CN111459184A - Unmanned aerial vehicle automatic carrier landing control method adopting segmented attack angle instruction - Google Patents

Unmanned aerial vehicle automatic carrier landing control method adopting segmented attack angle instruction Download PDF

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CN111459184A
CN111459184A CN202010296385.6A CN202010296385A CN111459184A CN 111459184 A CN111459184 A CN 111459184A CN 202010296385 A CN202010296385 A CN 202010296385A CN 111459184 A CN111459184 A CN 111459184A
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signal
angle
pitch angle
attack
aircraft
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赵红超
张友安
施建洪
曲东才
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Yantai Nanshan University
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Yantai Nanshan University
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    • G—PHYSICS
    • G05—CONTROLLING; REGULATING
    • G05D—SYSTEMS FOR CONTROLLING OR REGULATING NON-ELECTRIC VARIABLES
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    • G05D1/10—Simultaneous control of position or course in three dimensions
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Abstract

The invention relates to an automatic landing control method of an unmanned aerial vehicle by adopting a segmented attack angle instruction, belonging to the field of automatic landing control of aircrafts. Firstly, a four-section type sectional attack angle instruction with a constant value and height related is designed, the four-section type sectional attack angle instruction is compared with an attack angle signal to generate an attack angle error signal, then an attack angle equivalent control item is superposed to generate an accelerator deflection angle control law of the aircraft, and the tracking of a given attack angle is realized. And then establishing a deck motion compensation model, designing a height gliding instruction, generating a pitch angle expected signal from the height signal, comparing the pitch angle expected signal with the pitch angle to generate a pitch angle error signal, generating an expected signal of a pitch angle rate and a pitch angle rate error signal, and superposing an elevator equivalent signal to realize the height control of automatic landing. The method has the advantages that the decoupling control of the accelerator deflection angle and the elevator can be realized, and the attack angle and the pitch angle can be accurately tracked, so that the dynamic effect of automatic carrier landing is better.

Description

Unmanned aerial vehicle automatic carrier landing control method adopting segmented attack angle instruction
Technical Field
The invention relates to the field of aircraft automatic control landing control, in particular to a method for realizing automatic landing control of an unmanned aircraft by adopting a segmented attack angle instruction.
Background
The landing process of the carrier-based aircraft and the unmanned aircraft is a very complex process, and as the speed of the aircraft is continuously reduced, the system stability margin caused by stall is reduced, and the control difficulty is increased. Meanwhile, due to the limited space of the deck of the naval vessel, the aircraft must safely and accurately complete the tail landing control task, and the aircraft height control system must have higher control precision and better anti-interference capability. Meanwhile, in order to ensure the stability of the aircraft, the thrust system must be continuously adjusted according to the landing process, so that the stability of the speed and the attack angle of the aircraft in the landing process of the aircraft is ensured. Because the height control system and the thrust control system are mutually hinged, the height control system and the thrust control system are very easy to generate resonance flutter, and the quality of aircraft carrier landing is difficult to ensure. Based on the reasons, the decoupling of the sectional attack angle instruction, the attack angle equivalent control and the elevator equivalent control is realized by adopting the method, so that the accurate tracking control of the attack angle and the pitch angle can be realized, the resonance of the attack angle and the pitch angle is avoided, and the precision of automatic carrier landing is higher.
It is to be noted that the information invented in the above background section is only for enhancing the understanding of the background of the present invention, and therefore, may include information that does not constitute prior art known to those of ordinary skill in the art.
Disclosure of Invention
The invention aims to provide an automatic carrier landing control method of an unmanned aerial vehicle by adopting a segmented attack angle instruction, and further solves the problem of poor carrier landing effect caused by serious resonance coupling between an accelerator and an elevator in the automatic carrier landing process due to the limitations and defects of the related technology at least to a certain extent.
According to one aspect of the invention, an automatic carrier landing control method of an unmanned aerial vehicle adopting a segmented attack angle instruction is provided, and comprises the following steps:
s10, installing a JC-KYW28A type radio altimeter on the unmanned aerial vehicle, measuring the altitude of the unmanned aerial vehicle, segmenting according to the measured altitude and time, and designing a segmented unmanned aerial vehicle attack angle instruction signal;
step S20, mounting an SMV-1 type attack angle sensor on the unmanned aerial vehicle, measuring an aircraft attack angle signal, comparing the measured aircraft attack angle signal with an attack angle command signal to obtain an attack angle error signal, and integrating the attack angle error signal to obtain an attack angle error integral signal;
step S30, mounting an HPS-1H type speed sensing device on the unmanned aerial vehicle, measuring an aerial vehicle speed signal, mounting BWD-VG300 type inertia measurement equipment, measuring a pitch angle and a pitch angle rate of the aerial vehicle, and forming an attack angle equivalent feedback signal according to an aerial vehicle elevator deflection angle signal;
step S40, according to the attack angle equivalent feedback signal, the attack angle error signal and the attack angle error integral signal, filtering nonlinear transformation and linear superposition are carried out to obtain an accelerator deflection angle signal of the aircraft;
step S50, establishing a deck pitch and heave motion model according to the sea situation parameters to obtain a deck motion compensation instruction, and then designing a deck motion compensator to generate a deck motion height compensation signal;
step S60, according to the initial height of the unmanned aerial vehicle during landing, setting a height downward sliding instruction signal, overlapping a deck movement height compensation signal, then comparing the height downward sliding instruction signal with an aircraft height signal measured by a JC-KYW28A type radio altimeter to obtain a height error signal, and performing integration to obtain a height error integration signal;
step S70, generating a pitch angle time-varying instruction signal and a pitch angle constant instruction signal in a combined manner according to the height error signal and the height error integral signal, and generating a final pitch angle instruction signal through height signal switching;
and step S80, constructing an equivalent signal of the aircraft elevator according to the aircraft speed measuring signal, the attack angle measuring signal, the depression elevation angle speed measuring signal and the aircraft accelerator deflection angle signal.
Step S90, comparing the pitch angle command signal with the pitch angle measurement signal to obtain a pitch angle error signal, constructing an expected signal of the pitch angle rate of the aircraft, and comparing the expected signal with the pitch angle rate signal of the aircraft measured by BWD-VG300 type inertia measurement equipment to obtain the pitch angle rate error signal;
and step S110, comparing the pitch angle rate command signal with the pitch angle rate signal to obtain a pitch angle rate error signal, integrating to obtain a pitch angle rate error integral signal, and overlapping the aircraft elevator equivalent signal to form a final aircraft elevator deflection angle command signal which is transmitted to an aircraft elevator system to realize automatic landing of the aircraft.
In an exemplary embodiment of the invention, segmenting based on measured altitude and time, designing a segmented UAV angle of attack command signal comprises:
Figure BDA0002452341030000041
Figure BDA0002452341030000042
wherein h is an altitude signal obtained by measuring the altitude of the aircraft by using a radio altimeter of JC-KYW28A type. DeltaiIs a time-varying angle of attack signal that varies with altitude, where i is 1,2,3, 4. k is a radical ofiAndit is a time signal timed by taking the automatic landing start time of the aircraft as a zero point, αdIs an angle of attack command signal, where ti、αiThe parameters i are constant positive parameters, i is 1,2,3 and 4, and the detailed design is implemented in the following cases.
In an exemplary embodiment of the invention, comparing the measured signal of the angle of attack of the aircraft with the angle of attack command signal to obtain an angle of attack error signal, and integrating the angle of attack error signal to obtain an angle of attack error integrated signal comprises:
eα＝α-αd；
s1＝∫eαdt；
α is aircraft angle of attack signal measured by SMV-1 type angle of attack sensor, αdFor angle of attack command signals, eαIs angle of attackError signal, s1For the angle of attack error integral signal, dt represents the integral over time.
In an exemplary embodiment of the invention, forming an angle of attack equivalent feedback signal based on a velocity measurement signal of the aircraft and a pitch angle rate of the aircraft and based on an aircraft elevator yaw angle signal comprises:
p1＝a21v+a22α+a23ωz+a24 z；
v is the speed signal of the aircraft measured by HPS-1H type speed sensor and is recorded as the signal, α is the angle of attack signal of the aircraft measured by SMV-1 type angle of attack sensor and omegazFor the pitch angle rate signal of the aircraft obtained by the BWD-VG300 type inertia measurement device,zin order to be an aircraft elevator deflection signal,p1for an angle of attack equivalent feedback signal, where a2i(i is 1,2,3,4) is a constant proportional parameter signal, the detailed design of which will be described later in the examples.
In an exemplary embodiment of the present invention, the obtaining the throttle deflection angle signal of the aircraft by performing filtering nonlinear transformation and linear superposition according to the attack angle equivalent feedback signal, the attack angle error signal, and the attack angle error integral signal includes:
Figure BDA0002452341030000051
Figure BDA0002452341030000052
p＝k11eα+k12eα1+k13eα2+k14s1+p1；
wherein eαAs angle of attack error signal, eα1Filtering the error signal for angle of attack, eα2For an angle of attack nonlinear error signal, s represents the differential operator of the transfer function,
Figure BDA0002452341030000053
is a filter, omega1、TaThe signal is a constant positive signal, and the detailed design thereof is described in the following examples.p1For equivalent feedback signal of angle of attack, s1In order to integrate the signal for the angle of attack error,pis a throttle deflection angle signal of the aircraft, where k1i(i is 1,2,3,4) is a constant proportional parameter signal, the detailed design of which will be described later in the examples.
In an exemplary embodiment of the present invention, establishing a deck pitch and heave motion model based on the sea situation parameters, obtaining deck motion compensation instructions, and then designing a deck motion compensator to generate a deck motion height compensation signal comprises:
θw＝b1sin(w1t)+b2sin(w2t)+b3；
hs＝b4sin(w3t)+b4sin(w4t)；
hb＝hs-0.5xfθw；
hb1＝G1(s)hb；
Figure BDA0002452341030000061
wherein theta iswIs the longitudinal rocking angle of deck, hsFor the deck heave height, bi(i＝1,2,3,4,5)，w1、w2、w3、 w4Is a constant sea state related parameter, ci(i ═ 1, …,9) is a constant parameter of the transfer function, the detailed design of which is described in the examples hereinafter. x is the number offThe distance from an ideal landing point to a stern projection of a flight deck. h isbThe height is initially compensated for the deck. h isb1For deck height compensation signals, G1(s) is the transfer function of the deck compensator and s is the differential operator of the transfer function.
In an exemplary embodiment of the present invention, the setting an altitude glide command signal according to an initial altitude at which the unmanned aerial vehicle starts landing, and superimposing a deck movement altitude compensation signal to obtain an altitude error signal and an altitude error integral signal includes:
hc＝(h0-ha)e-τt+ha；
eh＝h-hc-hb1；
s2＝∫ehdt；
wherein h is0Initial altitude, h, for starting unmanned aerial vehicle on board a vesselcFor a high glide command signal, hb1For deck compensation of height, haτ is the time constant of the exponential glide for deck height, which is designed in detail for the examples hereinafter. h is an aircraft altitude signal measured by a JC-KYW28A type radio altimeter, ehFor the height error signal, s2For the height error integrated signal dt represents the integration of the time signal.
In an exemplary embodiment of the present invention, the generating a time-varying pitch angle command signal and a constant pitch angle command signal in combination according to the altitude error signal and the altitude error integral signal, and the generating a final pitch angle command signal by switching the altitude signals comprises:
Figure BDA0002452341030000071
Figure BDA0002452341030000072
wherein
Figure BDA0002452341030000073
Is a pitch angle time-varying command signal, k21、k22The detailed design of the parameter is described in the following examples.
Figure BDA0002452341030000074
The pitch angle constant value command signal and the constant value parameter are designed in detail and implemented in the following case.
Figure BDA0002452341030000075
For the final pitch angle command signal, h2The height is switched for the pitch angle command, which is a constant parameter, the detailed design of which is implemented in the examples hereinafter.
In an exemplary embodiment of the present invention, obtaining an aircraft elevator equivalent signal according to the aircraft speed measurement signal, the attack angle measurement signal, the pitch angle measurement signal, and the aircraft accelerator drift angle signal includes:
z1＝a41v+a42α+a43ωz+a44 p；
where v is the aircraft speed measurement signal, α is the angle of attack measurement signal, ωzIn order to be the pitch angle rate signal,pis an aircraft throttle deflection angle signal,z1for aircraft elevator equivalent signals, a41、 a42、a43And a44The feedback parameter is a constant value and the detailed design thereof is implemented in the examples which follow the document.
In an exemplary embodiment of the invention, the step of obtaining a pitch angle error signal by comparing the pitch angle command signal with the pitch angle measurement signal, and the step of constructing the desired signal of the pitch angle rate of the aircraft and the pitch angle rate error signal comprises:
Figure BDA0002452341030000076
Figure BDA0002452341030000077
Figure BDA0002452341030000078
wherein
Figure BDA0002452341030000079
In order to be the pitch angle command signal,
Figure BDA00024523410300000710
to obtain a pitch angle signal using a BWD-VG300 type inertial measurement device,
Figure BDA00024523410300000711
in order to be the pitch angle error signal,
Figure BDA00024523410300000712
is a pitch angle error non-linear signal, where Θ is a constant parameter, the detailed design of which is described in the examples below.
Figure BDA00024523410300000713
For pitch angle rate command signals, k31、k32The detailed design of the constant parameter signal is described in the following embodiments.
In an exemplary embodiment of the present invention, comparing the pitch angle rate command signal with the pitch angle rate signal to obtain a pitch angle rate error signal, integrating the pitch angle rate error signal to obtain a pitch angle rate error integral signal, and superimposing the aircraft elevator equivalent signal to form a final aircraft elevator yaw angle command signal includes:
Figure BDA0002452341030000081
s3＝∫eωdt；
z＝k41eω+k42s3+z1；
wherein
Figure BDA0002452341030000082
For said pitch angle rate command signal, ωzFor pitch angle rate signals, eωFor pitch angle rate error signal, s3For the pitch rate error integrated signal,z1in order to be the elevator equivalent signal,zfor the final elevator yaw angle command signal, where k41、k42The detailed design of the constant parameter signal is described in the following embodiments.
And finally, transmitting the elevator deflection angle instruction signal to an aircraft elevator system to drive the aircraft to realize automatic landing on the deck.
Advantageous effects
The unmanned aerial vehicle automatic landing control method adopting the segmented attack angle instruction has the advantages that firstly, a four-segment type highly-coupled attack angle instruction mode is adopted, and the problem of attack angle instruction jumping in the two processes of height gliding and height level flying is softened. Secondly, the coupling resonance problem between the deflection angle of the accelerator and the deflection angle of the elevator is effectively realized by adopting the mode of equivalent control of the attack angle and the equivalent control of the elevator, so that the decoupling of the attack angle control and the pitch angle control is realized. Then, the elevation angle error is driven through the altitude error, the elevation angle rate is driven through the elevation angle error, and the elevator deflection angle is generated through the elevation angle rate error signal to form a three-layer control method, so that a good dynamic effect of the elevation control is achieved, and the method has good theoretical innovation value and engineering application value.
It is to be understood that both the foregoing general description and the following detailed description are exemplary and explanatory only and are not restrictive of the invention, as claimed.
Drawings
The accompanying drawings, which are incorporated in and constitute a part of this specification, illustrate embodiments consistent with the invention and together with the description, serve to explain the principles of the invention. It is obvious that the drawings in the following description are only some embodiments of the invention, and that other drawings can be derived from those drawings by a person skilled in the art without inventive effort.
FIG. 1 is a flow chart of an automatic landing control method for an unmanned aerial vehicle, which adopts a sectional attack angle instruction, according to the invention;
FIG. 2 is a graph of an aircraft angle of attack command signal (in degrees) according to a method provided by an embodiment of the present invention;
FIG. 3 is a plot of an aircraft altitude signal (in meters) in accordance with a method provided by an embodiment of the present invention;
FIG. 4 is a plot of aircraft angle of attack signals (in degrees) according to a method provided by an embodiment of the invention;
FIG. 5 is a plot of an aircraft angle of attack error signal (in degrees) for a method provided by an embodiment of the present invention;
FIG. 6 is a graph of aircraft velocity signals (in meters/second) in accordance with a method provided by an embodiment of the present invention
FIG. 7 is a plot of an aircraft pitch angle signal (in degrees) in accordance with a method provided by an embodiment of the present invention;
FIG. 8 is a plot of an aircraft pitch angle rate signal (in radians/second) in accordance with a method provided by an embodiment of the present invention;
FIG. 9 is a plot of aircraft throttle deflection angle signal (in degrees) according to a method provided by an embodiment of the present invention;
FIG. 10 is a plot of an aircraft altitude error signal (in meters) in accordance with a method provided by an embodiment of the present invention;
FIG. 11 is a plot of aircraft pitch angle desired signal versus actual signal (in degrees) for a method provided by an embodiment of the present invention;
FIG. 12 is a plot of aircraft elevator yaw angle command signals (in degrees) for a method provided by an embodiment of the present invention;
fig. 13 is a comparison curve (unit: meter) of an automatic landing height signal and a height instruction signal of the method provided by the embodiment of the invention.
Detailed Description
Example embodiments will now be described more fully with reference to the accompanying drawings. Example embodiments may, however, be embodied in many different forms and should not be construed as limited to the examples set forth herein; rather, these embodiments are provided so that this disclosure will be thorough and complete, and will fully convey the concept of example embodiments to those skilled in the art. The described features, structures, or characteristics may be combined in any suitable manner in one or more embodiments. In the following description, numerous specific details are provided to provide a thorough understanding of embodiments of the invention. One skilled in the relevant art will recognize, however, that the invention may be practiced without one or more of the specific details, or with other methods, components, devices, steps, and so forth. In other instances, well-known technical solutions have not been shown or described in detail to avoid obscuring aspects of the invention.
The invention provides an automatic landing control method of an unmanned aerial vehicle by adopting a segmented attack angle instruction. Meanwhile, an attitude angle instruction signal is generated through an altitude error signal and an integral signal, the attitude angle instruction signal is compared with an attitude angle signal to obtain an attitude angle error signal and perform nonlinear change, an attitude angle rate expected signal is generated through the error signal and is compared with the attitude angle rate signal to obtain an error signal of the attitude angle rate, finally a control quantity of the elevating and steering system is generated through a pitch angle rate error signal, an equivalent control signal is superposed to realize convergence of altitude errors, and therefore automatic landing control of the aircraft is completed.
The method for controlling automatic landing of an unmanned aerial vehicle using a segmented attack angle command according to the present invention will be further explained and explained with reference to the accompanying drawings. Referring to fig. 1, the method for controlling automatic carrier landing of the unmanned aerial vehicle by using the segmented attack angle instruction comprises the following steps:
s10, installing a JC-KYW28A type radio altimeter on the unmanned aerial vehicle, measuring the altitude of the unmanned aerial vehicle, segmenting according to the measured altitude and time, and designing a segmented unmanned aerial vehicle attack angle instruction signal;
specifically, firstly, a JC-KYW28A type radio altimeter is installed on the aircraft, and the flying height of the aircraft is measured and recorded as h.
Then, from the measured height signal, a time-varying angle of attack signal is designed which varies with height, denoted ΔiWherein i is 1,2,3, 4. The design method is as follows:
Figure BDA0002452341030000121
wherein k isiAndithe detailed design of the parameter is described in the following examples.
And finally, timing by taking the starting moment of the automatic carrier landing of the aircraft as a zero point, recording a time signal as t, dividing the automatic carrier landing process into four stages, designing an attack angle instruction signal of each stage, and recording the attack angle instruction signal as αd。
The design process is as follows:
Figure BDA0002452341030000122
wherein t isi、αiThe parameters i are constant positive parameters, i is 1,2,3 and 4, and the detailed design is implemented in the following cases.
Step S20, mounting an SMV-1 type attack angle sensor on the unmanned aerial vehicle, measuring an aircraft attack angle signal, comparing the measured aircraft attack angle signal with an attack angle command signal to obtain an attack angle error signal, and integrating the attack angle error signal to obtain an attack angle error integral signal;
specifically, an SMV-1 type attack angle sensor is firstly installed on the aircraft, the attack angle of the aircraft is measured and recorded as α, and then the attack angle command signal α is sent to the aircraftdComparing to obtain an error signal, and recording as eαThe calculation method is as follows:
eα＝α-αd；
finally, integrating the attack angle error signal to obtain an attack angle error integral signal which is recorded as s1The calculation method is as follows:
s1＝∫eαdt；
where dt represents the integral of the time signal.
And step S30, mounting an HPS-1H type speed sensing device on the unmanned aerial vehicle, measuring an aerial vehicle speed signal, mounting BWD-VG300 type inertia measurement equipment, measuring a pitch angle and a pitch angle rate of the aerial vehicle, and forming an attack angle equivalent feedback signal according to an aerial vehicle elevator deflection angle signal.
In particularFirstly, mounting an HPS-1H type speed sensing device on an aircraft, measuring a speed signal of the aircraft and recording the speed signal as v; secondly, installing BWD-VG300 type inertia measuring equipment on the aircraft, measuring a pitch angle signal of the aircraft, and recording the pitch angle signal as
Figure BDA0002452341030000134
Simultaneously measuring pitch angle rate signal, recording as omegaz。
Then, introducing an aircraft elevator deflection signal, and recording the signal asz. Finally, proportional feedback is carried out through the signals to form an attack angle equivalent feedback signal which is recorded asp1The calculation method is as follows:
p1＝a21v+a22α+a23ωz+a24 z；
wherein a is2i(i is 1,2,3,4) is a constant proportional parameter signal, the detailed design of which will be described later in the examples.
And step S40, performing filtering nonlinear transformation and linear superposition according to the attack angle equivalent feedback signal, the attack angle error signal and the attack angle error integral signal to obtain an accelerator deflection angle signal of the aircraft.
Specifically, the angle of attack error signal e is first determinedαFiltering and non-linear variation are performed to obtain a filtered error signal denoted as eα1While obtaining a nonlinear error signal of angle of attack, denoted as eα2The calculation method is as follows:
Figure BDA0002452341030000131
Figure BDA0002452341030000132
where s represents the differential operator of the transfer function,
Figure BDA0002452341030000133
is a filter, omega1、TaIs a constant positive signal, which is detailedThe design is described in the following example implementation.
Then, according to the attack angle equivalent feedback signalp1Angle of attack error integral signal s1Filtering the error signal eα1Error signal e nonlinear with angle of attackα2Linear combination is carried out to obtain the throttle deflection angle signal of the aircraft and the signal is recorded aspThe calculation method is as follows:
p＝k11eα+k12eα1+k13eα2+k14s1+p1；
wherein k is1i(i is 1,2,3,4) is a constant proportional parameter signal, the detailed design of which will be described later in the examples.
Step S50, establishing a deck pitch and heave motion model according to the sea situation parameters to obtain a deck motion compensation instruction, and then designing a deck motion compensator to generate a deck motion height compensation signal;
specifically, firstly, a deck pitch model is established according to the sea situation parameters as follows:
θw＝b1sin(w1t)+b2sin(w2t)+b3；
wherein theta iswIs the deck pitch angle, bi(i＝1,2,3)，w1、w2The detailed design of the constant sea-state related parameters is described in the following examples.
Secondly, establishing a deck heave motion model according to the sea condition parameters as follows:
hs＝b4sin(w3t)+b4sin(w4t)；
wherein h issFor the deck heave height, bi(i＝4,5)，w3、w4The detailed design of the constant sea-state related parameters is described in the following examples.
Then, the deck preliminary compensation height is calculated as follows:
hb＝hs-0.5xfθw；
wherein xfFor ideal landing point to the stern of flight deckDistance. h isbThe height is initially compensated for the deck.
Finally, the following deck compensator is established, and the deck compensation height is obtained as follows:
hb1＝G1(s)hb；
wherein h isb1For deck compensation of height, G1(s) is the transfer function of the deck compensator and s is the differential operator of the transfer function. The transfer function of the deck compensator is designed as follows:
Figure BDA0002452341030000151
wherein c isi(i ═ 1, …,9) is a constant parameter of the transfer function, the detailed design of which is described in the examples below.
And S60, setting a height downward sliding instruction signal according to the initial height of the unmanned aerial vehicle during landing, superposing a deck movement height compensation signal, comparing the height downward sliding instruction signal with an aircraft height signal measured by a JC-KYW28A type radio altimeter to obtain a height error signal, and integrating to obtain a height error integral signal.
Specifically, firstly, the initial height of the unmanned aerial vehicle during landing start is recorded as h0(ii) a Then setting the height of the gliding instruction signal and recording the gliding instruction signal as hcThe calculation method is as follows:
hc＝(h0-ha)e-τt+ha；
wherein h isaτ is the time constant of the exponential glide for deck height, the detailed design of which is described in the examples below.
Then, a deck motion altitude compensation signal is superposed, and then h comparison is carried out on the deck motion altitude compensation signal and an aircraft altitude signal measured by a JC-KYW28A type radio altimeter to obtain an altitude error signal which is recorded as ehThe calculation method is as follows:
eh＝h-hc-hb1；
finally, the height error signal is integrated to obtain a height error integrated signal which is recorded as s2Which isThe calculation method is as follows:
s2＝∫ehdt；
where dt represents the integral of the time signal.
Step S70, generating a pitch angle time-varying instruction signal and a pitch angle constant instruction signal in a combined manner according to the height error signal and the height error integral signal, and generating a final pitch angle instruction signal through height signal switching;
specifically, firstly, a time-varying command signal of the pitch angle is designed according to the height error signal and the height error integral signal and recorded as
Figure BDA0002452341030000161
The calculation method is as follows:
Figure BDA0002452341030000162
wherein k is21、k22The detailed design of the parameter is described in the following examples.
Secondly, designing a constant pitch angle command signal and recording the constant pitch angle command signal as
Figure BDA0002452341030000163
It is a constant parameter, and the detailed design is described in the following examples.
Finally, according to the altitude signal of the aircraft, the pitch angle instruction is switched to obtain a pitch angle instruction signal which is recorded as
Figure BDA0002452341030000164
The calculation method is as follows:
Figure BDA0002452341030000165
wherein h is2The height is switched for the pitch angle command, and is a constant parameter, and the detailed design of the constant parameter is implemented in the following case.
And step S80, constructing an equivalent signal of the aircraft elevator according to the aircraft speed measuring signal, the attack angle measuring signal, the depression elevation angle speed measuring signal and the aircraft accelerator deflection angle signal.
Specifically, according to the aircraft speed measurement signal v, the attack angle measurement signal α and the pitch angle rate signal omegazAircraft throttle declination signalpCarrying out proportional feedback transformation to construct equivalent signals of the aircraft elevator, and recording the equivalent signals asz1The calculation method is as follows:
z1＝a41v+a42α+a43ωz+a44 p；
wherein a is41、a42、a43And a44The feedback parameter is a constant proportional feedback parameter, and the detailed design thereof is shown in the implementation of the following case.
And step S90, comparing the pitch angle command signal with the pitch angle measurement signal to obtain a pitch angle error signal, constructing an expected signal of the pitch angle rate of the aircraft, and comparing the expected signal with the pitch angle rate signal of the aircraft measured by BWD-VG300 type inertia measurement equipment to obtain the pitch angle rate error signal.
Specifically, the pitch angle command signal is firstly used
Figure BDA0002452341030000171
The pitch angle signal obtained by the BWD-VG300 type inertia measurement equipment
Figure BDA0002452341030000172
Comparing to obtain a pitch angle error signal, and recording the pitch angle error signal as
Figure BDA0002452341030000173
The calculation method is as follows:
Figure BDA0002452341030000174
secondly, according to the pitch angle error signal, carrying out nonlinear transformation to obtain a pitch angle error nonlinear signal which is recorded as
Figure BDA0002452341030000175
The calculation method is as follows:
Figure BDA0002452341030000176
wherein Θ is a constant parameter, and the detailed design thereof is shown in the examples and examples hereinafter.
Finally, according to the pitch angle error signal
Figure BDA0002452341030000177
Non-linear signal of pitch angle error
Figure BDA0002452341030000178
Linear combination is performed to obtain a pitch angle rate command signal which is recorded as
Figure BDA0002452341030000179
The calculation method is as follows:
Figure BDA00024523410300001710
wherein k is31、k32The detailed design of the constant parameter signal is described in the following embodiments.
And S100, comparing the pitch angle rate command signal with the pitch angle rate signal to obtain a pitch angle rate error signal, integrating to obtain a pitch angle rate error integral signal, and overlapping the aircraft elevator equivalent signal to form a final aircraft elevator deflection angle command signal which is transmitted to an aircraft elevator system to realize automatic landing of the aircraft.
Specifically, firstly, according to the pitch angle rate command signal
Figure BDA0002452341030000181
And pitch angle rate signal omegazComparing to obtain pitch angle rate error signal recorded as eωThe calculation method is as follows:
Figure BDA0002452341030000182
secondly, integrating the pitch angle rate error signal to obtain a pitch angle rate error integrated signal s3The calculation method is as follows:
s3＝∫eωdt；
then, aiming at the pitch angle rate error signal eωIntegral signal s of pitch angle rate error3Equivalent signal to elevatorz1Linear superposition is carried out to obtain an elevator deflection angle command signal which is recorded aszThe calculation method is as follows:
z＝k41eω+k42s3+z1；
wherein k is41、k42The detailed design of the constant parameter signal is described in the following embodiments.
And finally, transmitting the elevator deflection angle instruction signal to an aircraft elevator system to drive the aircraft to realize automatic landing on the deck.
Case implementation and computer simulation result analysis
In order to verify the correctness and the effectiveness of the method provided by the invention, the following case simulation is provided for simulation.
In step S10, t is designed1＝10，t2＝20，t3＝30，k1＝k2＝k3＝k4＝0.0087，Δ1＝100，Δ2＝50，Δ3＝20，Δ4The aircraft angle of attack command signal is obtained as shown in fig. 2, 10. Meanwhile, an elevation signal of the aircraft is measured by a radio altimeter of JC-KYW28A type as shown in figure 3.
In step S20, an SMV-1 type angle of attack sensor is installed on the unmanned aerial vehicle, and an angle of attack signal is measured as shown in fig. 4, resulting in an angle of attack error signal as shown in fig. 5.
In step S30, a is selected21＝-0.13，a22＝-10，a23＝29，a24-2.1. The speed signal of the aircraft is measured by installing an HPS-1H type speed sensing device as shown in figure 6. The BWD-VG300 type inertia measurement device is installed on the aircraft, and the pitch angle signal of the aircraft is measured as shown in FIG. 7, and the pitch angle rate signal is measured as shown in FIG. 8.
In step S40, Ω is selected1＝0.05，Ta＝1，k11＝-47，k12＝-52，k13＝-42， k14The throttle angle signal is obtained as shown in fig. 9 at-8.
In step S50, b is selected1＝0.0087，b2＝0.0052，b3＝0.0026，b4＝1.3，b5＝0.25， w1＝0.6，w2＝0.65，w3＝0.62，w4＝0.22。xf＝200，hb＝5，c1＝1.66，c2＝3.2， c3＝2.2，c4＝0.62，c5＝0.0042，c6＝0.07，c7＝0.422，c8＝1.12，c9＝1.02。
In step S60, τ is selected to be 0.4, and the height error signal is obtained as shown in fig. 10.
In step S70, k is selected21＝-0.087、k22＝-0.0017，
Figure BDA0002452341030000191
h2A comparison of the aircraft pitch angle desired signal to the actual signal is obtained as shown in fig. 11.
In step S80, a is selected41＝0.0016、a42＝-0.6425、a43-0.1402 and a44＝0.0012。
In step S90, Θ is selected to be 0.12, k31＝-1.5、k32＝-0.7。
In step S100, k is designed41＝5、k42The elevator yaw angle command signal is shown in fig. 12 at 0.2. The comparison curve of the finally realized automatic landing height signal and the height command signal is shown in fig. 13.
In fig. 2, it can be seen that the attack angle command designed by the present invention changes in a stepwise manner, mainly to avoid sudden change of the command from gliding to level flight during landing, so that compared with the conventional attack angle command design with constant value, the command design of the present invention can better adjust the attitude change of the aircraft. Fig. 4 and 5 show that the design of the invention can quickly realize the instruction tracking of the attack angle, the tracking error is quickly converged to 0, and the influence of the hinge is small. Fig. 3 and 6 show the altitude change and the speed change of the aircraft, respectively. Fig. 7 and 11 show the pitch angle versus pitch angle desired value curves of the aircraft, and it can be seen that the design of the present invention enables fast tracking of pitch angle versus desired command signal. Fig. 8 shows a pitch angle rate curve of the aircraft, fig. 9 and fig. 12 show curves of the accelerator deflection angle and the elevator deflection angle designed by the invention, and it can be seen that the curves are all within a normal range, and particularly, the flutter between the curves is less, so that no serious coupling effect is generated. Fig. 10 shows a height error curve of the aircraft, and it can be seen that the height error fluctuates with deck fluctuation, but the error fluctuation range is small. While figure 13 gives a comparison of the altitude of the aircraft with the desired altitude, it can be seen that the basic trends coincide at the extremities. Therefore, the implementation results of the cases show that the design method realizes the decoupling of the accelerator and the elevator system, avoids the resonance problem caused by the coupling between the accelerator and the elevator system, and ensures that the final automatic landing height control effect is better, thereby ensuring that the design method has high engineering practical value.
Other embodiments of the invention will be apparent to those skilled in the art from consideration of the specification and practice of the invention disclosed herein. This application is intended to cover any variations, uses, or adaptations of the invention following, in general, the principles of the invention and including such departures from the present disclosure as come within known or customary practice within the art to which the invention pertains. It is intended that the specification and examples be considered as exemplary only, with a true scope and spirit of the invention being indicated by the following claims.

Claims (10)

1. An automatic carrier landing control method of an unmanned aerial vehicle by adopting a segmented attack angle instruction is characterized by comprising the following steps:
s10, installing a JC-KYW28A type radio altimeter on the unmanned aerial vehicle, measuring the altitude of the unmanned aerial vehicle, segmenting according to the measured altitude and time, and designing a segmented unmanned aerial vehicle attack angle instruction signal;
step S20, mounting an SMV-1 type attack angle sensor on the unmanned aerial vehicle, measuring an aircraft attack angle signal, comparing the aircraft attack angle signal with an attack angle command signal to obtain an attack angle error signal, and integrating the attack angle error signal to obtain an attack angle error integral signal;
step S30, mounting an HPS-1H type speed sensing device on the unmanned aerial vehicle, measuring an aerial vehicle speed signal, mounting BWD-VG300 type inertia measurement equipment, measuring a pitch angle and a pitch angle rate of the aerial vehicle, and forming an attack angle equivalent feedback signal according to an aerial vehicle elevator deflection angle signal;
step S40, according to the attack angle equivalent feedback signal, the attack angle error signal and the attack angle error integral signal, filtering nonlinear transformation and linear superposition are carried out to obtain an accelerator deflection angle signal of the aircraft;
step S50, establishing a deck pitch and heave motion model according to the sea situation parameters to obtain a deck motion compensation instruction, and then designing a deck motion compensator to generate a deck motion height compensation signal;
step S60, according to the initial height of the unmanned aerial vehicle during landing, setting a height glide command signal, overlapping a deck movement height compensation signal, comparing the height glide command signal with an aircraft height signal measured by a JC-KYW28A type radio altimeter to obtain a height error signal, and integrating the height error signal to obtain a height error integral signal;
step S70, generating a pitch angle time-varying instruction signal and a pitch angle constant instruction signal in a combined manner according to the height error signal and the height error integral signal, and generating a final pitch angle instruction signal through height signal switching;
and step S80, constructing an equivalent signal of the aircraft elevator according to the aircraft speed measuring signal, the attack angle measuring signal, the pitch angle speed measuring signal and the aircraft accelerator drift angle signal.
Step S90, comparing the pitch angle instruction signal with the pitch angle measurement signal to obtain a pitch angle error signal, constructing an expected signal of the pitch angle rate of the aircraft, and comparing the expected signal with the pitch angle rate signal of the aircraft measured by BWD-VG300 type inertia measurement equipment to obtain the pitch angle rate error signal;
and S100, comparing the pitch angle rate command signal with the pitch angle rate signal to obtain a pitch angle rate error signal, integrating to obtain a pitch angle rate error integral signal, and overlapping the aircraft elevator equivalent signal to form a final aircraft elevator deflection angle command signal which is transmitted to an aircraft elevator system to realize automatic landing of the aircraft.
2. The method for controlling the automatic landing of the unmanned aerial vehicle by adopting the segmented attack angle command as claimed in claim 1, wherein the step of segmenting according to the measured altitude and the time is carried out, and the step of designing the segmented attack angle command signal of the unmanned aerial vehicle comprises the following steps:
Figure FDA0002452341020000021
Figure FDA0002452341020000031
wherein h is an altitude signal obtained by measuring the altitude of the aircraft by using a JC-KYW28A type radio altimeter. DeltaiIs a time-varying angle of attack signal that varies with altitude, where i is 1,2,3, 4. k is a radical ofiAndit is a time signal which takes the automatic landing starting moment of the aircraft as zero point for timing, αdIs an angle of attack command signal, where ti、αiIs a constant positive parameter, i is 1,2,3, 4.
3. The method for controlling the automatic landing of the unmanned aerial vehicle by adopting the segmented attack angle instruction according to claim 1, wherein the step of forming an attack angle equivalent feedback signal according to a speed measurement signal of the unmanned aerial vehicle, a pitch angle and a pitch angle rate of the unmanned aerial vehicle and an elevator deflection angle signal of the unmanned aerial vehicle comprises the following steps:
p1＝a21v+a22α+a23ωz+a24 z；
v is the speed signal of the aircraft measured by HPS-1H type speed sensor and is recorded as the signal, α is the angle of attack signal of the aircraft measured by SMV-1 type angle of attack sensor and omegazFor the pitch angle rate signal of the aircraft obtained by the BWD-VG300 type inertia measurement device,zin order to be an aircraft elevator deflection signal,p1for an angle of attack equivalent feedback signal, where a2iAnd (i is 1,2,3 and 4) is a constant value proportional parameter signal.
4. The method for controlling the automatic landing of the unmanned aerial vehicle by adopting the segmented attack angle instruction according to claim 1, wherein the step of performing filtering nonlinear transformation and linear superposition according to the attack angle equivalent feedback signal, the attack angle error signal and the attack angle error integral signal to obtain the throttle deflection angle signal of the unmanned aerial vehicle comprises the following steps:
eα＝α-αd；
s1＝∫eαdt；
Figure FDA0002452341020000041
Figure FDA0002452341020000042
p＝k11eα+k12eα1+k13eα2+k14s1+p1；
wherein eαAs angle of attack error signal, eα1Filtering for angle of attackError signal, eα2For an angle of attack nonlinear error signal, s represents the differential operator of the transfer function,
Figure FDA0002452341020000043
is a filter, omega1、TaThe signal is a constant positive signal, and the detailed design thereof is described in the following examples.p1For equivalent feedback signal of angle of attack, s1In order to integrate the signal for the angle of attack error,pis a throttle deflection angle signal of the aircraft, where k1i(i is 1,2,3,4) is a constant proportional parameter signal, the detailed design of which will be described later in the examples.
5. The unmanned aerial vehicle automatic landing control method adopting the segmented attack angle instruction according to claim 1, wherein a deck pitch and heave motion model is established according to sea condition parameters to obtain a deck motion compensation instruction, then a deck motion compensator is designed, and generating a deck motion height compensation signal comprises:
θw＝b1sin(w1t)+b2sin(w2t)+b3；
hs＝b4sin(w3t)+b4sin(w4t)；
hb＝hs-0.5xfθw；
hb1＝G1(s)hb；
Figure FDA0002452341020000044
wherein theta iswIs the longitudinal rocking angle of deck, hsFor the deck heave height, bi(i＝1,2,3,4,5)，w1、w2、w3、w4Is a constant sea state related parameter, ci(i ═ 1, …,9) is a constant parameter of the transfer function, the detailed design of which is described in the examples below. x is the number offThe distance from an ideal landing point to a stern projection of a flight deck. h isbThe height is initially compensated for the deck. h isb1Is the height of deckCompensation signal, G1(s) is the transfer function of the deck compensator and s is the differential operator of the transfer function.
6. The method for controlling automatic landing of the unmanned aerial vehicle by using the segmented attack angle command according to claim 1, wherein the step of setting a glide command signal of the altitude according to the initial altitude of the unmanned aerial vehicle at which the landing starts and superimposing a deck motion altitude compensation signal to obtain the altitude error signal and the altitude error integral signal comprises the steps of:
hc＝(h0-ha)e-τt+ha；
eh＝h-hc-hb1；
s2＝∫ehdt；
wherein h is0Initial altitude, h, for starting unmanned aerial vehicle on board a vesselcFor a high glide command signal, hb1For deck height compensation signal, haτ is the time constant of the exponential glide for deck height, the detailed design of which is described in the examples below. h is an aircraft altitude signal measured by a JC-KYW28A type radio altimeter, ehFor the height error signal, s2For a high error integrated signal, dt represents the integration of the time signal.
7. The method for controlling automatic landing of the unmanned aerial vehicle by using the segmented attack angle command as claimed in claim 1, wherein the step of generating the time-varying pitch angle command signal and the constant pitch angle command signal by combining the altitude error signal and the altitude error integral signal, and generating the final pitch angle command signal by switching the altitude signals comprises:
Figure FDA0002452341020000052
Figure FDA0002452341020000051
wherein
Figure FDA0002452341020000053
Is a pitch angle time-varying command signal, k21、k22Is a constant parameter.
Figure FDA0002452341020000054
The pitch angle constant value command signal is a constant value parameter.
Figure FDA0002452341020000055
For the final pitch angle command signal, h2And switching the height for the pitch angle instruction, wherein the height is a constant parameter.
8. The method for controlling the automatic landing of the unmanned aerial vehicle by adopting the segmented attack angle instruction according to claim 1, wherein the step of obtaining the equivalent signal of the aircraft elevator according to the speed measurement signal of the aircraft, the attack angle measurement signal, the pitch angle speed measurement signal and the throttle deflection signal of the aircraft comprises the following steps:
z1＝a41v+a42α+a43ωz+a44 p；
where v is the aircraft speed measurement signal, α is the angle of attack measurement signal, ωzIn order to be the pitch angle rate signal,pis an aircraft throttle deflection angle signal,z1for aircraft elevator equivalent signals, a41、a42、a43And a44Is a constant proportional feedback parameter.
9. The method for controlling automatic landing of the unmanned aerial vehicle by using the segmented attack angle command as claimed in claim 1, wherein the step of obtaining a pitch angle error signal by comparing the pitch angle command signal with the pitch angle measurement signal according to the pitch angle command signal, and the step of constructing the pitch angle rate error signal and the expected pitch angle rate signal of the unmanned aerial vehicle comprises the steps of:
Figure FDA0002452341020000064
Figure FDA0002452341020000061
Figure FDA0002452341020000062
wherein
Figure FDA0002452341020000065
In order to be the pitch angle command signal,
Figure FDA0002452341020000066
to obtain a pitch angle signal using a BWD-VG300 type inertial measurement device,
Figure FDA0002452341020000067
in order to be the pitch angle error signal,
Figure FDA0002452341020000068
is a pitch angle error nonlinear signal, where Θ is a constant parameter.
Figure FDA0002452341020000063
For pitch angle rate command signal, k31、k32Is a constant parameter signal.
10. The method for controlling the automatic landing of the unmanned aerial vehicle by adopting the segmented attack angle command according to claim 1, wherein the step of comparing the pitch angle rate command signal with the pitch angle rate signal to obtain a pitch angle rate error signal, the step of integrating to obtain a pitch angle rate error integral signal, and the step of superposing the aircraft elevator equivalent signal to form a final aircraft elevator deflection angle command signal comprises the steps of:
Figure FDA0002452341020000071
s3＝∫eωdt；
z＝k41eω+k42s3+z1；
wherein
Figure FDA0002452341020000072
For said pitch angle rate command signal, ωzFor pitch angle rate signals, eωFor pitch angle rate error signal, s3For the pitch rate error integrated signal,z1in order to be an elevator equivalent signal,zfor the final elevator yaw angle command signal, where k41、k42Is a constant parameter signal.
CN202010296385.6A 2020-04-15 2020-04-15 Unmanned aerial vehicle automatic carrier landing control method adopting segmented attack angle instruction Withdrawn CN111459184A (en)

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Cited By (3)

* Cited by examiner, † Cited by third party
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CN115562314A (en) * 2022-10-19 2023-01-03 航天科工火箭技术有限公司 Carrier rocket sublevel landing area control method, system, medium and computer equipment
CN115617057A (en) * 2022-10-31 2023-01-17 南京航空航天大学 Synchronous control method for longitudinal pitch angle of landing tail end of four-tilt rotor aircraft
CN116700358A (en) * 2023-08-08 2023-09-05 成都飞机工业（集团）有限责任公司 Nonlinear height-fixing compensation control method for unmanned aerial vehicle in turning stage

Cited By (6)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN115562314A (en) * 2022-10-19 2023-01-03 航天科工火箭技术有限公司 Carrier rocket sublevel landing area control method, system, medium and computer equipment
CN115562314B (en) * 2022-10-19 2024-06-07 航天科工火箭技术有限公司 Carrier rocket sublevel landing zone control method, system, medium and computer equipment
CN115617057A (en) * 2022-10-31 2023-01-17 南京航空航天大学 Synchronous control method for longitudinal pitch angle of landing tail end of four-tilt rotor aircraft
CN115617057B (en) * 2022-10-31 2024-03-29 南京航空航天大学 Method for synchronously controlling longitudinal pitch angle of landing tail end of four-tilting rotor aircraft
CN116700358A (en) * 2023-08-08 2023-09-05 成都飞机工业（集团）有限责任公司 Nonlinear height-fixing compensation control method for unmanned aerial vehicle in turning stage
CN116700358B (en) * 2023-08-08 2023-12-08 成都飞机工业（集团）有限责任公司 Nonlinear height-fixing compensation control method for unmanned aerial vehicle in turning stage

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