CN114578857A - Guidance aircraft autonomous control method, device and system based on full trajectory information - Google Patents
Guidance aircraft autonomous control method, device and system based on full trajectory information Download PDFInfo
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Abstract
The invention discloses a method, a device and a system for autonomously controlling a guided vehicle based on full trajectory information. Wherein, the method comprises the following steps: performing auxiliary calibration on an inertial sensor by using ballistic information in an ascending phase of the aircraft, and performing combined navigation with satellite information, geomagnetic information and the inertial sensor, wherein the ascending phase is a phase from the launching point to a ballistic vertex; and performing trajectory prediction and drop point evaluation according to the measured data in an aircraft descending stage, and performing composite control on the aircraft to control the aircraft to hit a target, wherein the descending stage is a stage from the trajectory vertex to the drop point. The invention solves the technical problem of low navigation precision and hit precision of the aircraft in the related technology.
Description
Technical Field
The invention relates to the field of navigation, in particular to a guidance aircraft autonomous control method, device and system based on full trajectory information.
Background
In the related art, the control method of the guidance aircraft only depending on inertia, geomagnetic/inertia information combination navigation cannot measure, compensate and calibrate sensor errors, so that the sensor errors are large, and the navigation precision is low.
And for the control method of the guidance aircraft which depends on the satellite/geomagnetic/inertia information combination navigation, the control method is easily influenced by satellite interference, and particularly the control precision is low and the hit precision is poor in the descending stage of the aircraft when the aircraft enters the upper space of a target range.
The method can not bind trajectory information in advance according to tactical requirements, relative positions between the target and meteorological conditions of the flying field, control the whole process of the aircraft by using the trajectory information in the flying process, and perform autonomous navigation control according to drop point prediction.
Disclosure of Invention
The embodiment of the invention provides a method, a device and a system for autonomously controlling a guided aircraft based on full trajectory information, which are used for at least solving the technical problem that the navigation precision and the hit precision of the guided aircraft are not high in the related technology.
According to one aspect of the embodiment of the invention, a guidance aircraft autonomous control method based on full ballistic information is provided, and comprises the following steps: performing auxiliary calibration on an inertial sensor by using ballistic information in an ascending phase of the aircraft, and performing combined navigation with satellite information, geomagnetic information and the inertial sensor, wherein the ascending phase is a phase from the launching point to a ballistic vertex; and in the aircraft descending stage, performing trajectory prediction and drop point evaluation according to the measured data, and performing composite control on the aircraft to control the aircraft to hit a target, wherein the descending stage is a stage from the trajectory peak to the drop point.
According to another aspect of the embodiment of the invention, there is also provided a guidance aircraft autonomous control device based on full ballistic information, including: an ascent stage control module configured to perform auxiliary calibration on an inertial sensor using ballistic information in an ascent stage of the aircraft, the ascent stage being a stage from the launch point to a ballistic vertex, and performing combined navigation with satellite information, geomagnetic information, and the inertial sensor; and the descending stage control module is configured to perform ballistic prediction and drop point evaluation according to the measured data in an aircraft descending stage, and perform composite control on the aircraft to control the aircraft to hit a target, wherein the descending stage is a stage from the ballistic vertex to the drop point.
In the embodiment of the invention, the whole process of the movement of the aircraft is controlled by using the trajectory information of the whole flying process of the aircraft, and the combined navigation is carried out with the satellite information, the geomagnetic information and the inertia information, so that the technical problems of low navigation precision and low hit precision of the aircraft in the related technology are solved, the navigation precision and the hit precision of the aircraft are further improved, and the autonomy of the aircraft is enhanced.
Drawings
The accompanying drawings, which are included to provide a further understanding of the invention and are incorporated in and constitute a part of this application, illustrate embodiment(s) of the invention and together with the description serve to explain the invention without limiting the invention. In the drawings:
fig. 1 is a flowchart of a method for autonomously controlling a guided vehicle based on full ballistic information according to a first embodiment of the present invention.
FIG. 2A is a schematic illustration of stages in a process for autonomous control of a guided vehicle based on full ballistic information in accordance with an embodiment of the invention.
Fig. 2B is a flowchart of a guided vehicle autonomous control method based on full ballistic information according to a second embodiment of the present invention.
Figure 2C is a schematic diagram of a ballistic model according to an embodiment of the invention.
Fig. 3 is a flowchart of a method for autonomously controlling a guided vehicle based on full ballistic information according to a third embodiment of the present invention.
Fig. 4 is a flowchart of a method for autonomously controlling a guided vehicle based on full ballistic information according to a fourth embodiment of the present invention.
Fig. 5 is a schematic structural diagram of a guided vehicle autonomous control device based on full ballistic information according to an embodiment of the invention.
Fig. 6 is a schematic structural diagram of another autonomous control device of the guided vehicle based on the full ballistic information according to the embodiment of the invention.
Detailed Description
In order to make the technical solutions of the present invention better understood, the technical solutions in the embodiments of the present invention will be clearly and completely described below with reference to the drawings in the embodiments of the present invention, and it is obvious that the described embodiments are only a part of the embodiments of the present invention, and not all of the embodiments. All other embodiments, which can be derived by a person skilled in the art from the embodiments given herein without making any creative effort, shall fall within the protection scope of the present invention.
It should be noted that the terms "first," "second," and the like in the description and claims of the present invention and in the drawings described above are used for distinguishing between similar elements and not necessarily for describing a particular sequential or chronological order. It is to be understood that the data so used is interchangeable under appropriate circumstances such that the embodiments of the invention described herein are capable of operation in sequences other than those illustrated or described herein. Furthermore, the terms "comprises," "comprising," and "having," and any variations thereof, are intended to cover a non-exclusive inclusion, such that a process, method, system, article, or apparatus that comprises a list of steps or elements is not necessarily limited to those steps or elements expressly listed, but may include other steps or elements not expressly listed or inherent to such process, method, article, or apparatus.
Example 1
According to an embodiment of the invention, a method for autonomously controlling a guided vehicle based on full ballistic information is provided, as shown in fig. 1, and comprises the following steps:
step S102, in the ascending stage of the aircraft, performing auxiliary calibration on an inertial sensor by using trajectory information, and performing combined navigation with satellite information, geomagnetic information and the inertial sensor, wherein the ascending stage is a stage from the launching point to a trajectory peak;
and S104, performing trajectory prediction and drop point evaluation according to the measured data in the aircraft descending stage, and performing composite control on the aircraft to control the aircraft to hit a target, wherein the descending stage is a stage from the trajectory vertex to the drop point.
In one exemplary embodiment, the method further comprises: before the aircraft is launched, trajectory information is bound in advance according to tactical requirements, relative positions with targets and meteorological conditions in the flying field.
In one exemplary embodiment, the auxiliary calibration of the inertial sensor using ballistic information, combined with satellite information, geomagnetic information, and inertial sensor navigation, comprises: under the condition that satellite signals are not captured, constructing a Kalman filter based on a trajectory model and an attitude estimation algorithm of geomagnetic information and comprehensively considering the attitude motion characteristics of a projectile body and geomagnetic field measurement information to estimate the attitude, the attitude change speed and the attitude angular acceleration information of the projectile body so as to generate continuous navigation positioning information; under the condition that the satellite signal is captured, trajectory calculation data, geomagnetic calculation data and satellite calculation data are utilized to carry out trajectory information coarse alignment and satellite information accurate alignment, the satellite calculation data, inertial navigation calculation data, geomagnetic calculation data, trajectory calculation data and meteorological data are intelligently and autonomously fused to construct a Kalman filter, and the position, the speed and the posture of the projectile are measured to generate continuous navigation positioning information.
In an exemplary embodiment, a kalman filter is constructed based on a trajectory model and an attitude estimation algorithm of geomagnetic information and by comprehensively considering the attitude motion characteristics of a projectile and geomagnetic field measurement information, and comprises: measuring the rotating speed of the aircraft in real time through geomagnetic resolution, and carrying out real-time attitude estimation on a pitch angle and a roll angle of a projectile body by combining meteorological information, a trajectory model and geomagnetic information; selecting an estimated state vector by taking the pitch angle of the projectile body, the rolling angle, the pitch angle of the projectile body, the angular velocity of the rolling angle and the angular acceleration as estimators, and establishing a state equation; measuring geomagnetic information, and establishing a measurement equation of the system according to a conversion relation between the measured geomagnetic information in a launching coordinate system and a projectile coordinate system; and performing Taylor expansion on a nonlinear measurement equation based on the state equation and the measurement equation, neglecting high-order terms, and constructing extended Kalman filtering, wherein a measurement noise variance matrix in the measurement equation is estimated by the measurement compensation precision of a geomagnetic sensor.
In an exemplary embodiment, ballistic information coarse alignment and satellite information precise alignment are performed by using ballistic solution data, geomagnetic solution data and satellite solution data, and the satellite solution data, inertial navigation solution data, geomagnetic solution data, ballistic solution data and meteorological data are intelligently and autonomously fused to construct a kalman filter, which includes: selecting a plurality of state variables to construct a state equation, wherein the state variables comprise attitude angle errors, speed errors, position errors, accelerometer zero offset and gyroscope drift errors; selecting a speed measurement value, a position measurement value and an attitude measurement value of the aircraft as measurement values, wherein the position measurement value is a difference value between a position measured by the inertial navigation resolving module and an output position of the satellite resolving module; the velocity measurement value is the difference value between the velocity calculated by the inertial navigation resolving module and the output velocity of the inertial navigation resolving module; the attitude measurement value is a difference value between a geomagnetic vector obtained by the geomagnetic model and a geomagnetic vector obtained by the geomagnetic resolving module; constructing a measurement equation based on the measurement values; and constructing a Kalman filter based on the state equation and the measurement equation.
In one exemplary embodiment, a ballistic prediction and a landing point assessment are performed based on measured data to perform a composite control of an aircraft, comprising: performing trajectory prediction and drop point evaluation by using a navigation result obtained by fusing satellite solution data, trajectory solution data and inertial navigation solution data; controlling the aircraft to make a large angle attitude adjustment at the end of the aircraft descent phase based on the trajectory prediction and the drop point assessment so that the aircraft hits a target.
In one exemplary embodiment, performing the drop point assessment includes: comparing the model trajectory prediction with the binding trajectory information, and predicting a drop point.
The embodiment utilizes the trajectory information of the whole flying process of the aircraft to control the whole moving process of the aircraft, and the combined navigation with the satellite information, the geomagnetic information and the inertia information improves the navigation precision and the hit precision of the aircraft and enhances the autonomy of the aircraft.
Example 2
According to the embodiment of the invention, a guidance aircraft autonomous control method based on full trajectory information is provided.
In this embodiment, as shown in fig. 2A, the process of autonomous control of the aircraft may be divided into 5 phases, a phase of binding ballistic information before launching, a satellite signal capturing phase, a multi-source fusion combined navigation phase, a ballistic prediction landing point evaluation phase, and a pose adjustment guidance control phase.
As shown in fig. 2B, the method includes:
step S202, the pre-staple trajectory information is transmitted.
Before launching, the theoretical trajectory is accurately calculated and is rapidly bound wirelessly according to tactical requirements, relative positions of the target and meteorological conditions of the flying field.
And step S204, navigating the aircraft in the ascending stage without acquiring the satellite signals.
And after the aircraft is launched, the thermal battery of the aircraft is activated to establish voltage so as to supply power to electronic equipment such as an autonomous navigation control assembly on the aircraft.
In the flight process, the position equation of the center of mass of the projectile relative to a launching coordinate system OXYZ is established as follows:
wherein, the first and the second end of the pipe are connected with each other,
a transformation matrix from a transmitting coordinate system to a quasiplasty coordinate system; thetavIs trajectory pitch angle; psivIs the ballistic yaw angle; v is the initial test speed of projectile body launching, and the direction points to the quasi-projectile body coordinate system ObXbAnd (4) direction.
The equation of the projectile attitude change and the projectile attitude angle established under the projectile coordinate system is as follows:
wherein the content of the first and second substances,
respectively a roll angle, a yaw angle and a pitch angle under the projectile coordinate system; gamma and theta are respectively a rolling angle and a pitching angle of the projectile coordinate system relative to the launching coordinate system in the initial state;
is the variation of the projectile attitude.
According to the motion rule of the rotary aircraft, when an initial launching value condition of the aircraft is given, a ballistic position model and a projectile attitude model can be solved. The geomagnetic calculation module measures the rotating speed of the aircraft in real time, real-time attitude estimation is carried out on the pitch angle and the roll angle of the projectile body by combining meteorological information, a ballistic model shown in figure 2C and geomagnetic information, and meanwhile, self-checking and initial alignment are carried out by the inertial navigation calculation module.
By combining the flight characteristics of the ballistic model and the rotary aircraft, the yaw angle of the aircraft at this stage is basically unchanged and is approximately zero, the pitch angle is gradually reduced under the action of gravity, the roll rate is gradually attenuated, and the resolving attitude of the ballistic model is regarded as a quadratic curve of time change:
wherein a, b and c are a quadratic term, a primary term and a constant term coefficient corresponding to the change of the attitude angle. Roll rate compliance
Wherein, ω is0The self-rotation angular rate of the projectile body at the moment of launching is shown, t is launching time, L is the length of the projectile body, D is the diameter of the projectile body, A is the moment of inertia of the projectile body, and k is a correction coefficient.
Taking the trajectory calculation aircraft pitch angle, roll angle, angular velocity and angular acceleration as estimators, and selecting the estimation state vector as:
establishing a state equation:
wherein, the first and the second end of the pipe are connected with each other,
for the pitch rate of the projectile body,
is the second derivative of the pitch angle of the projectile,
as to the angular rate of roll of the projectile,
is the second derivative of the roll angle of the projectile,
in the form of a first differential of the state vector, phi (t) is a matrix consisting of the state transition matrices for pitch and roll angles, phiθIs long bentElevation state transition matrix, phiγIs a roll angle state transition matrix. X (t) is the state vector at time t, w (t) is the zero mean Gaussian white noise vector, and G (t) is the state transition matrix of the noise vector.
The discretization form is: xk＝Φk/k-1Xk-1+ηk-1. Wherein, XkIs a state vector in discrete form, phik/k-1To estimate the state transition matrix of the state vector at time k from time k-1, Xk-1Is the state vector at time k-1, ηk-1Is a discrete vector of the noise vector.
Obtaining system observed quantity H according to measured geomagnetic informationbAnd according to the conversion relation of the measured geomagnetic information between the launching coordinate system and the projectile coordinate system, the system measurement equation is expressed as follows:
wherein, the first and the second end of the pipe are connected with each other,
respectively are the measured values of the three-axis geomagnetic information in the carrier coordinate system,
respectively are the measured values of the geomagnetic information of the three axes in the emission coordinate system.
The discretization equation is:
wherein n iskIs a zero mean white Gaussian noise vector, f [ X ]k]Representing a function related to the system state vector.
According to the nonlinear system linearization theory, Taylor expansion is carried out on a nonlinear measurement equation, high-order terms are ignored, an extended Kalman filter is constructed, and the nonlinear measurement equation is linearized as follows:
wherein, JkFor a transformation matrix from the state vector to the observed values, HbFor systematic observation, XkIs a state vector.
Obtain the state vector XkThe extended kalman filter recursion equation for the estimator is:
wherein, EKF filtering initial value X0、P0Can be determined according to the initial value of the attitude and the transmitting data; the filtering system noise covariance matrix Q is obtained according to the projectile motion rule; the measurement noise covariance matrix R is estimated from the magnetic sensor measurement compensation accuracy,
is the system state vector at time k,
is the system state vector at time k-1,
is a state vector at time k, phi, estimated from time k-1k/k-1To estimate the state transition matrix for the state vector at time k based on time k-1,
for transposition of the state transition matrix, Pk/k-1For estimating the covariance matrix, P, a priori at time k based on the estimated time k-1k-1Estimating covariance matrix, P, for the posteriori at time k-1kEstimating a covariance matrix, Γ, for the posteriors at the k instantsk-1A state transition matrix that is a system noise covariance matrix,
transposing the state transition matrix, Q, for the covariance matrix of the system noisek-1Is the covariance matrix of the system noise at time K-1, KkFilter gain at time k, JkFor the transition matrix from the state vector to the observation at time k,
for the transposition of the transformation matrix from the state vector to the observation at time k, RkThe covariance matrix of the noise is measured for time k,
for the system observed quantity under the carrier coordinate system at the moment k,
is a transformation matrix from the transmission coordinate system to the carrier coordinate system,
and I is an identity matrix, namely the observed quantity of the system under the emission coordinate system at the moment k.
The EKF attitude estimation algorithm based on the ballistic model and the geomagnetic information comprehensively considers the attitude motion characteristics of the projectile body and the geomagnetic field measurement information, can simultaneously estimate the attitude, the attitude change speed and the attitude angular acceleration information of the projectile body, and meets the condition that the yaw angle is approximately zero in the initial flight stage.
Step S206, the aircraft is navigated during the ascent phase of the acquisition of the satellite signals.
After the satellite receiver searches and captures the satellite signal, the satellite navigation positioning is realized; after the inertia calculation module performs self-checking, the inertia calculation module can comprehensively utilize the ballistic calculation module data, the geomagnetic calculation module data and the satellite calculation module data to perform ballistic information coarse alignment and satellite information accurate alignment.
According to the satellite resolving module information, the trajectory resolving module information and the inertial navigation resolving module information, based on sensor availability and information credibility analysis, intelligent and autonomous fusion of information such as satellite + inertial navigation/geomagnetism/trajectory/meteorology is achieved, a Kalman filter is built, information such as the position, speed and attitude of a projectile is measured, high-precision continuous available credible navigation positioning information is generated, and the influence of satellite navigation signal interference and deception is effectively inhibited.
Specifically, the attitude angle error phi is selectedE,φN,φUVelocity error δ VE,δVD,δVUPosition errors delta L, delta B, delta H, accelerometer zero offset
Drift error of gyroscope
And 15 state variables in total, wherein the state variables satisfy the following conditions:
the state equation is:
wherein the content of the first and second substances,
F1is a system matrix composed of an attitude angle error matrix, a speed error matrix and a position error matrix,
FIMU＝diag[-σx -σy -σz -βx -βy -βy]，
sigma, beta are the reciprocal of the time associated with the accelerometer and gyroscope, respectively,
in the form of a first differential of the state vector, X (t) being the state vector at time t, G (t)Is a matrix composed of a state transition matrix of state vectors, W (t) is a zero-mean Gaussian white noise vector, and G (t) is a state transition matrix of noise vectors.
Selecting speed, position and attitude as system observed values, wherein the position observed values are the difference values of the positions measured by the inertial navigation resolving module and the output positions of the satellite resolving module; the velocity measurement value is the difference value between the velocity calculated by the inertial navigation resolving module and the output velocity of the inertial navigation resolving module; the attitude measurement value is a difference value between a geomagnetic vector obtained by the geomagnetic model and a geomagnetic vector obtained by the geomagnetic calculation module.
The positions output by the inertial navigation resolving module and the satellite resolving module are respectively as follows:
wherein L, B and H are longitude, latitude and height positions of the projectile body, and N isN,NE,NDPosition error, L, output for satellite resolver moduleINSMeasuring the longitude of the projectile body for an inertial navigation resolving moduleINSMeasuring projectile latitude H for inertial navigation resolving moduleINSMeasuring projectile height, L, for inertial navigation resolving moduleGNSSMeasuring the longitude, B of the projectile for the satellite resolver moduleGNSSMeasuring the latitude, H, of the projectile for a satellite resolver moduleGNSSMeasuring the projectile body height for a satellite resolving module, wherein delta L is a longitude position error, delta B is a latitude position error, delta H is a height position error, and R isMh、RNhRespectively the radius of curvature of the meridian circle and the radius of curvature of the unitary mortise circle.
The speeds output by the inertial navigation resolving module and the satellite resolving module are respectively as follows:
wherein, VN,VE,VDAs the true velocity of the projectile, MN,ME,MDVelocity error, V, output for satellite resolver moduleNINSMeasuring projectile north direction velocity, V, for inertial navigation resolving moduleEINSOriental velocity, V, of projectile measured by inertial navigation resolving moduleUINSMeasuring the projectile upward velocity, V, for an inertial navigation solution moduleNGNSSMeasuring projectile directional velocity, V, for a satellite resolver moduleEGNSSMeasuring projectile directional velocity, V, for satellite solver moduleUGNSSAnd delta is the direction speed of the projectile body measured by the satellite resolving module.
The position measurement equation is:
wherein HpFor the position state transition matrix, Hp＝[0 diag(RMH,RNHcosL,1) 0]，ZpAs a result of observation of the position of the projectile, XpIs a system position state vector, VpFor vectors consisting of position errors output by the satellite solver module, Vp＝[NN NEND]T。
The velocity measurement equation is:
wherein HvFor the speed state transition matrix, Hv＝[0 diag(1,1,1) 0]，ZvAs an observation of projectile velocity, XvIs a system velocity state vector, VpFor vectors consisting of speed errors output by the satellite resolver module, Vp＝[MN ME MD]T。
The attitude measurement equation is:
wherein the content of the first and second substances,
for geomagnetic data obtained from an earth magnetic field model, Ax，Ay，AzAnd obtaining a triaxial geomagnetic measurement value under a navigation coordinate system through a conversion matrix according to the output of the geomagnetic solution module. Z is a linear or branched memberMAnd the observed value of the projectile attitude is obtained.
The overall observation equation of the system is:
wherein Z is a system observed value, ZvObservation of velocity of projectile, ZpObservation of the position of the projectile, ZMAnd the observed value of the projectile attitude is obtained.
And finally, constructing an extended Kalman filtering algorithm:
Pk,k-1＝fk,k-1Pk-1fT k,k-1+Qk-1 (16)
Pk＝(I-KkJk)Pk,k-1 (19)
in the formula (I), the compound is shown in the specification,
is an a posteriori system state vector estimate at time k,
is an a posteriori system state vector estimate at time k-1,
is a prior state vector at the k time, P, estimated from the k-1 timek/k-1For estimating the covariance matrix, P, a priori at time k based on the estimated time k-1k-1Estimating covariance matrix, P, for the posteriori at time k-1kEstimating the covariance matrix for the posteriori at time k, fk,k-1To estimate the state transition matrix of the state vector at time k from time k-1, fT k,k-1For transposing state transition matrices, QkCovariance, Q, of system noise at time kk-1Is the covariance of the system noise at time k-1, RkMeasuring the covariance of the noise for time k, Rk-1Covariance of the measured noise at time K-1, KkFilter gain at time k, JkFor the transition matrix from the state vector to the observation at time k,
transposing the transformation matrix from the state vector to the observation for time k, ZkFor the system observation at time k,
i is an identity matrix which is an observed value of a k-1 time system obtained according to k time.
Step S208, the aircraft is navigated in the descent phase.
And the aircraft enters a descent stage after passing through a trajectory vertex, trajectory prediction and drop point evaluation are rapidly performed by using a navigation result realized by fusion of information based on a satellite calculation module, information based on a trajectory calculation module and information based on an inertial navigation calculation module, so that autonomous composite control and aircraft pose adjustment in the descent stage are guided.
And step S210, controlling the aircraft to hit the target.
The autonomous navigation control assembly controls the guided aircraft, and the position and the attitude of the guided aircraft are adjusted by controlling an actuating mechanism such as a pulse engine and the like to control the landing point of the guided aircraft. By utilizing the fusion navigation information, the guidance aircraft is controlled by the autonomous navigation control component to carry out large-angle attitude adjustment in the tail stage of the aircraft descent, so that the accurate target and accurate attack capability of the guidance aircraft is greatly enhanced, and the target accuracy under the condition of complex electromagnetic interference of the guidance aircraft is improved.
The embodiment solves the following technical problems:
1. before launching, binding ballistic trajectory information in advance according to tactical requirements, relative positions between the target and meteorological conditions of a flying field, and accurately calculating the whole real-time ballistic trajectory in the flying process.
2. In the ascending stage of the aircraft, the inertial sensor is calibrated in an auxiliary mode by using trajectory information, and combined navigation is carried out on the inertial sensor, satellite information, geomagnetic information and the inertial sensor.
3. And in the descending stage of the aircraft, trajectory prediction and drop point evaluation are rapidly performed according to the measured data, the aircraft is subjected to coincidence composite control, and the target hit precision is improved.
According to the embodiment, the whole process of the movement of the aircraft is controlled by using the trajectory information of the whole flying process of the aircraft, and the combined navigation with the satellite information, the geomagnetic information and the inertia information improves the navigation precision and the hit precision of the aircraft, and enhances the autonomy of the aircraft.
Example 3
According to the embodiment of the invention, a method for autonomously controlling a guided vehicle based on full ballistic information is provided, and as shown in fig. 3, the method for controlling the full ballistic guided vehicle, which improves hit precision, provided by the embodiment of the invention comprises the following steps:
and S302, performing error compensation and calibration on other sensors by using the ballistic information.
Before the guided aircraft is launched, ballistic trajectory information is bound in advance according to tactical requirements, relative positions between the guided aircraft and targets and meteorological conditions of a flying field, and the whole-course real-time ballistic trajectory can be accurately calculated in the flying process.
In the ascending stage, after the satellite captures and finishes positioning, the guidance aircraft and the satellite information can carry out auxiliary alignment on the inertial sensor together to measure the position, the speed, the attitude and other information of the projectile body.
And S304, carrying out accurate navigation on the aircraft by utilizing the fusion of the ballistic information and the combined navigation information.
According to the satellite navigation information, the real-time trajectory information and the projectile body pose measurement information, the intelligent autonomous fusion of the trajectory information, the satellite information, the inertial navigation information, the geomagnetic information and other integrated navigation information is realized based on the sensor availability and information credibility analysis, and the high-precision navigation positioning of the guided aircraft is realized.
In an exemplary embodiment, if the satellite resolving module cannot normally capture the satellite signal due to the interference of the satellite signal, or the satellite resolving module does not capture the satellite signal before reaching the flight peak, in the ascent stage of the aircraft, the satellite information can be accurately aligned without depending on the satellite resolving module, and the inertial navigation resolving module can be compensated, corrected and initially aligned only by depending on the ballistic information coarse alignment, so that the overall control method can still achieve the basic functions.
And S306, predicting the drop point by using the ballistic information, and controlling the aircraft.
In the descending stage of the guided aircraft, trajectory prediction and drop point evaluation are quickly carried out on the basis of actual measurement information such as satellite information, inertial navigation information and geomagnetic information, the guided aircraft is controlled by the autonomous navigation control assembly, the pose of the guided aircraft is adjusted, and the guided aircraft is controlled to hit a target.
In an exemplary embodiment, in the drop point evaluation phase, only trajectory calculation module type trajectory prediction can be used, comparison with trajectory information before launching is carried out, drop point prediction is carried out, and guidance control is carried out.
It should be noted that, for simplicity of description, the above-mentioned method embodiments are described as a series of acts or combination of acts, but those skilled in the art will recognize that the present invention is not limited by the order of acts, as some steps may occur in other orders or concurrently in accordance with the invention. Further, those skilled in the art should also appreciate that the embodiments described in the specification are preferred embodiments and that the acts and modules referred to are not necessarily required by the invention.
Through the above description of the embodiments, those skilled in the art can clearly understand that the method according to the above embodiments can be implemented by software plus a necessary general hardware platform, and certainly can also be implemented by hardware, but the former is a better implementation mode in many cases. Based on such understanding, the technical solutions of the present invention or portions thereof contributing to the prior art may be embodied in the form of a software product, which is stored in a storage medium (such as ROM/RAM, magnetic disk, optical disk) and includes instructions for enabling a terminal device (which may be a mobile phone, a computer, a server, or a network device) to execute the method according to the embodiments of the present invention.
Example 4
According to the embodiment of the invention, a guidance aircraft autonomous control method based on full ballistic information is also provided, and as shown in fig. 4, the method comprises the following steps:
in step S402, the geomagnetic calculation module calculates based on the geomagnetic rotation speed.
The geomagnetic calculation module measures a geomagnetic rotation speed through a geomagnetic sensor and performs calculation based on the geomagnetic rotation speed.
In step S404, the trajectory calculation module performs trajectory calculation.
The trajectory calculation module binds trajectory information, and performs trajectory real-time calculation based on meteorological geomagnetic measurement and bound trajectory information of the geomagnetism calculation module to obtain trajectory positions, trajectory speeds and projectile postures.
In step S406, the satellite calculation module performs a composite calculation.
The satellite calculation module performs composite calculation based on the longitude and latitude, the speed and the flight time of the aircraft to obtain the trajectory position, the trajectory speed and the projectile attitude.
And step S408, aligning by the inertial navigation resolving module.
The inertial navigation resolving module performs self-checking on inertial devices such as an accelerometer and a gyroscope to obtain a self-checking result, performs ballistic information coarse alignment based on the self-checking result and the ballistic position, the ballistic velocity and the projectile attitude output by the ballistic resolving module, and performs satellite information accurate alignment based on the self-checking result and the ballistic position, the ballistic velocity and the projectile attitude output by the satellite resolving module.
And step S410, multi-source fusion combined navigation.
In the aircraft ascending phase, performing auxiliary calibration on an inertial sensor by using trajectory information, and performing combined navigation with satellite information, geomagnetic information and the inertial sensor, wherein the ascending phase is a phase from the launching point to a trajectory vertex; and performing trajectory prediction and drop point evaluation according to the measured data in an aircraft descending stage, and performing composite control on the aircraft to control the aircraft to hit a target, wherein the descending stage is a stage from the trajectory vertex to the drop point.
Example 5
According to an embodiment of the invention, there is also provided an apparatus for implementing the above-mentioned method for autonomously controlling a guided vehicle based on full ballistic information, as shown in fig. 5, the apparatus includes: an up phase control module 52 and a down phase control module 54.
The ascent phase control module 52 is configured to perform an auxiliary calibration of the inertial sensors with ballistic information during an ascent phase of the aircraft, the ascent phase being a phase from the launch point to a ballistic vertex, combined with satellite information, geomagnetic information, and inertial sensors;
the descent phase control module 54 is configured to perform a ballistic prediction and a drop point evaluation based on the measured data to perform a composite control of the aircraft to control the aircraft to hit a target during a descent phase of the aircraft, wherein the descent phase is a phase from the ballistic vertex to the drop point.
In an exemplary embodiment, the aircraft may further comprise a binding ballistic information module configured to bind ballistic information in advance according to tactical requirements, relative position to a target, and meteorological conditions of a flying field before launching the aircraft.
In an exemplary embodiment, the ramp-up phase control module 52 is further configured to: under the condition that satellite signals are not captured, a Kalman filter is constructed based on a ballistic model and an attitude estimation algorithm of geomagnetic information and comprehensively considering the attitude motion characteristics of the projectile body and geomagnetic field measurement information, so that the attitude, the attitude change speed and the attitude angular acceleration information of the projectile body are estimated, and the continuous navigation positioning information is generated.
For example, the ascent stage control module 52 may measure the rotation speed of the aircraft in real time through geomagnetic solution, and perform real-time attitude estimation of a pitch angle and a roll angle of a projectile body by combining meteorological information, a trajectory model, and geomagnetic information; selecting an estimated state vector by taking the pitch angle of the projectile body, the roll angle, the pitch angle of the projectile body, the angular velocity of the roll angle and the angular acceleration as estimators, and establishing a state equation; measuring geomagnetic information, and establishing a measurement equation of the system according to a conversion relation between the measured geomagnetic information in a launching coordinate system and a projectile coordinate system; and performing Taylor expansion on a nonlinear measurement equation based on the state equation and the measurement equation, neglecting high-order terms, and constructing extended Kalman filtering, wherein a measurement noise variance matrix in the measurement equation is estimated by the measurement compensation precision of a geomagnetic sensor.
In an exemplary embodiment, the ramp-up phase control module 52 is further configured to: under the condition that the satellite signal is captured, trajectory calculation data, geomagnetic calculation data and satellite calculation data are utilized to carry out trajectory information coarse alignment and satellite information accurate alignment, the satellite calculation data, inertial navigation calculation data, geomagnetic calculation data, trajectory calculation data and meteorological data are intelligently and autonomously fused to construct a Kalman filter, and the position, the speed and the posture of the projectile are measured to generate continuous navigation positioning information.
For example, the ascent stage control module 52 may select a plurality of state variables to construct a state equation, the plurality of state variables including attitude angle error, velocity error, position error, accelerometer null bias, and gyroscope drift error; selecting a speed measurement value, a position measurement value and an attitude measurement value of the aircraft as measurement values, wherein the position measurement value is a difference value between a position measured by the inertial navigation resolving module and an output position of the satellite resolving module; the velocity measurement value is the difference value between the velocity calculated by the inertial navigation resolving module and the output velocity of the inertial navigation resolving module; the attitude measurement value is a difference value between a geomagnetic vector obtained by the geomagnetic model and a geomagnetic vector obtained by the geomagnetic resolving module; constructing a measurement equation based on the measurement values; and constructing a Kalman filter based on the state equation and the measurement equation.
In an exemplary embodiment, the descent phase control module 54 is configured to perform trajectory prediction and drop point assessment using navigation results based on a fusion of satellite solution data, trajectory solution data, and inertial navigation solution data; controlling the aircraft to make a large angle attitude adjustment at the end of the aircraft descent phase based on the trajectory prediction and the drop point evaluation so that the aircraft hits a target.
In another exemplary embodiment, the descent phase control module 54 may be further configured to compare the model trajectory prediction to the binding trajectory information to predict a drop point.
Example 6
According to an embodiment of the invention, there is also provided an apparatus for implementing the above-mentioned method for autonomously controlling a guided vehicle based on full ballistic information, as shown in fig. 6, the apparatus includes: the system comprises a power supply module 61, an inertial navigation resolving module 63, a geomagnetic resolving module 64, a ballistic resolving module 65, a data recording module 66, a satellite resolving module 67, a gyroscope 68, an accelerometer 69, a magnetic sensor 70, an upper computer 71, a communication serial port 72 and an ARM microprocessor 62(NavOS system).
The gyroscope 68 and the accelerometer 69 are used for acquiring inertial navigation data of the aircraft and sending the inertial navigation data to the inertial navigation solution module 63. The magnetic sensor 70 is configured to collect a geomagnetic signal and send the geomagnetic signal to the geomagnetic solution module 64.
The geomagnetism resolving module 64 is configured to perform resolving based on the rotation speed of geomagnetism. The trajectory calculation module 65 binds trajectory information, and performs trajectory real-time calculation based on meteorological geomagnetic measurement and bound trajectory information of the geomagnetic calculation module 64 to obtain trajectory position, trajectory speed and projectile attitude. And transmits the ballistic position, ballistic velocity, and projectile attitude to the upper computer 71. The upper computer 71 is communicated with the data recording module 66 through a communication serial port 72.
The satellite calculation module 67 performs composite calculation based on the longitude and latitude, the speed and the flight time of the aircraft to obtain a trajectory position, a trajectory speed and a projectile attitude.
The inertial navigation calculating module 63, the geomagnetic calculating module 64, the ballistic calculating module 65 and the satellite calculating module 67 can all communicate with the ARM microprocessor 62, and transmit the calculated data to the ARM microprocessor 62. The ARM microprocessor 62 performs data fusion based on the received solution data for navigation. The specific navigation method is as described in embodiments 1 to 4, and is not described herein again.
Example 7
The embodiment of the invention also provides a storage medium. Alternatively, in the present embodiment, the storage medium stores a program thereon, and when the program is executed, the program may cause a computer to execute the method described in any one of embodiments 1 to 4.
Optionally, in this embodiment, the storage medium may include, but is not limited to: a U-disk, a Read-Only Memory (ROM), a Random Access Memory (RAM), a removable hard disk, a magnetic or optical disk, and other various media capable of storing program codes.
The above-mentioned serial numbers of the embodiments of the present invention are merely for description and do not represent the merits of the embodiments.
The integrated unit in the above embodiments, if implemented in the form of a software functional unit and sold or used as a separate product, may be stored in the above computer-readable storage medium. Based on such understanding, the technical solution of the present invention may be embodied in the form of a software product, which is stored in a storage medium and includes several instructions for causing one or more computer devices (which may be personal computers, servers, network devices, etc.) to execute all or part of the steps of the method according to the embodiments of the present invention.
In the above embodiments of the present invention, the descriptions of the respective embodiments have respective emphasis, and for parts that are not described in detail in a certain embodiment, reference may be made to related descriptions of other embodiments.
In the several embodiments provided in the present application, it should be understood that the disclosed client may be implemented in other manners. The above-described embodiments of the apparatus are merely illustrative, and for example, the division of the units is only one type of division of logical functions, and there may be other divisions when actually implemented, for example, a plurality of units or components may be combined or may be integrated into another system, or some features may be omitted, or not executed. In addition, the shown or discussed coupling or direct coupling or communication connection between each other may be an indirect coupling or communication connection through some interfaces, units or modules, and may be electrical or in other forms.
The units described as separate parts may or may not be physically separate, and parts displayed as units may or may not be physical units, may be located in one place, or may be distributed on a plurality of network units. Some or all of the units can be selected according to actual needs to achieve the purpose of the solution of the embodiment.
In addition, functional units in the embodiments of the present invention may be integrated into one processing unit, or each unit may exist alone physically, or two or more units are integrated into one unit. The integrated unit can be realized in a form of hardware, and can also be realized in a form of a software functional unit.
The foregoing is only a preferred embodiment of the present invention, and it should be noted that, for those skilled in the art, various modifications and decorations can be made without departing from the principle of the present invention, and these modifications and decorations should also be regarded as the protection scope of the present invention.
Claims (10)
1. A guidance aircraft autonomous control method based on full ballistic information is characterized by comprising the following steps:
performing auxiliary calibration on an inertial sensor by using ballistic information in an ascending phase of the aircraft, and performing combined navigation with satellite information, geomagnetic information and the inertial sensor, wherein the ascending phase is a phase from the launching point to a ballistic vertex;
and performing trajectory prediction and drop point evaluation according to the measured data in an aircraft descending stage, and performing composite control on the aircraft to control the aircraft to hit a target, wherein the descending stage is a stage from the trajectory vertex to the drop point.
2. The method of claim 1, further comprising: before the aircraft is launched, trajectory information is bound in advance according to tactical requirements, relative positions with targets and meteorological conditions in the flying field.
3. The method of claim 1, wherein using ballistic information to assist in calibration of inertial sensors, combined with satellite information, geomagnetic information, and inertial sensors, comprises:
under the condition that satellite signals are not captured, constructing a Kalman filter based on a trajectory model and an attitude estimation algorithm of geomagnetic information and comprehensively considering the attitude motion characteristics of a projectile body and geomagnetic field measurement information to estimate the attitude, the attitude change speed and the attitude angular acceleration information of the projectile body so as to generate continuous navigation positioning information;
under the condition that the satellite signal is captured, trajectory calculation data, geomagnetic calculation data and satellite calculation data are utilized to carry out trajectory information coarse alignment and satellite information accurate alignment, the satellite calculation data, inertial navigation calculation data, geomagnetic calculation data, trajectory calculation data and meteorological data are intelligently and autonomously fused to construct a Kalman filter, and the position, the speed and the posture of the projectile are measured to generate continuous navigation positioning information.
4. The method of claim 3, wherein a Kalman filter is constructed based on a posture estimation algorithm of a ballistic model and geomagnetic information and comprehensively considering the posture and motion characteristics of the projectile body and geomagnetic field measurement information, and comprises the following steps:
measuring the rotating speed of the aircraft in real time through geomagnetic resolution, and carrying out real-time attitude estimation on a pitch angle and a roll angle of a projectile body by combining meteorological information, a trajectory model and geomagnetic information;
selecting an estimated state vector by taking the pitch angle of the projectile body, the roll angle, the pitch angle of the projectile body, the angular velocity of the roll angle and the angular acceleration as estimators, and establishing a state equation;
measuring geomagnetic information, and establishing a measurement equation of the system according to a conversion relation between the measured geomagnetic information in a launching coordinate system and a projectile coordinate system;
and performing Taylor expansion on a nonlinear measurement equation based on the state equation and the measurement equation, neglecting high-order terms, and constructing extended Kalman filtering, wherein a measurement noise variance matrix in the measurement equation is estimated by the measurement compensation precision of a geomagnetic sensor.
5. The method according to claim 3, wherein the ballistic information coarse alignment and the satellite information precise alignment are performed by using the ballistic solution data, the geomagnetic solution data and the satellite solution data, the inertial navigation solution data, the geomagnetic solution data, the ballistic solution data and the meteorological data are intelligently and autonomously fused to construct a Kalman filter, comprising:
selecting a plurality of state variables to construct a state equation, wherein the state variables comprise attitude angle errors, speed errors, position errors, accelerometer zero offset and gyroscope drift errors;
selecting a speed measurement value, a position measurement value and an attitude measurement value of the aircraft as measurement values, wherein the position measurement value is a difference value between a position measured by the inertial navigation resolving module and an output position of the satellite resolving module; the velocity measurement value is the difference value between the velocity calculated by the inertial navigation resolving module and the output velocity of the inertial navigation resolving module; the attitude measurement value is a difference value between a geomagnetic vector obtained by the geomagnetic model and a geomagnetic vector obtained by the geomagnetic resolving module;
constructing a measurement equation based on the measurement values;
and constructing a Kalman filter based on the state equation and the measurement equation.
6. The method of claim 1, wherein ballistic prediction and drop point assessment are performed based on measured data to provide composite control of the aircraft, comprising:
performing trajectory prediction and drop point evaluation by using a navigation result obtained by fusing satellite solution data, trajectory solution data and inertial navigation solution data;
controlling the aircraft to make a large angle attitude adjustment at the end of the aircraft descent phase based on the trajectory prediction and the drop point evaluation so that the aircraft hits a target.
7. The method of claim 6, wherein performing a drop point assessment comprises: comparing the model trajectory prediction with the binding trajectory information, and predicting a drop point.
8. A guided vehicle autonomous control device based on full ballistic information, comprising:
an ascent stage control module configured to perform auxiliary calibration on an inertial sensor using ballistic information in an ascent stage of the aircraft, the ascent stage being a stage from the launch point to a ballistic vertex, and performing combined navigation with satellite information, geomagnetic information, and the inertial sensor;
and the descending stage control module is configured to perform ballistic prediction and drop point evaluation according to the measured data in an aircraft descending stage, and perform composite control on the aircraft to control the aircraft to hit a target, wherein the descending stage is a stage from the ballistic vertex to the drop point.
9. A guided vehicle autonomous control system based on full ballistic information, comprising:
the apparatus of claim 8;
an aircraft configured to hit a target under control of the device.
10. A computer-readable storage medium, on which a program is stored, which, when executed, causes a computer to carry out the method according to any one of claims 1 to 7.
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