A method of launch alignment for a ground based aircraft

Abstract

The application relates to a ground-based aircraft launching alignment method, which comprises the following steps: during the erecting process of a ground-based aircraft, updating a Kalman filter of a sub-inertial group of the ground-based aircraft according to navigation parameters of a vehicle-mounted reference inertial group installed on a launching frame, and correcting the navigation parameters of the sub-inertial group, so that the vehicle-mounted reference inertial group implements transfer alignment on the sub-inertial group of the ground-based aircraft; after the erecting is completed, the transfer alignment result of the sub-inertial group of the ground-based aircraft is taken as an initial attitude at a launching moment, and a Kalman filter for a flight stage is initialized; after the ground-based aircraft is launched, speed and position information provided by a loaded GNSS device is used for Kalman filtering combined navigation, the attitude deviation of the sub-inertial group of the ground-based aircraft is corrected, and when filtering estimation meets an estimation threshold, alignment correction is completed. The application guarantees rapidity and high precision, and solves the technical problem that a ground-based aircraft cannot obtain high alignment precision in a short time in the related art.

Description

Technical Field

[0001] This application relates to the field of aerospace navigation technology, specifically to a method for aligning a ground-based aircraft launch. Background Technology

[0002] Currently, ground-based aircraft require the establishment of a precise initial attitude matrix and initial alignment before launch. Furthermore, with increasingly demanding application requirements, rapid initial alignment within a limited timeframe is essential for successful rapid launch. Launch platforms typically experience various vibrations and interferences, significantly impacting initial alignment. Therefore, achieving rapid and high-precision initial alignment is a critical challenge that must be addressed before launch.

[0003] In related technologies, the initial alignment process of ground-based aircraft is typically completed using a strapdown inertial navigation system (SINS) for self-alignment. The requirements for initial alignment of a strapdown inertial navigation system generally include two aspects: accuracy and speed. However, these two requirements are often contradictory, and achieving the best alignment accuracy in the shortest possible time is a difficult problem to solve.

[0004] Therefore, it is necessary to design a new launch alignment method for ground-based aircraft to overcome the above problems. Summary of the Invention

[0005] This application provides a launch alignment method for ground-based aircraft, which can solve the technical problem in related technologies that ground-based aircraft cannot obtain high alignment accuracy in a short time.

[0006] In a first aspect, embodiments of this application provide a method for aligning a ground-based aircraft launch, which includes the following steps: During the erection of the ground-based aircraft, the Kalman filter of the ground-based aircraft sub-inertial group is updated and the navigation parameters of the sub-inertial group are corrected according to the navigation parameters set by the vehicle-mounted reference inertial group on the launcher, so that the vehicle-mounted reference inertial group can perform transfer alignment with the ground-based aircraft sub-inertial group. After the vehicle is erected, the alignment result transmitted by the ground-based aircraft sub-inertial navigation system is used as the initial attitude at the time of launch, and the Kalman filter during the flight phase is initialized. After the ground-based aircraft is launched, it uses the velocity and position information provided by the onboard GNSS equipment to perform Kalman filter-based navigation, corrects the attitude deviation of the ground-based aircraft's sub-inertial navigation system, and completes the alignment correction when the filter estimation meets the estimation threshold.

[0007] In conjunction with the first aspect, in one embodiment, during the erection process of the ground-based aircraft, updating the Kalman filter of the ground-based aircraft's sub-inertial navigation system and correcting the navigation parameters of the sub-inertial navigation system according to the navigation parameters set by the vehicle-mounted reference inertial navigation system on the launch pad, so as to enable the vehicle-mounted reference inertial navigation system to perform transfer alignment with the ground-based aircraft's sub-inertial navigation system, includes: At the start of the ground-based aircraft's erection, the ground-based aircraft's sub-inertial navigation system and Kalman filter are initialized according to the navigation parameters set by the vehicle-mounted reference inertial navigation system. The navigation calculations are performed to update the attitude, velocity, and position of the ground-based aircraft sub-inertial group. During the erection process, when the data from the vehicle-mounted reference inertial group is valid, the data is transferred and aligned, the Kalman filter of the ground-based aircraft sub-inertial group is updated, and the navigation parameters of the ground-based aircraft sub-inertial group are corrected.

[0008] In conjunction with the first aspect, in one implementation, during the erection process, when the onboard reference inertial group data is valid, alignment is performed, the Kalman filter of the ground-based aircraft sub-inertial group is updated, and the navigation parameters of the ground-based aircraft sub-inertial group are corrected, including: Calculate the Kalman filter state transition matrix, Kalman filter measurement vector, and Kalman filter measurement matrix based on the navigation parameters of the vehicle-mounted reference inertial group; Update the Kalman filter based on the Kalman filter state transition matrix, Kalman filter measurement vector, and Kalman filter measurement matrix; The attitude parameters of the ground-based aircraft sub-inertial navigation system are corrected based on the updated Kalman filter.

[0009] In conjunction with the first aspect, in one implementation, after the erection is completed, the alignment result transmitted by the ground-based aircraft sub-inertial navigation system is used as the initial attitude at launch time, and the Kalman filter for the flight phase is initialized, including: After the vehicle is erected and in place, the navigation quaternion calculated at the time of alignment completion will be used as the initial value of the attitude quaternion at the launch time of the ground-based aircraft's sub-inertial navigation system. The geographic latitude, longitude, and altitude of the alignment completion time will be used as the geographic latitude, longitude, and altitude of the launch time of the ground-based aircraft's sub-inertial navigation system. The initial state vector of the Kalman filter during the flight phase is set to a 21-dimensional zero-element vector. The initial covariance matrix and process noise equation matrix of the Kalman filter are both set to 21-dimensional diagonal matrices. The measurement noise equation matrix of the Kalman filter is set to a 6-dimensional diagonal matrix. The observation matrix of the Kalman filter is initialized to a 6×21-dimensional matrix.

[0010] In conjunction with the first aspect, in one implementation, the step of performing Kalman filter-based integrated navigation using velocity and position information provided by the onboard GNSS equipment after launch of the ground-based aircraft, correcting attitude deviations of the ground-based aircraft's sub-inertial navigation system, and completing alignment correction when the filter estimation meets the estimation threshold includes: During the flight phase of the ground-based aircraft, navigation calculations are performed on the ground-based aircraft's inertial navigation system to obtain attitude, velocity, and position. When the data from the GNSS equipment is valid, Kalman filter combined navigation is performed based on the speed and position information of the GNSS equipment to correct the attitude deviation of the ground-based aircraft's sub-inertial navigation. Based on a pre-set estimation threshold, the alignment correction process is completed when the filtered estimation meets the estimation threshold, and subsequent navigation calculations are performed based on the alignment results.

[0011] In conjunction with the first aspect, in one implementation, the step of performing navigation calculations on the ground-based aircraft's sub-inertial navigation system during the flight phase of the ground-based aircraft to obtain attitude, velocity, and position includes: The attitude of the ground-based aircraft's sub-inertial navigation system is updated using a quaternion chain multiplication rule, with the input being the gyroscope increment and the Earth's rotation angular rate being subtracted.

[0012] In conjunction with the first aspect, in one implementation, the velocity of the ground-based aircraft sub-inertial navigation system is updated using an incremental addition of the table, while simultaneously subtracting the rotational error vector, gravitational acceleration vector, Gothic acceleration vector, and centripetal acceleration vector.

[0013] In conjunction with the first aspect, in one implementation, the position of the ground-based aircraft sub-inertial navigation system is updated using a trapezoidal integral algorithm based on the velocity vector.

[0014] In conjunction with the first aspect, in one implementation, the step of performing Kalman filter-based integrated navigation based on the velocity and position information of the GNSS device when the GNSS device data is valid, and correcting the attitude deviation of the ground-based aircraft sub-inertial navigation system, includes: The discretized state transition matrix is ​​calculated based on the attitude, velocity, and position obtained from the sub-inertial navigation solution of the ground-based aircraft. The measurement vector of the Kalman filter is constructed based on the velocity and position information of the GNSS equipment; The Kalman filter is updated based on the measurement vector, and the navigation parameters of the ground-based aircraft sub-inertial navigation system are corrected based on the measurement update results of the Kalman filter.

[0015] In conjunction with the first aspect, in one implementation, based on a pre-set estimation threshold, the alignment correction process is completed when the filtered estimation meets the estimation threshold, and subsequent navigation calculations are performed based on the alignment results, including: Calculate the first and second convergence variables of the Kalman filter; Determine whether each element of the first convergent variable is less than the first set value during a consecutive preset number of measurement update cycles; and determine whether the absolute value of the difference between the maximum and minimum values ​​of each row element of the second convergent variable is less than the second set value during a consecutive preset number of measurement update cycles. If both the first and second set values ​​are met simultaneously, the Kalman filter is determined to have converged, the attitude correction condition is satisfied, and alignment is complete.

[0016] The beneficial effects of the technical solutions provided in this application include: By using the vehicle-mounted reference inertial navigation system to perform transfer alignment of the ground-based aircraft during the erection process, the alignment process is embedded into the inherent steps before launch, achieving initial alignment with zero additional time and greatly shortening the alignment time. Furthermore, after launch, data from GNSS equipment is used for further correction. The two-stage alignment before and after launch ensures high alignment accuracy. In other words, the ground-based aircraft launch alignment method in this embodiment ensures both speed and high accuracy, solving the technical problem in related technologies that ground-based aircraft cannot obtain high alignment accuracy in a short time. Attached Figure Description

[0017] To more clearly illustrate the technical solutions in the embodiments of this application, the accompanying drawings used in the description of the embodiments will be briefly introduced below. Obviously, the accompanying drawings described below are only some embodiments of this application. For those skilled in the art, other drawings can be obtained based on these drawings without creative effort.

[0018] Figure 1 This is a flowchart of a ground-based aircraft launch alignment method provided in an embodiment of this application. Detailed Implementation

[0019] To enable those skilled in the art to better understand the present application, the technical solutions in the embodiments of the present application will be clearly and completely described below with reference to the accompanying drawings. Obviously, the described embodiments are only some embodiments of the present application, and not all embodiments. Based on the embodiments in the present application, all other embodiments obtained by those of ordinary skill in the art without creative effort are within the scope of protection of the present application.

[0020] This application provides a method for aligning a ground-based aircraft launch, which can solve the technical problem in related technologies that ground-based aircraft cannot obtain high alignment accuracy in a short time.

[0021] See Figure 1 As shown, an embodiment of this application provides a method for aligning a ground-based aircraft launch, which includes the following steps: S1: During the erection of the ground-based aircraft, the Kalman filter of the ground-based aircraft sub-inertial group is updated and the navigation parameters of the sub-inertial group are corrected according to the navigation parameters set by the vehicle-mounted reference inertial group on the launcher, so that the vehicle-mounted reference inertial group can perform transfer alignment with the ground-based aircraft sub-inertial group.

[0022] S2: After the vehicle is erected, the alignment result transmitted by the ground-based aircraft sub-inertial navigation system is used as the initial attitude at the time of launch, and the Kalman filter during the flight phase is initialized.

[0023] S3: After the ground-based aircraft is launched, it uses the velocity and position information provided by the onboard GNSS equipment to perform Kalman filter-based navigation, corrects the attitude deviation of the ground-based aircraft's sub-inertial navigation system, and completes the alignment correction when the filter estimation meets the estimation threshold.

[0024] In this embodiment, the vehicle-mounted reference inertial group is installed on the launcher. Before launch, during the erection of the ground-based aircraft, the vehicle-mounted reference inertial group is used to transfer and align the sub-inertial groups of the ground-based aircraft. At this time, the vehicle-mounted reference inertial group is erected together with the launcher. After being erected in place, the attitude transfer from the vehicle-mounted reference inertial group to the sub-inertial groups is completed. The sub-inertial groups use the result of the transfer alignment as the initial value to start the navigation calculation after launch. After launch, the ground-based aircraft uses the velocity and position information provided by the onboard GNSS equipment to estimate the attitude error of the initial alignment before launch. When the estimation threshold is met, the second attitude correction is completed, that is, the two-stage alignment before and after launch is completed.

[0025] This embodiment utilizes an onboard reference inertial navigation system (INS) to perform transfer alignment of the ground-based aircraft during the erection process, embedding the alignment process into the pre-launch steps to achieve initial alignment with zero additional time, significantly shortening the alignment time. Furthermore, after launch, attitude correction is further performed using data from GNSS equipment. This two-stage alignment before and after launch ensures high alignment accuracy. Compared to a single initial alignment method, this method utilizes data from two different stages, avoiding the time-consuming and limited accuracy of a single alignment method. By comprehensively using effective data from before and after launch, it improves alignment accuracy. In other words, the ground-based aircraft launch alignment method of this embodiment ensures both speed and high accuracy, solving the technical problem in related technologies where ground-based aircraft cannot achieve high alignment accuracy in a short time.

[0026] Furthermore, in one embodiment, during the erection process of the ground-based aircraft, updating the Kalman filter of the ground-based aircraft's sub-inertial navigation system and correcting the navigation parameters of the sub-inertial navigation system according to the navigation parameters set by the vehicle-mounted reference inertial navigation system on the launch pad, so that the vehicle-mounted reference inertial navigation system performs transfer alignment with the ground-based aircraft's sub-inertial navigation system, may include: S11: When the ground-based aircraft begins to be erected, the sub-inertial navigation system of the ground-based aircraft is initialized according to the navigation parameters set by the vehicle-mounted reference inertial navigation system, and the Kalman filter is initialized.

[0027] S12: Perform navigation calculations to update the attitude, velocity, and position of the ground-based aircraft sub-inertial group. During the erection process, when the onboard reference inertial group data is valid, transfer and align it, update the Kalman filter of the ground-based aircraft sub-inertial group, and correct the navigation parameters of the ground-based aircraft sub-inertial group.

[0028] In some alternative embodiments, step S11 may include: S111: Ground-based aircraft sub-inertial group initialization: The ground-based aircraft sub-inertial group is directly assigned values ​​based on the navigation parameters of the vehicle-mounted reference inertial group, as shown in the following formula.

[0029]

[0030] In the formula: —Attitude quaternion of vehicle-mounted reference inertial navigation system; —Initial values ​​of attitude quaternions for ground-based aircraft sub-inertial navigation systems; —Latitude, longitude, and altitude; transmit the corresponding values ​​of the onboard reference inertial group at the start of alignment. —Northward speed, upward speed and eastward speed, start transmitting the corresponding values ​​of the onboard reference inertial group at the alignment time; —Initial values ​​for the geographic latitude, longitude, altitude, northward velocity, celestial velocity, and eastward velocity of the sub-inertial navigation system.

[0031] S112: Sub-inertial navigation Kalman filter initialization, calculated as follows.

[0032] The initial state vector of the Kalman filter is set to an 18-dimensional vector with zero elements, that is:

[0033] Kalman filter initial covariance matrix It is an 18-dimensional diagonal matrix. The diagonal elements are shown in the table below, and the other off-diagonal elements are 0.

[0034] Table 1. Definition of the initial covariance matrix P0 for Kalman filtering

[0035] Kalman filtering process noise equation matrix It is an 18-dimensional diagonal matrix. The diagonal elements are shown in the table below, and the other off-diagonal elements are 0.

[0036] Table 2 Definition of the noise variance matrix Qk in the Kalman filtering process

[0037] Kalman filter measurement noise equation matrix It is a 6-dimensional diagonal matrix. The diagonal elements are shown in the table below, and the other off-diagonal elements are 0.

[0038] Table 3. Definition of Kalman filter measurement noise variance matrix Rk

[0039] In one embodiment, in step S12, navigation calculations are performed to complete the attitude update, velocity update, and position update of the ground-based aircraft sub-inertial navigation system, including: S121: Sub-inertial navigation parameter velocity update, the calculation formula is as follows.

[0040] First, calculate the velocity increment:

[0041] Then perform the speed update calculation:

[0042] In the formula: —The apparent velocity increments in the XYZ directions of the sub-inertial navigation system; —Angular increments in the XYZ directions of the sub-inertial navigation system; —Attitude matrix of the navigation system in the previous update cycle The transpose of is obtained by quaternion calculation; —Sub-inertial navigation group sampling period; —Current update cycle gravitational acceleration; —The current update cycle is accelerating the connection between the North and East; —Current update cycle: North-South East-West acceleration; —The northward, celestial, and eastward speeds of the previous update cycle.

[0043] S122: Sub-inertial navigation parameter position update, the calculation formula is as follows.

[0044]

[0045] In the formula: —The northward, celestial, and eastward speeds of the previous update cycle; —The speed increments obtained from the speed update for the north, sky, and east directions; —Latitude, longitude, and altitude values ​​from the previous update period; —The radius of the meridian circle obtained in the current update cycle; —The radius of the Mao-You circle obtained in the current update cycle.

[0046] S123: Sub-INS navigation parameter attitude update, the calculation formula is as follows.

[0047]

[0048] In the formula: —Angular increments in the XYZ directions of the sub-inertial navigation system; Represents the magnitude of the three components of a vector; —The pose quaternion from the previous update cycle; —The attitude quaternion updated recursively.

[0049] In this step, the calculation formulas for gravitational acceleration, entrapment acceleration, Coriolis acceleration, meridian radius, and zonal radius involved in updating Earth parameters are well known in the field, and the specific calculations are not described in detail.

[0050] Furthermore, in some embodiments, during the erection process, when the onboard reference inertial group data is valid, the process of transferring and aligning the data, updating the Kalman filter of the ground-based aircraft sub-inertial group, and correcting the navigation parameters of the ground-based aircraft sub-inertial group may include: S124: Calculate the Kalman filter state transition matrix, Kalman filter measurement vector, and Kalman filter measurement matrix based on the navigation parameters of the vehicle-mounted reference inertial group.

[0051] S125: Update the Kalman filter based on the Kalman filter state transition matrix, Kalman filter measurement vector, and Kalman filter measurement matrix.

[0052] S126: Correct the attitude parameters of the ground-based aircraft sub-inertial navigation system based on the updated Kalman filter.

[0053] Specifically, step S124 includes: S1241: Calculate the state transition matrix of the Kalman filter. The calculation formula is as follows.

[0054] First, calculate the apparent velocity increment of the navigation system:

[0055] State transition matrix Initialized as an 18-dimensional identity matrix, the other non-zero related terms of the real-time calculated state transition matrix are as follows:

[0056]

[0057] In the formula: —Additional angular velocity to the north-eastern horizon, calculated using the following formula:

[0058] S1242: Calculate the moving base transfer alignment Kalman filter measurement vector. The calculation formula is as follows.

[0059]

[0060] In the formula: —The attitude transformation matrix from the current system to the navigation system during the current update cycle; —Coordinate transformation matrix from the reference system to the reference navigation system; —The velocity of the reference inertial navigation system in the three directions of north, sky, and east; —The velocity of the inertial navigation system in the three directions of north, sky, and east.

[0061] S1243: Calculate the Kalman filter measurement matrix using the following formula.

[0062] Observation matrix Initialized as a 6×18 dimensional matrix, with the following non-zero elements:

[0063]

[0064] S125: Calculate the standard Kalman filter update using the following formula.

[0065]

[0066]

[0067] In the formula: —The state vector of the previous filtering cycle, an 18*1 vector; —The covariance matrix of the previous filtering cycle, an 18*18 vector.

[0068] S126: Correct the sub-INS attitude parameters based on the Kalman filter results. The calculation formula is as follows.

[0069] Will Misalignment angle Extract the data and perform feedback correction on the attitude used in the next calculation:

[0070] In the formula: —The attitude quaternion of the sub-inertial navigation system to be corrected; — This represents quaternion multiplication.

[0071] Will Extract the northward velocity, the upward velocity, and the eastward velocity from the data, and correct the current velocity accordingly:

[0072] Furthermore, in one embodiment, after the erection is completed, the alignment result transmitted by the ground-based aircraft sub-inertial navigation system is used as the initial attitude at the launch moment, and the Kalman filter during the flight phase is initialized, which may include: S21: After the vehicle is erected and in position, the navigation quaternion calculated at the time of alignment completion will be used as the initial value of the attitude quaternion at the launch time of the ground-based aircraft's sub-inertial navigation system.

[0073] S22: Transmit the geographic latitude, longitude, and altitude of the alignment completion time as the geographic latitude, longitude, and altitude of the launch time of the ground-based aircraft's sub-inertial navigation system.

[0074] S23: Set the initial state vector of the Kalman filter during the flight phase to a 21-dimensional zero-element vector, set the initial covariance matrix and process noise equation matrix of the Kalman filter to 21-dimensional diagonal matrices, set the measurement noise equation matrix of the Kalman filter to a 6-dimensional diagonal matrix, and initialize the observation matrix of the Kalman filter to a 6×21-dimensional matrix.

[0075] In this embodiment, in step S21 above, the flight attitude is initialized, and the calculation formula is as follows:

[0076] In the formula: —Transmit the navigation quaternion calculated at the alignment completion time.

[0077] S22: Initialization of velocity and position during the flight phase, calculated using the following formula:

[0078] In the formula: —This indicates the initial latitude, longitude, and altitude of the sub-inertial navigation system (INS). —This indicates the initial northeast-sky velocity of the sub-inertial navigation system (INS). —Transmit the geographical latitude, longitude, and altitude at the time of completion of the binding.

[0079] S23: Kalman filter initialization during flight phase, calculated as follows: The initial state vector for the Kalman filter is set to a 21-dimensional vector with zero elements, i.e.: .

[0080] Kalman filter initial covariance matrix It is a 21-dimensional diagonal matrix. The diagonal elements are shown in the table below, and the other off-diagonal elements are 0.

[0081] Table 4 Initial Covariance Matrix of Kalman Filter definition

[0082] Kalman filtering process noise equation matrix It is a 21-dimensional diagonal matrix. The diagonal elements are shown in the table below, and the other off-diagonal elements are 0.

[0083] Table 5 Noise variance matrix of Kalman filtering process definition

[0084] Kalman filter measurement noise equation matrix It is a 6-dimensional diagonal matrix. The diagonal elements are shown in the table below, and the other off-diagonal elements are 0.

[0085] Table 6. Kalman Filter Measurement Noise Variance Matrix definition

[0086] Kalman filter observation matrix Initialize as a 6×21 dimensional matrix, with the following non-zero elements.

[0087] .

[0088] Further, in one embodiment, in step S3, the step of performing Kalman filter-based integrated navigation using the velocity and position information provided by the onboard GNSS equipment after the launch of the ground-based aircraft, correcting the attitude deviation of the ground-based aircraft's sub-inertial navigation system, and completing alignment correction when the filter estimation meets the estimation threshold, may include: S31: During the flight phase of the ground-based aircraft, navigation calculations are performed on the ground-based aircraft's inertial navigation system to obtain attitude, velocity, and position.

[0089] S32: When the data from the GNSS device is valid, perform Kalman filter-based navigation based on the speed and position information of the GNSS device to correct the attitude deviation of the ground-based aircraft's sub-inertial navigation system.

[0090] S33: Based on the preset estimation threshold, when the filtered estimation meets the estimation threshold, the alignment correction process is completed, and subsequent navigation calculations are performed based on the alignment results.

[0091] Based on the above technical solution, the step of performing navigation calculations on the ground-based aircraft's sub-inertial navigation system during the flight phase to obtain attitude, velocity, and position may include: S311: The attitude of the ground-based aircraft's sub-inertial navigation system is updated using a quaternion chain multiplication rule. The input is the gyroscope increment, while the Earth's rotation angular rate is subtracted. The formula is as follows:

[0092] in: —The attitude quaternion updated recursively; —The pose quaternion before the recursive update.

[0093] Indicates the sampling period. This represents the gyroscope increment. Represents the angular rate of Earth's rotation, symbol This represents quaternion multiplication.

[0094] function The formula for solving quaternions from a rotating vector is as follows.

[0095]

[0096] function The formula for multiplying a quaternion by a three-dimensional vector is as follows.

[0097]

[0098] Furthermore, in step S31, the velocity of the ground-based aircraft sub-inertial navigation system is updated using the table increment, while the rotational error vector, gravitational acceleration vector, Gothic acceleration vector, and centripetal acceleration vector are subtracted.

[0099] S312: Velocity Update: Velocity update uses the addition of table increments, while simultaneously subtracting the rotational error vector, gravitational acceleration vector, Gothic acceleration vector, and centripetal acceleration vector, as shown in the following formula.

[0100]

[0101]

[0102] in: This represents the updated velocity vector. This represents the velocity vector before the update. This indicates the increment of the table. It represents the acceleration due to gravity. Represents the Earth's rotational angular rate. This represents the angular rate of the navigation frame relative to the Earth frame.

[0103] Furthermore, in step S31, the position of the ground-based aircraft sub-inertial navigation system is updated using a trapezoidal integral algorithm based on the velocity vector.

[0104] S313: Position Update: Position update is based on the velocity vector and uses the trapezoidal integral algorithm, as shown in the following formula.

[0105]

[0106] in: Indicates latitude, longitude, and altitude. Indicates the current update portfolio period. This indicates the next update cycle. This represents the radius of the Earth's meridian. This represents the radius of the Earth's geocentric circle.

[0107] The calculations for updating Earth parameters involved in this step are well-known in the field, and the specific calculations will not be described in detail.

[0108] In some embodiments, when the GNSS device data is valid, performing Kalman filter-based integrated navigation based on the speed and position information of the GNSS device to correct the attitude deviation of the ground-based aircraft sub-inertial navigation system may include: S321: Calculate the discretized state transition matrix based on the attitude, velocity, and position calculated by the ground-based aircraft's inertial navigation sub-group.

[0109] S322: Construct the measurement vector of the Kalman filter based on the velocity and position information of the GNSS device.

[0110] S323: Update the Kalman filter based on the measurement vector, and correct the navigation parameters of the ground-based aircraft sub-inertial navigation system based on the measurement update results of the Kalman filter.

[0111] Specifically, S321: Calculate the discretized state transition matrix. The calculation formula is the strapdown inertial navigation error equation, which is well known in the field, and the specific calculation is not described in detail.

[0112] S322: Construct the measurement vectors for the Kalman filter as follows:

[0113] In the formula: —The northeastern navigation speed of the aircraft's inertial navigation system.

[0114] —The aircraft's inertial navigation system's northeast celestial navigation position, consisting of latitude, longitude, and altitude.

[0115] — Indicates the speed and location of the GNSS device.

[0116] Following step S322, the method further includes: performing Kalman filter extrapolation. .

[0117] S323: Update the Kalman filter based on the measurement vector: .

[0118] Then, the navigation parameters are corrected based on the Kalman filter results: First, obtain the Kalman filter measurement update results. The first three elements of the vector are the misalignment angles. It is used to provide feedback and correction of attitude information. .

[0119] In the formula: — Convert the rotating vector into a quaternion function, see S311.

[0120] —The quaternion calculated during the current navigation cycle before correction.

[0121] —The quaternion calculated for the current navigation cycle after correction.

[0122] — The misalignment angle calculated during the current filter update cycle.

[0123] Secondly, take the Kalman filter measurement update results. The 4th to 6th elements in the vector represent the velocity error. This is used to provide feedback and correction for the navigation speed information in the northeast sky:

[0124] In the formula: —The speed calculated during the current navigation cycle before correction.

[0125] — The corrected speed.

[0126] Secondly, take the Kalman filter measurement update results. The 7th to 9th elements in the vector represent the positional error. It is used to provide feedback and correction for the navigation position information in Northeast China.

[0127] In the formula: —The position calculated during the current navigation cycle before correction.

[0128] — The corrected position.

[0129] Further, in one embodiment, based on a pre-set estimation threshold, the alignment correction process is completed when the filtered estimation meets the estimation threshold, and subsequent navigation calculations are performed based on the alignment results, including: S331: Calculate the first and second convergence variables of the Kalman filter.

[0130] S332: Determine whether each element of the first convergent variable is less than the first set value during a consecutive preset number of measurement update cycles; and determine whether the absolute value of the difference between the maximum and minimum values ​​of each row element of the second convergent variable is less than the second set value during a consecutive preset number of measurement update cycles; if the requirements of the first set value and the second set value are met simultaneously, then the Kalman filter is determined to be converged, the attitude correction condition is met, and the alignment is completed.

[0131] Specifically, in step S331 above, the Kalman filter convergence variable is calculated using the following formula.

[0132] After each measurement update, the following variables are calculated:

[0133] S332: Determine the convergence variable as follows: 10 consecutive measurement update cycles correspond to (That is, the first convergent variable) Each element is less than 3; 10 consecutive measurement update cycles (That is, the second convergent variable, the attitude error estimated by the Kalman filter) The absolute value of the difference between the maximum and minimum values ​​of each row element is less than 30; If both of the above conditions are met, the Kalman filter is determined to have converged, the attitude correction condition is satisfied, the second alignment is completed, and the two-stage alignment process proposed in this application is thus completed.

[0134] Compared to a single initial alignment method, this application utilizes data from two different stages, avoiding the problems of long alignment times and limited alignment accuracy associated with a single alignment method. By comprehensively using effective data before and after launch, the alignment accuracy is improved.

[0135] Compared to some algorithms in related technologies that divide the alignment process into two steps, nonlinear and linear, this application uses a linear model for both stages, which reduces computational complexity and combines speed and continuity.

[0136] This application addresses the challenge of simultaneously meeting accuracy and speed requirements during the initial alignment process of ground-based aircraft. It proposes a two-stage alignment method, utilizing data from auxiliary navigation sensors before and after launch, without disrupting the launch procedure. Before launch, during the ground-based aircraft's erection process, a vehicle-mounted reference inertial group (INS) performs attitude transfer alignment to the missile-borne INS. The vehicle-mounted INS is erected along with the launch pad, and once in position, the attitude transfer from the vehicle-mounted INS to the ground-based aircraft's sub-INS is complete. The ground-based aircraft's sub-INS uses the transferred alignment results as initial values ​​for post-launch navigation calculations. After launch, the attitude error from the initial alignment is estimated using velocity and position information provided by GNSS equipment. Once the estimated threshold is met, a second attitude correction is performed, completing the two-stage alignment process before and after launch.

[0137] The method proposed in this application utilizes comprehensive data, and the two-stage alignment process is characterized by speed and continuity. It has high accuracy and is easy to implement, thus possessing good engineering application value.

[0138] In the description of this application, it should be noted that the terms "upper," "lower," etc., indicating the orientation or positional relationship are based on the orientation or positional relationship shown in the accompanying drawings, and are only for the convenience of describing this application and simplifying the description, and do not indicate or imply that the device or element referred to must have a specific orientation, or be constructed and operated in a specific orientation, and therefore should not be construed as a limitation of this application. Unless otherwise expressly specified and limited, the terms "installed," "connected," and "linked" should be interpreted broadly. For example, they can refer to a fixed connection, a detachable connection, or an integral connection; they can refer to a mechanical connection or an electrical connection; they can refer to a direct connection or an indirect connection through an intermediate medium; they can refer to the internal communication between two elements. For those skilled in the art, the specific meaning of the above terms in this application can be understood according to the specific circumstances.

[0139] It should be noted that in this application, relational terms such as "first" and "second" are used merely to distinguish one entity or operation from another, and do not necessarily require or imply any such actual relationship or order between these entities or operations. Furthermore, the terms "comprising," "including," or any other variations thereof are intended to cover non-exclusive inclusion, such that a process, method, article, or apparatus that comprises a list of elements includes not only those elements but also other elements not expressly listed, or elements inherent to such a process, method, article, or apparatus. Without further limitations, an element defined by the phrase "comprising one..." does not exclude the presence of other identical elements in the process, method, article, or apparatus that includes said element.

[0140] The above description is merely a specific embodiment of this application, enabling those skilled in the art to understand or implement this application. Various modifications to these embodiments will be readily apparent to those skilled in the art, and the general principles defined herein may be implemented in other embodiments without departing from the spirit or scope of this application. Therefore, this application is not to be limited to the embodiments shown herein, but is to be accorded the widest scope consistent with the principles and novel features claimed herein.

Claims

1. A method for aligning a ground-based aircraft launch, characterized in that, It includes the following steps: During the erection of the ground-based aircraft, the Kalman filter of the ground-based aircraft sub-inertial group is updated and the navigation parameters of the sub-inertial group are corrected according to the navigation parameters set by the vehicle-mounted reference inertial group on the launcher, so that the vehicle-mounted reference inertial group can perform transfer alignment with the ground-based aircraft sub-inertial group. After the vehicle is erected, the alignment result transmitted by the ground-based aircraft sub-inertial navigation system is used as the initial attitude at the time of launch, and the Kalman filter during the flight phase is initialized. After the ground-based aircraft is launched, it uses the velocity and position information provided by the onboard GNSS equipment to perform Kalman filter combined navigation, corrects the attitude deviation of the ground-based aircraft's sub-inertial navigation system, and completes the alignment correction when the filter estimation meets the estimation threshold. During the erection process of the ground-based aircraft, the Kalman filter of the ground-based aircraft's sub-inertial navigation system is updated and the navigation parameters of the sub-inertial navigation system are corrected according to the navigation parameters set by the vehicle-mounted reference inertial navigation system on the launch pad, so that the vehicle-mounted reference inertial navigation system performs transfer alignment with the ground-based aircraft's sub-inertial navigation system, including: At the start of the ground-based aircraft's erection, the ground-based aircraft's sub-inertial navigation system and Kalman filter are initialized according to the navigation parameters set by the vehicle-mounted reference inertial navigation system. The navigation calculation is performed to update the attitude, velocity and position of the ground-based aircraft sub-inertial group. During the erection process, when the data of the vehicle-mounted reference inertial group is valid, it is transferred and aligned, the Kalman filter of the ground-based aircraft sub-inertial group is updated and the navigation parameters of the ground-based aircraft sub-inertial group are corrected. After the erection is completed, the alignment result transmitted by the ground-based aircraft's sub-inertial navigation system is used as the initial attitude at launch time. Simultaneously, the Kalman filter for the flight phase is initialized, including: After the vehicle is erected and in place, the navigation quaternion calculated at the time of alignment completion will be used as the initial value of the attitude quaternion at the launch time of the ground-based aircraft's sub-inertial navigation system. The geographic latitude, longitude, and altitude of the alignment completion time will be used as the geographic latitude, longitude, and altitude of the launch time of the ground-based aircraft's sub-inertial navigation system. The initial state vector of the Kalman filter during the flight phase is set to a 21-dimensional zero-element vector. The initial covariance matrix and process noise equation matrix of the Kalman filter are both set to 21-dimensional diagonal matrices. The measurement noise equation matrix of the Kalman filter is set to a 6-dimensional diagonal matrix. The observation matrix of the Kalman filter is initialized to a 6×21-dimensional matrix.

2. The ground-based aircraft launch alignment method as described in claim 1, characterized in that, During the erection process, when the onboard reference inertial group data is valid, it is transmitted and aligned, the Kalman filter of the ground-based aircraft sub-inertial group is updated, and the navigation parameters of the ground-based aircraft sub-inertial group are corrected, including: Calculate the Kalman filter state transition matrix, Kalman filter measurement vector, and Kalman filter measurement matrix based on the navigation parameters of the vehicle-mounted reference inertial group; Update the Kalman filter based on the Kalman filter state transition matrix, Kalman filter measurement vector, and Kalman filter measurement matrix; The attitude parameters of the ground-based aircraft sub-inertial navigation system are corrected based on the updated Kalman filter.

3. The ground-based aircraft launch alignment method as described in claim 1, characterized in that, After launch of the ground-based aircraft, Kalman filter-based integrated navigation is performed using velocity and position information provided by the onboard GNSS equipment to correct attitude deviations of the ground-based aircraft's sub-inertial navigation system. Alignment correction is completed when the filter estimation meets the estimation threshold, including: During the flight phase of the ground-based aircraft, navigation calculations are performed on the ground-based aircraft's inertial navigation system to obtain attitude, velocity, and position. When the data from the GNSS equipment is valid, Kalman filter combined navigation is performed based on the speed and position information of the GNSS equipment to correct the attitude deviation of the ground-based aircraft's sub-inertial navigation. Based on a pre-set estimation threshold, the alignment correction process is completed when the filtered estimation meets the estimation threshold, and subsequent navigation calculations are performed based on the alignment results.

4. The ground-based aircraft launch alignment method as described in claim 3, characterized in that, During the flight phase of the ground-based aircraft, navigation calculations are performed on the ground-based aircraft's inertial navigation system (INS) to obtain attitude, velocity, and position, including: The attitude of the ground-based aircraft's sub-inertial navigation system is updated using a quaternion chain multiplication rule, with the input being the gyroscope increment and the Earth's rotation angular rate being subtracted.

5. The ground-based aircraft launch alignment method as described in claim 4, characterized in that, The velocity of the ground-based aircraft sub-inertial navigation system is updated by adding table increments, while simultaneously subtracting the rotational error vector, gravitational acceleration vector, Gothic acceleration vector, and centripetal acceleration vector.

6. The ground-based aircraft launch alignment method as described in claim 4, characterized in that, Based on the velocity vector, the position of the ground-based aircraft's sub-inertial navigation system is updated using a trapezoidal integral algorithm.

7. The ground-based aircraft launch alignment method as described in claim 3, characterized in that, When the GNSS equipment data is valid, Kalman filter-based integrated navigation is performed based on the speed and position information of the GNSS equipment to correct the attitude deviation of the ground-based aircraft's sub-inertial navigation system, including: The discretized state transition matrix is ​​calculated based on the attitude, velocity, and position obtained from the sub-inertial navigation solution of the ground-based aircraft. The measurement vector of the Kalman filter is constructed based on the velocity and position information of the GNSS equipment; The Kalman filter is updated based on the measurement vector, and the navigation parameters of the ground-based aircraft sub-inertial navigation system are corrected based on the measurement update results of the Kalman filter.

8. The ground-based aircraft launch alignment method as described in claim 3, characterized in that, Based on a pre-set estimation threshold, the alignment correction process is completed when the filtered estimate meets the threshold, and subsequent navigation calculations are performed based on the alignment results, including: Calculate the first and second convergence variables of the Kalman filter; Determine whether each element of the first convergent variable is less than the first set value during a consecutive preset number of measurement update cycles; and determine whether the absolute value of the difference between the maximum and minimum values ​​of each row element of the second convergent variable is less than the second set value during a consecutive preset number of measurement update cycles. If both the first and second set values ​​are met simultaneously, the Kalman filter is determined to have converged, the attitude correction condition is satisfied, and alignment is complete.

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