CN112327908B - Stable control method suitable for low rudder effect separation state - Google Patents
Stable control method suitable for low rudder effect separation state Download PDFInfo
- Publication number
- CN112327908B CN112327908B CN202011156237.0A CN202011156237A CN112327908B CN 112327908 B CN112327908 B CN 112327908B CN 202011156237 A CN202011156237 A CN 202011156237A CN 112327908 B CN112327908 B CN 112327908B
- Authority
- CN
- China
- Prior art keywords
- rudder
- control
- channel
- low
- filter
- Prior art date
- Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
- Active
Links
- 230000000694 effects Effects 0.000 title claims abstract description 46
- 238000000926 separation method Methods 0.000 title claims abstract description 44
- 238000000034 method Methods 0.000 title claims abstract description 23
- 238000001914 filtration Methods 0.000 claims abstract description 17
- 238000005096 rolling process Methods 0.000 claims description 19
- 238000013016 damping Methods 0.000 claims description 11
- 238000010008 shearing Methods 0.000 claims description 3
- 238000012546 transfer Methods 0.000 claims description 3
- 230000000087 stabilizing effect Effects 0.000 claims description 2
- 238000013461 design Methods 0.000 abstract description 4
- 230000005764 inhibitory process Effects 0.000 abstract description 2
- 238000010586 diagram Methods 0.000 description 4
- 238000005452 bending Methods 0.000 description 1
- 230000009286 beneficial effect Effects 0.000 description 1
- 230000007547 defect Effects 0.000 description 1
- 230000002401 inhibitory effect Effects 0.000 description 1
- 238000005259 measurement Methods 0.000 description 1
- 238000012986 modification Methods 0.000 description 1
- 230000004048 modification Effects 0.000 description 1
- 238000012545 processing Methods 0.000 description 1
- 230000001629 suppression Effects 0.000 description 1
Images
Classifications
-
- 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
- G05D1/107—Simultaneous control of position or course in three dimensions specially adapted for missiles
Landscapes
- Engineering & Computer Science (AREA)
- Aviation & Aerospace Engineering (AREA)
- Radar, Positioning & Navigation (AREA)
- Remote Sensing (AREA)
- Physics & Mathematics (AREA)
- General Physics & Mathematics (AREA)
- Automation & Control Theory (AREA)
- Feedback Control In General (AREA)
- Control Of Position, Course, Altitude, Or Attitude Of Moving Bodies (AREA)
Abstract
The invention discloses a stable control method suitable for a low-rudder-effect separation state, belongs to the field of stable control of aircrafts, and improves the stability and the high-frequency signal inhibition capability of a system in the low-rudder-effect separation state through the coordinated design of stable control loop control parameters and filter parameters, so that the aircrafts are stably controlled in the low-rudder-effect state. The invention comprises the following steps: (1) Adjusting control parameters according to flight time, speed and dynamic pressure adaptability, and reducing the control parameters in a low rudder effect separation state; (2) Under the low rudder effect separation state, adding amplitude limiting protection to the control parameters; (3) And parameters of the filter are adjusted according to the flight time adaptability, and the filtering depth and the filtering width of the filter are increased under the low rudder effect separation state.
Description
Technical Field
The invention belongs to the field of aircraft stability control, and particularly relates to an aircraft stability control method in low rudder effect state separation. The device is suitable for the stable control of the aircraft separated from the carrier or the feeder under the high-altitude low-rudder-effect state, or the stable control of the aircraft separated from the booster, and is also suitable for the stable control of the aircraft separated from the ground launcher under the low-rudder-effect state.
Background
The stability control system is an important component of the aircraft, and is used for effectively inhibiting initial attitude angular velocity interference and aerodynamic interference, responding to an attitude instruction and controlling the attitude of the aircraft to meet certain requirements in the aircraft separation process.
In the aircraft separation process, the initial attitude angular velocity exists due to the acting force of the separation mechanism; and the aircraft is close to the delivery device/carrier or the aircraft is close to the booster stage, so that complex flow field interference exists. And the low rudder effect state is separated, the rudder effect of the aircraft is too low, and the control capability is limited. On the other hand, in a low rudder effect state, the aircraft is impacted by a separating mechanism, and elastic vibration is easily excited.
In the prior art, the coordination design of the control parameters of the stable control loop and the parameters of the filter is not performed aiming at the low rudder effect separation state: (1) In a full airspace flight state, the filtering depth and the filtering width of a filter are not changed generally, a larger high-frequency rudder instruction signal may appear in a low rudder effect separation state, the working condition of a rudder system is deteriorated, and instability can be caused in a serious condition; (2) Control parameter amplitude limiting protection is not added aiming at a low-rudder-effect separation state, generally, control parameters of a control loop are obviously higher than those of a high-rudder-effect flight state in a low-rudder-effect flight state, a large rudder deflection requirement can be generated under small instruction input or angular velocity interference, the requirements on a rudder deflection angle and a rudder deflection angular velocity are large, a rudder system is in a nonlinear working state in a short time, the stability margin of a stability control system can be reduced, and instability can be caused in a serious situation.
Disclosure of Invention
The invention aims to provide a stable control method suitable for a low-rudder-efficiency separation state, which ensures that an aircraft is stably controlled when the initial interference is large and the elastic vibration magnitude is large in the low-rudder-efficiency separation state. According to the technical scheme, the stability of the system and the high-frequency signal inhibition capability in the low-rudder-effect separation state are improved through the coordinated design of the control parameters of the stable control loop and the parameters of the filter, and the stable control of the aircraft in the low-rudder-effect state is ensured.
In order to achieve the above object, the present invention provides a stable control method suitable for a low rudder effect separation state, comprising the steps of:
firstly, adjusting control parameters according to flight time, speed and dynamic pressure adaptability, and reducing the control parameters in a low rudder effect separation state;
secondly, adding amplitude limiting protection to the control parameters in a low rudder effect separation state;
and step three, adjusting filter parameters according to the flight time adaptability, and increasing the filtering depth and the filtering width of the filter in a low rudder effect separation state.
In the above stable control method for the low rudder effect separation state, in the first step, in the low rudder effect separation state, the reducing of the control parameter specifically includes: the open-loop shearing frequency designed in the low-rudder-effect state is lower than that in the high-rudder-effect state, and the system rapidity designed in the low-rudder-effect state is lower than that in the high-rudder-effect state.
In the second step, firstly, the amplitude limiting value of the damping loop control parameter is determined, and then the value of the attitude control loop control parameter after amplitude limiting is obtained; the control parameters comprise a pitching channel control parameter, a yawing channel control parameter and a rolling channel control parameter.
The above stable control method suitable for the low rudder effect separation state, wherein in the second step, amplitude limiting protection is added to the pitch channel control parameter and the yaw channel control parameter, and specifically:
km’＝km×Kb 1 ×Kb 1
in the formula: ksf is a damping loop control parameter, km' is an attitude control loop control parameter after amplitude limiting,
is the clipping value of the control parameter ksf. Kb 1 The intermediate variable represents the ratio of the clipping value of ksf to the pre-clipping ksf.
The above stable control method suitable for the low rudder effect separation state, wherein in the second step, amplitude limiting protection is added to the control parameters of the rolling channel, and specifically:
g 3 ’＝g 3 ×Kb 2 ×Kb 2
in the formula: g 2 Damping loop control parameter for rolling channel, g 3 ' are rolling channel attitude control loop control parameters after clipping,
to control the parameter g 2 The clipping value of (1). Kb 2 Is an intermediate variable, representing g 2 Amplitude limit value and pre-amplitude limit g 2 Is measured in the measurement.
The above stable control method suitable for the low rudder effect separation state, wherein in step three, the filter parameter is adaptively adjusted according to the flight time, and specifically:
taking a notch filter as an example, the filter is in the form of
In the formula, T n 、T d Is a parameter, ξ, determining the center frequency of the filter n 、ξ d Are parameters that determine the notch width and depth of the filter. Under the condition of low rudder effect separation (i.e. t < t) 1 Time) increase xi d The method increases the filtering depth and width of the filter, and specifically comprises the following steps:
in which ξ 1 ＞ξ 2 。
The stable control method suitable for the low rudder effect separation state is characterized in that a rudder deflection instruction is calculated according to the control parameters and the filter parameters obtained in the first step, the third step, the control surface is driven to deflect, control force and control moment are applied, and the missile attitude is stabilized.
u dR ＝g 2 ×u gyR -g 3 ×u ΔR
u dP ＝(ksf×u gyP -km×u ΔP )×W(z)
u dY ＝(-ksf×u gyY -km×u ΔY )×W(z)
In the formula u dR 、u dP 、u dY The rudder deflection instructions u of a rolling channel, a pitching channel and a yawing channel ΔR 、u ΔP 、u ΔY Attitude deviation instructions u for the roll channel, pitch channel, and yaw channel, respectively gyR 、u gyP 、u gy Y is the output quantity of the rate gyro of the rolling channel, the pitching channel and the yawing channel respectively, and W (z) is the transfer function after discretization of the pitching channel filter.
Compared with the prior art, the invention has the technical beneficial effects that:
(1) By reducing the control parameters of the low rudder effect separation state, the large rudder deflection requirement under the angular velocity disturbance in the low rudder effect separation state is avoided on the one hand, and on the other hand, the rudder deflection requirement caused by the elastic vibration signal is reduced.
(2) And in a low-rudder-effect separation state, the control parameters are reduced, and meanwhile, amplitude limiting protection is added to the control parameters, so that the bandwidth of a stability control loop is reduced, the stability margin of the system is improved, and the defect of insufficient stability margin after the filter is deepened and widened is avoided.
(3) The filtering depth and the filtering width of the filter are adjusted according to the flight time adaptability, and the filtering depth and the filtering width of the filter are increased under the low rudder effect separation state, so that the suppression effect on high-frequency vibration signals is greatly improved.
Drawings
The present invention provides a stable control method for a low rudder effect separation state, which is provided by the following embodiments and accompanying drawings.
FIG. 1 is a block diagram of the control principle of the pitch channel;
FIG. 2 is a block diagram of the control principle of the yaw channel;
fig. 3 is a control schematic block diagram of the scroll channel.
Detailed Description
The following will describe in further detail a stable control method for low rudder effect disengaged state according to the present invention with reference to the accompanying drawings.
The method is explained by taking a rolling stable axisymmetric three-channel control aircraft as an example, and control schematic diagrams of three channels can be respectively shown in figures 1-3, wherein ksf is a damping loop control parameter, km is an attitude control loop control parameter, and g is 2 Damping loop control parameter for rolling channel, g 3 Control loop control parameters for the scrolling channel attitude. The following steps are described:
1. adjusting control parameters according to flight time, speed and dynamic pressure adaptability, and reducing the control parameters of the low rudder effect separation state (it should be noted that the design method for reducing the control parameters in the low rudder effect state is that the open-loop shearing frequency designed in the low rudder effect state is smaller than that in the high rudder effect state, the designed system rapidity is lower than that in the high rudder effect state, and the control parameters of the low rudder effect are not smaller than that in the high rudder effect), the formula is as follows:
ksf＝f 1 (t,q,Vm)
km＝f 2 (t,q,Vm)
g2＝f 3 (t,q,Vm)
g3＝f 4 (t,q,Vm)
in the formula, t is flight time, q is dynamic pressure, and Vm is velocity.
2. Under the low rudder effect separation state, amplitude limiting protection is added to the control parameters, and the amplitude limiting value of the damping loop is determined firstly, so that the amplitude-limited value of the attitude control loop is obtained;
2.1, carrying out amplitude limiting processing on the deflection channel control parameters ksf and km according to the following formula:
km’＝km×Kb 1 ×Kb 1
in the formula: ksf is a damping loop control parameter, km' is an attitude control loop control parameter after amplitude limiting,
is the clipping value of the control parameter ksf. Kb 1 The intermediate variable represents the ratio of the clipping value of ksf to the pre-clipping ksf.
2.2 controlling the parameter g for the rolling channel according to the following formula 2 、g 3 Carrying out amplitude limiting treatment:
g 3 ’＝g 3 ×Kb 2 ×Kb 2
in the formula: g 2 Damping loop control parameter for rolling channel, g 3 ' rolling after clippingThe channel attitude control loop controls the parameters of the channel,
to control the parameter g 2 The clipping value of (1). Kb 2 Is an intermediate variable, representing g 2 Amplitude limit value and pre-amplitude limit g 2 The ratio of (a) to (b).
3. And parameters of the filter are adjusted according to the flight time adaptability, and the filtering depth and the filtering width of the filter are increased under the low rudder effect separation state.
Taking a typical notch filter as an example, the filter is in the form of
Increasing filter parameter xi in low rudder effect airspace d And the filtering depth and width of the filter are improved.
Filter parameter xi d The calculation formula of (c) is as follows:
wherein ξ 1 ＞ξ 2 。
4. And (4) calculating a rudder deflection instruction according to the control parameters and the filter parameters obtained in the steps 1-3, driving the control surface to deflect, applying a control force and a control moment, and stabilizing the posture of the missile.
u dR ＝g 2 ×u gyR -g 3 ×u ΔR
u dP ＝(ksf×u gyP -km×u ΔP )×W(z)
u dY ＝(-ksf×u gyY -km×u ΔY )×W(z)
In the formula u dR 、u dP 、u dY The rudder deflection instructions u of a rolling channel, a pitching channel and a yawing channel ΔR 、u ΔP 、u ΔY Attitude deviation commands u for the roll channel, pitch channel, yaw channel, respectively gyR 、u gyP 、u gyY The rate gyro output quantities of the rolling channel, the pitching channel and the yawing channel are respectively. W (z) is a transfer function after discretization of the bending deflection channel filter.
Those skilled in the art will appreciate that those matters not described in detail in the present specification are well known in the art.
Although the present invention has been described with reference to the preferred embodiments, it is not intended to limit the present invention, and those skilled in the art can make variations and modifications of the present invention without departing from the spirit and scope of the present invention by using the methods and technical contents disclosed above.
Claims (4)
1. A stable control method suitable for a low rudder effect separation state is characterized by comprising the following steps:
firstly, adjusting control parameters according to flight time, speed and dynamic pressure adaptability, and reducing the control parameters in a low rudder effect separation state;
secondly, adding amplitude limiting protection to the control parameters in a low rudder effect separation state;
thirdly, adjusting filter parameters according to flight time adaptability, and increasing the filtering depth and the filtering width of the filter in a low rudder effect separation state;
in the first step, in a low rudder effect separation state, the control parameters are reduced, specifically: the open-loop shearing frequency designed in the low-rudder-effect state is lower than that in the high-rudder-effect state, and the system rapidity designed in the low-rudder-effect state is lower than that in the high-rudder-effect state;
in the second step, firstly, the amplitude limiting value of the damping loop control parameter is determined, and then the value of the attitude control loop control parameter after amplitude limiting is obtained; the control parameters comprise a pitching channel control parameter, a yawing channel control parameter and a rolling channel control parameter;
in the third step, the filter parameter adjusting method is as follows:
taking a notch filter as an example, the filter is in the form of
In the formula, T n 、T d Is a parameter, ξ, that determines the center frequency of the filter n 、ξ d Is a parameter that determines the notch width and depth of the filter;
under the condition of low rudder effect separation (i.e. t < t) 1 Time) increase xi d The method increases the filtering depth and width of the filter, and specifically comprises the following steps:
in which ξ 1 ＞ξ 2 。
2. The stability control method suitable for the low rudder effect separation state according to claim 1, wherein in the second step, amplitude limiting protection is added to the pitch channel control parameter and the yaw channel control parameter, specifically:
km’＝km×Kb 1 ×Kb 1
in the formula: ksf is a damping loop control parameter, km' is an attitude control loop control parameter after amplitude limiting,
for controlling the amplitude limit of the parameter ksf, kb 1 The intermediate variable represents the ratio of the clipping value of ksf to ksf before clipping.
3. A stability control method suitable for a low rudder effect separation state according to claim 2, wherein in the second step, amplitude limiting protection is added to the control parameters of the rolling channel, specifically:
g 3 ’＝g 3 ×Kb 2 ×Kb 2
in the formula: g 2 Damping loop control parameter for rolling channel, g 3 ' are rolling channel attitude control loop control parameters after clipping,
to control the parameter g 2 Amplitude limit value of Kb 2 Is an intermediate variable, representing g 2 Amplitude limit value and pre-amplitude limit g 2 The ratio of (a) to (b).
4. A stability control method for a low rudder effect decoupling state according to claim 3, calculating a rudder deflection instruction according to the control parameters and the filter parameters obtained in the first step, the third step, driving the control surface to deflect, applying a control force and a control moment, and stabilizing the posture of the missile:
u dR ＝g 2 ×u gyR -g 3 ×u ΔR
u dP ＝(ksf×u gyP -km×u ΔP )×W(z)
u dY ＝(-ksf×u gyY -km×u ΔY )×W(z)
in the formula u dR 、u dP 、u dY The rudder deflection instructions u of a rolling channel, a pitching channel and a yawing channel ΔR 、u ΔP 、u ΔY Attitude deviation instructions u for the roll channel, pitch channel, and yaw channel, respectively gyR 、u gyP 、u gyY The output quantities of the rate gyros of the rolling channel, the pitching channel and the yawing channel are respectively, and W (z) is a transfer function after discretization of the pitch-yaw channel filter.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202011156237.0A CN112327908B (en) | 2020-10-26 | 2020-10-26 | Stable control method suitable for low rudder effect separation state |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202011156237.0A CN112327908B (en) | 2020-10-26 | 2020-10-26 | Stable control method suitable for low rudder effect separation state |
Publications (2)
| Publication Number | Publication Date |
|---|---|
| CN112327908A CN112327908A (en) | 2021-02-05 |
| CN112327908B true CN112327908B (en) | 2023-01-17 |
Family
ID=74311899
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CN202011156237.0A Active CN112327908B (en) | 2020-10-26 | 2020-10-26 | Stable control method suitable for low rudder effect separation state |
Country Status (1)
| Country | Link |
|---|---|
| CN (1) | CN112327908B (en) |
Families Citing this family (1)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN115235444B (en) * | 2022-07-19 | 2023-02-24 | 青岛哈尔滨工程大学创新发展中心 | Method for measuring control loop bandwidth of full-angle hemispherical resonator gyroscope |
Family Cites Families (5)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN106406096B (en) * | 2016-10-26 | 2019-04-26 | 北京航空航天大学 | A kind of coupling suitable for the horizontal sidestep maneuver of aircraft utilizes control method for coordinating |
| CN107167822B (en) * | 2017-07-13 | 2019-12-10 | 北京理工大学 | method for simulating satellite navigation positioning of four-axis aircraft |
| RU2662576C1 (en) * | 2017-09-11 | 2018-07-26 | Акционерное общество "Российская самолетостроительная корпорация "МиГ" (АО "РСК "МиГ") | Aircraft side movement at landing approach automatic control system |
| CN107607112A (en) * | 2017-09-13 | 2018-01-19 | 哈尔滨工业大学 | Aircraft inexpensive pose measuring apparatus and measuring method |
| CN110411289B (en) * | 2019-06-13 | 2021-10-15 | 上海航天控制技术研究所 | Separation stability control method for inhibiting strong missile interference |
-
2020
- 2020-10-26 CN CN202011156237.0A patent/CN112327908B/en active Active
Non-Patent Citations (3)
| Title |
|---|
| 旋转导弹舵指令限幅方法研究;楼朝飞;《现代防御技术》;20171231;全文 * |
| 考虑舵机动力学的旋转导弹指令限幅方法研究;李想;《现代防御技术》;20190430;全文 * |
| 舵偏角限幅及角速率约束情况下高超声速飞行器的模糊小波神经网络抗饱和控制;梁捷;《MDAD2017》;20171231;全文 * |
Also Published As
| Publication number | Publication date |
|---|---|
| CN112327908A (en) | 2021-02-05 |
Similar Documents
| Publication | Publication Date | Title |
|---|---|---|
| CN107844123B (en) | Nonlinear aircraft track control method | |
| CN104155988B (en) | Multichannel attitude controller of aircraft | |
| EP0742142B1 (en) | Method and apparatus for reducing unwanted sideways motion in the aft cabin and roll-yaw upsets of an airplane due to atmospheric turbulence and wind gusts | |
| CN102707624B (en) | Design method of longitudinal controller region based on conventional aircraft model | |
| CN110411289B (en) | Separation stability control method for inhibiting strong missile interference | |
| CN105159308B (en) | A kind of Reusable launch vehicles landing phase guides coupling design method integrated with control law | |
| CN111123967A (en) | Fixed-wing unmanned aerial vehicle carrier landing control method based on adaptive dynamic inversion | |
| CN112327908B (en) | Stable control method suitable for low rudder effect separation state | |
| CN107977009A (en) | A kind of airbreather attitude control law design method for considering coupling | |
| CN106970633B (en) | Inhibit the flight control method of control input saturation | |
| CN110007683A (en) | A kind of control method of the anti-cross wind landing of low aspect ratio all-wing aircraft unmanned plane | |
| US5102072A (en) | Adaptive gain and phase controller for autopilot for a hypersonic vehicle | |
| CN109460055B (en) | Aircraft control capability determining method and device and electronic equipment | |
| CN112000127B (en) | Reverse-step-method-based aircraft lateral combined control method | |
| CN110609567A (en) | Satellite inertia combined navigation terminal deception method for quad-rotor unmanned aerial vehicle | |
| CN104656659B (en) | Shipboard aircraft ski-jump take-off automatic flight control method | |
| CN116045744A (en) | Control method and device for solid carrier rocket separator remains falling area | |
| CN108437980A (en) | A kind of Vehicular yaw stable control method adaptive based on saturation | |
| CN109992003A (en) | Robustness roll angle method of rate control and system | |
| WO2023240862A1 (en) | Rocket-boosted launch and takeoff control method for unmanned aerial vehicle having flying-wing layout | |
| CN108828941B (en) | Separation control method based on parameter identification | |
| CN115657458A (en) | Aircraft climbing track control method based on energy matching | |
| CN111273678B (en) | Boundary protection method for large lift-drag ratio unmanned aerial vehicle | |
| CN111949043B (en) | On-line extraction method for start control time based on attitude angular speed discrimination | |
| CN110426955B (en) | Hypersonic control surface manipulation efficiency prediction method based on coupling utilization |
Legal Events
| Date | Code | Title | Description |
|---|---|---|---|
| PB01 | Publication | ||
| PB01 | Publication | ||
| SE01 | Entry into force of request for substantive examination | ||
| SE01 | Entry into force of request for substantive examination | ||
| GR01 | Patent grant | ||
| GR01 | Patent grant |