CN113212733B - Large-aspect-ratio conventional-layout unmanned aerial vehicle gust load alleviation method - Google Patents
Large-aspect-ratio conventional-layout unmanned aerial vehicle gust load alleviation method Download PDFInfo
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- CN113212733B CN113212733B CN202110479502.7A CN202110479502A CN113212733B CN 113212733 B CN113212733 B CN 113212733B CN 202110479502 A CN202110479502 A CN 202110479502A CN 113212733 B CN113212733 B CN 113212733B
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- B—PERFORMING OPERATIONS; TRANSPORTING
- B64—AIRCRAFT; AVIATION; COSMONAUTICS
- B64C—AEROPLANES; HELICOPTERS
- B64C3/00—Wings
- B64C3/10—Shape of wings
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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
- B64C13/24—Transmitting means
- B64C13/38—Transmitting means with power amplification
- B64C13/50—Transmitting means with power amplification using electrical energy
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- B—PERFORMING OPERATIONS; TRANSPORTING
- B64—AIRCRAFT; AVIATION; COSMONAUTICS
- B64C—AEROPLANES; HELICOPTERS
- B64C9/00—Adjustable control surfaces or members, e.g. rudders
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- B—PERFORMING OPERATIONS; TRANSPORTING
- B64—AIRCRAFT; AVIATION; COSMONAUTICS
- B64C—AEROPLANES; HELICOPTERS
- B64C9/00—Adjustable control surfaces or members, e.g. rudders
- B64C9/32—Air braking surfaces
- B64C9/323—Air braking surfaces associated with wings
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- B—PERFORMING OPERATIONS; TRANSPORTING
- B64—AIRCRAFT; AVIATION; COSMONAUTICS
- B64C—AEROPLANES; HELICOPTERS
- B64C3/00—Wings
- B64C3/10—Shape of wings
- B64C3/14—Aerofoil profile
- B64C2003/147—Aerofoil profile comprising trailing edges of particular shape
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- B—PERFORMING OPERATIONS; TRANSPORTING
- B64—AIRCRAFT; AVIATION; COSMONAUTICS
- B64C—AEROPLANES; HELICOPTERS
- B64C9/00—Adjustable control surfaces or members, e.g. rudders
- B64C2009/005—Ailerons
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Abstract
The application belongs to the technical field of unmanned aerial vehicles, and particularly relates to a sudden wind load alleviation method for an unmanned aerial vehicle with a large aspect ratio and a conventional layout, which comprises the following steps: after the wind is met, starting a sudden wind slowing mode, placing the ailerons positioned at the trailing edge of the wings in a loose floating mode, and rapidly reducing sudden wind load increment generated by high-frequency short-period wind on the wings by utilizing the floating action of the ailerons along with the wind under the influence of the sudden wind; and then deflecting the elevator to generate a moment opposite to the gust to control the attitude and the track of the airplane. The application provides a sudden wind load reduction design method for a large-aspect-ratio conventional-layout unmanned aerial vehicle, the redesign of a main wing surface (wing) structure is not needed, the structure is simple, the weight is light, the cost is low, the descent of local loads is brought only by the fact that ailerons at the tail edges of the wings are free to float and deflect under the action of sudden wind, and the sudden wind overload increase of the wings can be effectively reduced in a short time.
Description
Technical Field
The application belongs to the technical field of unmanned aerial vehicles, and particularly relates to a sudden wind load alleviation method for an unmanned aerial vehicle with a large aspect ratio and a conventional layout.
Background
When an aircraft is subjected to gust disturbances during flight, particularly vertical gusts and turbulence, it is common practice to deflect the elevators to produce a pitching moment to turn the aircraft, with the change in pitch attitude controlling the longitudinal movement. The result is: on one hand, a certain delay time is generated; on the other hand, the influence of overload generated by gust on the airplane cannot be effectively inhibited, so that the airplane jolts for a long time and the fatigue life of the airplane is influenced.
In order to avoid the problem, a direct lift control concept appears, namely, a closed-loop control system is adopted, the acceleration deviation of the machine body is used as an excitation signal, the additional control surface is driven to deflect to complete direct lift control, and gust load alleviation is carried out. The essence of the method is that the overload generated by gust is restrained by active control of direct lift force, and the influence of aerodynamic load on the airplane is relieved. For example, when encountering a disturbance in the updraft in flight, the flight control system actively steers additional control surfaces (e.g., two side ailerons) while deflecting upward to reduce positive lift; when the aircraft is disturbed by downwash, the ailerons on the two sides deflect downwards to reduce the negative lift force, and the influence of overload generated by gust on the aircraft can be quickly and effectively inhibited. However, the method has high response requirements on a control system, the wind changes are not regular, the scale and the strength combination are not complex, the requirement of real-time control surface manipulation for reducing the sudden wind load is met, high resolving requirements are provided for a flight control computer, once the frequency of real-time control surface manipulation is coupled with attitude response and the frequency of the machine body structure, the sudden wind load can be further increased, and the design difficulty and the risk are high.
The high aspect ratio unmanned aerial vehicle generally arranges multiunit control surface on the wing, provides multiple possibility for load slow down mode when meeting wind.
The patent of the prior related improved patent, such as patent application number 201310643782.6, named as 'a full-motion wingtip gust retarding device applicable to a high-aspect-ratio airplane', discloses a full-motion wingtip gust retarding device applicable to a high-aspect-ratio airplane, and belongs to the field of wing structure design of high-aspect-ratio airplanes. The device consists of a main wing, a full-motion wing tip, a torsional elastic connecting shaft, a locking mechanism, a gust detection sensor and an adjusting device. The full-motion wingtip accounts for 10% -20% of the span length of the wing and is connected with the main wing through the torsional elastic connecting shaft and the locking mechanism. The wing tip is provided with a gust detection sensor for detecting vertical gust, and the adjusting device consisting of the lead screw, the lead screw nut, the motor and the slide rail can adjust the position of the torsional elastic connecting shaft, so that the full-motion wing tip gust retarding device has the best gust retarding effect under different gust intensities. The above patent requires redesign of the wing structure, the result is complex and the actual use effect is not ideal.
Disclosure of Invention
In order to overcome the defects of the prior art, the invention provides a sudden wind load alleviation method of an unmanned aerial vehicle with a large aspect ratio and a conventional layout, which adopts a method of combining passive direct force with active moment control to fulfill the aim of alleviating sudden wind load.
In order to achieve the purpose, the technical scheme of the application is as follows:
a sudden wind load alleviation method for a large-aspect-ratio conventional-layout unmanned aerial vehicle comprises the following steps:
after the wind is met, starting a sudden wind slowing mode, placing the ailerons positioned at the trailing edge of the wings in a loose floating mode, and rapidly reducing sudden wind load increment generated by high-frequency short-period wind on the wings by utilizing the floating action of the ailerons along with the wind under the influence of the sudden wind; and then deflecting the elevator to generate a moment opposite to the gust to control the attitude and the track of the airplane.
Further, the specific steps are as follows:
step 1) determination of a sudden wind alleviation requirement
Continuous multi-beat judgment of whether vertical acceleration ay exceeds safety threshold aySecureIf so, judging that a slowing measure needs to be taken, and starting a sudden wind slowing mode; otherwise, the sudden wind retarding mode is not started, and the aircraft flies normally;
step 2) loosening and floating of wing control surface
And (3) redistributing the functions of all groups of control surfaces of the wings in different modes: a) when the vertical acceleration ay does not exceed the safety threshold aySecureWhen the airplane is used, the control surface on the wing is operated according to the normal function; b) when the vertical acceleration ay exceeds the safety threshold aySecureWhen the airplane is in use, the inner control surface of the reserved part is used as an aileron, the rest outer control surfaces are placed in a loose and floating state, the loose and floating control surface is passively and quickly deflected to a low-load state after being influenced by gust in an uncontrolled state, and the aerodynamic load and the airplane are in a pneumatic load stateThe gust load increment directions of the wing stabilizing surfaces are opposite; the inner aileron is reserved for the aileron operation function, and both the outer ailerons can be used for reducing the gust load and are placed in a loose and floating state;
step 3) elevator control
The elevator channel adopts a control function, the vertical acceleration increment delta ay, the pitch angle delta phi and the height parameter delta H which exceed a threshold value are used as the input of the elevator, the elevator instruction needing to be operated is obtained through resolving of the control function, then the instruction signal is converted into a mechanical signal through a servo loop, the elevator is driven by a servo actuator to deflect, and therefore all attitude and trajectory parameters are improved.
Further, the multiple beats in the step 1) are 5-10 beats, and the interval time of each beat is 0.02s-0.05 s.
Further, the vertical acceleration ay is an acceleration perpendicular to the horizontal plane of the unmanned aerial vehicle body axis, and is positive upwards.
Further, the floating state in step 2) means that no force is applied to the control surface, and the control surface is allowed to float with the wind.
Further, the elevator deflection feedback increment control function in the step 3) is as follows:
△δZ = K△ay×△ay + K△Φ×△Φ + K△H×△H (1)
in the above formula:
K△aythe magnitude is optimally set according to a control target;
K△Φfeeding back parameters for the change of the pitching attitude angle of the unmanned aerial vehicle, and optimally setting the quantity value according to a control target;
K△Hthe flight height change feedback parameter is used, and the value is optimally set according to a control target;
△ay= ay -10。
the invention has the following beneficial effects:
the application provides a sudden wind load alleviation design method for a large-aspect-ratio conventional-layout unmanned aerial vehicle, the redesign of the structure of a main wing surface (wing) is not needed, the structure is simple, the weight is light, the cost is low, the reduction of local load is brought only by the free floating deflection of an aileron at the rear edge of the wing under the action of sudden wind, and the sudden wind overload increase of the wing can be effectively alleviated in a short time; and then, the elevators are used in a matched manner, so that the effective control of long-period physical quantities such as the attitude and the track of the airplane can be met. The method reduces the complicated control law flight control resolving link, reduces the coupling risk of the control system and parameters such as attitude response frequency, body structure frequency and the like, does not increase the burden of a flight control computer, and is low in risk, simple and effective.
Drawings
Fig. 1 is a high aspect ratio drone gust load mitigation logic block diagram.
Fig. 2 is a schematic diagram of the high aspect ratio conventional layout unmanned aerial vehicle gust load mitigation control surface function allocation.
Fig. 3 is a schematic of elevator lane control logic.
FIG. 4 is a schematic diagram of deviation variation of ailerons in a floating state. Wherein the vertical axis is angle and the horizontal axis is time.
Fig. 5 is a schematic view of rudder angle change. Wherein the vertical axis is angle and the horizontal axis is time.
FIG. 6 is a graph comparing vertical acceleration. Where the vertical axis is speed and the horizontal axis is time.
Fig. 7 is a schematic view of pitch angle comparison. Wherein the vertical axis is angle and the horizontal axis is time.
FIG. 8 is a schematic diagram comparing height variations. Wherein the vertical axis is height and the horizontal axis is time.
Detailed Description
Example 1
A sudden wind load alleviation method for a large-aspect-ratio conventional-layout unmanned aerial vehicle comprises the following steps:
after the wind is met, starting a sudden wind slowing mode, placing the ailerons positioned at the trailing edge of the wings in a loose floating mode, and rapidly reducing sudden wind load increment generated by high-frequency short-period wind on the wings by utilizing the floating action of the ailerons along with the wind under the influence of the sudden wind; and then deflecting the elevator to generate a moment opposite to the gust to control the attitude and the track of the airplane.
Further, the specific steps are as follows:
step 1) determination of a sudden wind alleviation requirement
Continuous multi-beat judgment of whether vertical acceleration ay exceeds safety threshold aySecureIf so, judging that a slowing measure needs to be taken, and starting a sudden wind slowing mode; otherwise, the sudden wind retarding mode is not started, and the aircraft flies normally;
the multiple beats are 5-10 beats, and the interval time of each beat is 0.02-0.05 s.
The vertical acceleration ay is the acceleration perpendicular to the horizontal plane of the unmanned aerial vehicle body axis, and is positive upwards.
Step 2) loosening and floating of wing control surface
And (3) redistributing the functions of all groups of control surfaces of the wings in different modes: a) when the vertical acceleration ay does not exceed the safety threshold aySecureWhen the airplane is used, the control surface on the wing is operated according to the normal function; b) when the vertical acceleration ay exceeds the safety threshold aySecureWhen the airplane is in a non-control state, the loose control surface is passively and quickly deflected to a low-load state after being influenced by gust, and the pneumatic load of the loose control surface is opposite to the gust load increment direction of the wing stabilizing surface; the inner aileron is reserved for the aileron operation function, and both the outer ailerons can be used for reducing the gust load and are placed in a loose and floating state; the floating state means that no force is applied to the control surface, and the control surface is allowed to float with the wind.
Step 3) elevator control
The elevator channel adopts a control function, the vertical acceleration increment delta ay, the pitch angle delta phi and the height parameter delta H which exceed a threshold value are used as the input of the elevator, the elevator instruction needing to be operated is obtained through the calculation of the control function, then an instruction signal is converted into a mechanical signal through a servo loop, the elevator is driven by a servo actuator to deflect, and further, all the attitude and track parameters are improved.
The elevator deflection feedback increment control function is as follows:
△δZ = K△ay×△ay + K△Φ×△Φ + K△H×△H (1)
in the above formula:
K△aythe magnitude is optimally set according to a control target;
K△Φfeeding back parameters for the change of the pitching attitude angle of the unmanned aerial vehicle, and optimally setting the quantity value according to a control target;
K△Hthe flight height change feedback parameter is used, and the value is optimally set according to a control target;
△ay= ay -10。
the application provides a sudden wind load alleviation design method for a large-aspect-ratio conventional-layout unmanned aerial vehicle, the redesign of the structure of a main wing surface (wing) is not needed, the structure is simple, the weight is light, the cost is low, the reduction of local load is brought only by the free floating deflection of an aileron at the rear edge of the wing under the action of sudden wind, and the sudden wind overload increase of the wing can be effectively alleviated in a short time; and then, the elevators are used in a matched manner, so that the effective control of long-period physical quantities such as the attitude and the track of the airplane can be met. The method reduces the complicated control law flight control resolving link, reduces the coupling risk of the control system and parameters such as attitude response frequency, body structure frequency and the like, does not increase the burden of a flight control computer, and is low in risk, simple and effective.
Example 2
The embodiment of the size parameter and gust load alleviation implementation method of the high-aspect-ratio unmanned aerial vehicle is provided in the example. The wing has a half span length of 6 m and a wing chord length of 1 m; three groups of ailerons are arranged on the wings, and a group of elevators is arranged on the horizontal tail. The maximum vertical gust speed is 4.2m/s, and the airplane flying speed is 100 m/s; the ailerons initially slow down the yaw and the elevator starts the yaw after 50m of height reduction.
In this example, the response conditions of parameters such as vertical acceleration, pitch angle, trajectory and the like with and without gust load alleviation control are compared, as shown in fig. 4-8, the comparison shows that: the vertical acceleration is reduced by more than 20 percent (the corresponding gust load is also reduced by more than 20 percent), and the pitch angle and the track can be effectively controlled. In actual use, the parameter values in the control function can be adjusted according to the target optimization.
Claims (5)
1. A sudden wind load alleviation method for a large-aspect-ratio conventional-layout unmanned aerial vehicle is characterized by comprising the following steps of: the method comprises the following steps:
after the wind is met, starting a sudden wind slowing mode, placing the ailerons positioned at the trailing edge of the wings in a loose floating mode, and rapidly reducing sudden wind load increment generated by high-frequency short-period wind on the wings by utilizing the floating action of the ailerons along with the wind under the influence of the sudden wind; then, the elevator is deflected to generate a moment opposite to the gust, and the attitude and the track of the airplane are controlled;
the method comprises the following specific steps:
step 1) determination of a sudden wind alleviation requirement
Continuous multi-beat judgment of whether vertical acceleration ay exceeds safety threshold aySecureIf so, judging that a slowing measure needs to be taken, and starting a sudden wind slowing mode; otherwise, the sudden wind retarding mode is not started, and the aircraft flies normally;
step 2) loosening and floating of wing control surface
And (3) redistributing the functions of all groups of control surfaces of the wings in different modes: a) when the vertical acceleration ay does not exceed the safety threshold aySecureWhen the airplane is used, the control surface on the wing is operated according to the normal function; b) when the vertical acceleration ay exceeds the safety threshold aySecureWhen the airplane is in a non-control state, the loose control surface is passively and quickly deflected to a low-load state after being influenced by gust, and the pneumatic load of the loose control surface is opposite to the gust load increment direction of the wing stabilizing surface; the inner aileron is reserved for the aileron operation function, and both the outer ailerons can be used for reducing the gust load and are placed in a loose and floating state;
step 3) elevator control
The elevator channel adopts a control function, the vertical acceleration increment delta ay, the pitch angle delta phi and the height parameter delta H which exceed a threshold value are used as the input of the elevator, the elevator instruction needing to be operated is obtained through resolving of the control function, then the instruction signal is converted into a mechanical signal through a servo loop, the elevator is driven by a servo actuator to deflect, and therefore all attitude and trajectory parameters are improved.
2. The method of claim 1, wherein the method comprises the following steps: the multiple beats in the step 1) are 5-10 beats, and the interval time of each beat is 0.02-0.05 s.
3. The method of claim 1, wherein the method comprises the following steps: the vertical acceleration ay is the acceleration perpendicular to the horizontal plane of the unmanned aerial vehicle body axis, and is positive upwards.
4. The method of claim 1, wherein the method comprises the following steps: the floating state in the step 2) means that no force is applied to the control surface, and the control surface is allowed to float along with wind.
5. The method of claim 1, wherein the method comprises the following steps: the elevator deflection feedback increment control function in the step 3) is as follows:
△δZ = K△ay×△ay + K△Φ×△Φ + K△H×△H (1)
in the above formula:
K△aythe magnitude is optimally set according to a control target;
K△Φfeeding back parameters for the change of the pitching attitude angle of the unmanned aerial vehicle, wherein the quantity value is optimally set according to a control target;
K△Hthe flight height change feedback parameter is used, and the value is optimally set according to a control target;
△ay= ay -10。
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Citations (5)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| GB564408A (en) * | 1941-12-30 | 1944-09-27 | Phillips & Powis Aircraft Ltd | Improvements in means for actuating aircraft control surfaces |
| US6161801A (en) * | 1998-04-30 | 2000-12-19 | Daimlerchrysler Aerospace Airbus Gmbh | Method of reducing wind gust loads acting on an aircraft |
| EP2778397A1 (en) * | 2013-03-15 | 2014-09-17 | Frontier Wind, LLC | Actuation mechanisms for load management devices on aerodynamic blades |
| CN109902330A (en) * | 2018-11-12 | 2019-06-18 | 中航通飞研究院有限公司 | A kind of aerofoil and its strength calculation method of no stringer and rib web |
| CN112698565A (en) * | 2020-12-28 | 2021-04-23 | 北京航空航天大学 | Aircraft auxiliary piloting system and method for asymmetric thrust compensation in approach phase |
Family Cites Families (10)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN103640687B (en) * | 2013-12-03 | 2015-12-30 | 北京航空航天大学 | A kind of full dynamic formula Full-motion wingtip gust alleviation device being applicable to high-aspect-ratio aircraft |
| US20170021914A1 (en) * | 2013-12-11 | 2017-01-26 | Moog Inc. | Fiber optic sensing and control system |
| GB2546246A (en) * | 2016-01-05 | 2017-07-19 | Airbus Operations Ltd | An aircraft wing with a movable wing tip device for load alleviation |
| US10710702B2 (en) * | 2016-08-12 | 2020-07-14 | Aviation Partners, Inc. | Shape adaptive airfoil |
| JP7097052B2 (en) * | 2018-04-04 | 2022-07-07 | 国立研究開発法人宇宙航空研究開発機構 | Airplane gust response mitigation system and airplane gust response mitigation method |
| CN108516101B (en) * | 2018-04-28 | 2020-02-14 | 北京航空航天大学 | Active and passive combined control method for reducing gust of fixed-wing aircraft |
| GB2574603A (en) * | 2018-06-11 | 2019-12-18 | Airbus Operations Ltd | Vertical stabilizer for an aircraft |
| US10889369B2 (en) * | 2018-08-29 | 2021-01-12 | Textron Innovations Inc. | Passive gust alleviation systems for aircraft devices |
| CN109131833A (en) * | 2018-09-28 | 2019-01-04 | 成都飞机工业（集团）有限责任公司 | A kind of high aspect ratio wing of high lift-rising |
| CN112607054B (en) * | 2020-12-29 | 2023-01-13 | 中国航空工业集团公司西安飞机设计研究所 | Model is slowed down in wing gust |
-
2021
- 2021-04-30 CN CN202110479502.7A patent/CN113212733B/en active Active
Patent Citations (5)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| GB564408A (en) * | 1941-12-30 | 1944-09-27 | Phillips & Powis Aircraft Ltd | Improvements in means for actuating aircraft control surfaces |
| US6161801A (en) * | 1998-04-30 | 2000-12-19 | Daimlerchrysler Aerospace Airbus Gmbh | Method of reducing wind gust loads acting on an aircraft |
| EP2778397A1 (en) * | 2013-03-15 | 2014-09-17 | Frontier Wind, LLC | Actuation mechanisms for load management devices on aerodynamic blades |
| CN109902330A (en) * | 2018-11-12 | 2019-06-18 | 中航通飞研究院有限公司 | A kind of aerofoil and its strength calculation method of no stringer and rib web |
| CN112698565A (en) * | 2020-12-28 | 2021-04-23 | 北京航空航天大学 | Aircraft auxiliary piloting system and method for asymmetric thrust compensation in approach phase |
Non-Patent Citations (1)
| Title |
|---|
| 气动伺服弹性研究的进展与挑战;杨超，黄超，吴志刚，唐长红;《航空学报》;20150425;第36卷(第4期);全文 * |
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