CN109696090B - Online single-shot thrust identification method for carrier rocket/missile - Google Patents
Online single-shot thrust identification method for carrier rocket/missile Download PDFInfo
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Abstract
The invention provides an online single-shot thrust identification method for a carrier rocket/missile, and belongs to the technical field of aircraft control. The online single-shot thrust identification method comprises the following steps: step one, establishing a carrier rocket/missile health work model; step two, primarily diagnosing thrust loss faults of the carrier rocket/missile; and step three, correcting the thrust loss coefficient. The online single-shot thrust identification method for the carrier rocket/missile has the characteristics of simple structure and concise design process.
Description
Technical Field
The invention relates to an online single-shot thrust identification method for a carrier rocket/missile, and belongs to the technical field of aircraft control.
Background
The composition of a carrier rocket/missile system is complex, the flying environment is severe, various unknown faults can occur, and the flying performance and the control performance of the carrier rocket/missile are influenced. In many failure modes of ballistic missiles, the failure of a power system is a failure source which has high occurrence probability of carrier rockets/missiles and has the most serious consequences. During the flight process, the normal work of an engine is directly influenced by the conditions of turbo pump failure, thrust chamber failure, pipeline valve failure and the like of a carrier rocket/missile, so that thrust loss or direct shutdown of a certain single-engine is caused. On one hand, the carrier rocket/missile power system influences the total thrust after a single engine fault occurs, and the actual output thrust is lower than a preset value, so that the launching task is possibly influenced; on the other hand, the missile engine arrangement is asymmetric after the single-shot thrust fails, great interference torque is generated to destroy the stability of the carrier rocket/missile, and missile instability can be seriously caused, so that a carrier rocket/missile single-shot thrust identification method with high identification speed and accurate result needs to be researched, and the method has important research significance for enhancing the reliability of the carrier rocket/missile and improving the success rate of launching and hitting tasks.
By carrying out retrieval and analysis on documents in the prior art, the research on engine single-shot thrust identification is not carried out by directly utilizing the dynamic characteristics of a carrier rocket/missile at present, and the research on a power system mainly focuses on two aspects of engine fault diagnosis and total thrust estimation. In the aspect of engine fault diagnosis, a cloud model and a BP neural network are organically combined in a serial connection mode in the conventional method, firstly, a cloud transformation method is utilized to identify the structure of the network and extract the characteristics of the cloud model, and meanwhile, a unit delay link is introduced into an output layer to describe the dynamic characteristics of the working process of an engine, so that a liquid rocket engine fault diagnosis method based on a dynamic cloud BP network is researched and provided, the engine fault mode diagnosis is realized, but the method has the advantages of low identification speed, incapability of accurately obtaining the single thrust of a carrier rocket/missile and no engineering practice significance. In the aspect of missile total thrust estimation, the conventional method successfully applies a state and parameter joint estimation method based on a strong tracking filter mainly by establishing a mathematical model of a carrier rocket to realize correct estimation of the thrust parameters of the carrier rocket, but the method can only realize the estimation of the total thrust of an engine and cannot meet the requirements of rapid diagnosis and positioning of a failure mode of a single-engine.
Disclosure of Invention
The invention aims to solve the problem of online thrust identification of a carrier rocket/missile in the flight process, and provides an online single-shot thrust identification method for the carrier rocket/missile. The online single-shot thrust identification method for the carrier rocket/missile is a single-shot thrust identification method combining a servo mechanism-swing angle mapping relation and a maximum likelihood criterion, and adopts the following technical scheme:
an online single-shot thrust identification method for a carrier rocket/missile, the online single-shot thrust identification method comprises the following steps:
the method comprises the following steps: establishing a carrier rocket/missile health working model, describing the flight state of the carrier rocket/missile under the condition of no fault, predicting state information at the future moment and the swing angle information of a servo mechanism through dynamic integration, and laying a foundation for single-engine thrust fault identification of an engine; wherein, the carrier rocket \ missile health work model comprises: a missile system state equation, a missile system dynamics integrator and a missile control system;
step two: carrying out initial diagnosis on thrust loss faults of the carrier rocket/missile, initially realizing fault location by analyzing an actual servo mechanism swing angle and a servo mechanism swing angle in a fault-free mode in the real-time flight process of the carrier rocket/missile, determining a fault engine, initially obtaining a fault loss coefficient, and providing an initial value for correcting the thrust loss coefficient;
step three: and correcting the thrust loss coefficient, recording attitude information in each delta t time interval as observed quantity, taking the thrust loss coefficient as the quantity to be identified, correcting the thrust loss coefficient by utilizing maximum likelihood estimation, acquiring the accurate thrust loss coefficient, and finally finishing single thrust identification.
Further, the establishment process of the health work model comprises the following steps:
firstly, establishing a missile system state equation according to a missile mass center kinetic equation and a mass center kinematic equation, wherein the missile mass center kinetic equation and the mass center kinematic equation are respectively as follows:
the missile system state equation is established in the form of:
wherein the content of the first and second substances,
the system state vector is derived from the sensor information of the missile system at the current moment, namely the system state vector of the missile engine in the healthy working state;
secondly, acquiring a derivative of a missile system vector according to a prestored missile pneumatic parameter model, a prestored structural parameter model and a prestored engine model by combining with the missile system state equation in the first step, and performing dynamic integration on the missile system vector by combining with a missile system dynamic integrator to acquire a missile system vector under the healthy working state of the engine at the future moment;
thirdly, resolving the missile speed, the attitude and the attitude angular speed information at the future moment by using a missile control system according to the kinetic integral to obtain a virtual engine swing angle for maintaining the attitude stability of the missile control system.
Further, the missile control system comprises a PID controller and a swing angle distribution controller; the PID controller is the same as a controller actually loaded by the missile in structure and parameters, and the swing angle distribution controller is the same as the swing angle distribution controller actually loaded by the missile in structure and parameters.
Further, the specific process of the initial diagnosis of the thrust loss fault of the carrier rocket/missile in the step two comprises the following steps:
step 1, establishing a lost thrust-swing angle deviation model base in an off-line manner;
2, under the condition of healthy working of the missile system, simulating and calculating an engine virtual swing angle at a future moment by utilizing a carrier rocket/missile healthy working model and combining a loss thrust-swing angle deviation model library, and preliminarily judging whether the missile has a thrust loss fault at a corresponding moment by utilizing the comparison between the engine virtual swing angle and the engine swing angle calculated by a missile actual control system when the missile system flies to the corresponding moment;
step 3, in any flight process under the fault-free condition, the missile calculates the flight state at the future moment through the current flight state and calculates the virtual engine swing angle at the future working moment through a missile control system; meanwhile, the swing angle of the engine at the moment is compared with the virtual swing angle calculated by the health work simulation model at the last moment, and a swing angle comparison result is obtained; and carrying out preliminary engine fault diagnosis according to the swing angle comparison result.
Further, the thrust loss coefficient correction process in step three includes:
step 1, recording the missile range T ═ T1Second to T ═ T1The attitude angle and the attitude angular velocity information within + delta t seconds are used as observed quantity measured values;
step 2, not considering process noise, converting the missile six-degree-of-freedom typing kinetic equation into a general form, wherein the general form is as follows:
wherein K represents a thrust loss coefficient and is a to-be-identified quantity;
and 3, determining an observation equation for describing the relation between the observation vector and the state vector according to the general form, wherein the observation equation is expressed as:
y(ti)＝h(x(ti),K；ti)+(ti)i＝1,2...N (5)
wherein, y (t)i) To observe the vector and to take the attitude angle and attitude angular velocity as the observation vector, i.e.
where v (i) is the output error:
v(i)＝ym(ti)-h(x(ti),K；ti) (7)
ym(i) is the measured value of the observed quantity; r is a covariance matrix of observation noise, when the statistical characteristics of the observation noise are unknown, the optimal estimation of R is adopted, and the optimal estimation of R is as follows:
step 5, solving the estimation value of the parameter K to be identified by adopting a Newton-Raphson iterative algorithm
And (5) enabling the criterion function J (K) to reach the minimum value, finally obtaining an accurate thrust loss coefficient, and completing single thrust identification.
The invention has the beneficial effects that:
the invention provides an online single-shot thrust identification method for a carrier rocket/missile. The method comprises the steps of firstly establishing a carrier rocket/missile health work model for describing the flight state of the carrier rocket/missile under the condition of no fault, predicting state information at the future moment and swing angle information of a servo mechanism through dynamic integration, and laying a foundation for identifying the single-engine thrust fault of an engine; then establishing a single-engine thrust fault-swing angle mapping relation, analyzing the servo mechanism change relation after different single-engine engines have faults through off-line mathematical simulation, establishing a model base, and then initially realizing fault location by analyzing the actual servo mechanism swing angle and the servo mechanism swing angle in a fault-free mode in the real-time flight process of the carrier rocket/missile, determining a fault engine and initially obtaining a fault loss coefficient so as to provide an initial value for correcting the thrust loss coefficient. And finally, carrying out thrust loss coefficient correction research, recording attitude information in each delta t period as observed quantity, taking the thrust loss coefficient as the quantity to be identified, correcting the thrust loss coefficient by utilizing maximum likelihood estimation, obtaining an accurate thrust loss coefficient, and finally completing single thrust identification. The identification method only utilizes attitude angle and attitude angular velocity information of the carrier rocket/missile system, completes single-shot thrust identification based on the dynamic characteristics of the carrier rocket/missile and the swing angle output value of the servo mechanism, does not need to add redundant monitoring elements, and has the characteristics of simple structure and concise design process; after the system fault is preliminarily determined, the thrust loss coefficient correction is carried out every delta t moment, the identification result has higher precision, and the actual engineering requirement is met; therefore, the method has wide application prospect in the identification of the single-shot thrust of the carrier rocket/missile.
Drawings
FIG. 1 is a block diagram of an engine swing angle generation process;
FIG. 2 is a diagram of a carrier rocket/missile health operation model;
FIG. 3 is a diagram of a preliminary missile thrust loss diagnosis process;
FIG. 4 is a diagram of a process of generating a pivot angle in a thrust failure mode;
FIG. 5 is a flow chart of thrust loss coefficient correction;
Detailed Description
The present invention will be further described with reference to the following specific examples, but the present invention is not limited to these examples.
Example 1:
an online single-shot thrust identification method for a carrier rocket/missile is shown in fig. 1 to 5, and comprises the following steps:
the method comprises the following steps: establishing a carrier rocket/missile health working model, describing the flight state of the carrier rocket/missile under the condition of no fault, predicting state information at the future moment and the swing angle information of a servo mechanism through dynamic integration, and laying a foundation for single-engine thrust fault identification of an engine; wherein, the carrier rocket \ missile health work model comprises: a missile system state equation, a missile system dynamics integrator and a missile control system;
step two: carrying out initial diagnosis on thrust loss faults of the carrier rocket/missile, initially realizing fault location by analyzing an actual servo mechanism swing angle and a servo mechanism swing angle in a fault-free mode in the real-time flight process of the carrier rocket/missile, determining a fault engine, initially obtaining a fault loss coefficient, and providing an initial value for correcting the thrust loss coefficient;
step three: and correcting the thrust loss coefficient, recording attitude information in each delta t time interval as observed quantity, taking the thrust loss coefficient as the quantity to be identified, correcting the thrust loss coefficient by utilizing maximum likelihood estimation, acquiring the accurate thrust loss coefficient, and finally finishing single thrust identification.
The principle and the process for establishing the carrier rocket/missile health working model are as follows:
when the system is not in fault, the engine works healthily, and at the moment, the control system obtains real-time projectile data such as the attitude angle, the attitude angular velocity and the like of the projectile through the output of the projectile sensor for resolving. And calculating corresponding engine equivalent swing angles through a control algorithm, then distributing the equivalent swing angles to actual swing angles of the four engines by adopting a control distribution principle, and controlling the spray pipes of the engines to swing so as to keep the posture of the projectile body stable.
In order to obtain the engine swing angle under the condition of healthy operation of the engine in real time and compare the engine swing angle with the swing angle under the condition of engine failure, a system healthy operation simulation model needs to be established in the missile flight process. The carrier rocket/missile health work model comprises a missile system state equation, a missile system dynamics integrator and a missile control system, and then the establishment process of the carrier rocket/missile health work model comprises the following steps:
firstly, establishing a missile system state equation according to a missile mass center kinetic equation and a mass center kinematic equation, wherein the missile mass center kinetic equation and the mass center kinematic equation are respectively as follows:
the missile system state equation is established in the form of:
wherein the content of the first and second substances,
the system state vector is derived from the sensor information of the missile system at the current moment, namely the system state vector of the missile engine in the healthy working state;
secondly, acquiring a derivative of a missile system vector according to a prestored missile pneumatic parameter model, a prestored structural parameter model and a prestored engine model by combining with the missile system state equation in the first step, and performing dynamic integration on the missile system vector by combining with a missile system dynamic integrator to acquire a missile system vector under the healthy working state of the engine at the future moment;
thirdly, resolving the missile speed, the attitude and the attitude angular speed information at the future moment by using a missile control system according to the kinetic integral to obtain a virtual engine swing angle for maintaining the attitude stability of the missile control system.
The missile control system is used for resolving according to the information such as the speed, the attitude and the like of the missile at the future moment calculated in the dynamic integral, and comprises a PID (proportion integration differentiation) controller and a swing angle distribution controller; the PID controller is the same as a controller actually loaded by the missile in structure and parameters, and the swing angle distribution controller is the same as the swing angle distribution controller actually loaded by the missile in structure and parameters.
Step two, the specific process of the initial diagnosis of the thrust loss fault of the carrier rocket/missile comprises the following steps:
step 1, establishing a lost thrust-swing angle deviation model base in an off-line manner;
2, under the condition of healthy working of the missile system, simulating and calculating an engine virtual swing angle at a future moment by utilizing a carrier rocket/missile healthy working model and combining a loss thrust-swing angle deviation model library, and preliminarily judging whether the missile has a thrust loss fault at a corresponding moment by utilizing the comparison between the engine virtual swing angle and the engine swing angle calculated by a missile actual control system when the missile system flies to the corresponding moment;
step 3, in any flight process under the fault-free condition, the missile calculates the flight state at the future moment through the current flight state and calculates the virtual engine swing angle at the future working moment through a missile control system; meanwhile, the swing angle of the engine at the moment is compared with the virtual swing angle calculated by the health work simulation model at the last moment, and a swing angle comparison result is obtained; and carrying out preliminary engine fault diagnosis according to the swing angle comparison result.
Step two, the working principle of the initial diagnosis of the thrust loss fault of the carrier rocket/missile is as follows: under the condition that the missile system works healthily, the virtual simulation model calculates the virtual swing angle of the engine at the future moment, if the system does not break down, the engine swing angle calculated by the actual control system when the missile system flies to the corresponding moment is the same as the virtual swing angle generated by calculation, and the missile system works healthily; when thrust loss is caused by engine faults in the missile system, the actual missile is deviated between the engine swing angle calculated by the control system at the corresponding moment and the engine swing angle obtained by the healthy work simulation model, and the missile is judged to have the thrust loss fault preliminarily.
In any flight process under the condition of no fault, the missile calculates the flight state at the future moment through the current flight state, calculates the virtual swing angle of the engine at the future working moment through the virtual control system, and simultaneously compares the swing angle of the engine at the moment with the virtual swing angle calculated by the health working simulation model at the previous moment. And carrying out primary engine fault diagnosis according to the swing angle result.
The interference caused by the thrust loss of the engine to the system can be divided into two parts, one part is due to the asymmetry of the thrust caused by the thrust loss of the jet pipe, the thrust of the engine which is symmetrical to the center of the fault engine is larger than that of the fault engine at the moment, and the attitude stability of the projectile body is influenced by the occurrence of constant interference torque; the other part is represented as an interference torque term caused by control torque deviation due to thrust loss, and the control torque coefficient is reduced due to engine thrust loss, so that the structural coefficient of a control system is influenced.
The change of the control moment coefficient caused by thrust loss can cause the change of the engine swing angle, but the influence on the actual swing angle change of the engine is small, so the influence is ignored firstly in the primary thrust identification process of the missile. At the moment, compared with the missile system in a healthy state, the missile system in a fault mode is subjected to a constant interference moment, and in order to maintain the stable posture of the missile body, the extra deflection angle of the engine generates a control moment to counteract the influence of the interference moment.
In order to calculate thrust loss according to the engine pivot angle in a healthy state and the engine pivot angle deviation in a fault state, a lost thrust-pivot angle deviation model base needs to be established offline, simulation under the condition that different single engines lose different thrust is respectively carried out in the dynamic integral, the engine pivot angle in the fault state is obtained, and the pivot angle deviation can be obtained by making a difference with the pivot angle obtained in the dynamic integral in the fault-free state. And establishing a mapping relation between the simulation condition and the swing angle deviation under the fault condition, and finishing a thrust loss-swing angle deviation model library. In the actual flight process, the thrust of a failed single-engine and the loss of the single-engine can be reflected by the swing angle deviation, and the thrust identification is preliminarily completed.
However, since the influence of the change of the control moment coefficient caused by thrust loss on the engine pivot angle change is omitted in the above calculation, the estimated value of the thrust loss coefficient K is inaccurate, and in order to obtain a more accurate thrust observed value of the engine nozzle, the thrust loss coefficient value needs to be further corrected, where the correction is the thrust loss coefficient correction in step three, specifically, the thrust loss coefficient correction in step three includes:
step 1, assuming that T is T at time T1And detecting that the missile power system has faults through a thrust-swing angle mapping relation at the second, preliminarily positioning a fault engine, and then developing a thrust loss coefficient correction design based on the maximum likelihood estimation. Recording missile range T ═ T1Second to T ═ T1The attitude angle and the attitude angular velocity information within + delta t seconds are used as observed quantity measured values;
step 2, not considering process noise, converting the missile six-degree-of-freedom typing kinetic equation into a general form, wherein the general form is as follows:
wherein K represents a thrust loss coefficient and is a to-be-identified quantity;
and 3, determining an observation equation for describing the relation between the observation vector and the state vector according to the general form, wherein the observation equation is expressed as:
y(ti)＝h(x(ti),K；ti)+(ti)i＝1,2...N (5)
wherein, y (t)i) To observe the vector and to take the attitude angle and attitude angular velocity as the observation vector, i.e.
where v (i) is the output error:
v(i)＝ym(ti)-h(x(ti),K；ti) (7)
ym(i) is the measured value of the observed quantity; r is a covariance matrix of observation noise, when the statistical characteristics of the observation noise are unknown, the optimal estimation of R is adopted, and the optimal estimation of R is as follows:
step 5, solving the estimation value of the parameter K to be identified by adopting a Newton-Raphson iterative algorithm
And (5) enabling the criterion function J (K) to reach the minimum value, finally obtaining an accurate thrust loss coefficient, and completing single thrust identification.
The Δ t period is determined according to engineering practical application.
By adopting the method, the missile six-degree-of-freedom flight dynamics equation is used for performing dynamics integration, and the simulation is carried out from T to T1Second to T ═ T1The missile flying process in the + delta T period is carried out according to the T ═ T1Second to T ═ T1And (4) taking actual attitude information of the missile within + delta t seconds as an actual measurement value of the observed quantity to correct the thrust loss coefficient of the engine, and acquiring an accurate thrust loss coefficient.
To prevent fromThe thrust loss coefficient is not accurately identified, the thrust loss coefficient is changed, and the like, from T to T1Recording attitude information in each delta t time period at the moment as observed quantity, and performing maximum likelihood estimation once every delta t time to correct the thrust loss coefficient once
The thrust loss coefficient is corrected every delta t time, the accuracy and the real-time performance of the coefficient are guaranteed, the proper delta t can be determined according to the actual task requirement and the precision requirement, only one parameter to be identified is in the estimation object, and the model has extremely high convergence speed and anti-interference performance and extremely high engineering practice significance.
The invention provides an online single-shot thrust identification method for a carrier rocket/missile, which comprises the steps of firstly establishing a carrier rocket/missile health working model for describing the flight state of the carrier rocket/missile under the fault-free condition, predicting the state information at the future moment and the swing angle information of a servo mechanism through the dynamic integral, and laying a foundation for single-shot thrust fault identification of an engine; then establishing a single-engine thrust fault-swing angle mapping relation, analyzing the servo mechanism change relation after different single-engine engines have faults through off-line mathematical simulation, establishing a model base, and then initially realizing fault location by analyzing the actual servo mechanism swing angle and the servo mechanism swing angle in a fault-free mode in the real-time flight process of the carrier rocket/missile, determining a fault engine and initially obtaining a fault loss coefficient so as to provide an initial value for correcting the thrust loss coefficient. And finally, carrying out thrust loss coefficient correction research, recording attitude information in each delta t period as observed quantity, taking the thrust loss coefficient as the quantity to be identified, correcting the thrust loss coefficient by utilizing maximum likelihood estimation, obtaining an accurate thrust loss coefficient, and finally completing single thrust identification. The identification method only utilizes attitude angle and attitude angular velocity information of the carrier rocket/missile system, completes single-shot thrust identification based on the dynamic characteristics of the carrier rocket/missile and the swing angle output value of the servo mechanism, does not need to add redundant monitoring elements, and has the characteristics of simple structure and concise design process; after the system fault is preliminarily determined, the thrust loss coefficient correction is carried out every delta t moment, the identification result has higher precision, and the actual engineering requirement is met; therefore, the method has wide application prospect in the identification of the single-shot thrust of the carrier rocket/missile.
Although the present invention has been described with reference to the preferred embodiments, it should be understood that various changes and modifications can be made therein by those skilled in the art without departing from the spirit and scope of the invention as defined in the appended claims.
Claims (2)
1. An online single-shot thrust identification method for a carrier rocket/missile is characterized by comprising the following steps:
the method comprises the following steps: establishing a carrier rocket/missile health working model, describing the flight state of the carrier rocket/missile under the condition of no fault, predicting state information at the future moment and the swing angle information of a servo mechanism through dynamic integration, and laying a foundation for single-engine thrust fault identification of an engine; wherein, the carrier rocket \ missile health work model comprises: a missile system state equation, a missile system dynamics integrator and a missile control system;
step two: carrying out initial diagnosis on thrust loss faults of the carrier rocket/missile, initially realizing fault location by analyzing an actual servo mechanism swing angle and a servo mechanism swing angle in a fault-free mode in the real-time flight process of the carrier rocket/missile, determining a fault engine, initially obtaining a fault loss coefficient, and providing an initial value for correcting the thrust loss coefficient;
step three: correcting the thrust loss coefficient, recording attitude information in each delta t time period as observed quantity, taking the thrust loss coefficient as the quantity to be identified, correcting the thrust loss coefficient by utilizing maximum likelihood estimation, obtaining an accurate thrust loss coefficient, and finally completing single thrust identification;
the establishment process of the health work model comprises the following steps:
firstly, establishing a missile system state equation according to a missile mass center kinetic equation and a mass center kinematic equation, wherein the missile mass center kinetic equation and the mass center kinematic equation are respectively as follows:
wherein
Three-channel rotational inertia;
three-channel attitude angular velocity; q is dynamic pressure; s is a pneumatic reference area;
α respectively representing an attack angle and a sideslip angle, and l is rocket reference length;
is a pneumatic damping moment coefficient;
ψ, γ denote a pitch angle, a yaw angle, and a roll angle, respectively; j. the design is a squareRThe length of the swinging of the spray pipe is; m isRThe nozzle swing mass; x is the number ofRRepresents the hinge-to-rocket vertex distance; x is the number ofTRepresenting the distance from the center of mass to the top point of the rocket; r iszRepresenting an engine mounting radius; v represents rocket speed;γ,ψ,
respectively representing three-channel equivalent swing angles of the engine; t isi(i ═ 1,2,3,4) for each engine thrust; b ═ M1M2M3M4]To control the matrix, MBX,MBY,MBZ,MKY,MKZAll are disturbance moments;
the missile system state equation is established in the form of:
wherein the content of the first and second substances,
the system state vector is derived from the sensor information of the missile system at the current moment, namely the system state vector of the missile engine in the healthy working state;
secondly, acquiring a derivative of a missile system vector according to a prestored missile pneumatic parameter model, a prestored structural parameter model and a prestored engine model by combining with the missile system state equation in the first step, and performing dynamic integration on the missile system vector by combining with a missile system dynamic integrator to acquire a missile system vector under the healthy working state of the engine at the future moment;
thirdly, resolving missile speed, attitude and attitude angular speed information at a future moment by using a missile control system according to a dynamic integral to obtain a virtual engine swing angle for maintaining the attitude stability of the missile control system;
step two, the specific process of the initial diagnosis of the thrust loss fault of the carrier rocket/missile comprises the following steps:
step 1, establishing a lost thrust-swing angle deviation model base in an off-line manner;
2, under the condition of healthy working of the missile system, simulating and calculating an engine virtual swing angle at a future moment by utilizing a carrier rocket/missile healthy working model and combining a loss thrust-swing angle deviation model library, and preliminarily judging whether the missile has a thrust loss fault at a corresponding moment by utilizing the comparison between the engine virtual swing angle and the engine swing angle calculated by a missile actual control system when the missile system flies to the corresponding moment;
step 3, in any flight process under the fault-free condition, the missile calculates the flight state at the future moment through the current flight state and calculates the virtual engine swing angle at the future working moment through a missile control system; meanwhile, the swing angle of the engine at the moment is compared with the virtual swing angle calculated by the health work simulation model at the last moment, and a swing angle comparison result is obtained; performing preliminary engine fault diagnosis according to the swing angle comparison result;
step three, the thrust loss coefficient correction process comprises:
step 1, recording the missile range T ═ T1Second to T ═ T1The attitude angle and the attitude angular velocity information within + delta t seconds are used as observed quantity measured values;
step 2, not considering process noise, converting the missile six-degree-of-freedom typing kinetic equation into a general form, wherein the general form is as follows:
wherein K represents a thrust loss coefficient and is a to-be-identified quantity;
and 3, determining an observation equation for describing the relation between the observation vector and the state vector according to the general form, wherein the observation equation is expressed as:
y(ti)＝h(x(ti),K；ti)+(ti)i＝1,2...N (5)
wherein, y (t)i) To observe the vector and to take the attitude angle and attitude angular velocity as the observation vector, i.e.
Step 4, taking the identification criterion as a likelihood criterion function, and obtaining the following criterion function as follows:
where v (i) is the output error:
v(i)＝ym(ti)-h(x(ti),K；ti) (7)
ym(i) is the measured value of the observed quantity; r is a covariance matrix of observation noise, when the statistical characteristics of the observation noise are unknown, the optimal estimation of R is adopted, and the optimal estimation of R is as follows:
step 5, solving the estimation value of the parameter K to be identified by adopting a Newton-Raphson iterative algorithm
And (5) enabling the criterion function J (K) to reach the minimum value, finally obtaining an accurate thrust loss coefficient, and completing single thrust identification.
2. The on-line single-shot thrust identification method according to claim 1, wherein the missile control system comprises a PID controller and a swing angle distribution controller; the PID controller is the same as a controller actually loaded by the missile in structure and parameters, and the swing angle distribution controller is the same as the swing angle distribution controller actually loaded by the missile in structure and parameters.
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