Missile erection filtering method, system and device and readable storage medium

Abstract

The invention discloses a missile erection filtering method, system and device and a readable storage medium, and relates to the technical field of inertial navigation and the field of initial alignment. Based on the reference inertial unit pitch angle before erection, the reference inertial unit pitch angle at the current moment, the inertial unit angular velocity on the missile and the trajectory launching inclination angle, whether the missile is in any one of the non-target states of the non-erection state, the just erection state, the to-be-erected in-place state and the erection completion state or not is judged; if yes, measurement updating is not carried out in the Kalman filtering calculation process of the current moment, and time updating is carried out based on the state transition matrix of the current moment, the state vector of the previous moment, the mean square error matrix of the previous moment and the process noise variance matrix; and attitude information output by the inertial navigation system at the current moment is used as a missile attitude result for calculation at the next moment. According to the method, shaking between the guided missile and the launching box in the erecting process before launching is considered, and the attitude precision of initial alignment of the guided missile is improved.

Description

Technical Field

[0001] The present application relates to the field of initial alignment in the field of inertial navigation technology, and in particular to a filtering method, system, device and readable storage medium for missile erection. Background Art

[0002] With the increasing demand for precision guidance in modern weapon systems, the initial alignment technology of missiles plays a vital role in ensuring the navigation accuracy of missiles during flight. In order to ensure the precision guidance capability of missiles, launch platforms are usually equipped with high-precision inertial navigation systems. However, for cost control reasons, missiles themselves mostly use medium- and low-precision inertial navigation systems. In this context, the transfer alignment algorithm is particularly important. This algorithm effectively utilizes key information such as attitude and speed in the high-precision inertial navigation system, and combines Kalman filtering processing with specific maneuvers such as missile erection, acceleration and deceleration of the launch platform, and steering to estimate and correct the attitude error of the on-board inertial navigation system and complete the initial alignment of the missile.

[0003] At present, conventional transfer alignment algorithms are usually based on an assumption that the high-precision inertial navigation system on the launch platform is completely strapdown with the inertial navigation system on the missile, which means that the relative position of the two remains unchanged during the maneuver. However, in the actual application scenario of box / tube-mounted missile weapon systems, due to the structural accuracy of the launch box / tube, the missile often has a certain amount of shaking with the launch box / tube during the erection process before launch. This shaking condition will directly affect the accuracy of the transfer alignment algorithm, thereby affecting the initial attitude accuracy of the missile.

[0004] Therefore, when there is shaking between the missile and the launch box / tube during the erection process before launch, how to avoid the influence of this condition on the transfer alignment algorithm and thus improve the accuracy of the missile's initial attitude is a problem that needs to be solved urgently. Summary of the invention

[0005] The present application provides a filtering method, system, device and readable storage medium for missile erection, which can effectively take into account the shaking between the missile and the launch box / tube during the erection process before launch, thereby improving the attitude accuracy of the initial alignment of the missile.

[0006] In a first aspect, an embodiment of the present application provides a filtering method for missile erection, the filtering method for missile erection comprising:

[0007] Before performing the measurement update of the Kalman filter, judging whether the missile is in a non-target state based on the reference inertial group pitch angle before erection, the reference inertial group pitch angle at the current moment, the inertial group angular velocity on the missile, and the ballistic launch inclination angle, wherein the non-target state is any one of the following: the state of not erecting, the state of just erecting, the state of about to erect in place, and the state of completed erection;

[0008] If so, during the Kalman filter calculation at the current moment, measurement update is not performed. Time update is carried out based on the state transition matrix at the current moment, the state vector at the previous moment, the mean square error matrix at the previous moment, and the process noise variance matrix. And the attitude information output by the inertial navigation system at the current moment is used as the missile attitude result, where the missile attitude result is used for the calculation of the transfer alignment algorithm at the next moment, and the time update result is used for the time update in the next cycle.

[0009] Combined with the first aspect, in an implementation, the determination of whether the missile is in a non-target state based on the pitch angle of the reference inertial unit before erection, the pitch angle of the reference inertial unit at the current moment, the angular velocity of the on-board inertial unit, and the ballistic launch inclination angle includes:

[0010] If it is detected that the absolute value of the difference between the pitch angle of the reference inertial unit at the current moment and the pitch angle of the reference inertial unit before erection is less than a preset angle threshold, the absolute value of the difference between the pitch angle of the reference inertial unit at the current moment and the ballistic launch inclination angle is less than a preset angle threshold, or the angular velocity of the on-board inertial unit is less than a preset angular velocity threshold, then it is determined that the missile is in a non-target state;

[0011] If it is detected that the absolute value of the difference between the pitch angle of the reference inertial unit at the current moment and the pitch angle of the reference inertial unit before erection is not less than a preset angle threshold, the absolute value of the difference between the pitch angle of the reference inertial unit at the current moment and the ballistic launch inclination angle is not less than a preset angle threshold, and the angular velocity of the on-board inertial unit is not less than a preset angular velocity threshold, then it is determined that the missile is not in a non-target state.

[0012] Combined with the first aspect, in an implementation, after the step of determining whether the missile is in a non-target state based on the pitch angle of the reference inertial unit before erection, the pitch angle of the reference inertial unit at the current moment, the angular velocity of the on-board inertial unit, and the ballistic launch inclination angle, it further includes:

[0013] If the missile is not in a non-target state, a measurement vector is determined based on the on-board inertial unit velocity, the reference inertial unit velocity, the lever arm velocity, the reference inertial navigation attitude quaternion at the current moment, and the on-board inertial navigation attitude quaternion at the current moment;

[0014] A measurement matrix at the current moment is determined based on the earth's angular velocity, the attitude transfer matrix of the on-board inertial unit, the sampling period of the on-board inertial unit, and the angle increment output by the on-board inertial unit in the current cycle;

[0015] Measurement residual check is carried out based on the measurement vector, the measurement noise variance matrix, the measurement matrix at the current moment, the predicted state vector at the current moment, and the predicted mean square error matrix at the current moment;

[0016] Based on the result of the measurement residual check, it is determined whether to perform measurement update on the missile.

[0017] In combination with the first aspect, in one embodiment, determining whether to perform measurement update on the missile based on the result of measurement residual check further includes:

[0018] If so, perform measurement update on the missile based on the measurement vector, measurement noise variance matrix, predicted mean square error matrix at the current moment, measurement matrix at the current moment, and predicted state vector at the current moment;

[0019] If not, do not perform measurement update, and only execute the step of performing time update on the missile based on the state transition matrix at the current moment, state vector at the previous moment, mean square error matrix at the previous moment, and process noise variance matrix.

[0020] In combination with the first aspect, in one embodiment, performing measurement update on the missile based on the measurement vector, measurement noise variance matrix, predicted mean square error matrix at the current moment, measurement matrix at the current moment, and predicted state vector at the current moment includes:

[0021] Determine the filter gain matrix based on the measurement noise variance matrix, predicted mean square error matrix at the current moment, and measurement matrix at the current moment;

[0022] Determine the estimated mean square error matrix based on the filter gain matrix, predicted mean square error matrix at the current moment, and measurement matrix at the current moment;

[0023] Determine the state vector estimate value based on the filter gain matrix, measurement vector, predicted state vector at the current moment, and measurement matrix at the current moment;

[0024] Take the filter gain matrix, estimated mean square error matrix, and state vector estimate value as the measurement update result.

[0025] In combination with the first aspect, in one embodiment, after the step of determining the state vector estimate value based on the filter gain matrix, measurement vector, predicted state vector at the current moment, and measurement matrix at the current moment, it further includes:

[0026] Perform feedback correction on the navigation parameters corresponding to the inertial navigation system based on the state vector estimate value to obtain the corrected navigation parameters, and perform the calculation for the next cycle based on the corrected navigation parameters;

[0027] Zero out the elements in the state vector estimate value that have completed feedback correction.

[0028] In combination with the first aspect, in one embodiment, performing time update on the missile based on the state transition matrix at the current moment, state vector at the previous moment, mean square error matrix at the previous moment, and process noise variance matrix includes:

[0029] Determine the predicted state vector at the current moment based on the state transition matrix at the current moment and the state vector at the previous moment;

[0030] Determine the predicted mean square error matrix at the current moment based on the state transition matrix at the current moment, the mean square error matrix at the previous moment, and the process noise variance matrix;

[0031] Take the predicted state vector at the current moment and the predicted mean square error matrix at the current moment as the time update result.

[0032] In a second aspect, an embodiment of the present application provides a filtering system for missile erection. The filtering system for missile erection includes:

[0033] A first processing module, which is used to judge whether the missile is in a non-target state based on the pitch angle of the reference inertial unit before erection, the pitch angle of the reference inertial unit at the current moment, the angular velocity of the on-board inertial unit, and the ballistic launch inclination angle before performing the measurement update of the Kalman filter. The non-target state is any one of the non-erected state, the just-erected state, the about-to-be-erected state, and the completed-erection state;

[0034] A second processing module, which is used to, if so, not perform the measurement update during the Kalman filter calculation at the current moment, perform time update based on the state transition matrix at the current moment, the state vector at the previous moment, the mean square error matrix at the previous moment, and the process noise variance matrix, and use the attitude information output by the inertial navigation system at the current moment as the missile attitude result. The missile attitude result is used for the calculation of the transfer alignment algorithm at the next moment, and the time update result is used for the time update of the next cycle.

[0035] In a third aspect, an embodiment of the present application provides a filtering device for missile erection. The filtering device for missile erection includes a processor, a memory, and a filtering program for missile erection stored on the memory and executable by the processor. When the filtering program for missile erection is executed by the processor, the steps of the filtering method for missile erection as described in any one of the foregoing are implemented.

[0036] In a fourth aspect, an embodiment of the present application provides a computer-readable storage medium, on which a filtering program for missile erection is stored. When the filtering program for missile erection is executed by a processor, the steps of the filtering method for missile erection as described in any one of the foregoing are implemented.

[0037] The beneficial effects brought by the technical solutions provided by the embodiments of the present application include:

[0038] Before the measurement update of the Kalman filter, it is determined whether the missile is in any of the non-target states such as the non-erected state, the just-erected state, the state about to be erected in place, and the completed erection state based on the pitch angle of the reference inertial unit before erection, the pitch angle of the reference inertial unit at the current moment, the angular velocity of the on-missile inertial unit, and the ballistic launch inclination angle; if the missile is in the just-erected state or the state about to be erected in place, at this time, the erection angular velocity of the missile changes greatly. If the measurement update is directly performed in these states, the accuracy of the transfer alignment algorithm may decrease due to the shaking between the missile and the launch box / cylinder during the erection process; if the missile is in the non-erected state or the completed erection state, at this time, the missile is completely stationary, and the Kalman filter has a poor estimation effect on some attitude errors. Therefore, the measurement update of the missile is not performed in the non-target state, and only the time update is performed based on the state transition matrix at the current moment, the state vector at the previous moment, the mean square error matrix at the previous moment, and the process noise variance matrix. The time update result is used for the time update of the next cycle; at the same time, since the attitude parameters of the inertial navigation system at this time are more accurate than the attitude parameters obtained after the measurement update of the missile without considering the shaking between the missile and the launch box / cylinder during the erection process before launch in the prior art, the calculation of the next moment will be directly based on the attitude parameters of the inertial navigation system at the current moment, thereby effectively improving the accuracy of the attitude update calculation. In this application, the erection state of the missile is identified before the measurement update, considering the shaking between the missile and the launch box / cylinder during the erection process before launch; and when the missile is in the non-target state, the measurement update of the missile is not performed, but the calculation of the next moment is based on the attitude parameters of the inertial navigation system at the current moment, that is, the attitude parameters of the inertial navigation system are not feedback-corrected based on the attitude error estimation result of the Kalman filter, avoiding the adverse impact of the Kalman filter algorithm on the attitude update of the inertial navigation system when the estimation effect is poor, thereby improving the attitude accuracy of the missile transfer alignment algorithm. BRIEF DESCRIPTION OF THE DRAWINGS

[0039] Figure 1 It is a schematic flowchart of an embodiment of the filtering method for missile erection in this application;

[0040] Figure 2 For this application Figure 1 It is a detailed flowchart of step S10 in this application;

[0041] Figure 3 For this application Figure 1 It is a detailed flowchart of step S20 in this application;

[0042] Figure 4 It is a schematic hardware structure diagram of the filtering device for missile erection involved in the embodiment solution of this application. DETAILED DESCRIPTION OF THE EMBODIMENTS

[0043] To enable those skilled in the art to better understand the solution of this application, the following will clearly and completely describe the technical solution in the embodiments of this application in conjunction with the accompanying drawings in the embodiments of this application. Obviously, the described embodiments are only a part of the embodiments of this application, rather than all the embodiments. Based on the embodiments in this application, all other embodiments obtained by those of ordinary skill in the art without creative efforts belong to the scope of protection of this application.

[0044] To make the purpose, technical solution, and advantages of this application clearer, the following will further describe the embodiments of this application in detail in conjunction with the accompanying drawings.

[0045] In a first aspect, an embodiment of this application provides a filtering method for missile erection.

[0046] In one embodiment, referring to Figure 1 , Figure 1 is a schematic flowchart of an embodiment of the filtering method for missile erection in this application. As Figure 1 shown, the filtering method for missile erection includes:

[0047] Step S10: Before the measurement update of the Kalman filter, based on the pitch angle of the reference inertial unit before erection, the pitch angle of the reference inertial unit at the current moment, the angular velocity of the on-board inertial unit, and the ballistic launch inclination angle, determine whether the missile is in a non-target state, where the non-target state is any one of the non-erected state, the just-erected state, the about-to-be-erected state, and the completed-erected state.

[0048] Exemplarily, in the embodiments of this application, four coordinate systems, namely the inertial coordinate system, the navigation coordinate system, the high-precision reference inertial unit body coordinate system, and the on-board inertial unit body coordinate system, will be defined. The following will describe the above four coordinate systems in detail:

[0049] Inertial coordinate system (I system): The origin O of the inertial coordinate system is located at the center of the earth, and the axis directions remain unchanged in inertial space (which can be briefly recorded as ECI). It is defined that the X axis points to the vernal equinox point and the Z axis coincides with the earth's rotation axis, and the Y axis forms a right-handed system with the X and Z axes.

[0050] Navigation coordinate system (N system): The origin O of the navigation coordinate system is located at the center of the inertial unit. It is defined that the Z axis points to the sky along the local vertical line, XOY is in the local horizontal plane, the X axis points to the east, and the Y axis points to the north (east-north-up, which can be briefly recorded as ENU). Among them, the inertial unit refers to the Inertial Measurement Unit (IMU).

[0051] High-precision reference inertial unit body coordinate system (B MSystem): The origin O of the body coordinate system of the high-precision reference inertial unit is taken at the center of the reference inertial unit; the Y-axis coincides with the longitudinal axis of the reference inertial unit, and the positive direction is forward along the longitudinal axis; the Z-axis is perpendicular to the Y-axis and lies in the vertical plane, and the positive direction is upward; the X-axis forms a right-handed coordinate system (right-front-up, abbreviated as RFU) with the Y-axis and Z-axis. Here, the reference inertial unit refers to a high-precision inertial measurement unit.

[0052] The body coordinate system of the inertial unit on the missile (Bs system): The origin O of the body coordinate system of the inertial unit on the missile is taken at the center of the inertial unit on the missile; the X-axis coincides with the longitudinal axis of the missile body, and the positive direction is toward the head; the Y-axis is in the longitudinal symmetry plane of the missile body, and the positive direction is upward; the Z-axis is perpendicular to the XOY plane, forming a right-handed coordinate system (front-up-right, abbreviated as FUR). Here, the inertial unit on the missile refers to a medium- and low-precision inertial measurement unit.

[0053] It should be understood that the transfer alignment algorithm includes two major parts: the inertial navigation algorithm and the Kalman filter algorithm. The main purpose is to obtain the initial attitude information of the inertial unit on the missile before the Inertial Navigation System (INS) starts to work; the inertial navigation algorithm updates navigation information such as attitude, velocity, and position based on inertial sensors. The Kalman filter algorithm includes two steps: time update and measurement update, which are used to estimate and correct the errors of the navigation information output by the inertial navigation algorithm; before performing the Kalman filter, it also includes initializing navigation parameters, initializing filter parameters, and calculating the inertial navigation system. These are the preconditions for the Kalman filter to perform state estimation. Among them, the initialization of navigation parameters provides a starting point for subsequent inertial navigation calculations, that is, it defines the initial state; the initialization of filter parameters defines the initial values for the recursive calculation of the Kalman filter, and reasonable filter parameters help to enhance the robustness of the filter.

[0054] Specifically, the initialization of navigation parameters (i.e., the initial attitude matrix, velocity, and position) is as follows:

[0055] Calculate the attitude matrix according to the attitude information of the reference inertial unit Calculate the attitude transformation matrix according to the nominal installation relationship between the reference inertial unit and the inertial unit on the missile Multiply the two to obtain the initial attitude matrix of the inertial unit on the missile Then convert it to a quaternion The initialization of velocity and position can be obtained by directly assigning the velocity, latitude, longitude, and altitude in the N system calculated by the reference inertial unit. The principle and method of this initialization are conventional technical means in this field. For the sake of simplicity of description, it will not be elaborated here.

[0056] Among them, the filter parameters include the state vector x k , the estimated mean square error matrix P k , the process noise variance matrix Q k, measurement noise variance matrix R k and state transition matrix Φ k , the initialization of the filtering parameters is as follows:

[0057] Select the attitude misalignment angle φ in the 3 directions (N system) N , velocity error δv in the 3 directions (N system) N (δ represents error), gyro drift of the 3 axes (B S system) ([[]] represents the projection of the angular velocity error of the Bs system relative to the I system caused by gyro drift in the Bs system), bias of the 3-axis accelerometer (Bs system) ([[]] represents the specific force error output in the Bs system caused by the accelerometer bias), installation deviation angle in the 3 directions (Bs system) arm error in the 3 directions (Bs system) ([[]] represents the arm size in the Bs system), flexure deformation angle in the 3 directions (Bs system) (f refers to flexure deformation) and flexure deformation angular velocity in the 3 directions (Bs system) as the state vector, that is, the 24-dimensional state vector x of the transfer alignment filter k , where the 3 directions refer to the X-axis direction, Y-axis direction, and Z-axis direction, and the 3 axes refer to the X-axis, Y-axis, and Z-axis. Therefore, x k is expressed as follows:

[0058]

[0059] The state vector x k , estimated mean square error matrix P k , process noise variance matrix Q k and measurement noise variance matrix R k can be initialized through prior knowledge or experimental data to ensure that the system has reasonable estimated values at the beginning. The principles and methods of the above parameter initialization are common knowledge in the field and will not be elaborated here for the sake of simplicity of description.

[0060] The initialization of the non-zero elements in the state transition matrix Φ k is as follows:

[0061] Φ k (1:21,1:21) = I 21×21

[0062] Φ k (19:21,22:24) = t s I 3×3

[0063]

[0064] In the formula, I 21×21 is a 21-dimensional identity matrix, I 3×3 is a 3-dimensional identity matrix, Φ k (1:21, 1:21) are the elements corresponding to the 1-21st rows and 1-21st columns of the state transition matrix; β x = β y = β z = 2.146 / τ, where τ is the correlation time of the random process of the flexural deformation angle, and its specific value can be determined according to actual requirements and is not limited here; t s is the sampling period of the inertial unit on the missile.

[0065] The calculations of the inertial navigation system include (1) Earth parameter update, (2) attitude update calculation, (3) velocity update calculation, (4) position update calculation, and (5) lever arm velocity calculation, etc., which are described in detail below:

[0066] (1) Earth parameter update:

[0067] Update the radius of the meridian circle the radius of the prime vertical the acceleration due to gravity the angular velocity of the N-system's rotation about its own axis the displacement angular velocity of the N-system

[0068] (2) Attitude update calculation:

[0069]

[0070] In the formula, t k is the current time (i.e., the k-th moment); is the attitude quaternion of the inertial unit on the missile before recursive update; is the attitude quaternion of the inertial unit on the missile after recursive update; is the quaternion from the Bs-system at the k-th moment to the Bs-system at the k + 1-th moment; is the angular increment output by the inertial unit; is the projection of the angular velocity of the N-system relative to the I-system in the N-system; the symbol is quaternion multiplication; F q×v represents the function of multiplying a quaternion by a vector; F v→q (v) represents the function of converting a rotation vector into a quaternion, and the calculation formula is:

[0071]

[0072] (3) Velocity update calculation:

[0073]

[0074] In the formula, is the velocity increment output by the inertial measurement unit; is the angular increment output by the inertial measurement unit; is the velocity compensation amount caused by specific force; is the velocity of the N coordinate system at time k; is the gravitational acceleration of the N coordinate system at time k; is the velocity of the N coordinate system at time k + 1; is the projection of the rotation angular velocity of the N coordinate system relative to the E coordinate system (where the E coordinate system refers to the geodetic coordinate system) in the N coordinate system at time k; is the projection of the rotation angular velocity of the E coordinate system relative to the I coordinate system in the N coordinate system at time k; is the attitude quaternion of the on - missile inertial measurement unit before recursive update; t s is the sampling period of the on - missile inertial measurement unit.

[0075] (4) Position update calculation:

[0076]

[0077] In the formula, L k is the latitude at time k, with the unit of rad; λ k is the longitude at time k, with the unit of rad; h k is the altitude at time k, with the unit of m; L k+1 is the latitude at time k + 1, with the unit of rad; λ k+1 is the longitude at time k + 1, with the unit of rad; h k+1 is the altitude at time k + 1, with the unit of m; is the velocity in the Y - axis direction of the N coordinate system at time k; is the velocity in the Y - axis direction of the N coordinate system at time k + 1; is the velocity in the X - axis direction of the N coordinate system at time k; is the velocity in the X - axis direction of the N coordinate system at time k + 1; is the velocity in the Z - axis direction of the N coordinate system at time k; is the velocity in the Z - axis direction of the N coordinate system at time k + 1; is the radius of the meridian at time k; is the altitude of the reference inertial measurement unit at time k; is the radius of the prime vertical at time k.

[0078] (5) Lever - arm velocity calculation:

[0079]

[0080] In the formula, is the lever - arm dimension in the X - axis direction, is the lever - arm dimension in the Y - axis direction, is the lever arm dimension in the Z-axis direction, is the lever arm dimension, that is, the projection of the position of the on-board inertial assembly relative to the reference inertial assembly in the Bs system, with the unit of m; is the lever arm velocity; is the attitude transfer matrix of the on-board inertial assembly from the Bs system to the N system; is the attitude transfer matrix of the on-board inertial assembly from the N system to the Bs system; t s is the sampling period of the on-board inertial assembly; is the angular increment output by the inertial assembly; is the projection of the rotational angular velocity of the E system relative to the I system in the N system.

[0081] It should be noted that the non-target state is any one of the non-erected state, just-erected state, about-to-be-erected state, and completed-erected state. Before the measurement update of the Kalman filter, by comparing the pitch angle of the reference inertial assembly before erection with the pitch angle of the reference inertial assembly at the current moment, and combining the angular velocity of the on-board inertial assembly and the ballistic launch inclination angle, it can be judged whether the missile is in the non-target state. Specifically, the difference between the pitch angle of the reference inertial assembly before erection and the pitch angle of the reference inertial assembly at the current moment, the difference between the ballistic launch inclination angle and the pitch angle of the reference inertial assembly at the current moment, and the magnitude of the angular velocity of the on-board inertial assembly can be obtained to identify the erection state of the missile, so as to provide a judgment basis for the measurement update of the Kalman filter and ensure that the filtering process is carried out under the state of stable erection.

[0082] Step S20: If so, no measurement update is performed during the Kalman filter calculation at the current moment. Time update is performed based on the state transfer matrix at the current moment, the state vector at the previous moment, the mean square error matrix at the previous moment, and the process noise variance matrix, and the attitude information output by the inertial navigation system at the current moment is used as the missile attitude result. Among them, the missile attitude result is used for the calculation of the transfer alignment algorithm at the next moment, and the time update result is used for the time update of the next cycle.

[0083] Exemplarily, in the embodiment of the present application, if the current moment is in the non-target state, the predicted state vector is calculated based on the state transfer matrix at the current moment and the state vector at the previous moment, and the predicted mean square error matrix is calculated based on the state transfer matrix at the current moment, the process noise variance matrix, and the mean square error matrix at the previous moment; at this time, the results of the time update (i.e., the predicted mean square error matrix and the predicted state vector) are used as the input for the time update of the next cycle; meanwhile, the attitude parameters at the current moment are provided by the inertial navigation system.

[0084] In this application, by identifying the missile erection state before measurement update, the shaking situation between the missile and the launch box / cylinder during the erection process before launch is considered; and when the missile is in a non-target state, the missile is not measured and updated, but the attitude parameters at the current moment are output based on the inertial navigation system, that is, the attitude parameters of the inertial navigation system are not corrected based on the Kalman filter algorithm, avoiding the influence of inaccurate measurement information on the Kalman filter estimation effect, thereby improving the attitude accuracy of the transfer alignment algorithm.

[0085] Further, in an embodiment, referring to Figure 2 As shown, determining whether the missile is in a non-target state based on the pitch angle of the reference inertial unit before erection, the pitch angle of the reference inertial unit at the current moment, the angular velocity of the on-board inertial unit, and the ballistic launch inclination angle includes:

[0086] Step S101: If it is detected that the absolute value of the difference between the pitch angle of the reference inertial unit at the current moment and the pitch angle of the reference inertial unit before erection is less than a preset angle threshold, the absolute value of the difference between the pitch angle of the reference inertial unit at the current moment and the ballistic launch inclination angle is less than a preset angle threshold, or the angular velocity of the on-board inertial unit is less than a preset angular velocity threshold, then it is determined that the missile is in a non-target state;

[0087] Step S102: If it is detected that the absolute value of the difference between the pitch angle of the reference inertial unit at the current moment and the pitch angle of the reference inertial unit before erection is not less than a preset angle threshold, the absolute value of the difference between the pitch angle of the reference inertial unit at the current moment and the ballistic launch inclination angle is not less than a preset angle threshold, and the angular velocity of the on-board inertial unit is not less than a preset angular velocity threshold, then it is determined that the missile is not in a non-target state.

[0088] Exemplarily, in the embodiment of this application, the specific values of the preset angle threshold and the preset angular velocity threshold can be determined according to actual needs and are not limited herein. For example, the preset angle threshold can preferably be taken as 5°, and the preset angular velocity threshold can preferably be taken as 0.5° / s. Specifically, if it is detected that the absolute value of the difference between the pitch angle of the reference inertial unit at the current moment and the pitch angle of the reference inertial unit before erection is less than the preset angle threshold the absolute value of the difference between the pitch angle of the reference inertial unit at the current moment and the ballistic launch inclination angle is less than the preset angle threshold the angular velocity of the on-board inertial unit is less than the preset angular velocity threshold any one of them indicates that the missile has not completed the erection process, has completed the erection process, or is in an unstable erection state, then it can be determined that the missile is in a non-target state; if it is detected that the absolute value of the difference between the pitch angle of the reference inertial unit at the current moment and the pitch angle of the reference inertial unit before erection is not less than the preset angle threshold the absolute value of the difference between the pitch angle of the reference inertial unit at the current moment and the ballistic launch inclination angle is not less than the preset angle threshold and the angular velocity of the on-board inertial unit is not less than the preset angular velocity threshold It is explained that the erection state of the missile has reached certain stability requirements, so it can be determined that the missile is not in a non-target state.

[0089] Further, in one embodiment, after the step of determining whether the missile is in a non-target state based on the pitch angle of the reference inertial unit before erection, the pitch angle of the reference inertial unit at the current moment, the angular velocity of the on-board inertial unit, and the ballistic launch inclination angle, the following is further included:

[0090] If the missile is not in a non-target state, then a measurement vector is determined based on the on-board inertial unit velocity, the reference inertial unit velocity, the lever arm velocity, the quaternion of the reference inertial navigation attitude at the current moment, and the quaternion of the on-board inertial navigation attitude at the current moment;

[0091] Based on the angular velocity of the earth's rotation, the attitude transfer matrix of the on-board inertial unit, the sampling period of the on-board inertial unit, and the angle increment output by the on-board inertial unit in the current period, the measurement matrix at the current moment is determined;

[0092] Measurement residual verification is performed based on the measurement vector, the measurement noise variance matrix, the measurement matrix at the current moment, the predicted state vector at the current moment, and the predicted mean square error matrix at the current moment;

[0093] Based on the result of the measurement residual verification, it is determined whether to perform measurement update on the missile.

[0094] Exemplarily, in the embodiment of the present application, if the missile is not in a non-target state, that is, the missile is in a stable erection state, then the measurement vector z can be obtained by substituting the on-board inertial unit velocity, the reference inertial unit velocity, the lever arm velocity, the quaternion of the reference inertial navigation attitude angle at the current moment, and the quaternion of the on-board inertial navigation attitude angle at the current moment into the following calculation formula k , and the calculation formula is as follows:

[0095]

[0096] In the formula, is the quaternion of the reference inertial navigation attitude angle at the current moment; is the quaternion of the on-board inertial navigation attitude angle at the current moment, is the conjugate quaternion of; is the quaternion for rotating the Bs system to the B S ′ system at the kth moment, is the on-board inertial unit velocity in the X-axis direction of the N system, is the on-board inertial unit velocity in the Y-axis direction of the N system, is the on-board inertial unit velocity in the Z-axis direction of the N system; is the reference inertial unit velocity in the X-axis direction of the N system, is the reference inertial unit velocity in the Y-axis direction of the N system, is the reference inertial group velocity in the Z-axis direction of the N system; is the lever arm velocity in the X-axis direction of the N system, is the lever arm velocity in the Y-axis direction of the N system, is the lever arm velocity in the Z-axis direction of the N system; z k is the measurement vector.

[0097] Substitute the angular velocity of the Earth's rotation, the attitude transfer matrix of the on-board inertial group, the sampling period of the on-board inertial group, and the angle increment output by the on-board inertial group in the current period into the following calculation formula to obtain the measurement matrix at the current moment. The calculation formula is:

[0098]

[0099] H k (4:6, 4:6) = I 3×3

[0100]

[0101] In the formula, is the attitude transfer matrix of the on-board inertial group from the N system to the Bs system, is the attitude transfer matrix of the on-board inertial group from the Bs system to the N system; is the angle increment output by the on-board inertial group in the current period, t s is the sampling period of the on-board inertial group; is the projection of the angular velocity of the rotation of the Bs system relative to the E system in the Bs system, is the projection of the angular velocity of the rotation of the E system relative to the I system in the N system, that is, the angular velocity of the Earth's rotation in the N system. Among them, the E system refers to the geodetic coordinate system; H k is the measurement matrix at the current moment; is expressed as the skew-symmetric matrix of.

[0102] Substitute the measurement noise variance matrix, the measurement matrix at the current moment, and the predicted mean square error matrix at the current moment into the following calculation formula to obtain matrix A. The calculation formula is as follows:

[0103]

[0104] In the formula, R k is the measurement noise variance matrix; H k is the measurement matrix at the current moment, is the transpose of the measurement matrix at the current moment; P k / k-1 is the predicted mean square error matrix at the current moment.

[0105] Take the square root of the diagonal elements of A to obtain vector A'. If the measurement vector z k has any element that does not satisfy |zk -H k x k / k-1 |≤5*A', it indicates that there is a large error in the measurement information at the current moment, so no measurement update is performed, and only time update is carried out; if all elements in the measurement vector z k all satisfy |z k -H k x k / k-1 |≤5*A', it represents that the error of the measurement information at the current moment is small, and the measurement update calculation can continue.

[0106] Further, in one embodiment, determining whether to perform measurement update on the missile based on the result of measurement residual check further includes:

[0107] If so, perform measurement update on the missile based on the measurement vector, measurement noise variance matrix, predicted mean square error matrix at the current moment, measurement matrix at the current moment, and predicted state vector at the current moment;

[0108] If not, do not perform measurement update, and only execute the step of performing time update on the missile based on the state transition matrix at the current moment, state vector at the previous moment, mean square error matrix at the previous moment, and process noise variance matrix.

[0109] Exemplarily, in the embodiment of the present application, if all elements in the measurement vector satisfy the judgment condition of measurement residual check, it is determined that the result of measurement residual check is: perform measurement update on the missile. At this time, the attitude, speed, installation deviation angle error, etc. of the missile can be estimated based on the time update result and measurement information at the current moment, and the error estimation result is used for feedback correction of the inertial navigation system. Finally, the updated state vector and estimated mean square error matrix will provide a basis for prediction and update in the next cycle; if at least one element in the measurement vector does not satisfy the judgment condition of measurement residual check, it is determined that the result of measurement residual check is: do not perform measurement update on the missile, that is, only perform time update on the missile. After the time update is completed, the transfer alignment algorithm enters the next calculation cycle.

[0110] Further, in one embodiment, performing measurement update on the missile based on the measurement vector, measurement noise variance matrix, predicted mean square error matrix at the current moment, measurement matrix at the current moment, and predicted state vector at the current moment includes:

[0111] Determine the filter gain matrix based on the measurement noise variance matrix, predicted mean square error matrix at the current moment, and measurement matrix at the current moment;

[0112] Determine the estimated mean square error matrix based on the filter gain matrix, predicted mean square error matrix at the current moment, and measurement matrix at the current moment;

[0113] Determine the state vector estimated value based on the filtering gain matrix, the measurement vector, the predicted state vector at the current moment, and the measurement matrix at the current moment;

[0114] Use the filtering gain matrix, the estimated mean square error matrix, and the state vector estimated value as the measurement update result.

[0115] Exemplarily, in the embodiment of the present application, the measurement update of the missile includes updating the filtering gain matrix, the estimated mean square error matrix, and the state vector estimated value. Specifically, substitute the measurement noise variance matrix, the predicted mean square error matrix at the current moment, and the measurement matrix at the current moment into the following calculation formula to obtain the filtering gain matrix. The calculation formula is:

[0116]

[0117] In the formula, R k is the measurement noise variance matrix; P k / k-1 is the predicted mean square error matrix at the current moment; H k is the measurement matrix at the current moment, is the transpose of the measurement matrix at the current moment; K k is the filtering gain matrix.

[0118] Substitute the filtering gain matrix, the predicted mean square error matrix at the current moment, and the measurement matrix at the current moment into the following calculation formula to obtain the estimated mean square error matrix. The calculation formula is:

[0119] P k =(I - K k H k )P k / k-1

[0120] In the formula, K k is the filtering gain matrix; P k / k-1 is the predicted mean square error matrix at the current moment; H k is the measurement matrix at the current moment; P k is the estimated mean square error matrix; I is the identity matrix.

[0121] Substitute the filtering gain matrix, the measurement vector, the predicted state vector at the current moment, and the measurement matrix at the current moment into the following calculation formula to obtain the state vector estimated value. The calculation formula is:

[0122] x k =x k / k-1 +K k (z k -H k x k / k-1 )

[0123] In the formula, K k is the filtering gain matrix; zk is the measurement vector; x k / k-1 is the predicted state vector at the current moment; H k is the measurement matrix at the current moment; x k is the estimated value of the state vector.

[0124] Further, in one embodiment, after the step of determining the estimated value of the state vector based on the filter gain matrix, the measurement vector, the predicted state vector at the current moment, and the measurement matrix at the current moment, the method further includes:

[0125] Performing feedback correction on the navigation parameters corresponding to the inertial navigation system based on the estimated value of the state vector to obtain corrected navigation parameters, and performing calculations for the next cycle based on the corrected navigation parameters;

[0126] Performing a zeroing process on the estimated value of the state vector after the feedback correction is completed.

[0127] Exemplarily, in the embodiments of the present application, part of the navigation results in the inertial navigation system are compensated according to the estimated value of the state vector to correct its errors; after the feedback correction is completed, the system uses the corrected navigation results to perform calculations for the next cycle to ensure the accuracy of the navigation results; a zeroing process is performed on part of the estimated value of the state vector after the feedback correction is completed to ensure the correctness of the calculations for the next cycle.

[0128] Specifically, substituting the estimated value of the state vector and the attitude parameters corresponding to the inertial navigation system into the following calculation formula to obtain the corrected attitude parameters, and the calculation formula is:

[0129]

[0130] In the formula, φ is the estimated value of the state vector corresponding to the correction of the attitude parameters; represents quaternion multiplication; is the attitude parameter corresponding to the inertial navigation system at time k + 1; is the corrected attitude parameter at time k + 1; F v→q (φ) represents a function for converting a rotation vector into a quaternion, and the calculation formula is:

[0131]

[0132] It should be noted that after calculating the estimated value of the state vector, in addition to correcting the attitude parameters, the speed, gyro zero bias, accelerometer bias value, installation angle, and lever arm can also be corrected to reduce the measurement error of the sensor, improve the accuracy and reliability of the system, and ensure that the transfer alignment algorithm can provide more accurate results. The following gives the specific formulas for these corrections:

[0133] Speed correction:

[0134]

[0135] where x k (4), x k (5) and x k (6) are the state vector estimated values corresponding to the velocity correction; is the velocity in the X-axis direction of the N system output by the inertial navigation system, is the velocity in the Y-axis direction of the N system output by the inertial navigation system, is the velocity in the Z-axis direction of the N system output by the inertial navigation system; is the velocity in the corrected X-axis direction of the N system, is the velocity in the corrected Y-axis direction of the N system, is the velocity in the corrected Z-axis direction of the N system.

[0136] Gyro zero bias correction:

[0137]

[0138] where x k (7), x k (8), x k (9) are the state vector estimated values corresponding to the gyro zero bias correction; is the gyro zero bias in the X-axis direction of the B S system estimated by the Kalman filter algorithm, is the gyro zero bias in the Y-axis direction of the B S system estimated by the Kalman filter algorithm, is the gyro zero bias in the Z-axis direction of the B S system estimated by the Kalman filter algorithm; is the corrected gyro zero bias in the X-axis direction of the B S system, is the corrected gyro zero bias in the Y-axis direction of the B S system, is the corrected gyro zero bias in the Z-axis direction of the B S system.

[0139] Accelerometer bias correction:

[0140]

[0141] where x k (10), x k (11), x k (12) are the state vector estimated values corresponding to the accelerometer bias correction; is the accelerometer bias in the X-axis direction of the B S system estimated by the Kalman filter algorithm, B estimated by the Kalman filter algorithm S is the bias value of the accelerometer in the Y-axis direction, B estimated by the Kalman filter algorithm S is the bias value of the accelerometer in the Z-axis direction; B after calibration S is the bias value of the accelerometer in the X-axis direction, B after calibration S is the bias value of the accelerometer in the Y-axis direction, B after calibration S is the bias value of the accelerometer in the Z-axis direction.

[0142] Installation angle calibration:

[0143]

[0144] where μ is the estimated value of the state vector corresponding to the installation angle calibration; is the attitude quaternion after attitude calibration; is the attitude quaternion after installation angle calibration; F v→q func(μ) represents the function that converts the rotation vector into a quaternion.

[0145] Lever arm calibration:

[0146]

[0147] where x k (16), x k (17), x k (18) are the estimated values of the state vector corresponding to the lever arm calibration; B estimated by the Kalman filter algorithm S is the lever arm in the X-axis direction, B estimated by the Kalman filter algorithm S is the lever arm in the Y-axis direction, B estimated by the Kalman filter algorithm S is the lever arm in the Z-axis direction; B after calibration S is the lever arm in the X-axis direction, B after calibration S is the lever arm in the Y-axis direction, B after calibration S is the lever arm in the Z-axis direction. It should be noted that after completing the feedback calibration, the state vector corresponding to the feedback calibration needs to be set to zero. The formula for setting the state vector to zero is as follows:

[0148] x k (1:18) = 0 18×1

[0149] where x k (1:18) is the state vector of rows 1 - 18 after zeroing; 0 18×1 is the zero vector of rows 1 - 18.

[0150] Further, in one embodiment, referring to Figure 3 as shown, the time update of the missile based on the state transition matrix at the current moment, the state vector at the previous moment, the mean square error matrix at the previous moment, and the process noise variance matrix includes:

[0151] Step S201: Determine the predicted state vector at the current moment based on the state transition matrix at the current moment and the state vector at the previous moment;

[0152] Step S202: Determine the predicted mean square error matrix at the current moment based on the state transition matrix at the current moment, the mean square error matrix at the previous moment, and the process noise variance matrix;

[0153] Step S203: Use the predicted state vector at the current moment and the predicted mean square error matrix at the current moment as the time update result.

[0154] Exemplarily, in the embodiment of the present application, the time update of the missile includes updating the predicted state vector and the predicted mean square error matrix to obtain the predicted state vector at the current moment and the predicted mean square error matrix at the current moment. Specifically, the predicted state vector at the current moment can be obtained by substituting the state transition matrix at the current moment and the state vector at the previous moment into the following calculation formula, and the calculation formula is as follows:

[0155] x k / k-1 = Φ k / k-1 x k-1

[0156] where Φ k / k-1 is the state transition matrix at the current moment; x k-1 is the state vector at the previous moment; x k / k-1 is the predicted state vector at the current moment.

[0157] Substitute the state transition matrix at the current moment, the mean square error matrix at the previous moment, and the process noise variance matrix into the following calculation formula to obtain the predicted mean square error matrix at the current moment, and the calculation formula is:

[0158]

[0159] where Φ k / k-1 is the state transition matrix at the current moment, is the transpose of the state transition matrix at the current moment; P k-1 is the mean square error matrix at the previous moment; Qk is the process noise variance matrix; P k / k-1 is the predicted mean square error matrix at the current moment.

[0160] In a second aspect, an embodiment of the present application further provides a filtering system for missile erection. The filtering system for missile erection includes:

[0161] A first processing module, which is configured to determine whether the missile is in a non-target state based on the pitch angle of the reference inertial unit before erection, the pitch angle of the reference inertial unit at the current moment, the angular velocity of the on-board inertial unit, and the ballistic launch inclination angle before the measurement update of the Kalman filter. The non-target state is any one of the non-erection state, the just-erected state, the about-to-be-erected state, and the completed-erection state;

[0162] A second processing module, which is configured to, if so, not perform measurement update during the Kalman filter calculation at the current moment, perform time update based on the state transition matrix at the current moment, the state vector at the previous moment, the mean square error matrix at the previous moment, and the process noise variance matrix, and use the attitude information output by the inertial navigation system at the current moment as the missile attitude result, where the missile attitude result is used for the calculation of the transfer alignment algorithm at the next moment, and the time update result is used for the time update of the next cycle.

[0163] Further, in an embodiment, the first processing module is specifically configured to:

[0164] If it is detected that the absolute value of the difference between the pitch angle of the reference inertial unit at the current moment and the pitch angle of the reference inertial unit before erection is less than a preset angle threshold, the absolute value of the difference between the pitch angle of the reference inertial unit at the current moment and the ballistic launch inclination angle is less than a preset angle threshold, or the angular velocity of the on-board inertial unit is less than a preset angular velocity threshold, it is determined that the missile is in a non-target state;

[0165] If it is detected that the absolute value of the difference between the pitch angle of the reference inertial unit at the current moment and the pitch angle of the reference inertial unit before erection is not less than a preset angle threshold, the absolute value of the difference between the pitch angle of the reference inertial unit at the current moment and the ballistic launch inclination angle is not less than a preset angle threshold, and the angular velocity of the on-board inertial unit is not less than a preset angular velocity threshold, it is determined that the missile is not in a non-target state.

[0166] Further, in an embodiment, the first processing module is specifically further configured to:

[0167] If the missile is not in a non-target state, a measurement vector is determined based on the on-board inertial unit velocity, the reference inertial unit velocity, the lever arm velocity, the reference inertial navigation attitude quaternion at the current moment, and the on-board inertial navigation attitude quaternion at the current moment;

[0168] Determine the measurement matrix at the current moment based on the angular velocity of the Earth's rotation, the attitude transfer matrix of the inertial unit on the missile, the sampling period of the inertial unit on the missile, and the angle increment output by the inertial unit on the missile in the current period;

[0169] Perform measurement residual verification based on the measurement vector, the measurement noise variance matrix, the measurement matrix at the current moment, the predicted state vector at the current moment, and the predicted mean square error matrix at the current moment;

[0170] Determine whether to perform measurement update on the missile based on the result of the measurement residual verification.

[0171] Furthermore, in one embodiment, the first processing module is specifically further configured to:

[0172] If so, perform measurement update on the missile based on the measurement vector, the measurement noise variance matrix, the predicted mean square error matrix at the current moment, the measurement matrix at the current moment, and the predicted state vector at the current moment;

[0173] If not, do not perform measurement update, and only execute the step of performing time update on the missile based on the state transition matrix at the current moment, the state vector at the previous moment, the mean square error matrix at the previous moment, and the process noise variance matrix.

[0174] Furthermore, in one embodiment, the first processing module is specifically further configured to:

[0175] Determine the filter gain matrix based on the measurement noise variance matrix, the predicted mean square error matrix at the current moment, and the measurement matrix at the current moment;

[0176] Determine the estimated mean square deviation matrix based on the filter gain matrix, the predicted mean square error matrix at the current moment, and the measurement matrix at the current moment;

[0177] Determine the estimated value of the state vector based on the filter gain matrix, the measurement vector, the predicted state vector at the current moment, and the measurement matrix at the current moment;

[0178] Take the filter gain matrix, the estimated mean square deviation matrix, and the estimated value of the state vector as the measurement update result.

[0179] Furthermore, in one embodiment, the first processing module is specifically further configured to:

[0180] Perform feedback correction on the navigation parameters corresponding to the inertial navigation system based on the estimated value of the state vector to obtain the corrected navigation parameters, and perform calculations for the next cycle based on the corrected navigation parameters;

[0181] Perform a zeroing process on the estimated value of the state vector after completing the feedback correction.

[0182] Furthermore, in one embodiment, the second processing module is specifically configured to:

[0183] Determine the predicted state vector at the current moment based on the state transition matrix at the current moment and the state vector at the previous moment;

[0184] Determine the predicted mean square error matrix at the current moment based on the state transition matrix at the current moment, the mean square error matrix at the previous moment, and the process noise variance matrix;

[0185] Take the predicted state vector at the current moment and the predicted mean square error matrix at the current moment as the time update result.

[0186] In this application, by identifying the missile erection state before measurement update, the situation of the shaking between the missile and the launch box / cylinder during the erection process before launch is considered; and when the missile is in a non-target state, measurement update is not performed on the missile, but the calculation for the next moment is based on the attitude parameters of the inertial navigation system at the current moment, that is, the attitude parameters of the inertial navigation system are not feedback corrected based on the attitude error estimation result of the Kalman filter, avoiding the adverse effect of the Kalman filter algorithm on the attitude update of the inertial navigation system when the estimation effect is poor, thereby improving the attitude accuracy of the missile transfer alignment algorithm. Among them, the functional implementation of each module in the above missile erection filtering system corresponds to each step in the above missile erection filtering method embodiment, and its functions and implementation processes will not be elaborated here one by one.

[0187] In a third aspect, an embodiment of this application provides a filtering device for missile erection. The filtering device for missile erection can be a device with data processing functions such as a personal computer (PC), a laptop computer, a server, etc.

[0188] Refer to Figure 4 , Figure 4 which is a schematic diagram of the hardware structure of the filtering device for missile erection involved in the solution of the embodiment of this application. In the embodiment of this application, the filtering device for missile erection may include a processor, a memory, a communication interface, and a communication bus.

[0189] Among them, the communication bus can be of any type and is used to interconnect the processor, the memory, and the communication interface.

[0190] The communication interface includes interfaces such as input / output (I / O) interfaces, physical interfaces, and logical interfaces for interconnecting the components inside the filtering device for missile erection, as well as interfaces for interconnecting the filtering device for missile erection with other devices (such as other computing devices or user devices). The physical interface can be an Ethernet interface, an optical fiber interface, an ATM interface, etc.; the user device can be a display screen (Display), a keyboard (Keyboard), etc.

[0191] The memory can be various types of storage media, such as random access memory (RAM), read-only memory (ROM), non-volatile RAM (NVRAM), flash memory, optical memory, hard disk, programmable ROM (PROM), erasable PROM (EPROM), electrically erasable PROM (EEPROM), etc.

[0192] The processor can be a general-purpose processor, which can call the filtering program for missile erection stored in the memory and execute the filtering method for missile erection provided in the embodiments of the present application. For example, the general-purpose processor can be a central processing unit (CPU). Among them, the method executed when the filtering program for missile erection is called can refer to the various embodiments of the filtering method for missile erection in the present application, which will not be elaborated here.

[0193] Those skilled in the art can understand that Figure 4 the hardware structure shown in does not constitute a limitation to the present application, and may include more or fewer components than shown in the figure, or combine some components, or have different component arrangements.

[0194] In a fourth aspect, the embodiments of the present application further provide a readable storage medium.

[0195] The filtering program for missile erection is stored on the readable storage medium of the present application. When the filtering program for missile erection is executed by a processor, the steps of the filtering method for missile erection as described above are implemented.

[0196] Among them, the method implemented when the filtering program for missile erection is executed can refer to the various embodiments of the filtering method for missile erection in the present application, which will not be elaborated here.

[0197] The terms "including" and "having" and any variations thereof in the specification, claims and above-mentioned drawings of the present application are intended to cover non-exclusive inclusion. For example, a process, method, system, product or device that includes a series of steps or units is not limited to the listed steps or units, but optionally further includes steps or units not listed, or optionally further includes other steps or units inherent to these processes, methods, products or devices. The descriptions of the terms "first", "second", "third", etc. are used to distinguish different objects, etc., and do not represent a sequence, nor do they limit that "first", "second" and "third" are of different types.

[0198] In the description of the embodiments of the present application, words such as "exemplary", "for example" or "for instance" are used to indicate examples, illustrations or explanations. Any embodiment or design solution described as "exemplary", "for example" or "for instance" in the embodiments of the present application should not be construed as being more preferred or having more advantages than other embodiments or design solutions. Rather, the use of words such as "exemplary", "for example" or "for instance" is intended to present relevant concepts in a specific manner.

[0199] In the description of the embodiments of the present application, unless otherwise specified, " / " means "or". For example, A / B may mean A or B; "and / or" in the text is merely a description of the association relationship between associated objects, indicating that there can be three relationships. For example, A and / or B may represent: A exists alone, A and B exist simultaneously, and B exists alone. In addition, in the description of the embodiments of the present application, "a plurality of" means two or more than two.

[0200] In some processes described in the embodiments of the present application, there are multiple operations or steps that appear in a specific order. However, it should be understood that these operations or steps may not be executed in the order in which they appear in the embodiments of the present application or may be executed in parallel. The serial numbers of the operations are only used to distinguish different operations, and the serial numbers themselves do not represent any execution order. In addition, these processes may include more or fewer operations, and these operations or steps may be executed in sequence or in parallel, and these operations or steps may be combined.

[0201] It should be noted that the serial numbers of the above embodiments of the present application are only for description and do not represent the advantages or disadvantages of the embodiments.

[0202] Through the description of the above embodiments, those skilled in the art can clearly understand that the above-described embodiment methods can be implemented by means of software plus a necessary general hardware platform. Of course, they can also be implemented by hardware, but in many cases the former is a better implementation method. Based on such an understanding, the technical solution of the present application, in essence, or the part that contributes to the prior art, can be embodied in the form of a software product. This computer software product is stored in a storage medium as described above (such as ROM / RAM, magnetic disk, optical disc), and includes several instructions for causing a terminal device to execute the methods described in various embodiments of the present application.

[0203] The above are only the preferred embodiments of the present application, and do not limit the patent scope of the present application accordingly. Any equivalent structural or equivalent process transformation made by using the content of the specification and drawings of the present application, or directly or indirectly applied in other related technical fields, shall be similarly included in the patent protection scope of the present application.

Claims

1. A filtering method for missile erection, characterized in that, The filtering method for missile erection includes: Before the measurement update of Kalman filtering, it is judged whether the missile is in a non-target state based on the pitch angle of the reference inertial unit before erection, the pitch angle of the reference inertial unit at the current moment, the angular velocity of the on-board inertial unit, and the ballistic launch inclination angle. The non-target state is any one of the non-erected state, the just-erected state, the about-to-be-erected state, and the erected state. If so, during the Kalman filtering calculation at the current moment, the measurement update is not performed. The time update is carried out based on the state transition matrix at the current moment, the state vector at the previous moment, the mean square error matrix at the previous moment, and the process noise variance matrix. And the attitude information output by the inertial navigation system at the current moment is used as the missile attitude result. Among them, the missile attitude result is used for the calculation of the transfer alignment algorithm at the next moment, and the time update result is used for the time update of the next cycle.

2. The filtering method for missile erection as claimed in claim 1, wherein The judgment of whether the missile is in a non-target state based on the pitch angle of the reference inertial unit before erection, the pitch angle of the reference inertial unit at the current moment, the angular velocity of the on-board inertial unit, and the ballistic launch inclination angle includes: If it is detected that the absolute value of the difference between the pitch angle of the reference inertial unit at the current moment and the pitch angle of the reference inertial unit before erection is less than the preset angle threshold, the absolute value of the difference between the pitch angle of the reference inertial unit at the current moment and the ballistic launch inclination angle is less than the preset angle threshold, or the angular velocity of the on-board inertial unit is less than the preset angular velocity threshold, it is determined that the missile is in a non-target state. If it is detected that the absolute value of the difference between the pitch angle of the reference inertial unit at the current moment and the pitch angle of the reference inertial unit before erection is not less than the preset angle threshold, the absolute value of the difference between the pitch angle of the reference inertial unit at the current moment and the ballistic launch inclination angle is not less than the preset angle threshold, and the angular velocity of the on-board inertial unit is not less than the preset angular velocity threshold, it is determined that the missile is not in a non-target state.

3. The filtering method for missile erection as claimed in claim 1 or 2, characterized in that After the step of judging whether the missile is in a non-target state based on the pitch angle of the reference inertial unit before erection, the pitch angle of the reference inertial unit at the current moment, the angular velocity of the on-board inertial unit, and the ballistic launch inclination angle, it further includes: If the missile is not in a non-target state, the measurement vector is determined based on the on-board inertial unit velocity, the reference inertial unit velocity, the lever arm velocity, the quaternion of the reference inertial navigation attitude at the current moment, and the quaternion of the on-board inertial navigation attitude at the current moment. The measurement matrix at the current moment is determined based on the earth's angular velocity, the attitude transfer matrix of the on-board inertial unit, the sampling period of the on-board inertial unit, and the angle increment output by the on-board inertial unit in the current cycle. The measurement residual check is carried out based on the measurement vector, the measurement noise variance matrix, the measurement matrix at the current moment, the predicted state vector at the current moment, and the predicted mean square error matrix at the current moment. It is determined whether to perform measurement update on the missile based on the result of the measurement residual check.

4. The filtering method for missile erection according to claim 3, characterized in that The determination of whether to perform measurement update on the missile based on the result of the measurement residual check further includes: If so, the missile is subjected to measurement update based on the measurement vector, the measurement noise variance matrix, the predicted mean square error matrix at the current moment, the measurement matrix at the current moment, and the predicted state vector at the current moment. If not, measurement update is not performed, and only the step of time-updating the missile based on the state transition matrix at the current moment, the state vector at the previous moment, the mean square error matrix at the previous moment, and the process noise variance matrix is executed.

5. The filtering method for missile erection as claimed in claim 4, wherein The measurement update of the missile based on the measurement vector, the measurement noise variance matrix, the predicted mean square error matrix at the current moment, the measurement matrix at the current moment, and the predicted state vector at the current moment includes: Determining a filter gain matrix based on the measurement noise variance matrix, the predicted mean square error matrix at the current moment, and the measurement matrix at the current moment; Determining an estimated mean square error matrix based on the filter gain matrix, the predicted mean square error matrix at the current moment, and the measurement matrix at the current moment; Determining an estimated value of the state vector based on the filter gain matrix, the measurement vector, the predicted state vector at the current moment, and the measurement matrix at the current moment; Taking the filter gain matrix, the estimated mean square error matrix, and the estimated value of the state vector as the measurement update result.

6. The filtering method for missile erection according to claim 5, wherein After the step of determining the estimated value of the state vector based on the filter gain matrix, the measurement vector, the predicted state vector at the current moment, and the measurement matrix at the current moment, it further includes: Performing feedback correction on the navigation parameters corresponding to the inertial navigation system based on the estimated value of the state vector to obtain corrected navigation parameters, and performing calculations for the next cycle based on the corrected navigation parameters; Performing a zeroing process on the estimated value of the state vector after the feedback correction is completed.

7. The filtering method for missile erection as described in claim 1, characterized in that, The time update of the missile based on the state transition matrix at the current moment, the state vector at the previous moment, the mean square error matrix at the previous moment, and the process noise variance matrix includes: Determining the predicted state vector at the current moment based on the state transition matrix at the current moment and the state vector at the previous moment; Determining the predicted mean square error matrix at the current moment based on the state transition matrix at the current moment, the mean square error matrix at the previous moment, and the process noise variance matrix; Taking the predicted state vector at the current moment and the predicted mean square error matrix at the current moment as the time update result.

8. A filtering system for missile erection, characterized in that, The filtering system for missile erection includes: A first processing module, which is used to judge whether the missile is in a non-target state before the measurement update of the Kalman filter, based on the pitch angle of the reference inertial unit before erection, the pitch angle of the reference inertial unit at the current moment, the angular velocity of the on-board inertial unit, and the ballistic launch inclination angle, and the non-target state is any one of the non-erected state, the just-erected state, the about-to-be-erected state, and the completed-erected state; A second processing module, which is used to not perform measurement update during the Kalman filter calculation at the current moment, perform time update based on the state transition matrix at the current moment, the state vector at the previous moment, the mean square error matrix at the previous moment, and the process noise variance matrix, and take the attitude information output by the inertial navigation system at the current moment as the missile attitude result, where the missile attitude result is used for the calculation of the transfer alignment algorithm at the next moment, and the time update result is used for the time update of the next cycle.

9. A filtering device for missile erection, characterized in that, The filtering device for missile erection includes a processor, a memory, and a filtering program for missile erection stored on the memory and executable by the processor. When the filtering program for missile erection is executed by the processor, the steps of the filtering method for missile erection according to any one of claims 1 to 7 are implemented.

10. A computer-readable storage medium, characterized in that, A filtering program for missile erection is stored on the computer-readable storage medium. When the filtering program for missile erection is executed by a processor, the steps of the filtering method for missile erection according to any one of claims 1 to 7 are implemented.

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