CN109656262A - A kind of control method improving baby plane landing mission comfort of passenger - Google Patents
A kind of control method improving baby plane landing mission comfort of passenger Download PDFInfo
- Publication number
- CN109656262A CN109656262A CN201910045668.0A CN201910045668A CN109656262A CN 109656262 A CN109656262 A CN 109656262A CN 201910045668 A CN201910045668 A CN 201910045668A CN 109656262 A CN109656262 A CN 109656262A
- Authority
- CN
- China
- Prior art keywords
- comfort
- passenger
- state
- speed
- angle
- 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.)
- Granted
Links
- 238000000034 method Methods 0.000 title claims abstract description 41
- 238000010276 construction Methods 0.000 claims abstract description 21
- 238000013461 design Methods 0.000 claims abstract description 19
- 230000001276 controlling effect Effects 0.000 claims description 14
- 230000001105 regulatory effect Effects 0.000 claims description 11
- 239000011159 matrix material Substances 0.000 claims description 9
- 230000014509 gene expression Effects 0.000 claims description 8
- 238000004088 simulation Methods 0.000 claims description 8
- 238000005096 rolling process Methods 0.000 claims description 7
- 208000002173 dizziness Diseases 0.000 claims description 6
- 238000012360 testing method Methods 0.000 claims description 3
- 230000008569 process Effects 0.000 abstract description 11
- 238000005312 nonlinear dynamic Methods 0.000 abstract description 4
- 238000004364 calculation method Methods 0.000 abstract description 2
- 238000010001 crabbing Methods 0.000 description 7
- 238000010586 diagram Methods 0.000 description 4
- 238000006073 displacement reaction Methods 0.000 description 4
- 230000001629 suppression Effects 0.000 description 4
- 230000008878 coupling Effects 0.000 description 3
- 238000010168 coupling process Methods 0.000 description 3
- 238000005859 coupling reaction Methods 0.000 description 3
- 230000004044 response Effects 0.000 description 3
- 230000009466 transformation Effects 0.000 description 3
- 230000000694 effects Effects 0.000 description 2
- 239000004576 sand Substances 0.000 description 2
- RZVHIXYEVGDQDX-UHFFFAOYSA-N 9,10-anthraquinone Chemical compound C1=CC=C2C(=O)C3=CC=CC=C3C(=O)C2=C1 RZVHIXYEVGDQDX-UHFFFAOYSA-N 0.000 description 1
- 241000208340 Araliaceae Species 0.000 description 1
- 235000005035 Panax pseudoginseng ssp. pseudoginseng Nutrition 0.000 description 1
- 235000003140 Panax quinquefolius Nutrition 0.000 description 1
- 230000009286 beneficial effect Effects 0.000 description 1
- 230000008901 benefit Effects 0.000 description 1
- 230000008859 change Effects 0.000 description 1
- 238000013499 data model Methods 0.000 description 1
- 230000007423 decrease Effects 0.000 description 1
- 238000011161 development Methods 0.000 description 1
- 235000008434 ginseng Nutrition 0.000 description 1
- 230000002401 inhibitory effect Effects 0.000 description 1
- 238000013507 mapping Methods 0.000 description 1
- 238000013178 mathematical model Methods 0.000 description 1
- 230000007246 mechanism Effects 0.000 description 1
- 230000009467 reduction Effects 0.000 description 1
- 238000011160 research Methods 0.000 description 1
- 238000010998 test method Methods 0.000 description 1
- 230000003313 weakening effect Effects 0.000 description 1
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
- G05D1/08—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
- G05D1/0816—Control of attitude, i.e. control of roll, pitch, or yaw specially adapted for aircraft to ensure stability
- G05D1/0825—Control of attitude, i.e. control of roll, pitch, or yaw specially adapted for aircraft to ensure stability using mathematical models
-
- 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
Landscapes
- Engineering & Computer Science (AREA)
- General Physics & Mathematics (AREA)
- Physics & Mathematics (AREA)
- Automation & Control Theory (AREA)
- Remote Sensing (AREA)
- Radar, Positioning & Navigation (AREA)
- Aviation & Aerospace Engineering (AREA)
- Mathematical Analysis (AREA)
- Pure & Applied Mathematics (AREA)
- Mathematical Physics (AREA)
- Mathematical Optimization (AREA)
- Algebra (AREA)
- Feedback Control In General (AREA)
- Control Of Position, Course, Altitude, Or Attitude Of Moving Bodies (AREA)
Abstract
The present invention is to provide a kind of control methods for improving baby plane landing mission comfort of passenger.Step 1: obtaining baby plane state of flight amount and control amount；Step 2: estimating rough air quantity of state；Step 3: calculating comfort of passenger and determine compensation rate；Step 4: construction location and speed compensation control device；Step 5: construction carriage angle compensating controller；Step 6: construction aspect angular speed repays controller；Step 7: determining each controller parameter.The present invention is directed to baby plane landing mission, the present invention proposes to improve the control method of comfort of passenger, design disturbance observer estimates rough air velocity component, it is proposed the calculation method of comfort of passenger, construct comfort of passenger compensation model, in conjunction with the transverse and longitudinal comprehensive design thought based on nonlinear dynamic inverse, aircraft landing process status deviation is effectively controlled, realizes the purpose for improving baby plane landing mission comfort of passenger.
Description
Technical field
The present invention relates to a kind of aircraft autocontrol method, specifically a kind of control of baby plane landing mission
Method processed.
Background technique
With the development of national economy, the quantity of baby plane is more and more, will be used wider and wider, baby plane body
Product is small, and construction is relatively easy, and all many-sided by favor in leisure and travelling etc., passenger wants the comfort level for taking baby plane
Ask higher.But baby plane is since its quality is small, can passenger loading it is less, therefore be affected by rough air, especially
In baby plane landing mission, violent flow perturbation, causes aircraft vertical and horizontal lateral velocity variations easily cause and multiply
The even airsick situation of objective dizziness, meanwhile, in subaerial landing mission, aircraft state control accuracy decline causes aircraft
Aspect angular speed variation, also be easy to cause passenger's dizziness situation, seriously affect the comfort level of passenger above.For this purpose,
For baby plane landing mission, the influence and reduction flight status deviation of effective disturbance suppression air-flow are improving passenger
In terms of comfort level, there is important theoretical significance and practical application meaning.
It is driven by the end of currently, baby plane still uses pilot's manual control mode to complete landing operation, passing through manipulation
The state of flight of bar and throttle lever control aircraft, piloting strategies rely primarily on the experience of pilot.But manual control exists such as
Lower disadvantage: firstly, pilot manipulation can not timely and effectively inhibit the influence of uncertain rough air；Secondly, pilot is
Accurately control aircraft, often by it is longitudinal and it is horizontal be laterally separated control, this has ignored the longitudinal direction of aircraft and horizontal lateral state coupling
It closes, necessarily will affect the state in another direction to the control in one of direction.Therefore, it is influenced in landing mission rough air
Under, using pilot's manual control mode, it is difficult to the comfort level of the accurate control and raising passenger of aircraft.
Current related literatures often assumes that air-flow is disturbed when the rough air of research aircraft flight course influences
Momentum is known, but in real aircraft landing mission, rough air state is not directly measured, can be in aircraft mould
Design point observer on the basis of type, estimates rough air, and then analyzes the velocity component of rough air.In addition,
Aircraft is difficult to realize accurately control, and can be adopted due to containing a large amount of transverse and longitudinal nonlinear coupling characteristics in flight control system
With the mode of nonlinear dynamic inverse, nonlinear influencing factors therein are offset in aircraft portrait and horizontal lateral circuit.
Summary of the invention
The purpose of the present invention is to provide one kind can effectively control aircraft landing process status deviation, improves baby plane
The control method of the raising baby plane landing mission comfort of passenger of landing mission comfort of passenger.
The object of the present invention is achieved like this:
Step 1: obtaining baby plane state of flight amount and control amount；
Step 2: estimating rough air quantity of state；
Step 3: calculating comfort of passenger and determine compensation rate；
Step 4: construction location and speed compensation control device；
Step 5: construction carriage angle compensating controller；
Step 6: construction aspect angular speed repays controller；
Step 7: determining each controller parameter.
The present invention may also include:
1. the calculating formula of state of flight amount are as follows:
Wherein, x is aircraft state amount, and u is aircraft control amount, and f, g, h, k are the non-linear expressions of quantity of state, by aircraft
Quantity of state and aerodynamic parameter are constituted；
The control amount is the angle of rudder reflection and accelerator open degree that baby plane executing agency is obtained by sensor, the execution
Mechanism includes aileron, elevator, rudder and throttle.
It is estimated described in 2. in rough air quantity of state, baby plane landing reference model is by non-linear landing model in the phase
Hope that Taylor series expansion obtains at state point, landing reference model are as follows:
Wherein, A, B, C, D are that linear equation state matrix will be non-thread according to the quantity of state and control amount exported in step 1
Property landing model and reference model are configured to disturbance state observer, the rough air quantity of state of output estimation.
3. the calculating comfort of passenger simultaneously determines that compensation rate specifically includes: baby plane comfort of passenger model is established,
And fuzzy control theory is utilized, the non-linear input and output for establishing speed, attitude angle, attitude angular velocity and comfort of passenger are closed
System, comfort level function FcIt indicates are as follows:
Fc=ξ (Vxga,Vyga,Vzga,φ,θ,p,q)
Wherein, Vxga、VygaAnd VzgaFor velocity component of the rough air under earth axes, φ is roll angle, and θ is to bow
The elevation angle, p are angular velocity in roll, and q is rate of pitch, and ξ indicates to influence the non-linear table of the flight status of comfort of passenger
Up to formula.
4. the construction location and speed compensation control implement body include: to increase speed in the controlling unit of aircraft position
Independent control link, for maintaining aircraft to complete landing task with constant speed；Comfort level is introduced in speed compensation control device
The velocity compensation variable of model output, desin speed dynamic inverse introduce the output of comfort level model in the compensating controller of position
Velocity compensation variable, design position dynamic inverse inhibit nonlinear influencing factors and rough air using compensation variable and dynamic inverse
To the disturbance velocity of comfort of passenger, desired attitude angle and accelerator open degree are exported according to aircraft position and speed.
5. construction carriage angle compensating controller described in specifically includes: introducing comfort level in attitude angle controlling unit
The attitude angle compensation variable of model output designs roll angle and pitch angle dynamic inverse, and introducing sideslip angular motion state is inverse, according to attitude angle
Export desired attitude angular velocity.
6. construction aspect angular speed described in is repaid controller and is specifically included: introducing in attitude angular velocity controlling unit
The attitude angular velocity compensation variable of comfort level model output, design rolling rate, rate of pitch and yaw rate dynamic
It is inverse, inhibit the phenomenon that aircraft speed-changing rotation leads to passenger's dizziness, desired rudder face control amount is exported according to attitude angular velocity.
7. each controller parameter of determination described in specifically includes: according to attitude angular velocity compensating controller, posture angle compensation
The sequence of controller, position and speed compensating controller according to regulating time, is surpassed from time domain angle using l-G simulation test data
The index of tune amount and steady-state error establishes the pid parameter of each link one by one, forms the controller of parameter and structural integrity.
The present invention is directed to landing mission of the small-sized fixed wing aircraft from rough air area to airfield landing area, proposes one kind
Improve comfort of passenger automatic Landing control method, this method can make aircraft manipulate automatically each rudder face (aileron, elevator,
Rudder) and throttle to complete automatic Landing task, comfort level compensating module and be introduced into control process by building, effectively
Disturbance suppression air-flow etc. influences the factor of comfort of passenger, realizes the accurate control to aircraft state, improves the comfort level of passenger.
On the one hand, each controlling unit of this patent introduces the dynamic inverse that longitudinal state and horizontal lateral state are constituted, and fully considers two
The coupling condition in a direction；On the other hand, the comfort level compensating module that this patent proposes can estimate rough air state, count simultaneously
Rough air speed, the compensation rate of attitude angle and attitude angular velocity are calculated, is used in each controlling unit of aircraft, keeps aircraft logical
Cross body movement caused by rudder face control disturbance suppression air-flow, passenger itself speed and angular speed caused by weakening due to disturbance
Variation, improves the comfort level of passenger, especially causes passenger's dizziness and airsick degree to rough air.
Beneficial effects of the present invention: being directed to baby plane landing mission, and the present invention proposes to improve the control of comfort of passenger
Method, design disturbance observer estimate rough air velocity component, propose the calculation method of comfort of passenger, construct occupant comfort
Compensation model is spent, in conjunction with the transverse and longitudinal comprehensive design thought based on nonlinear dynamic inverse, effectively controls aircraft landing process status
Deviation realizes the purpose for improving baby plane landing mission comfort of passenger.
Detailed description of the invention
Fig. 1 is control flow chart of the invention.
Fig. 2 is rough air state observer schematic diagram.
Fig. 3 is aircraft fixed speed control schematic diagram.
Fig. 4 is aircraft position control principle drawing.
Fig. 5 is carriage angle control principle drawing.
Fig. 6 is carriage angle speed control schematic diagram.
Fig. 7 is step signal response curve schematic diagram.
Fig. 8 is air speed simulation comparison curve.
Fig. 9 is rate of pitch simulation comparison curve.
Figure 10 is roll angle simulation comparison curve.
Figure 11 is lateral position deviation simulation comparison curve.
Specific embodiment
It illustrates below and the present invention is described in more detail.
In conjunction with Fig. 1, the first embodiment of the invention includes the following steps:
Step 1 obtains flight status and control amount.
The quantity of state of aircraft can be used following formula and be calculated:
In above formula, x is aircraft state amount, and u is aircraft control amount, and f, g, h, k are the non-linear expressions of quantity of state, the table
It is made of up to formula aircraft state amount and aerodynamic parameter.Aircraft executing agency includes aileron, elevator, rudder and throttle, control
Amount obtains the angle of rudder reflection and accelerator open degree of the above executing agency by sensor.
Step 2 estimates rough air quantity of state.
Construct the model of aircraft landing reference model, model Taylor at expectation state point by non-linear landing model
Series expansion obtains, and model is as follows:
In above formula, A, B, C, D are linear equation state matrix.It, will be non-according to the quantity of state and control amount exported in step 1
Linear landing model and reference model are configured to disturbance state observer, the rough air quantity of state of output estimation.
Step 3 calculates comfort of passenger and determines compensation rate.
Baby plane comfort of passenger model is established, and utilizes fuzzy control theory, establishes speed, attitude angle, attitude angle
The non-linear input/output relation of speed and comfort of passenger, comfort level function FcIt is expressed as following formula:
Fc=ξ (Vxga,Vyga,Vzga,φ,θ,p,q) (3)
In above formula, Vxga、VygaAnd VzgaFor velocity component of the rough air under earth axes, φ is roll angle, θ is
Pitch angle, p is angular velocity in roll, q is rate of pitch, and ξ indicates to influence the non-linear of the flight status of comfort of passenger
Expression formula, wherein VyAnd VzIt is related to the predicted state value of rough air.Establish the Linear Mapping of comfort level and flight status
Relationship, and determine comfort of passenger compensating coefficient amount, which is exported, is used for subsequent control link.
Step 4, construction location and speed compensation control device.
The independent control link increased speed in traditional aircraft position controlling unit, for maintaining aircraft with constant speed
Degree completes landing task.The velocity compensation variable of comfort level model output is introduced in speed compensation control device, desin speed is dynamic
State is inverse, and the velocity compensation variable of comfort level model output is introduced in the compensating controller of position, and design position dynamic inverse utilizes benefit
Variable and dynamic inverse are repaid, inhibits nonlinear influencing factors and rough air to the disturbance velocity of comfort of passenger, according to aircraft position
It sets and exports desired attitude angle and accelerator open degree with speed.
Step 5, construction carriage angle compensating controller.
The attitude angle compensation variable that the output of comfort level model is introduced in attitude angle controlling unit, designs roll angle and pitching
Angle dynamic inverse, novelty introducing sideslip angular motion state is inverse, exports desired attitude angular velocity according to attitude angle.
Step 6, construction aspect angular speed repay controller.
The attitude angular velocity compensation variable that the output of comfort level model is introduced in attitude angular velocity controlling unit, designs rolling
Angular speed, rate of pitch and yaw rate dynamic inverse inhibit the phenomenon that aircraft speed-changing rotation leads to passenger's dizziness, according to appearance
State angular speed exports desired rudder face control amount.
Step 7 determines each controller parameter.
According to the sequence of attitude angular velocity compensating controller, attitude angle compensating controller, position and speed compensating controller,
From time domain angle, established one by one using l-G simulation test data according to the index of regulating time, overshoot and steady-state error
The pid parameter of each link forms the controller of parameter and structural integrity.
Second of embodiment of the invention specifically includes:
1, flight status and control amount are obtained
Baby plane executing agency of the present invention includes aileron, rudder, elevator and throttle, airplane motion equation
It is indicated with following formula are as follows:
In above formula, x is the state of flight of aircraft, including aircraft flight speed V, angle of attack α, yaw angle β, roll angle φ, bows
Elevation angle theta, yaw angle ψ, angular velocity in roll p, rate of pitch q, yaw rate r.U is the control amount of aircraft, including aileron, side
To rudder, elevator angle of rudder reflection and accelerator open degree.Control amount can directly obtain real value by sensor corresponding on aircraft, pass through
The nonlinear model can accurately calculate the state value of aircraft.
2, rough air quantity of state is estimated
The present invention proposes flying speed V, angle of attack, pitching angle theta, yaw angle ψ when aircraft is kept constant, while yaw angle
When β, roll angle φ, angular velocity in roll p, rate of pitch q, yaw rate r are maintained near 0 value, the comfort level of passenger is most
Good, the expectation state of flight of baby plane is as shown in table 1 below:
1 aircraft landing process expectation state of table
| Baby plane state of flight | Desired value |
| Flying speed V | Steady state value |
| Angle of attack α | Steady state value |
| Pitching angle theta | Steady state value |
| Yaw angle ψ | Steady state value |
| Yaw angle β | 0 |
| Roll angle φ | 0 |
| Angular velocity in roll p | 0 |
| Rate of pitch q | 0 |
| Yaw rate r | 0 |
Constant flying speed and the angle of attack can make baby plane maintain certain kinetic energy, and can make from physics angle
Personnel's uniform force is seized the opportunity, comfort is high.Constant pitch angle and yaw angle can make baby plane maintain certain posture, be convenient for
Control the transverse and longitudinal position of aircraft.Remaining state keeps 0 value as far as possible, guarantees that aircraft is smaller along three direction rotations, improves passenger
Comfort level, can especially weaken the airsick phenomenon of passenger.
Baby plane keeps it is expected state of flight shown in upper table in landing mission, according to data shown in upper table, by aircraft
Landing nonlinear model carries out Taylor series expansion, and linear mathematical model can be obtained, specific as follows shown:
In above formula, A, B, C, D are the state matrix of linear data model, reference mould of the present invention as baby plane
Type will estimate in real time the velocity component of rough air using the reference model.
3, it calculates comfort of passenger and determines compensation rate
The factor for influencing comfort of passenger includes: rough air velocity component and aircraft roll angle, pitch angle, roll angle
Speed and rate of pitch are effective control of effective control disturbance air-flow, need to estimate the velocity component of rough air, this hair
Bright design rough air state observer, to predict the real-time status of rough air, the principle of rough air state observer
As shown in Figure 2.
In Fig. 2, the above nonlinear model and reference model are inputted into initial flight quantity of state and initial control amount respectively, point
Not Huo Qu reality output state and prediction output state, using coordinate converter by both the above state by aircraft body coordinate system
Be converted into earth axes, and can calculated gas flow in the velocity component of earth axes, coordinate converter is using following several
Transformation matrix and formula are realized:
By the transformation matrix T of earth axes to aircraft body coordinate systembgAre as follows:
By the transformation matrix T of aircraft body coordinate system to air path axis systemabAre as follows:
Landing speed component V of the aircraft in earth axesxg,Vygand VzgIt is represented by following formula:
According to above formula, earth axes it is longitudinal, horizontal it is lateral and vertical on, can get aircraft virtual condition and prediction shape
Corresponding velocity component [the V of statexgr,Vygr,Vzgr] and [Vxgf,Vygf,Vzgf], rough air is longitudinal, horizontal lateral in earth axes
With it is vertical on velocity component Vxga、VygaAnd VzgaIt is represented by following formula:
The present invention uses design of fuzzy control comfort level calculator, and calculator input is rough air velocity component, posture
Angular displacement and attitude angular velocity deviation, the description of input are divided into 7 classes: it is honest, in just, it is just small, zero, bear in small, negative, be negative big, it is defeated
The vague description entered are as follows: PB, PM, PS, ZE, NS, NM, NB.Output is occupant comfort angle value, and the description of output falls into 5 types: very
It is uncomfortable, more uncomfortable, without it is obvious feel, relatively comfortable, as snug as a bug in a rug, the vague description of output are as follows: PB, PM, ZE, NM, NB.It examines
Consider comfort level and its influence factor is complicated non-linear relation, comfort level is mapped as by the present invention by the way of linearisation
The linear relationship of its influence factor, comfort level function FcIt is expressed as following formula:
Coefficient matrix is penalty coefficient in above formula, which is determined by comfort level, vague description PB, PM of comfort level, ZE,
NM, NB respectively correspond kcIt is 0.4,0.3,0.2,0.1,0.The coefficient and corresponding comfort degree factor product as compensation rate,
It is introduced in subsequent controller design link.
4, construction location and speed compensation control device
The present invention constructs speed dynamic inverse according to longitudinal horizontal crabbing state, and introduces the disturbance of comfort level module for compensating
Air velocity compensation rate calculates desired Throttle Opening Control amount T, and Throttle Opening Control rule is as follows:
In above formula, RVTo assist regulated variable, the variable can be calculated by PID controller.Aircraft landing process flight speed
It is as shown in Figure 3 to spend control principle.
In Fig. 3, the present invention designs landing speed dynamic inverse, shown in form:
The control signal delta V of PID controller 1 is desired speed V in Fig. 3cmd, real-time speed V, rough air close speed Vf
Difference, be shown below:
Δ V=Vcmd-V-Vf
The relevant parameter of PID controller 1 can determine that parameter chooses principle will according to the individually designed adjusting of different model aircraft
It is described below.The Throttle Opening Control amount of the aircraft of control process final output shown in Fig. 3.
In the controlling unit of aircraft position, the present invention calculates expectation pitch command according to lengthwise position, according to horizontal lateral
Position calculates expectation yaw angle instruction, then calculates expectation roll angle instruction by yaw angle deviation again, that is, passes through manipulation aircraft
The control to the horizontal lateral position of aircraft is realized in yaw and rolling.Lengthwise position is constructed using longitudinal and horizontal crabbing state
Dynamic inverse, lateral position dynamic inverse and yaw angle dynamic inverse, and controller is designed, aircraft rolling is calculated according to transverse and longitudinal position deviation
The desired control amount of corner, pitch angle and yaw angle, desired attitude angle control amount expression formula are shown below:
In above formula, Rψ、To assist regulated variable, it will be controlled by PID controller and calculate each auxiliary adjustment change
Amount.Control principle such as Fig. 4 in aircraft position shows.
In Fig. 4, do not consider that auxiliary adjustment becomes in yaw angle dynamic inverse, lateral position dynamic inverse and lengthwise position dynamic inverse
The part of amount is expressed, and for each dynamic inverse by longitudinal and horizontal crabbing state integrating representation, rough air is vertical and lateral
Velocity compensation amount introduces position control link, auxiliary variable R as position compensation amount via integratorψ、According to
The deviation for yawing angular displacement and transverse and longitudinal position and position real value and offset utilizes PID controller 2, controller 3, control
Device 4 processed, which calculates, to be obtained, and the relevant parameter of these three PID controllers can be determined according to the individually designed adjusting of different model aircraft, ginseng
Number is chosen principle and will be introduced later.The desired roll angle of control process final output and pitch angle described in Fig. 4.
5, carriage angle compensating controller is constructed
The present invention proposes to be determined as rolling by control object to the control thought for introducing yaw angle in attitude angle control process
Corner, pitch angle and yaw angle.Using longitudinal and horizontal crabbing state, attitude angle dynamic inverse, the phase of corresponding angular speed are constructed
Control amount expression formula is hoped to be shown below:
In above formula, Rφ、Rθ、RβTo assist regulated variable, each auxiliary adjustment variable will be calculated by PID controller.Control is former
Reason is as shown in Figure 5.
In Fig. 5, attitude angle dynamic inverse is by longitudinal and horizontal crabbing state integrating representation, as shown in above formula.Auxiliary adjustment
Variable Rφ、Rθ、RβAccording to three corresponding posture angular displacement, is calculated and obtained using PID controller 5, controller 6, controller 7.This
The relevant parameter of three PID controllers can determine that parameter chooses principle will be rear according to the individually designed adjusting of different model aircraft
Text is introduced.The above-mentioned desired attitude angle control amount of control process final output.
6, construction aspect angular speed repays controller
The present invention is compensated using the attitude angular velocity that longitudinal and horizontal crabbing state and comfort level compensating module calculate
Component constructs attitude angular velocity dynamic inverse, and designs controller, aileron δa, rudder δr, elevator δsDesired control amount,
It is shown below:
In above formula, matrix M and N are the mathematic(al) representation of aircraft air power and posture torque and control rudder face, can be passed through
Reversed operation can derive acquisition, not repeat here.Rp、Rq、RrTo assist regulated variable, will be calculated by PID controller each auxiliary
Help regulated variable.Aileron, rudder, the control principle of elevator are as shown in Figure 6.
In Fig. 6, attitude angular velocity dynamic inverse expression formula, the dynamic inverse by longitudinal and horizontal crabbing state integrating representation,
Three angular rate compensation amounts such as angular velocity in roll are provided by comfort level compensating module in Fig. 6.Auxiliary adjustment variable Rp、Rq、RrAccording to
The deviation of three attitude angular velocity desired values and actual value and offset is counted using PID controller 8, controller 9, controller 10
It calculates and obtains.The relevant parameter of these three PID controllers can determine that parameter is chosen former according to the individually designed adjusting of different model aircraft
Reason is described below.The control amount of above-mentioned control process final output aileron, rudder, elevator.
7, each controller parameter is determined
10 PID controllers are shared in controlling unit of the invention, since different aircrafts have different aerodynamic parameters and knot
Structure constant etc., the present invention propose the principle that controller parameter is formulated using test method(s): in each controlling unit, by inputting rank
Jump signal, judges regulating time, the overshoot, steady-state error of the above links response curve, meets certain condition it, with
This formulates the parameter of the above corresponding PID controller, and control system is made to have preferable dynamic characteristic and steady-state characteristic.Each link pair
The response curve form of step signal is as shown in Figure 7.
In Fig. 7, TsFor regulating time, δsFor overshoot, esFor steady-state error.TsAnd δsIt can reflect that the dynamic of system is special
Property, esIt can indicate steady-state characteristic, the control effect of system controller parameter can be reflected by these variables.For each control ring
Section, the present invention is to input step signal initial value and final value and Ts、δsAnd esBoundary require setting rule as shown in table 2 below:
2 time domain index boundary rule list of table
Verify example and explanation
For above-mentioned theory, the present invention carries out simulating, verifying work, by taking a baby plane as an example, desired landing speed
For 70m/s, initial lateral deviation is that 5m chooses do not consider Nonlinear Dynamic here for the control performance for comparing the method for the present invention
Conventional PID control method method as a comparison that is inverse and inhibiting flow-disturbing function, simulation comparison result as illustrated in figs. 8-11, wherein this
It invents the method proposed and control methods is marked with " this patent method " and " control methods " respectively.
In figures 8-11, first in the longitudinal direction, this patent method can faster release rate deviation, make aircraft faster
The desired speed for maintaining 70m/s on, compare commonsense method, can better disturbance suppression air-flow to a certain extent；This is specially
The rate of pitch of sharp method is restored to 0 value of expectation more quickly by way of increasing overshoot, provides for system stronger
Increase steady effect；Secondly horizontally, in face of opposite rolling angular displacement, the roll angle regulating time of this patent is shorter, overshoot
It is smaller, while lateral position deviation can be faster eliminated, it is more more superior than commonsense method.
To sum up, the present invention can quickly eliminate the state of flight deviation of landing transverse and longitudinal, while pass through comfort level compensating module
The influence of rough air and attitude angular velocity to birdman's comfort level can be effectively suppressed, improve baby plane landing mission and multiply
The comfort level of visitor.
Claims (8)
1. a kind of control method for improving baby plane landing mission comfort of passenger, it is characterized in that:
Step 1: obtaining baby plane state of flight amount and control amount；
Step 2: estimating rough air quantity of state；
Step 3: calculating comfort of passenger and determine compensation rate；
Step 4: construction location and speed compensation control device；
Step 5: construction carriage angle compensating controller；
Step 6: construction aspect angular speed repays controller；
Step 7: determining each controller parameter.
2. the control method according to claim 1 for improving baby plane landing mission comfort of passenger, it is characterized in that flying
The calculating formula of row quantity of state are as follows:
Wherein, x is aircraft state amount, and u is aircraft control amount, and f, g, h, k are the non-linear expressions of quantity of state, by aircraft state
Amount and aerodynamic parameter are constituted；
The control amount is the angle of rudder reflection and accelerator open degree that baby plane executing agency is obtained by sensor.
3. the control method according to claim 1 for improving baby plane landing mission comfort of passenger, it is characterized in that: institute
It states and estimates in rough air quantity of state, baby plane landing reference model is safe at expectation state point by non-linear landing model
It strangles series expansion to obtain, landing reference model are as follows:
Wherein, A, B, C, D are linear equation state matrix, will be non-linear according to the quantity of state and control amount exported in step 1
Land model and reference model are configured to disturbance state observer, the rough air quantity of state of output estimation.
4. the control method according to claim 1 for improving baby plane landing mission comfort of passenger, it is characterized in that institute
It states and calculates comfort of passenger and determine that compensation rate specifically includes: establishing baby plane comfort of passenger model, and utilize Fuzzy Control
System is theoretical, establishes the non-linear input/output relation of speed, attitude angle, attitude angular velocity and comfort of passenger, comfort level function
FcIt indicates are as follows:
Fc=ξ (Vxga,Vyga,Vzga,φ,θ,p,q)
Wherein, Vxga、VygaAnd VzgaFor velocity component of the rough air under earth axes, φ is roll angle, and θ is pitch angle,
P is angular velocity in roll, and q is rate of pitch, and ξ indicates to influence the non-linear expressions of the flight status of comfort of passenger.
5. the control method according to claim 1 for improving baby plane landing mission comfort of passenger, it is characterized in that institute
It states construction location and speed compensation control implement body includes: the independent control ring increased speed in the controlling unit of aircraft position
Section, for maintaining aircraft to complete landing task with constant speed；The output of comfort level model is introduced in speed compensation control device
Velocity compensation variable, desin speed dynamic inverse, the velocity compensation that the output of comfort level model is introduced in the compensating controller of position become
Amount, design position dynamic inverse inhibit nonlinear influencing factors and rough air to occupant comfort using compensation variable and dynamic inverse
The disturbance velocity of degree exports desired attitude angle and accelerator open degree according to aircraft position and speed.
6. the control method according to claim 1 for improving baby plane landing mission comfort of passenger, it is characterized in that institute
The construction carriage angle compensating controller stated specifically includes: the appearance of comfort level model output is introduced in attitude angle controlling unit
State angle compensation variable designs roll angle and pitch angle dynamic inverse, and introducing sideslip angular motion state is inverse, exports desired appearance according to attitude angle
State angular speed.
7. the control method according to claim 1 for improving baby plane landing mission comfort of passenger, it is characterized in that institute
The construction aspect angular speed stated is repaid controller and is specifically included: it is defeated that comfort level model is introduced in attitude angular velocity controlling unit
Attitude angular velocity compensation variable out, design rolling rate, rate of pitch and yaw rate dynamic inverse inhibit aircraft to become
The phenomenon that speed rotation leads to passenger's dizziness exports desired rudder face control amount according to attitude angular velocity.
8. the control method according to claim 1 for improving baby plane landing mission comfort of passenger, it is characterized in that institute
Each controller parameter of the determination stated specifically includes: according to attitude angular velocity compensating controller, attitude angle compensating controller, position and
The sequence of speed compensation control device is missed from time domain angle using l-G simulation test data according to regulating time, overshoot and stable state
The index of difference establishes the pid parameter of each link one by one, forms the controller of parameter and structural integrity.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN201910045668.0A CN109656262B (en) | 2019-01-17 | 2019-01-17 | Control method for improving passenger comfort level of small airplane in landing process |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN201910045668.0A CN109656262B (en) | 2019-01-17 | 2019-01-17 | Control method for improving passenger comfort level of small airplane in landing process |
Publications (2)
| Publication Number | Publication Date |
|---|---|
| CN109656262A true CN109656262A (en) | 2019-04-19 |
| CN109656262B CN109656262B (en) | 2021-10-01 |
Family
ID=66119828
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CN201910045668.0A Active CN109656262B (en) | 2019-01-17 | 2019-01-17 | Control method for improving passenger comfort level of small airplane in landing process |
Country Status (1)
| Country | Link |
|---|---|
| CN (1) | CN109656262B (en) |
Cited By (1)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN111392051A (en) * | 2019-11-14 | 2020-07-10 | 南京航空航天大学 | Self-adaptive landing deck control system and method for rotor type aircraft |
Citations (5)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US4131252A (en) * | 1976-06-15 | 1978-12-26 | The Boeing Company | Apparatus for compensating for airflow blockage in aircraft slotted flap assemblies |
| CN102393630A (en) * | 2011-09-26 | 2012-03-28 | 南京航空航天大学 | Carrier aircraft landing guide and control system for inhibiting airflow disturbance of stern and control method for system |
| CN103700286A (en) * | 2013-12-11 | 2014-04-02 | 南京航空航天大学 | Automatic carrier-landing guiding method of carrier-borne unmanned aircraft |
| CN105182984A (en) * | 2015-07-16 | 2015-12-23 | 孙明玮 | Linear active disturbance rejection control (ADRC) design and parameter tuning of aircraft pitch attitude |
| CN109062237A (en) * | 2018-09-17 | 2018-12-21 | 南京航空航天大学 | A kind of unmanned tiltrotor active disturbance rejection attitude control method |
-
2019
- 2019-01-17 CN CN201910045668.0A patent/CN109656262B/en active Active
Patent Citations (5)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US4131252A (en) * | 1976-06-15 | 1978-12-26 | The Boeing Company | Apparatus for compensating for airflow blockage in aircraft slotted flap assemblies |
| CN102393630A (en) * | 2011-09-26 | 2012-03-28 | 南京航空航天大学 | Carrier aircraft landing guide and control system for inhibiting airflow disturbance of stern and control method for system |
| CN103700286A (en) * | 2013-12-11 | 2014-04-02 | 南京航空航天大学 | Automatic carrier-landing guiding method of carrier-borne unmanned aircraft |
| CN105182984A (en) * | 2015-07-16 | 2015-12-23 | 孙明玮 | Linear active disturbance rejection control (ADRC) design and parameter tuning of aircraft pitch attitude |
| CN109062237A (en) * | 2018-09-17 | 2018-12-21 | 南京航空航天大学 | A kind of unmanned tiltrotor active disturbance rejection attitude control method |
Non-Patent Citations (5)
| Title |
|---|
| FERIT ÇAKICI，ETC: "Control System Design of a Vertical Take-off and Landing Fixed-Wing UAV", 《IFAC-PAPERSONLINE》 * |
| 张勇，等: "基于扩张状态观测器的舰载机俯仰自适应滑模控制", 《海军航空工程学院学报》 * |
| 王奇，等: "抗尾流干扰的自动着舰非线性控制研究", 《飞行力学》 * |
| 王日俊，等: "基于扰动观测器的多旋翼无人机机载云台模糊自适应跟踪控制", 《浙江大学学报(工学版)》 * |
| 齐浩然，等: "基于LADRC的四旋翼无人机自适应姿态跟踪控制", 《基于LADRC的四旋翼无人机自适应姿态跟踪控制》 * |
Cited By (2)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN111392051A (en) * | 2019-11-14 | 2020-07-10 | 南京航空航天大学 | Self-adaptive landing deck control system and method for rotor type aircraft |
| CN111392051B (en) * | 2019-11-14 | 2022-08-05 | 南京航空航天大学 | Self-adaptive landing deck control system and method for rotor type aircraft |
Also Published As
| Publication number | Publication date |
|---|---|
| CN109656262B (en) | 2021-10-01 |
Similar Documents
| Publication | Publication Date | Title |
|---|---|---|
| Zhang et al. | Adaptive constrained backstepping controller with prescribed performance methodology for carrier-based UAV | |
| CN110908281A (en) | Finite-time convergence reinforcement learning control method for attitude motion of unmanned helicopter | |
| Schumacher et al. | Adaptive control of UAVs in close-coupled formation flight | |
| CN108717265B (en) | Unmanned aerial vehicle cruise tracking control system and control method based on control variable parameterization | |
| CN106774361A (en) | A kind of aircraft carrier based on feedforward and feedback complex control the warship stern stream suppressing method of warship | |
| CN109703768B (en) | Soft air refueling docking method based on attitude/trajectory composite control | |
| CN109703769B (en) | Air refueling docking control method based on preview strategy | |
| CN103761899B (en) | Method for simulating strength of reversible control loading system of flight simulator | |
| Bao et al. | Design of a fixed-wing UAV controller based on adaptive backstepping sliding mode control method | |
| Mystkowski | 721. An application of mu-synthesis for control of a small air vehicle and simulation results | |
| CN114721266B (en) | Self-adaptive reconstruction control method under condition of structural failure of control surface of airplane | |
| Torabi et al. | Intelligent pitch controller identification and design | |
| CN113051662B (en) | Pneumatic modeling and performance evaluation method for folding wingtip variant aircraft based on CFD and DATCOM | |
| CN109656262A (en) | A kind of control method improving baby plane landing mission comfort of passenger | |
| CN116736723A (en) | Modeling and fuzzy self-adaptive sliding mode control method for elastic hypersonic aircraft under aerodynamic heat influence | |
| CN116088311A (en) | Unmanned helicopter second-order uncertain sliding mode control method based on adaptive neural network extended state observer | |
| CN116088549B (en) | Tailstock type vertical take-off and landing unmanned aerial vehicle attitude control method | |
| Snyder | Autopilot design with learn-to-fly | |
| Siddarth et al. | Global tracking control structures for nonlinear singularly perturbed aircraft systems | |
| Hecker et al. | Advanced gust load alleviation system for large flexible aircraft | |
| CN112506046B (en) | Stability augmentation control method for wingtip hinged combined type flight platform | |
| Larsson | System identification of flight mechanical characteristics | |
| CN114036628A (en) | Method for collaborative design of wingspan and control strategy of morphing aircraft | |
| Pollack et al. | Multi-objective Design and Performance Analysis of Incremental Control Allocation-based Flight Control Laws | |
| CN109376420A (en) | A kind of airplane nose down, which changes, falls high algorithm |
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 |