CN116203840A - Adaptive gain scheduling control method for reusable carrier - Google Patents

Adaptive gain scheduling control method for reusable carrier Download PDF

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CN116203840A
CN116203840A CN202211683045.4A CN202211683045A CN116203840A CN 116203840 A CN116203840 A CN 116203840A CN 202211683045 A CN202211683045 A CN 202211683045A CN 116203840 A CN116203840 A CN 116203840A
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yaw
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王永帅
孙明玮
陈增强
张婷玉
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Nankai University
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Abstract

The invention discloses a reusable carrier self-adaptive gain scheduling control method, which is based on the real-time estimation and compensation effects of ESO on uncertainty and disturbance, and realizes the accurate attitude control of RLV under the effects of unmodeled dynamic, unknown external disturbance and channel coupling; based on a classification compensation mechanism of the estimator for disturbance, realizing self-adaptive scheduling of control gain under cross-domain maneuver; and the adaptive compensation is carried out based on the nominal controller, the scheduling range is designed according to priori information, and the stability performance of the proposed controller is ensured. The invention improves the attitude control performance and the anti-interference capability of the cross-domain recovery of the RLV and has good adaptability and robustness.

Description

Adaptive gain scheduling control method for reusable carrier
Technical Field
The invention relates to the technical field of aerospace vehicle control, in particular to a reusable carrier self-adaptive gain scheduling control method based on estimation and compensation.
Background
The reusable carrier (Reusable Launch Vehicle, RLV) is a type of aircraft that can perform the firing mission again after the scheduled firing mission is completed, through service and fueling. The vertical take-off and landing scheme inherits the traditional carrier rocket configuration and has the advantages of being low in technical span and research and development cost and the like. However, RLV has large space and speed spans in vertical return flight, severe dynamic pressure variation, complex and changeable flight environment, strong internal and external uncertainties and disturbance such as pneumatic parameter deviation, structure and wind disturbance, and the like, and each channel presents serious nonlinear coupling characteristics, and particularly when internal parameters of a system are changed or serious external disturbance occurs, the traditional control theory is difficult to meet the high-performance control requirement under the special maneuver of a modern carrier rocket.
The accurate attitude control during vertical landing is the key for realizing the safety recovery of the carrier, and the prior engineering mostly adopts a gain scheduling method to carry out attitude design on the aircraft, for example, adopts a classical control method such as PID (proportion integration differentiation) and the like to carry out control gain scheduling design. However, the simple linear control method cannot cope with strong nonlinear coupling in the system and unknown disturbance and uncertainty, and it is difficult to obtain a satisfactory control effect. With the development of modern technology, advanced control theory such as robust control, sliding mode control, self-adaptive backstepping control, disturbance observer-based control methods and the like are developed in RLV attitude control.
Robust control allows the design controller to suppress the effects of uncertainty by quantitatively analyzing the uncertainty. Based on H ∞ The robust control method can solve the problem of uncertainty caused by uncertain pneumatic parameters, liquid fuel shaking and the like, thereby realizing the robust attitude control design of the RLV. The sliding mode control designs the switching hyperplane of the system according to the expected dynamic characteristics of the system, so that the parameter uncertainty and external disturbance are effectively treated. Because the traditional sliding mode control method based on the linear sliding mode surface is difficult to avoid the shake problem of the control quantity, the terminal sliding mode control is gradually developed, the rapid convergence of a control system is realized, and meanwhile, the shake problem which is concerned is solved. The nonsingular terminal sliding mode control further solves the singular problem of the traditional terminal sliding mode, and promotes the application of the sliding mode control technology in aircraft control. In addition, the self-adaptive backstepping control decomposes the nonlinear control system into a plurality of subsystems which are not more than the system order, virtual controllers which enable all the subsystems to be stable are respectively designed, and the virtual controllers are gradually backstepped by utilizing the Lyapunov function, so that the controllers which enable the whole system to be stable are finally obtained. The controller designed is extremely complex due to the multi-layer recursion of this approach.
In contrast, control ideas based on disturbance estimation and compensation are more intuitive and have significant advantages in terms of handling system disturbances and uncertainties, such as unknown input observers, disturbance observers, generalized proportional integral observers, uncertainty and disturbance estimators, extended state observers (Extended State Observer, ESO), etc. The ESO design needs minimum dynamic system information, has a simple structure and is easy to design, and the system state can be estimated while disturbance is estimated, so that the ESO design has a great deal of attention in the control field.
However, when the reusable vehicle returns to re-atmosphere, the cross-domain maneuver of the RLV in large airspace causes the aircraft's own characteristics and external flight environment to vary with airspace, speed domain, resulting in complex and varying total disturbance types for ESO estimation. The influence of different disturbance types on the system is different, and the compensation mechanism is inappropriate and can be the opposite, so that the control performance and the stability margin are affected. Currently, related control methods specifically directed to cross-domain maneuver characteristics remain relatively lacking.
Disclosure of Invention
The invention aims at solving the problems of wide space crossing, serious coupling effect among channels, serious fuel consumption, great shaking, unmodeled dynamics such as aircraft structural change and the like of RLV reentry flight, and provides a reusable carrier self-adaptive gain scheduling attitude control method based on a combined strategy of an observer and an estimator, which is applied to attitude control of a reusable carrier reentry atmosphere so as to solve the control gain scheduling problem of an RLV cross-domain maneuvering process in a large space domain, thereby realizing safe recovery of the RLV.
The technical scheme adopted for realizing the purpose of the invention is as follows:
a method of adaptive gain scheduling control for a reusable carrier, comprising the steps of:
s1, designing a guidance law for the RLV flight attitude according to the track change requirement of the RLV reentry flight process so as to generate a corresponding attitude tracking instruction; the guidance instructions of the three channels of pitching, rolling and yawing are as follows:
Figure BDA0004019796180000031
θ is the ballistic dip angle, α r In order for the guidance law of angle of attack,
Figure BDA0004019796180000034
is pitch angle command, beta r For sideslip angle command, gamma r Is a roll angle command; set sideslip angle command beta r =0 for realizing yaw angle control ψ r ＝Ψ v ，Ψ r Is yaw angle, psi v Is the ballistic deflection angle;
s2, designing a six-degree-of-freedom dynamic model of the RLV reentry flight process, and establishing an RLV attitude control model under disturbance and uncertainty by considering unmodeled dynamics, uncertainty and disturbance including air density, flight speed and dynamic characteristic changes under cross-domain maneuver, coupling effect among channels, fuel consumption and large-scale shaking and aircraft structural changes;
the six-degree-of-freedom dynamic model is:
kinetic equation of RLV centroid movement:
Figure BDA0004019796180000032
the kinetic equation for the rotation of the RLV around the centroid is:
Figure BDA0004019796180000033
kinematic equation of RLV centroid motion:
Figure BDA0004019796180000041
kinematic equation for RLV rotation about centroid
Figure BDA0004019796180000042
Wherein m is RLV mass, g is gravitational acceleration, alpha and beta are attack angle and sideslip angle respectively,
Figure BDA0004019796180000046
psi, gamma are pitch angle, yaw angle and roll angle, respectively, theta, psi V ,γ V Respectively, the ballistic inclination angle, the ballistic deflection angle and the roll angle, V is the displacement speed, x, y and z are the position coordinates, omega x ,ω y ,ω z For angular velocity, J x ,J y ,J z For moment of inertia, M x ,M y ,M z The force moment vector is respectively a component of the external moment vector on each axis of the projectile body coordinate system, and X, Y and Z are respectively resistance, lifting force and side force;
Figure BDA0004019796180000043
Figure BDA0004019796180000044
wherein ,
Figure BDA0004019796180000045
is dynamic pressure, ρ is air density of the flying height of RLV, S is characteristic area of RLV, L b ,L c The lateral and longitudinal reference lengths, c, of the RLV, respectively x ,c y ,c z Respectively a drag coefficient, a lift coefficient and a side force coefficient, m x ,m y ,m z Respectively representing a rolling moment coefficient, a yaw moment coefficient and a pitching moment coefficient;
taking into account the RLV characteristics and various uncertainties and disturbances of the return flight, for
Figure BDA0004019796180000047
And (3) carrying out secondary derivation on the phi and gamma, and establishing a following attitude control model:
Figure BDA0004019796180000051
Figure BDA0004019796180000052
Figure BDA0004019796180000053
wherein ,
Figure BDA0004019796180000054
b f ,b p ,b r control gain representing pitch, yaw, roll channel,/->
Figure BDA0004019796180000055
Moment coefficient components of pitch, yaw and roll channels, respectively, +.>
Figure BDA0004019796180000056
Is the static derivative, delta z ,δ y ,δ x Mathematical rudders to be designed for pitch, yaw and roll channels;
the state information in the pitch, yaw and roll channels other than rudder effectiveness is defined as disturbance f 1 ,f 2 ,f 3 And f 1 =f 01 +d 1 ,f 2 ＝f 02 +d 2 ,f 3 ＝f 03 +d 3 ，f 01 ,f 02 ,f 03 D is modeled dynamics, d 1 ,d 2 ,d 3 Including remaining unmodeled dynamics, uncertainty, and unknown external disturbances in the pitch, yaw, and roll channels;
Figure BDA0004019796180000057
s3, estimating integrated total disturbance of three channels of pitching, rolling and yawing in real time by adopting ESO and performing feedback compensation;
using angular velocity information omega for three channels, pitch, yaw and roll, respectively z ,ω y ,ω x Designing ESO to estimate total disturbance in real time, wherein ESO designed for a pitch channel, a yaw channel and a roll channel are respectively
Figure BDA0004019796180000058
Figure BDA0004019796180000061
Figure BDA0004019796180000062
wherein ,
Figure BDA0004019796180000063
estimated values of angular velocities of pitch, yaw and roll channels, respectively, i.e. +.>
Figure BDA00040197961800000610
z ψ1 ≈ω y ,z γ1 ≈ω x ，/>
Figure BDA0004019796180000064
Estimated total disturbance values for pitch, yaw and roll channels, respectively, i.e
Figure BDA00040197961800000611
z ψ2 ≈f 2 +(b p -b p0 )δ y ，z γ2 ≈f 3 +(b r -b r0 )δ x ，b f0 ,b p0 ,b r0 B is f ,b p ,b r Is determined by the set of parameters;
placing ESO poles at-omega oz ,-ω oy ,-ω ox Where omega oz ,ω oy ,ω ox Six observer gains l for pitch, yaw and roll channel bandwidths, respectively z1 ,l z2 ,l y1 ,l y2 ,l x1 ,l x2 Satisfy the following requirements
Figure BDA0004019796180000065
Figure BDA0004019796180000066
S4, considering continuous changes of system gain and uncertainty caused by dynamic characteristics and pneumatic parameters of the RLV under the cross-domain maneuver, classifying and identifying disturbance and uncertainty of the RLV under the cross-domain maneuver by adopting a symbol estimator, and performing respective compensation based on a nominal controller to realize accurate attitude control of the RLV reentry flight;
the symbol estimator is adopted to carry out analysis estimation and compensation on disturbance and uncertainty, and the established loss function is that
Figure BDA0004019796180000067
/>
Figure BDA0004019796180000068
Figure BDA0004019796180000069
wherein ,δbf ,δ bp ,δ br Pitch, yaw and roll channel system gains b, respectively f ,b p ,b r Is used to determine the uncertainty of the estimate of (c),
Figure BDA00040197961800000612
f ψ ,f r disturbance and uncertainty which are irrelevant to rudder efficiency; solving the established loss function by adopting a symbol projection gradient strategy to obtain:
Figure BDA0004019796180000071
Figure BDA0004019796180000072
Figure BDA0004019796180000073
wherein ,δbf ,δ bp ,δ br Estimated values of pitch, yaw and roll channel rudder uncertainty,
Figure BDA0004019796180000078
f ψ ,f γ estimated values of steering independent disturbances and uncertainties in pitch, yaw and roll channels, alpha 1 ,α 2 ,α 3 The update constants to be designed of pitch, yaw and roll channels are respectively more than 0;
based on the estimation results and the attitude angle and angular rate information of the three channels, the mathematical rudder control laws of the pitch, yaw and roll channels are designed as follows:
Figure BDA0004019796180000074
Figure BDA0004019796180000075
Figure BDA0004019796180000076
wherein ,kpf ,k pp ,k pr Proportional feedback gains, k, for pitch, yaw and roll channels, respectively df ,k dp ,k dr Differential feedback gain for three channels;
by mathematicsRudder delta x ,δ y ,δ z The equivalent grid rudders for three channels of pitch, yaw and roll are obtained as follows:
Figure BDA0004019796180000077
the three-channel equivalent grid rudder is applied to a six-degree-of-freedom model of the RLV to realize accurate tracking guidance law of pitch angle, sideslip angle and roll angle, namely
Figure BDA0004019796180000079
ψ＝ψ r ＝ψ V (β＝0),γ＝ r =0, and attitude control of RLV cross-domain maneuver is performed.
According to the invention, by designing the adaptive gain scheduling control strategy of the RLV, the disturbance such as strong coupling among channels in the cross-domain maneuver, uncertainty of parameters and models and the like is estimated and separated and compensated in real time, so that the gesture control precision of the cross-domain maneuver can be effectively improved.
The adaptive gain scheduling design is carried out based on the nominal controller, the control gain selection problem in the traditional ADRC controller is solved, the adaptive gain scheduling design is simple in structure and easy to design, the stability of gesture control is ensured under cross-domain maneuver, meanwhile, the stability margin of a control system is improved, and the adaptive gain scheduling control system has stronger robustness and adaptability.
Drawings
Fig. 1 is a schematic diagram of a reusable carrier adaptive gain scheduling control method of the present invention.
FIG. 2 is a cross-domain maneuver height profile view of the RLV of the present invention.
Fig. 3 is a velocity profile of the RLV cross-domain maneuver of the present invention.
Fig. 4 is a three-way attitude angle tracking curve of the present invention.
Fig. 5 is a three-way attitude angular velocity change curve of the present invention.
Fig. 6 is a control moment curve of three mathematical rudders of the invention.
Detailed Description
The invention is described in further detail below with reference to the drawings and the specific examples. It should be understood that the specific embodiments described herein are for purposes of illustration only and are not intended to limit the scope of the invention.
The invention can reuse the carrier self-adaptive gain scheduling control method, based on the real-time estimation and compensation effect of ESO on uncertainty and disturbance, can realize the accurate attitude control of RLV under the effects of unmodeled dynamic, unknown external disturbance and channel coupling. Based on a classification compensation mechanism of the estimator for disturbance, the self-adaptive scheduling of control gain under cross-domain maneuver can be realized. In order to ensure the stability performance of the proposed controller, adaptive compensation is performed based on the nominal controller, and the scheduling range is designed according to a priori information. The invention improves the attitude control performance and the anti-interference capability of the cross-domain recovery of the RLV and has good adaptability and robustness.
As shown in fig. 1, the method for controlling adaptive gain scheduling of a reusable carrier according to the embodiment of the present invention includes the following steps:
step one: according to the track change requirement of the RLV reentry flight process, designing a proper guidance law for the RLV flight attitude so as to generate a corresponding attitude tracking instruction;
in the step (one), RLV reenters the flight, the flight altitude and Mach number thereof are gradually reduced. Thus, the longitudinal height profile is designed to be h r ，
Figure BDA0004019796180000091
wherein ,t0 ,t 1 ,t 2 ,h 01 ,h 02 ,a 1 To be designed constant, t 0 For the initial time value, t 1 ,t 2 To segment the time of the function, h 01 To reenter the initial altitude of flight, h 02 To reenter the end altitude of the flight, a 1 As a function of the coefficients of the piecewise function,
Figure BDA0004019796180000092
root for pitch channelDesigning an attack angle guidance law alpha according to the current real altitude h of the RLV reentry flight process r The method comprises the following steps:
Figure BDA0004019796180000094
/>
wherein the saturation function is defined as
Figure BDA0004019796180000093
i＝1,2，e h ＝h r H is a height tracking error, h i ,k pα ,k dα To be designed constant, h i Is the height error, k pα Is a proportionality coefficient, k dα Is the differential coefficient alpha 0 Is the initial value of attack angle.
The inertia time constant considering the ballistic tilt angle θ is much larger than
Figure BDA0004019796180000097
And a time constant of alpha, according to a geometrical relationship
Figure BDA0004019796180000095
Guidance law of attack angle alpha r Conversion to approximately equivalent pitch angle command +.>
Figure BDA0004019796180000098
Figure BDA0004019796180000096
For yaw path, yaw command is set to be ψ r ＝ψ v I.e. the yaw angle tracks the ballistic yaw angle. When the high-precision instruction is tracked, beta zero control can be basically realized, so that the sideslip angle instruction is set to be beta r =0 to achieve yaw angle control ψ r ＝ψ v 。
For a roll channel, roll angle tracking is set to γ r ＝0。
Therefore, the three-channel guidance instruction is set as:
Figure BDA0004019796180000101
step two: designing a six-degree-of-freedom dynamic model of the RLV reentry flight process, fully considering air density, flight speed and dynamic characteristic change under cross-domain maneuver, strong coupling effect among channels, unmodeled dynamic, uncertainty and disturbance such as fuel consumption, large-amplitude shaking, aircraft structural change and the like, and establishing an RLV attitude control model under the disturbance and uncertainty;
the built RLV six-degree-of-freedom model is as follows:
kinetic equation of RLV centroid movement:
Figure BDA0004019796180000102
the kinetic equation for the rotation of the RLV around the centroid is:
Figure BDA0004019796180000103
kinematic equation of RLV centroid motion:
Figure BDA0004019796180000104
kinematic equation for RLV rotation around centroid:
Figure BDA0004019796180000111
wherein m is RLV mass, g is gravitational acceleration, alpha and beta are attack angle and sideslip angle respectively,
Figure BDA0004019796180000119
psi, gamma are pitch angle, yaw angle and roll angle, respectively, theta, psi V ,γ V Respectively, are the inclination angle of the trajectory and the bulletTrack deflection angle and roll angle, V is displacement velocity, x, y, z is position coordinate, ω x ,ω y ,ω z For angular velocity, J x ,J y ,J z For moment of inertia, M x ,M y ,M z The force moment vector is respectively the component of the external moment vector on each axis of the projectile body coordinate system, and X, Y and Z are respectively resistance, lifting force and side force. The pneumatic coefficient formula of the disclosed winsed-Cone model is adopted, and comprises the following steps:
Figure BDA0004019796180000112
Figure BDA0004019796180000113
wherein ,
Figure BDA0004019796180000114
is dynamic pressure (ρ is air density at the flying height of RLV), S is the characteristic area of RLV, L b ,L c The lateral and longitudinal reference lengths of the RLV, respectively. c x ,c y ,c z Respectively a drag coefficient, a lift coefficient and a side force coefficient, m x ,m y ,m z Respectively representing a rolling moment coefficient, a yaw moment coefficient and a pitching moment coefficient. All the 6 pneumatic coefficients can be calculated in real time by the published Winged-Cone model data.
Taking into account the inherent characteristics of RLV and various uncertainties and disturbances of return flight, in (6)
Figure BDA00040197961800001110
Performing secondary derivation, and establishing the following attitude control model:
Figure BDA0004019796180000115
Figure BDA0004019796180000116
Figure BDA0004019796180000117
wherein ,
Figure BDA0004019796180000118
b f ,b p ,b r control gain representing pitch, yaw, roll channel,/->
Figure BDA0004019796180000121
Major moment coefficient components of pitch, yaw and roll channels, respectively, +.>
Figure BDA0004019796180000122
Is static derivative, is related to attack angle and Mach number, delta z ,δ y ,δ x Mathematical rudders to be designed for pitch, yaw and roll channels. Furthermore, the state information in the pitch, yaw and roll channels, except for rudder performance, is defined as disturbance f 1 ,f 2 ,f 3 And f 1 =f 01 +d 1 ,f 2 =f 02 +d 2 ,f 3 =f 03 +d 3 ，f 01 ,f 02 ,f 03 D is modeled dynamics, d 1 ,d 2 ,d 3 Including residual unmodeled dynamics, uncertainty, and unknown external disturbances in pitch, yaw, and roll channels:
Figure BDA0004019796180000123
it can be seen that the modelable disturbance f of three channels pitch, yaw, roll 01 ,f 02 ,f 03 Including inter-channel coupling, pneumatic parameter uncertainty, etc. Uncertainty disturbance f 1 ,f 2 ,f 3 And system gain b f ,b p ,b r Is to be controlled by RLV three-channel gestureSolves the key problems.
Step three: for effectively processing the coupling effect among channels, unmodeled dynamics and unknown external disturbance in RLV attitude control, ESO is adopted to estimate the integrated total disturbance of pitch, roll and yaw channels in real time, and feedback compensation is carried out;
to weaken f 1 ,f 2 ,f 3 Inter-channel coupling (aerodynamic coupling, inertial coupling, steering amount coupling), parameter and model uncertainty, other unknown external disturbances, etc., utilizing angular velocity information ω of pitch, yaw and roll channels, respectively z ,ω y ,ω x ESO is designed to estimate the total disturbance in real time. Based on (7), the ESO for pitch, yaw and roll channel designs are:
Figure BDA0004019796180000131
Figure BDA0004019796180000132
Figure BDA0004019796180000133
wherein ,
Figure BDA0004019796180000134
estimated values of angular velocities of pitch, yaw and roll channels, respectively, i.e. +.>
Figure BDA0004019796180000139
z ψ1 ≈ω y ,z γ1 ≈ω x ，/>
Figure BDA0004019796180000135
Estimated total disturbance values for pitch, yaw and roll channels, respectively, i.e
Figure BDA0004019796180000138
z ψ2 ≈f 2 +(b p -b p0 )δ y ，z γ2 ≈f 3 +(b r -b r0 )δ x 。b f0 ,b p0 ,b r0 B is f ,b p ,b r Is set to a nominal value. />
To ensure observer convergence, ESO poles are placed at- ω oz ,-ω oy ,-ω ox Where omega oz ,ω oy ,ω ox Pitch, yaw and roll channel bandwidths, respectively. At this time six observer gains l z1 ,l z2 ,l y1 ,l y2 ,l x1 ,l x2 Satisfy the following requirements
Figure BDA0004019796180000136
Considering continuous change of system gain caused by coupling action among channels under RLV cross-domain maneuvering, uncertain aerodynamic parameters and the like, the traditional single control gain method is difficult to obtain a satisfactory attitude control effect. Although ADRC technology has strong robustness and immunity, the control gain b of three channels of pitch, yaw and roll f ,b p ,b r As a very critical parameter in ADRC, not only the compensation accuracy of the disturbance but also the stability margin of the control system is directly affected. Thus, adaptive gain scheduling strategies are employed to compensate for disturbances and gains, respectively, in real time.
Step four: taking the continuous changes of system gain and uncertainty caused by dynamic characteristics, aerodynamic parameters and the like of the RLV under the cross-domain maneuver into consideration, classifying and identifying the disturbance and uncertainty of the RLV under the cross-domain maneuver by adopting a symbol estimator, and performing respective compensation based on a nominal controller to realize the precise attitude control of the RLV reentry flight;
adopting a symbol estimator to analyze, estimate and compensate disturbance and uncertainty, and establishing a loss function as
Figure BDA0004019796180000137
Figure BDA0004019796180000141
Figure BDA0004019796180000142
wherein ,δbf ,δ bp ,δ br Pitch, yaw and roll channel system gains b, respectively f ,b p ,b r Is used to determine the uncertainty of the estimate of (c),
Figure BDA0004019796180000149
f ψ ,f r is disturbance and uncertainty independent of rudder efficiency. To achieve minimization of the loss function, the symbolic projection gradient strategy is used to solve (11) - (13), resulting in:
Figure BDA0004019796180000143
Figure BDA0004019796180000144
Figure BDA0004019796180000145
wherein ,δbf ,δ bp ,δ br Estimated values of pitch, yaw and roll channel rudder uncertainty,
Figure BDA00040197961800001410
f ψ ,f γ estimated values of steering independent disturbances and uncertainties in pitch, yaw and roll channels, alpha 1 ,α 2 ,α 3 And > 0 is the update constant to be designed for pitch, yaw and roll channels, respectively.
Based on the estimation results (14) - (16) and the three-channel attitude angle and angular rate information, the mathematical rudder control laws of the pitch, yaw and roll channels are designed as follows:
Figure BDA0004019796180000146
/>
Figure BDA0004019796180000147
Figure BDA0004019796180000148
wherein ,kpf ,k pp ,k pr Proportional feedback gains, k, for pitch, yaw and roll channels, respectively df ,k dp ,k dr For differential feedback gain of three channels, the design converts the three channels into a series integration form that is easy to control by estimating, compensating for the disturbance and controlling the gain, respectively.
In the RLV pneumatic deceleration stage, only the grid rudder works, and at the moment, the grid rudder delta is used x ,δ y ,δ z The equivalent grid rudders of the obtained pitching, yawing and rolling channels are
Figure BDA0004019796180000151
The three-channel equivalent grid rudder (20) acts on the six-degree-of-freedom model of the RLV, so that the precise tracking guidance law of pitch angle, sideslip angle and roll angle can be realized, namely
Figure BDA0004019796180000152
ψ＝ψ r ＝ψ V (β＝0),γ＝γ r ＝0。
Therefore, the RLV attitude control is performed based on adaptive gain scheduling control laws (17) - (19) of pitch, yaw and roll channels, and adaptive scheduling of system gain in a cross-domain maneuver can be realized, thereby improving the attitude control performance of the cross-domain maneuver.
The invention comprehensively considers wide cross-domain airspace, wide speed domain, large air density, high flying speed and dynamic characteristic change in the RLV reentry flight, serious coupling effect among channels, fuel consumption, great shaking, structural change of an aircraft and other unmodeled dynamic, uncertainty and disturbance, proposes a self-adaptive gain scheduling control strategy, adopts a symbol projection gradient strategy to adaptively schedule the system gain based on the active disturbance rejection controller, does not depend on RLV characteristic point extraction, has simple structure and easy design, and can effectively treat the uncertainty and disturbance of the RLV cross-domain maneuvering process, thereby improving the gesture control precision.
In order to verify the effectiveness of the self-adaptive gain scheduling control strategy provided by the invention, an RLV six-degree-of-freedom model is built by MATLAB software, and the self-adaptive gain scheduling control strategy is designed based on the attitude control model, so that the effectiveness of the self-adaptive gain scheduling control strategy in the aspects of disturbance and uncertainty treatment and the strong robustness and adaptability in the aspects of improving the attitude control precision are verified. During simulation, the data of the relevant parameters are as follows:
t 0 ＝0,t 1 ＝60,t 2 ＝120,h 01 ＝25000,h 02 =1000, height error h 1 ＝500,h 2 =50, scaling factor k pα Differential coefficient k=0.2 dα =1, initial value of attack angle α 0 =5. The total weight of the RLV aircraft is 136080kg, the longitudinal reference length is 24.384m, the lateral reference length is 18.288m, and the reference area is 334.73m 2 The initial moment of inertia in the x direction is 1355818 kg.m 2 Initial moment of inertia in y and z directions is 13558180 kg.m 2 。
ω of ＝10,b f0 ＝2.45，ω op ＝10,b p0 ＝0.5，ω or ＝10,b r0 ＝11.18。δ bf ,δ bp ,δ br And
Figure BDA0004019796180000161
f ψ ,f γ the initial values of (a) are all set to 0, alpha 1 ,α 2 ,α 3 ＝0.001。k pf ＝42,k df ＝20，k pp ＝0.5,k dp ＝20，k pr ＝62,k dr =38. The initial value of the control attitude angle is [5 0.1]Degree, initial value of position [0 25000 0 ]]m, the initial speed value is 4Ma, and the initial angular speed value is 0.
The simulation results are shown in fig. 2-6, and are divided into two parts:
a first part: guidance law
According to the reference value of the longitudinal section of the RLV re-entry atmospheric flight, the attack angle is designed, so that the pitch angle tracking control is performed according to the geometric relationship, the obtained reference value and the actual value of the longitudinal section are shown in figure 2, and the control strategy provided by the invention can well track the longitudinal section and has good tracking performance. Meanwhile, the RLV flight speed is shown in figure 3, so that the attenuation of the RLV flight speed in the air re-entry flight process is realized, the effectiveness of the guidance law adopted by the invention is further illustrated, and a good foundation is laid for realizing the precise control of the attitude.
A second part: attitude control
The invention provides an RLV cross-domain maneuver by adopting a self-adaptive gain scheduling strategy, wherein in simulation results, figure 4 shows a control result of an attitude angle, figure 5 shows a control result of three angular speeds, and figure 6 is a control mathematical rudder designed based on the invention. The result shows that the invention can effectively treat the uncertainty of coupling, pneumatic parameter uncertainty, air density change and the like among channels; the method realizes stable and accurate tracking of guidance laws, and proves the effectiveness and strong robustness and adaptability of the proposed algorithm.
While the fundamental and principal features of the invention and advantages of the invention have been shown and described, it will be apparent to those skilled in the art that the invention is not limited to the details of the foregoing exemplary embodiments, but may be embodied in other specific forms without departing from the spirit or essential characteristics thereof; the present embodiments are, therefore, to be considered in all respects as illustrative and not restrictive, the scope of the invention being indicated by the appended claims rather than by the foregoing description, and all changes which come within the meaning and range of equivalency of the claims are therefore intended to be embraced therein.
Furthermore, it should be understood that although the present disclosure describes embodiments, not every embodiment is provided with a separate embodiment, and that this description is provided for clarity only, and that the disclosure is not limited to the embodiments described in detail below, and that the embodiments described in the examples may be combined as appropriate to form other embodiments that will be apparent to those skilled in the art.

Claims (1)

1. A method for adaptive gain scheduling control of a reusable carrier, comprising the steps of:
s1, designing a guidance law for the RLV flight attitude according to the track change requirement of the RLV reentry flight process so as to generate a corresponding attitude tracking instruction; the guidance instructions of the three channels of pitching, rolling and yawing are as follows:
Figure FDA0004019796170000011
θ is the ballistic dip angle, α r For angle of attack guidance law, θ r Is pitch angle command, beta r For sideslip angle command, gamma r Is a roll angle command; set sideslip angle command beta r =0 for realizing yaw angle control ψ r ＝ψ v ，ψ r Is yaw angle, psi v Is the ballistic deflection angle;
s2, designing a six-degree-of-freedom dynamic model of the RLV reentry flight process, and establishing an RLV attitude control model under disturbance and uncertainty by considering unmodeled dynamics, uncertainty and disturbance including air density, flight speed and dynamic characteristic changes under cross-domain maneuver, coupling effect among channels, fuel consumption and large-scale shaking and aircraft structural changes;
the six-degree-of-freedom dynamic model is:
kinetic equation of RLV centroid movement:
Figure FDA0004019796170000012
the kinetic equation for the rotation of the RLV around the centroid is:
Figure FDA0004019796170000013
kinematic equation of RLV centroid motion:
Figure FDA0004019796170000021
kinematic equation for RLV rotation about centroid
Figure FDA0004019796170000022
Wherein m is RLV mass, g is gravitational acceleration, alpha and beta are attack angle and sideslip angle respectively,
Figure FDA0004019796170000023
respectively pitch angle, yaw angle and roll angle, theta, phi V ，γ V Respectively, the ballistic inclination angle, the ballistic deflection angle and the roll angle, V is the displacement speed, x, y and z are the position coordinates, omega x ，ω y ，ω z For angular velocity, J x ，J y ，J z For moment of inertia, M x ，M y ，M z The force moment vector is respectively a component of the external moment vector on each axis of the projectile body coordinate system, and X, Y and Z are respectively resistance, lifting force and side force;
Figure FDA0004019796170000024
Figure FDA0004019796170000025
wherein ,
Figure FDA0004019796170000026
is dynamic pressure, ρ is air density of the flying height of RLV, S is characteristic area of RLV, L b ，L c The lateral and longitudinal reference lengths, c, of the RLV, respectively x ，c y ，c z Respectively a drag coefficient, a lift coefficient and a side force coefficient, m x ，m y ，m z Respectively representing a rolling moment coefficient, a yaw moment coefficient and a pitching moment coefficient;
taking into account the RLV characteristics and various uncertainties and disturbances of the return flight, for
Figure FDA0004019796170000027
And (3) carrying out secondary derivation on the psi and gamma, and establishing a posture control model as follows:
Figure FDA0004019796170000031
Figure FDA0004019796170000032
Figure FDA0004019796170000033
wherein ,
Figure FDA00040197961700000311
b f ，b p ，b r control gain representing pitch, yaw, roll channel,/->
Figure FDA0004019796170000035
Figure FDA0004019796170000036
Moment coefficient components of pitch, yaw and roll channels, respectively, +.>
Figure FDA0004019796170000037
Is the static derivative, delta z ，δ y ，δ x Mathematical rudders to be designed for pitch, yaw and roll channels;
the state information in the pitch, yaw and roll channels other than rudder effectiveness is defined as disturbance f 1 ，f 2 ，f 3 And f 1 ＝f 01 +d 1 ，f 2 ＝f 02 +d 2 ，f 3 ＝f 03 +d 3 ，f 01 ，f 02 ，f 03 D is modeled dynamics, d 1 ，d 2 ，d 3 Including remaining unmodeled dynamics, uncertainty, and unknown external disturbances in the pitch, yaw, and roll channels;
Figure FDA0004019796170000038
s3, estimating integrated total disturbance of three channels of pitching, rolling and yawing in real time by adopting ESO and performing feedback compensation;
using angular velocity information omega for three channels, pitch, yaw and roll, respectively z ，ω y ，ω x Designing ESO to estimate total disturbance in real time, wherein ESO designed for a pitch channel, a yaw channel and a roll channel are respectively
Figure FDA0004019796170000039
Figure FDA00040197961700000310
Figure FDA0004019796170000041
wherein ,
Figure FDA0004019796170000042
estimated values of angular velocities of pitch, yaw and roll channels, respectively, i.e
Figure FDA0004019796170000043
Estimated total disturbance of pitch, yaw and roll channels, respectively, i.e. +.>
Figure FDA0004019796170000044
z ψ2 ≈f 2 +(b p -b p0 )δ y ，z γ2 ≈f 3 +(b r -b r0 )δ x ，b f0 ，b p0 ，b r0 B is f ，b p ，b r Is determined by the set of parameters;
placing ESO poles at-omega oz ，-ω oy ，-ω ox Where omega oz ，ω oy ，ω ox Six observer gains l for pitch, yaw and roll channel bandwidths, respectively z1 ，l z2 ，l y1 ，l y2 ，l x1 ，l x2 Satisfy l z1 ＝2ω oz ，
Figure FDA0004019796170000045
l y1 ＝2ω oy ，/>
Figure FDA0004019796170000046
l x1 ＝2ω ox ，/>
Figure FDA0004019796170000047
S4, considering continuous changes of system gain and uncertainty caused by dynamic characteristics and pneumatic parameters of the RLV under the cross-domain maneuver, classifying and identifying disturbance and uncertainty of the RLV under the cross-domain maneuver by adopting a symbol estimator, and performing respective compensation based on a nominal controller to realize accurate attitude control of the RLV reentry flight;
and adopting a symbol estimator to carry out analysis estimation and compensation on disturbance and uncertainty, wherein the established loss function is as follows:
Figure FDA0004019796170000048
Figure FDA0004019796170000049
Figure FDA00040197961700000410
wherein ,δbf ，δ bp ，δ br Pitch, yaw and roll channel system gains b, respectively f ，b p ，b r Is used to determine the uncertainty of the estimate of (c),
Figure FDA00040197961700000411
disturbance and uncertainty which are irrelevant to rudder efficiency; solving the established loss function by adopting a symbol projection gradient strategy to obtain:
Figure FDA00040197961700000412
Figure FDA00040197961700000413
Figure FDA0004019796170000051
wherein ,δbf ，δ bp ，δ br Estimated values of pitch, yaw and roll channel rudder uncertainty,
Figure FDA0004019796170000052
estimated values of steering independent disturbances and uncertainties in pitch, yaw and roll channels, alpha 1 ，α 2 ，α 3 >o is the update constant to be designed of the pitch, yaw and roll channels respectively;
based on the estimation results and the attitude angle and angular rate information of the three channels, the mathematical rudder control laws of the pitch, yaw and roll channels are designed as follows:
Figure FDA0004019796170000053
Figure FDA0004019796170000054
Figure FDA0004019796170000055
wherein ,kpf ，k pp ，k pr Proportional feedback gains, k, for pitch, yaw and roll channels, respectively df ，k dp ，k dr Differential feedback gain for three channels;
by a mathematical rudder delta x ，δ y ，δ z The equivalent grid rudders for three channels of pitch, yaw and roll are obtained as follows:
Figure FDA0004019796170000056
three-channel equivalent grid rudder is applied to a six-degree-of-freedom model of the RLV to realize pitch angle, sideslip angle and rollPrecise tracking of angle guidance law, i.e.
Figure FDA0004019796170000057
And performing attitude control of the RLV cross-domain maneuver. />
CN202211683045.4A 2022-12-27 2022-12-27 Adaptive gain scheduling control method for reusable carrier Pending CN116203840A (en)

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Publication number Priority date Publication date Assignee Title
CN117192726A (en) * 2023-09-07 2023-12-08 山东科技大学 Quick reflector control method and device based on improved active disturbance rejection control

Cited By (2)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN117192726A (en) * 2023-09-07 2023-12-08 山东科技大学 Quick reflector control method and device based on improved active disturbance rejection control
CN117192726B (en) * 2023-09-07 2024-03-15 山东科技大学 Quick reflector control method and device based on improved active disturbance rejection control

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