Carrier rocket navigation resolving method and system based on inertial measurement unit data

Abstract

The invention discloses a carrier rocket navigation resolving method and system based on inertial measurement unit data, and particularly relates to the field of resolving. A rocket strapdown inertial measurement unit is powered on and preheated, a navigation computer sends a'calibration parameter downloading 'request instruction to the inertial measurement unit, and the inertial measurement unit downloads calibration parameters pre-stored in FLASH to the navigation computer; then, the navigation computer calculates the apparent velocity increment and the angular increment of the rocket body relative to the inertial system in real time according to the total pulse number measured and output by the inertial unit and calibration parameters of the inertial unit; then, according to the initial launching point position information, an inertial measurement unit self-alignment process is started until the launching ignition moment is reached; generating initial non-levelness by inertial measurement unit self-alignment, and calculating an initial attitude matrix of the rocket by using the initial non-levelness at a launching ignition moment; constructing an SE (3) group according to the initial attitude matrix; and discretizing the SE (3) group by adopting a Picit iteration method, and updating a navigation solution equation.

Description

Technical Field

[0001] The present invention relates to the field of navigation solution technology, and more specifically, to a launch vehicle navigation solution method and system based on inertial group data. Background Art

[0002] During the flight of a launch vehicle, accurate navigation information is crucial to ensure that the rocket flies along the predetermined orbit. As an autonomous navigation system, the inertial navigation system does not rely on external information and can continuously and stably provide key motion parameters such as the position, velocity and attitude of the rocket relative to the inertial reference system throughout the flight of the rocket. These parameters are the basis of the rocket flight control system. The flight control system accurately adjusts the attitude and orbit of the rocket based on the information provided by the inertial navigation system to ensure that the rocket successfully sends the payload into the predetermined orbit and realizes the accurate deployment of satellites.

[0003] At present, the most widely used solution in the field of rocket inertial navigation is the navigation solution method based on quaternion update. Quaternion is used to represent the attitude information of the rocket. The quaternion is updated by the gyro output information of the inertial device to reflect the change of attitude, and other navigation information such as the position and speed of the rocket is further derived based on this.

[0004] However, it still has some shortcomings in actual use, such as the complex calculation process of the quaternion update navigation solution method, the instability of quaternion attitude control, and the inability to utilize multi-core and multi-threaded resources. Summary of the invention

[0005] In order to overcome the above-mentioned defects of the prior art, an embodiment of the present invention provides a launch vehicle navigation solution method and system based on inertial group data to solve the problems raised in the above-mentioned background technology.

[0006] To achieve the above object, the present invention provides the following technical solutions: Step A1: The rocket strapdown inertial group is powered on and preheated, and the navigation computer sends a "calibration parameter download" request command to the inertial group. The inertial group downloads the calibration parameters pre-stored in the FLASH to the navigation computer, and the navigation computer obtains the calibration parameters of the inertial group; Step A2: The navigation computer calculates in real time the apparent velocity increment and angular increment of the rocket body relative to the inertial system based on the total pulse number measured and output by the inertial group and the calibration parameters of the inertial group; Step A3: According to the initial launch point information, start the inertial group self-alignment process until the launch ignition time; Step A4: The inertial group self-alignment generates an initial non-level degree, and the initial non-level degree at the launch ignition time is used to calculate the rocket initial attitude matrix; Step A5: construct the SE(3) group according to the initial posture matrix; Step A6: Use the Picard iteration method to discretize the SE(3) group.

[0007] Preferably, in step A1, in the preparation stage before rocket launch, the strapdown inertial group will be powered on according to a predetermined timing; after power-on, the circuits, sensors and other components inside the inertial group will gradually enter a stable working state; in order to ensure that the internal components of the inertial group reach the best working state, preheating is performed; the length of the preheating time depends on the model of the inertial group and the working environment temperature; during the preheating process, the output of the inertial group is stable, reducing the error caused by temperature changes; When the IMU is preheated, the navigation computer sends a request command of "download calibration parameters" to the IMU through the communication interface; the command usually contains specific command code and check code information. After receiving the request command, the IMU performs check and confirmation operations; if the command is correct, the IMU is ready to download the calibration parameters, and the IMU sends confirmation information to the navigation computer; After confirming the download request, the inertial group reads the pre-stored calibration parameters from its built-in FLASH memory; the calibration parameters include the zero bias, scale factor, and installation error key information of the gyroscope and accelerometer; the inertial group transmits the read calibration parameters to the navigation computer through the communication interface.

[0008] Preferably, in step A2, the speed increment is calculated as follows:

[0009]

[0010] is the transformation matrix composed of the calibration parameters of the inertial group, is the difference between the full pulses of the accelerometer measured by the inertial group at two adjacent moments, T is the sampling period, is the accelerometer conversion coefficient of the inertial group calibration data, The accelerometer zero bias of the inertial group calibration data.

[0011] The calculation method of the angular increment is as follows:

[0012]

[0013] is the transformation matrix composed of the calibration parameters of the inertial group, is the difference between the full pulses of the gyroscope measured by the inertial group at two adjacent moments, T is the sampling period, is the accelerometer conversion coefficient of the inertial group calibration data, Gyro zero drift for IUG calibration data.

[0014] Preferably, in step A3, before the rocket is launched, accurate initial launch point information is obtained; the information includes the latitude and longitude of the launch point, the altitude and the component of the earth's rotation speed at this point; after the initial launch point information is obtained, the self-alignment process of the inertial group is started; During the inertial group self-alignment process, the measurement data of the inertial group is monitored in real time and adjusted as needed; the acceleration, angular velocity and other data output by the inertial group are monitored in real time to ensure the stability and accuracy of the data; error analysis is performed on the inertial group measurement data to identify and compensate for various error sources; based on the error analysis results, the calibration parameters of the inertial group and related parameters in the self-alignment algorithm are adjusted to improve the alignment accuracy; the measurement results of the inertial group are compared with the known high-precision navigation data to evaluate the alignment accuracy.

[0015] Preferably, in step A5, the method for calculating the posture transformation matrix is ​​specifically as follows: , in, It is represented as the posture transformation matrix, C is represented as the rotation part posture matrix, Expressed as the translational part vector, Represented as a 4×4 matrix space; The posture matrix C is a unit orthogonal matrix, and the posture matrix C satisfies the properties of the special orthogonal group SO(3); the posture transformation matrix is an orthogonal transformation matrix that satisfies the properties of the special Euclidean group SE(3), transforming the pose transformation matrix at any time The SE(3) group is: , in, Expressed as the attitude matrix of the rocket relative to the inertial system, Represented as an n×m matrix space, To represent n-dimensional vector space, Represented as a pose transformation matrix; Then SO(3) is: , Among them, SO(3) represents the special orthogonal group and C represents the posture matrix.

[0016] The specific calculation method of the differential equation of the SE(3) group is: ,in, Expressed as a differential equation, Expressed as a pose transformation matrix, Represented as a total vector, the operator Expressed as a Mapping; The specific calculation method is: , in, Indicated by The antisymmetric matrix of the elements; Expressed as the translational part vector, Represented as a 4×4 matrix space; The specific calculation method is: , Indicated by The antisymmetric matrix composed of elements, Represented as a 3×3 matrix space; The calculation method of the total vector is as follows: , in, Expressed as a total vector, Represents the velocity increment vector of the translation part , measured by the inertial group accelerometer and converted through calibration parameters; Represents the rotation part angle increment vector , measured by the inertial group gyroscope and converted through calibration parameters.

[0017] Preferably, in step A6, the differential equation calculation method of the discretized SE(3) group is specifically as follows: , in Expressed as the pose transformation matrix at time k, Represented as a matrix is the exponential function of the exponent, It is represented as the posture transformation matrix at time k+1; in, The specific method of dividing the form characteristics into blocks is as follows: exist Part, using Taylor series expansion, the expansion calculation method is as follows: , in, Represented as a matrix is the exponential function of the exponent; ,in for The module length, For The length of the overlap is 1 unit vector; Translational part is the increment of the arrow's apparent velocity from time k to time k+1. Since the translation update is after the rotation, it needs to be multiplied on the left. Make the conversion; Matrix The calculation method of the exponential function of exponent is as follows: , in, Represented as a matrix is the exponential function of the exponent, Expressed as The corresponding rotation part represents the rotation matrix, Expressed as the translational part vector, Transformation Matrix The update equation, that is, the navigation solution equation, then the calculation method of the navigation solution equation is as follows: , in, Expressed as the posture matrix at time k+1, Expressed as the translational part vector at time k+1, Expressed as the posture matrix at time k, Expressed as the translational part vector at time k, Expressed as The corresponding rotation part represents the rotation matrix; The navigation solution equation is updated, and the update method is as follows: Time 0 means navigation initialization. Initialize according to the rocket's initial attitude matrix; the translational vector uses the velocity increment vector measured by the inertial group at time 0. .

[0018] The present invention also provides a carrier rocket navigation solution system based on inertial group data, using the carrier rocket navigation solution method based on inertial group data as described above, the solution system comprises: Calibration parameter acquisition module: used for power-on preheating of the rocket strapdown inertial group. The navigation computer sends a "calibration parameter download" request command to the inertial group. The inertial group downloads the calibration parameters pre-stored in the FLASH to the navigation computer, and the navigation computer obtains the calibration parameters of the inertial group. Calculation increment module: used by the navigation computer to calculate the apparent velocity increment and angular increment of the rocket body relative to the inertial system in real time based on the total pulse number measured and output by the inertial group and the calibration parameters of the inertial group; Startup module: used to start the inertial group self-alignment process according to the initial launch point information until the launch ignition moment; Rocket initial attitude matrix calculation module: used for inertial group self-alignment to generate initial non-levelness, and uses the initial non-levelness at the launch ignition moment to calculate the rocket initial attitude matrix; SE(3) group construction module: constructs the SE(3) group according to the initial posture matrix; Discretization module: used to discretize the SE(3) group using the Picard iteration method.

[0019] Technical effects and advantages of the present invention: 1. The position and attitude integrated navigation solution method based on the special Euclidean group representation integrates the two-step calculation of the traditional strapdown inertial navigation attitude and position solution into one-step calculation, avoiding the complexity of the calculation process, and avoiding the quaternion unitization and attitude control dispersion phenomenon; 2. Pose transformation matrix Iterative updates use the navigation computer CPU multi-core and multi-threaded parallel computing to improve computing efficiency and real-time performance. BRIEF DESCRIPTION OF THE DRAWINGS

[0020] Figure 1 It is a schematic diagram of the method flow of the present invention.

[0021] Figure 2 It is a schematic diagram of multi-threaded parallel computing of the present invention.

[0022] Figure 3 It is a schematic diagram of module connection of the present invention. DETAILED DESCRIPTION

[0023] The following will be combined with the drawings in the embodiments of the present invention to clearly and completely describe the technical solutions in the embodiments of the present invention. Obviously, the described embodiments are only part of the embodiments of the present invention, not all of the embodiments. Based on the embodiments of the present invention, all other embodiments obtained by ordinary technicians in this field without creative work are within the scope of protection of the present invention.

[0024] See also Figure 1 As shown, the present invention provides a launch vehicle navigation solution method based on inertial group data, and the method is: Step A1: The rocket strapdown inertial group is powered on and preheated, and the navigation computer sends a "calibration parameter download" request command to the inertial group. The inertial group downloads the calibration parameters pre-stored in the FLASH to the navigation computer, and the navigation computer obtains the calibration parameters of the inertial group; In the step A1, in the preparation stage before the rocket launch, the strapdown inertial group will be powered on according to a predetermined timing; after powering on, the circuits, sensors and other components inside the inertial group will gradually enter a stable working state; in order to ensure that the internal components of the inertial group reach the best working state, preheating is performed; the length of the preheating time depends on the model of the inertial group and the working environment temperature; during the preheating process, the output of the inertial group is stable, reducing the error caused by temperature changes; When the IMU is preheated, the navigation computer sends a request command of "download calibration parameters" to the IMU through the communication interface; the command usually contains specific command code and check code information. After receiving the request command, the IMU performs check and confirmation operations; if the command is correct, the IMU is ready to download the calibration parameters, and the IMU sends confirmation information to the navigation computer; After confirming the download request, the inertial group reads the pre-stored calibration parameters from its built-in FLASH memory; the calibration parameters include the zero bias, scale factor, and installation error key information of the gyroscope and accelerometer; the inertial group transmits the read calibration parameters to the navigation computer through the communication interface.

[0025] Step A2: The navigation computer calculates in real time the apparent velocity increment and angular increment of the rocket body relative to the inertial system based on the total pulse number measured and output by the inertial group and the calibration parameters of the inertial group; In step A2, the speed increment is calculated as follows:

[0026]

[0027] is the transformation matrix composed of the calibration parameters of the inertial group, is the difference between the full pulses of the accelerometer measured by the inertial group at two adjacent moments, T is the sampling period, is the accelerometer conversion coefficient of the inertial group calibration data, The accelerometer zero bias of the inertial group calibration data.

[0028] The calculation method of the angular increment is as follows:

[0029]

[0030] is the transformation matrix composed of the calibration parameters of the inertial group, is the difference between the full pulses of the gyroscope measured by the inertial group at two adjacent moments, T is the sampling period, is the accelerometer conversion coefficient of the inertial group calibration data, Gyro zero drift for IUG calibration data.

[0031] Step A3: According to the initial launch point information, start the inertial group self-alignment process until the launch ignition time; In step A3, before the rocket is launched, accurate initial launch point information is obtained; the information includes the latitude and longitude of the launch point, the altitude, and the component of the earth's rotation speed at this point; after the initial launch point information is obtained, the self-alignment process of the inertial group is started; During the IMU self-alignment process, the IMU measurement data is monitored in real time and adjusted as needed; the acceleration, angular velocity and other data output by the IMU are monitored in real time to ensure the stability and accuracy of the data; error analysis is performed on the IMU measurement data to identify and compensate for various error sources; based on the error analysis results, the calibration parameters of the IMU and related parameters in the self-alignment algorithm are adjusted to improve the alignment accuracy; the IMU measurement results are compared with known high-precision navigation data to evaluate the alignment accuracy; Step A4: The inertial group self-alignment generates an initial non-level degree, and the initial non-level degree at the launch ignition time is used to calculate the rocket initial attitude matrix; In step A4, the calculation method of the rocket initial attitude matrix is ​​specifically as follows: , in, Expressed as the initial posture matrix, Expressed as initial non-levelness.

[0032] Step A5: construct the SE(3) group according to the initial posture matrix; In step A5, the method for calculating the posture transformation matrix is ​​specifically as follows: , in, It is represented as the posture transformation matrix, C is represented as the rotation part posture matrix, Expressed as the translational part vector, Represented as a 4×4 matrix space; The posture matrix C is a unit orthogonal matrix, and the posture matrix C satisfies the properties of the special orthogonal group SO(3); the posture transformation matrix is an orthogonal transformation matrix that satisfies the properties of the special Euclidean group SE(3), transforming the pose transformation matrix at any time The SE(3) group is: , in, Expressed as the attitude matrix of the rocket relative to the inertial system, Represented as an n×m matrix space, To represent n-dimensional vector space, Represented as a pose transformation matrix; Then SO(3) is: , Among them, SO(3) represents the special orthogonal group, and C represents the posture matrix; The specific calculation method of the differential equation of the SE(3) group is: ,in, Expressed as a differential equation, Expressed as a pose transformation matrix, Represented as a total vector, the operator Expressed as a Mapping; The specific calculation method is: , in, Indicated by The antisymmetric matrix of the elements; Expressed as the translational part vector, Represented as a 4×4 matrix space; The specific calculation method is: , Indicated by The antisymmetric matrix composed of elements, Represented as a 3×3 matrix space; The calculation method of the total vector is as follows: , in, Expressed as a total vector, Represents the velocity increment vector of the translation part , measured by the inertial group accelerometer and converted through calibration parameters; Represents the rotation part angle increment vector , measured by the inertial group gyroscope and converted through calibration parameters.

[0033] Step A6: discretize the SE(3) group using the Picard iteration method; In step A6, the calculation method of the differential equation of the discretized SE(3) group is specifically as follows: , in Expressed as the pose transformation matrix at time k, Represented as a matrix is the exponential function of the exponent, It is represented as the posture transformation matrix at time k+1; in, The specific method of dividing the form characteristics into blocks is as follows: exist Part, using Taylor series expansion, the expansion calculation method is as follows: , in, Represented as a matrix is the exponential function of the exponent; ,in for The module length, For The length of the overlap is 1 unit vector; For a unit vector The antisymmetric matrix It has the following properties:

[0034] According to the antisymmetric matrix Nature, will Further expansion, the calculation method is as follows:

[0035]

[0036]

[0037]

[0038]

[0039]

[0040]

[0041]

[0042] Substituting the expanded result into , we can get: ,in, The corresponding rotation part represents the rotation matrix, which is recorded as: ; Translational part is the increment of the arrow's apparent velocity from time k to time k+1. Since the translation update is after the rotation, it needs to be multiplied on the left. Make the conversion; Matrix The calculation method of the exponential function of exponent is as follows: , in, Represented as a matrix is the exponential function of the exponent, Expressed as The corresponding rotation part represents the rotation matrix, Expressed as the translational part vector, Transformation Matrix The update equation, that is, the navigation solution equation, then the calculation method of the navigation solution equation is as follows: , in, Expressed as the posture matrix at time k+1, Expressed as the translational part vector at time k+1, Expressed as the posture matrix at time k, Expressed as the translational part vector at time k, Expressed as The corresponding rotation part represents the rotation matrix; The navigation solution equation is updated, and the update method is as follows: Time 0 means navigation initialization. Initialize according to the rocket's initial attitude matrix; the translational vector uses the velocity increment vector measured by the inertial group at time 0. .

[0043] Example 1 Use the two CPU cores of the navigation computer to start a thread for calculation. Thread A in core A calculates Update, thread B in core B calculates After both threads have finished calculating, the calculation results of the two threads will be and Sent to the navigation calculation thread for subsequent navigation calculations. The timing of the calculation process is as follows Figure 2 shown.

[0044] Finish and Update, further by The attitude angle of the rocket body in the inertial system is derived from The velocity and position in the inertial system are derived by integration to complete an inertial navigation solution;

[0045] in, express The element in the mth row and nth column of represents the navigation cycle, Represents the gravitational acceleration.

[0046] See also Figure 2 As shown, the present invention also provides a carrier rocket navigation solution system based on inertial group data, using the carrier rocket navigation solution method based on inertial group data as described above, the solution system comprises: Calibration parameter acquisition module: used for power-on preheating of the rocket strapdown inertial group. The navigation computer sends a "calibration parameter download" request command to the inertial group. The inertial group downloads the calibration parameters pre-stored in the FLASH to the navigation computer, and the navigation computer obtains the calibration parameters of the inertial group. Calculation increment module: used by the navigation computer to calculate the apparent velocity increment and angular increment of the rocket body relative to the inertial system in real time based on the total pulse number measured and output by the inertial group and the calibration parameters of the inertial group; Startup module: used to start the inertial group self-alignment process according to the initial launch point information until the launch ignition moment; Rocket initial attitude matrix calculation module: used for inertial group self-alignment to generate initial non-levelness, and uses the initial non-levelness at the launch ignition moment to calculate the rocket initial attitude matrix; SE(3) group calculation module: calculates the SE(3) group based on the initial posture matrix; Discretization module: used to discretize the SE(3) group using the Picard iteration method.

[0047] Finally: The above description is only a preferred embodiment of the present invention and is not intended to limit the present invention. Any modifications, equivalent substitutions, improvements, etc. made within the spirit and principles of the present invention should be included in the protection scope of the present invention.

Claims

1. A launch vehicle navigation solution method based on inertial group data, characterized in that: include: Step A1: The rocket strapdown inertial group is powered on and preheated, and the navigation computer sends a "calibration parameter download" request command to the inertial group. The inertial group downloads the calibration parameters pre-stored in the FLASH to the navigation computer, and the navigation computer obtains the calibration parameters of the inertial group; Step A2: The navigation computer calculates in real time the apparent velocity increment and angular increment of the rocket body relative to the inertial system based on the total pulse number measured and output by the inertial group and the calibration parameters of the inertial group; Step A3: According to the initial launch point information, start the inertial group self-alignment process until the launch ignition time; Step A4: The inertial group self-alignment generates an initial non-level degree, and the initial non-level degree at the launch ignition time is used to calculate the rocket initial attitude matrix; Step A5: construct the SE(3) group according to the initial posture matrix; Step A6: Use the Picard iteration method to discretize the SE(3) group.

2. The method for solving launch vehicle navigation based on inertial group data according to claim 1, characterized in that: In step A1, during the preparation stage before rocket launch, the strapdown inertial group will be powered on according to a predetermined sequence; after power-on, the circuits, sensors and other components inside the inertial group will gradually enter a stable working state; in order to ensure that the internal components of the inertial group reach the best working state, preheating is performed; the length of the preheating time depends on the model of the inertial group and the working environment temperature; during the preheating process, the output of the inertial group is stable, reducing errors caused by temperature changes.

3. The method for solving launch vehicle navigation based on inertial group data according to claim 1, characterized in that: In step A2, the speed increment is calculated as follows: ； ； is the transformation matrix composed of the calibration parameters of the inertial group, is the difference between the full pulses of the accelerometer measured by the inertial group at two adjacent moments, T is the sampling period, is the accelerometer conversion coefficient of the inertial group calibration data, Calibrate the accelerometer bias data for the inertial group; The calculation method of the angular increment is as follows: ； ； is the transformation matrix composed of the calibration parameters of the inertial group, is the difference between the full pulses of the gyroscope measured by the inertial group at two adjacent moments, T is the sampling period, is the accelerometer conversion coefficient of the inertial group calibration data, Gyro zero drift for IUG calibration data.

4. The method for solving launch vehicle navigation based on inertial group data according to claim 1, characterized in that: In step A3, before the rocket is launched, accurate initial launch point information is obtained; the information includes the latitude and longitude of the launch point, the altitude, and the component of the earth's rotation speed at that point; After obtaining the initial launch point information, start the self-alignment process of the inertial group.

5. The method for solving launch vehicle navigation based on inertial group data according to claim 1, characterized in that: In step A4, the calculation method of the rocket initial attitude matrix is ​​specifically as follows: ， in, Expressed as the initial posture matrix, Expressed as initial non-levelness.

6. The method for solving launch vehicle navigation based on inertial group data according to claim 1, characterized in that: In step A5, the method for constructing the posture transformation matrix is ​​specifically as follows: ， in, It is represented as the posture transformation matrix, C is represented as the rotation part posture matrix, Expressed as the translational part vector, Represented as a 4×4 matrix space; The posture matrix C is a unit orthogonal matrix, and the posture matrix C satisfies the properties of the special orthogonal group SO(3); the posture transformation matrix is an orthogonal transformation matrix that satisfies the properties of the special Euclidean group SE(3), transforming the pose transformation matrix at any time The SE(3) group is: ， in, Expressed as the attitude matrix of the rocket relative to the inertial system, Represented as an n×m matrix space, To represent n-dimensional vector space, Represented as a pose transformation matrix; Then SO(3) is: ， Among them, SO(3) represents the special orthogonal group and C represents the posture matrix.

7. The method for calculating launch vehicle navigation based on inertial group data according to claim 6, characterized in that: The specific calculation method of the differential equation of the SE(3) group is: ,in, Expressed as a differential equation, Expressed as a pose transformation matrix, Represented as a total vector, the operator Expressed as a Mapping; The specific calculation method is: ， in, Indicated by The antisymmetric matrix of the elements; Expressed as the translational part vector, Represented as a 4×4 matrix space; The specific calculation method is: ， Indicated by The antisymmetric matrix composed of elements, Represented as a 3×3 matrix space; The calculation method of the total vector is as follows: ， in, Expressed as a total vector, Represents the velocity increment vector of the translation part , measured by the inertial group accelerometer and converted through calibration parameters; Represents the rotation part angle increment vector , measured by the inertial group gyroscope and converted through calibration parameters.

8. The method for solving launch vehicle navigation based on inertial group data according to claim 1, characterized in that: In step A6, the calculation method of the differential equation of the discretized SE(3) group is specifically as follows: ， in Expressed as the pose transformation matrix at time k, Represented as a matrix is the exponential function of the exponent, It is represented as the posture transformation matrix at time k+1; in, The specific method of dividing the form characteristics into blocks is as follows: exist Part, using Taylor series expansion, the expansion calculation method is as follows: ， in, Represented as a matrix is the exponential function of the exponent; ,in for The module length, For The length of the overlap is 1 unit vector; Translational part is the increment of the arrow's apparent velocity from time k to time k+1. Since the translation update is after the rotation, it needs to be multiplied on the left. Make the conversion; Matrix The calculation method of the exponential function of exponent is as follows: ， in, Represented as a matrix is the exponential function of the exponent, Expressed as The corresponding rotation part represents the rotation matrix, Expressed as the translational part vector, Transformation Matrix The update equation, that is, the navigation solution equation, then the calculation method of the navigation solution equation is as follows: ， in, Expressed as the posture matrix at time k+1, Expressed as the translational part vector at time k+1, Expressed as the posture matrix at time k, Expressed as the translational part vector at time k, Expressed as The corresponding rotation part represents the rotation matrix; The navigation solution equation is updated, and the update method is as follows: Time 0 means navigation initialization. Initialize according to the rocket's initial attitude matrix; the translational vector uses the velocity increment vector measured by the inertial group at time 0. .

9. A launch vehicle navigation solution system based on inertial group data, using a launch vehicle navigation solution method based on inertial group data as claimed in any one of claims 1 to 8, characterized in that: The solution system comprises: Calibration parameter acquisition module: used for power-on preheating of the rocket strapdown inertial group. The navigation computer sends a "calibration parameter download" request command to the inertial group. The inertial group downloads the calibration parameters pre-stored in the FLASH to the navigation computer, and the navigation computer obtains the calibration parameters of the inertial group. Calculation increment module: used by the navigation computer to calculate the apparent velocity increment and angular increment of the rocket body relative to the inertial system in real time based on the total pulse number measured and output by the inertial group and the calibration parameters of the inertial group; Startup module: used to start the inertial group self-alignment process according to the initial launch point information until the launch ignition moment; Rocket initial attitude matrix calculation module: used for inertial group self-alignment to generate initial non-levelness, and uses the initial non-levelness at the launch ignition moment to calculate the rocket initial attitude matrix; SE(3) group construction module: constructs the SE(3) group according to the initial posture matrix; Discretization module: used to discretize the SE(3) group using the Picard iteration method.

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