned aerial vehicle attitude angle control method - Google Patents
Three-axis static unstable flying wing unmanned aerial vehicle attitude angle control method Download PDFInfo
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- 238000011217 control strategy Methods 0.000 claims description 7
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- XOFYZVNMUHMLCC-ZPOLXVRWSA-N prednisone Chemical compound O=C1C=C[C@]2(C)[C@H]3C(=O)C[C@](C)([C@@](CC4)(O)C(=O)CO)[C@@H]4[C@@H]3CCC2=C1 XOFYZVNMUHMLCC-ZPOLXVRWSA-N 0.000 claims description 5
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- G—PHYSICS
- G05—CONTROLLING; REGULATING
- G05D—SYSTEMS FOR CONTROLLING OR REGULATING NON-ELECTRIC VARIABLES
- G05D1/00—Control of position, course, altitude or attitude of land, water, air or space vehicles, e.g. using automatic pilots
- G05D1/08—Control of attitude, i.e. control of roll, pitch, or yaw
- G05D1/0808—Control of attitude, i.e. control of roll, pitch, or yaw specially adapted for aircraft
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- G—PHYSICS
- G05—CONTROLLING; REGULATING
- G05D—SYSTEMS FOR CONTROLLING OR REGULATING NON-ELECTRIC VARIABLES
- G05D1/00—Control of position, course, altitude or attitude of land, water, air or space vehicles, e.g. using automatic pilots
- G05D1/10—Simultaneous control of position or course in three dimensions
- G05D1/101—Simultaneous control of position or course in three dimensions specially adapted for aircraft
- G05D1/106—Change initiated in response to external conditions, e.g. avoidance of elevated terrain or of no-fly zones
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Abstract
The invention discloses a three-axis static unstable flying wing unmanned aerial vehicle attitude angle control method, which comprises the following steps: 1) The stability augmentation system is used for improving the static stability and dynamic stability of the aircraft; 2) On the basis of stability enhancement of the attack angle and the sideslip angle, an attitude angle control loop of the flying wing unmanned aerial vehicle is designed; 3) And determining the control gain of the angle of attack stability enhancement and the pitch angle rate control gain by using a pole allocation strategy. The control method provided by the invention has the advantages that the natural longitudinal force arm of the flying-wing unmanned aerial vehicle is short, the longitudinal relaxation static stability is organically combined with the flying-wing unmanned aerial vehicle, the advantages of the flying-wing unmanned aerial vehicle are greatly exerted, and powerful technical support is provided for the aspects of smart flight, load lifting and the like of the subsequent flying-wing unmanned aerial vehicle; the disturbance rejection capability of the attitude angle control of the unmanned aerial vehicle is greatly improved, the unmanned aerial vehicle has excellent robust capability, the unmanned aerial vehicle with the flying wing layout can play a role in unsteady flight under the condition of controllable attitude, the maneuverability and the dexterity of the unmanned aerial vehicle with the flying wing are improved, and a solid foundation is laid for the excellent control of the attitude angle of the unmanned aerial vehicle with the flying wing.
Description
Technical Field
The invention relates to the technical field of aircraft flight, in particular to a three-axis static unstable flying wing unmanned aerial vehicle attitude angle control method.
Background
At present, the known unmanned plane with the flying wing layout is designed by integrating a machine body and wings, so that the influence of vortex and shock waves of a tail wing can be eliminated, and the friction resistance and the interference resistance of the plane are reduced. The wing body fusion increases the wing area and the aspect ratio, and greatly increases the lift force. However, the longitudinal control moment arm of the flying-wing unmanned plane is short, and the three-axis static instability is caused by the aerodynamic layout without a vertical tail, so that the control difficulty of the flying-wing unmanned plane is increased. At present, no control research on the flying wing unmanned aerial vehicle is seen.
Disclosure of Invention
The invention aims to provide a three-axis static unstable flying-wing unmanned aerial vehicle attitude angle control method which can enable a flying-wing layout unmanned aerial vehicle to exert static unstable flight performance under the condition of controllable attitude and improve the maneuverability and dexterity of the flying-wing unmanned aerial vehicle.
The invention is realized by the following technical scheme: a three-axis static unstable flying wing unmanned aerial vehicle attitude angle control method comprises the following steps:
(1) The stability augmentation system is used for improving the static stability and dynamic stability of the aircraft;
(2) On the basis of stability enhancement of the attack angle and the sideslip angle, an attitude angle control loop of the flying wing unmanned aerial vehicle is designed;
(3) And determining the control gain of the angle of attack stability enhancement and the pitch angle rate control gain by using a pole allocation strategy.
According to the technical scheme, the working principle is that according to the stability augmentation principle, for an aircraft with unstable longitudinal, transverse and longitudinal directions, the aircraft is difficult to control, and the stability augmentation system is required to be applied to improve the static stability and dynamic stability of the aircraft. After the stability enhancement of the angle of attack sideslip angle is completed, a pure angular rate control scheme is adopted for realizing the accurate control of the attitude angle, namely the rolling angle and the pitch angle, and the pure angular rate control scheme is adopted for the angular rate ring of the attitude control. The inner ring refers to the inner loop of the control loop and the outer ring refers to the outer loop of the control loop. The inner loop in this patent refers specifically to the angular rate control loop in the attitude control loop. The outer ring specifically refers to an attitude angle control loop in the attitude control loop. A strategy of pole allocation is adopted to determine the control gain of the angle of attack stability enhancement and the pitch angle rate control gain; the attitude control ring adopts the proportional control of the rolling angle and the pitch angle, the static-difference-free control strategy of the rolling angle and the pitch angle is realized by a mathematical formula, and the design of the strategy is finally realized by the formula. Robust capability refers to the ability of a control system to maintain certain performance characteristics thereof under perturbation of certain parameters. The pure angular velocity and airflow angle cross mixing method is adopted to greatly improve the anti-interference capability of the attitude angle of the unmanned aerial vehicle, so that the robustness is greatly improved, and the unmanned aerial vehicle has excellent robustness.
In order to better implement the method of the present invention, further, in the step (1), the specific process of improving the static stability and dynamic stability of the aircraft by using the stability enhancement system is as follows: and the aircraft unmanned aerial vehicle is characterized in that an incidence angle feedback is introduced into an elevator control channel, so that the longitudinal static stability is changed from static instability to static stability, sideslip angle signals are respectively introduced into an aileron rudder and a rudder, the roll static instability of the aircraft unmanned aerial vehicle is changed into roll static stability, and the yaw static instability is changed into yaw static stability.
In order to better implement the method of the present invention, in the step (1), an elevator airflow angle feedback form is obtained by a stabilizing system according to an original mathematical formula of stabilizing an elevator induced angle signal, which is specifically as follows:
the original mathematical formula of sideslip angle signal stability augmentation is introduced through the aileron rudder to obtain the airflow angle feedback form of the aileron rudder, and the airflow angle feedback form is specifically as follows:
The original mathematical formula of sideslip angle signal stability augmentation is introduced through a rudder to obtain a rudder airflow angle feedback form, and the method is specifically as follows:
delta E、δA、δR is the elevator deflection angle, the aileron deflection angle and the rudder deflection angle, P, Q, R is the roll angle rate, the pitch angle rate and the yaw angle rate, alpha and beta are the attack angle signal and the sideslip angle signal, and K is the proportional coefficient corresponding to each physical quantity.
In order to better realize the method of the invention, further, according to the original mathematical formula of the lift-induced angle-of-attack signal stability augmentation by the stability augmentation system, the specific process for obtaining the lift-air flow angle feedback form is as follows:
The original formula for the elevator to introduce the angle of attack signal is: delta E＝KαΔα+KEΔWE
The original formula of the sideslip angle signal introduced by the aileron rudder is as follows: delta a＝KβΔβ+KβΔWa
The original formula for the rudder to introduce the sideslip angle signal is: delta r＝KβΔβ+KβΔWr
Wherein ΔW E is a steering column command, and K is a scale factor of the physical quantity. The steering rod can operate the elevator, namely K EΔWE can be regarded as the original elevator deflection K EΔWE＝δE1, and the angular rate ring is used for proportional controlThen a feedback form of the elevator can be derived
By the same token, the formula for leading the aileron rudder to introduce sideslip angle signals is as follows
In order to better realize the method of the invention, further, in the step (2), on the basis of the stability enhancement of the attack angle and the sideslip angle, the specific process of designing the attitude angle control loop of the flying wing unmanned aerial vehicle is as follows: after the stability increase of the angle of attack sideslip angle is completed, in order to realize the accurate control of the attitude angle, namely the rolling angle and the pitch angle, a pure angular rate control strategy is adopted in the inner ring of the device.
To better implement the method of the present invention, further, the control strategy for the pure angular rate of the inner ring is:
Where Q g is the pitch rate given signal, P g is the roll angle rate given signal, And the control parameters corresponding to the integral terms are respectively adopted.
In order to better realize the method of the invention, in the step (3), a strategy of pole allocation is adopted to determine the control gain of the angle of attack stability enhancement and the pitch angle rate control gain, and the specific contents are as follows: the attitude control ring adopts the proportional control of the rolling angle and the pitch angle to realize the static-free control of the rolling angle and the pitch angle.
In order to better realize the method of the invention, further, the attitude control ring controls the rolling angle and the pitch angle according to the following strategies:
Respectively controlling pitch angle deviation theta-theta g and roll angle deviation phi-phi g to enable the pitch angle deviation theta-theta g and the roll angle deviation phi-phi g to meet control response static-error-free control;
Wherein, theta and phi are pitch angle and roll angle signals respectively, theta g、φg is pitch angle instruction set and roll angle instruction set respectively, Respectively corresponding proportional coefficients.
Compared with the prior art, the invention has the following advantages:
(1) The invention solves the defect of short natural longitudinal force arm of the flying-wing unmanned aerial vehicle, organically combines longitudinal relaxation static stability with the flying-wing unmanned aerial vehicle, greatly exerts the advantages of the flying-wing unmanned aerial vehicle, and provides powerful technical support for the aspects of smart flight, load lifting and the like of the subsequent flying-wing unmanned aerial vehicle;
(2) The invention adopts the air flow angle feedback to solve the problem of unstable triaxial static state of the flying wing unmanned aerial vehicle, and the technology solves the flying problem of the flying wing unmanned aerial vehicle with extremely low cost, thereby laying a solid foundation for the excellent control of the attitude angle of the flying wing unmanned aerial vehicle;
(3) The inner ring of the invention adopts a method of cross mixing of pure angular velocity and airflow angle, solves the control problem of the three-axis static unstable flying-wing unmanned aerial vehicle, greatly improves the disturbance rejection capability of the attitude angle control of the flying-wing unmanned aerial vehicle, has excellent robust capability, can lead the flying-wing layout unmanned aerial vehicle to exert static unstable flight performance under the condition of controllable attitude, and improves the maneuverability and dexterity of the flying-wing unmanned aerial vehicle.
Detailed Description
In order to make the objects, process conditions and advantages of the present invention more apparent, the present invention will be further described in detail with reference to the following examples, but the embodiments of the present invention are not limited thereto, and various substitutions and modifications according to the general knowledge and conventional means of the art without departing from the technical spirit of the present invention, should be included in the scope of the present invention, and the specific examples described herein are only for explaining the present invention and are not limited thereto.
Example 1:
the embodiment provides a three-axis static unstable flying wing unmanned aerial vehicle attitude angle control method, which specifically comprises the following steps:
step 1, introducing attack angle feedback into an elevator control channel to change longitudinal static stability from static instability to static stability, respectively introducing sideslip angle signals into an aileron rudder and a rudder, changing roll static instability of the flying wing unmanned aerial vehicle into roll static stability, changing yaw static instability into yaw static stability, and feeding back three-axis airflow angles as follows:
delta E、δA、δR is the elevator deflection angle, the aileron deflection angle and the rudder deflection angle, P, Q, R is the roll angle rate, the pitch angle rate and the yaw angle rate, alpha and beta are the attack angle signal and the sideslip angle signal, and K is the proportional coefficient corresponding to each physical quantity.
And 2, designing an attitude angle control loop of the flying-wing unmanned aerial vehicle on the basis of stability enhancement of the attack angle and the sideslip angle. After the stability enhancement of the angle of attack sideslip angle is completed, in order to realize the accurate control of the attitude angle, namely the roll angle and the pitch angle, a pure angular rate control scheme is adopted for the inner ring, and the control strategy of the pure angular rate of the inner ring is as follows:
Where Q g is the pitch rate given signal, P g is the roll angle rate given signal, And the control parameters corresponding to the integral terms are respectively adopted.
Step 3, determining control gain of angle of attack stability enhancement and pitch angle rate control gain by utilizing a pole allocation strategy; the attitude control ring adopts the proportional control of the rolling angle and the pitch angle to realize the static-free control of the rolling angle and the pitch angle. The control strategy is as follows:
Wherein theta and phi are pitch angle and roll angle signals respectively, theta g、φg is pitch angle instruction set and roll angle instruction set respectively, Respectively corresponding proportional coefficients.
While embodiments of the present invention have been shown and described, it will be understood by those of ordinary skill in the art that: many changes, modifications, substitutions and variations may be made to the embodiments without departing from the spirit and principles of the invention, the scope of which is defined by the claims and their equivalents.
Claims (6)
1. The three-axis static unstable flying wing unmanned aerial vehicle attitude angle control method is characterized by comprising the following steps of:
(1) The stability augmentation system is used for improving the static stability and dynamic stability of the aircraft; the elevator airflow angle feedback form is obtained by the stability augmentation system according to an original mathematical formula of the elevator introduced angle of attack signal stability augmentation, and is specifically as follows:
the original mathematical formula of sideslip angle signal stability augmentation is introduced through the aileron rudder to obtain the airflow angle feedback form of the aileron rudder, and the airflow angle feedback form is specifically as follows:
The original mathematical formula of sideslip angle signal stability augmentation is introduced through a rudder to obtain a rudder airflow angle feedback form, and the method is specifically as follows:
Delta E、δA、δR is an elevator deflection angle, an aileron deflection angle and a rudder deflection angle, P, Q, R is a roll angle rate, a pitch angle rate and a yaw angle rate, alpha and beta are attack angle signals and sideslip angle signals, and K is a proportional coefficient corresponding to each physical quantity;
According to the original mathematical formula of the stabilizing system, the specific process for obtaining the elevator airflow angle feedback form is as follows:
The original formula for the elevator to introduce the angle of attack signal is: delta E＝KαΔα+KEΔWE
Wherein delta E is the elevator deflection angle, K α is the proportionality coefficient of the angle of attack stability augmentation signal, delta alpha is the angle of attack stability augmentation signal, K E is the proportionality coefficient of the steering column control elevator command, and delta W E is the steering column control elevator command signal;
The original formula of the sideslip angle signal introduced by the aileron rudder is as follows: delta a＝KβΔβ+KaΔWa
Wherein Δδ a is an aileron deflection angle, K β is a proportionality coefficient of a sideslip angle stability augmentation signal, Δβ is a sideslip angle stability augmentation signal, K a is a proportionality coefficient of a steering column steering aileron command, and Δw a is a steering column steering aileron command signal;
The original formula for the rudder to introduce the sideslip angle signal is: delta r＝KβΔβ+KβΔWr
Wherein delta r is rudder deflection angle, delta W r is pedal quantity signal, a steering rod can operate the elevator, namely K EΔWE can be regarded as original elevator deflection K EΔWE＝δE1, and the angular rate ring is used for proportional controlThen a feedback form of the elevator can be derived
By the same token, the formula for leading the aileron rudder to introduce sideslip angle signals is as follows
(2) On the basis of stability enhancement of the attack angle and the sideslip angle, an attitude angle control loop of the flying wing unmanned aerial vehicle is designed;
(3) And determining the control gain of the angle of attack stability enhancement and the pitch angle rate control gain by using a pole allocation strategy.
2. The method for controlling the attitude angle of the three-axis static unstable flying wing unmanned aerial vehicle according to claim 1, wherein in the step (1), the specific process of improving the static stability and the dynamic stability of the aircraft by using the stability augmentation system is as follows: and the aircraft unmanned aerial vehicle is characterized in that an incidence angle feedback is introduced into an elevator control channel, so that the longitudinal static stability is changed from static instability to static stability, sideslip angle signals are respectively introduced into an aileron rudder and a rudder, the roll static instability of the aircraft unmanned aerial vehicle is changed into roll static stability, and the yaw static instability is changed into yaw static stability.
3. The method for controlling the attitude angle of the three-axis static unstable flying-wing unmanned aerial vehicle according to claim 1 or 2, wherein in the step (2), the specific process of designing the attitude angle control loop of the flying-wing unmanned aerial vehicle on the basis of the stability increase of the attack angle and the sideslip angle is as follows: after the stability increase of the angle of attack sideslip angle is completed, in order to realize the accurate control of the attitude angle, namely the rolling angle and the pitch angle, a pure angular rate control strategy is adopted in the inner ring of the device.
4. The method for controlling attitude angle of three-axis static unstable flying wing unmanned aerial vehicle according to claim 3, wherein the control strategy of pure angular rate of inner ring is:
Where Q g is the pitch rate given signal, P g is the roll angle rate given signal, And the control parameters corresponding to the integral terms are respectively adopted.
5. The method for controlling the attitude angle of the three-axis static unstable flying wing unmanned aerial vehicle according to claim 1 or 2, wherein in the step (3), a strategy of pole allocation is adopted to determine the control gain of the angle of attack for stability augmentation and the control gain of the pitch angle rate, and the specific contents are as follows: the attitude control ring adopts the proportional control of the rolling angle and the pitch angle to realize the static-free control of the rolling angle and the pitch angle.
6. The method for controlling the attitude angle of the three-axis static unstable flying wing unmanned aerial vehicle according to claim 1 or 2, wherein the attitude control ring controls the rolling angle and the pitch angle according to the following strategies:
Respectively controlling pitch angle deviation theta-theta g and roll angle deviation phi-phi g to enable the pitch angle deviation theta-theta g and the roll angle deviation phi-phi g to meet control response static-error-free control;
Wherein, theta and phi are pitch angle and roll angle signals respectively, theta g、φg is pitch angle instruction set and roll angle instruction set respectively, Respectively corresponding proportional coefficients.
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Citations (2)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN106828883A (en) * | 2015-12-07 | 2017-06-13 | 中航通飞研究院有限公司 | The longitudinally controlled augmentation control rule of amphibious aircraft |
| CN114637203A (en) * | 2022-03-15 | 2022-06-17 | 航天时代飞鸿技术有限公司 | Flight control system for medium-high speed and large-sized maneuvering unmanned aerial vehicle |
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| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| IT1392259B1 (en) * | 2008-12-11 | 2012-02-22 | Alenia Aeronautica Spa | ESTIMATION PROCEDURE OF THE INCIDENCE ANGLE AND THE DERAPATE CORNER OF AN AIRCRAFT |
| US10520389B2 (en) * | 2013-11-27 | 2019-12-31 | United States Of America As Represented By The Administrator Of Nasa | Aerodynamic modeling using flight data |
| US10604236B2 (en) * | 2016-06-01 | 2020-03-31 | Regents Of The University Of Minnesota | Fault-tolerant aircraft flight control using a subset of aerodynamic control surfaces |
| CN109407690A (en) * | 2018-12-27 | 2019-03-01 | 湖北航天飞行器研究所 | A kind of aircraft stable control method |
| CN110316358A (en) * | 2019-03-29 | 2019-10-11 | 南京航空航天大学 | Fighter plane High Angle of Attack control method based on dynamic inverse |
| CN110703788A (en) * | 2019-10-16 | 2020-01-17 | 北京航空航天大学 | A Stability Augmentation Control Method for a Micro Flapping-Wing Aircraft and Its Implementation |
| CN111045449A (en) * | 2019-11-26 | 2020-04-21 | 成都飞机工业（集团）有限责任公司 | Nonlinear compensation-based unmanned aerial vehicle roller power control method |
| CN111717372A (en) * | 2020-05-22 | 2020-09-29 | 成都飞机工业（集团）有限责任公司 | Large-overload disc-stabilizing maneuvering control method for flying-wing unmanned aerial vehicle |
| CN112433534B (en) * | 2020-11-13 | 2022-04-01 | 南京航空航天大学 | Flying wing unmanned aerial vehicle semi-rolling reverse maneuvering flight control method |
| CN112486193B (en) * | 2020-11-19 | 2022-04-22 | 南京航空航天大学 | Three-axis full-authority control method of flying-wing unmanned aerial vehicle based on self-adaptive augmentation control theory |
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Patent Citations (2)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN106828883A (en) * | 2015-12-07 | 2017-06-13 | 中航通飞研究院有限公司 | The longitudinally controlled augmentation control rule of amphibious aircraft |
| CN114637203A (en) * | 2022-03-15 | 2022-06-17 | 航天时代飞鸿技术有限公司 | Flight control system for medium-high speed and large-sized maneuvering unmanned aerial vehicle |
Non-Patent Citations (1)
| Title |
|---|
| 静不稳定飞翼无人机机动飞行控制技术研究;杨艺;《中国博士学位论文全文数据库 工程科技II辑》;20160715(第07(2016)期);C031-22 * |
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