A fast alignment method based on virtual gyroscope and accelerometer

Abstract

The invention discloses a rapid alignment method based on a virtual gyroscope and an accelerometer, comprising: a strapdown inertial navigation system performs coarse alignment under a carrier inertial navigation system according to initial position information of a bound aircraft, and constructs an initial attitude matrix of the strapdown inertial navigation system; a mathematical model of a virtual gyroscope and a virtual accelerometer is constructed, and virtual output values of the gyroscope and the accelerometer are obtained by using the collected output values of the gyroscope and the accelerometer through the mathematical model; the strapdown inertial navigation system performs inertial navigation solution using the virtual output values of the gyroscope and the accelerometer, including updating the attitude matrix, velocity, and position; a Kalman filter is used for fine alignment, and the attitude matrix is finely corrected using the fine alignment result; the finely corrected attitude matrix, the position, and the velocity are used for subsequent inertial navigation of the strapdown inertial navigation system. The method of the invention can significantly shorten the time for fine alignment and improve the alignment accuracy.

Description

Technical Field

[0001] The present invention relates to the field of strapdown inertial technology, and in particular to a rapid alignment method based on a virtual gyroscope and an accelerometer, which is suitable for rapid alignment of high-precision strapdown inertial navigation systems on various rocket-propelled aircraft before launch. Background Art

[0002] At present, the use of strapdown inertial navigation systems to perform attitude and full-course corrections on rocket-propelled aircraft can effectively increase their range and significantly improve their landing accuracy, becoming a major trend in the development of rocket-propelled aircraft.

[0003] With the advancement of technology, the cost of strapdown inertial navigation systems using medium and high precision fiber optic gyroscopes and accelerometers is becoming lower and lower, and the size is becoming smaller and smaller, making them more and more widely used in the field of airborne applications of rocket propulsion aircraft.

[0004] Strapdown inertial navigation systems require initial alignment during operation: determining the initial attitude matrix so that the mathematical platform coordinate system and the navigation coordinate system coincide. Initial alignment can be divided into coarse and fine phases; static and dynamic base alignment depending on the base's motion state; and autonomous and non-autonomous alignment depending on the degree of reliance on external information. Accuracy and time are the two most important technical indicators for initial alignment.

[0005] Strapdown inertial navigation systems (SINS) on rocket-propelled vehicles typically use a static base alignment scheme. This involves coarse alignment followed by fine alignment using a Kalman filter. This alignment requires real-time data collection from gyroscopes and accelerometers, and typically takes over 150 seconds. With the increasing demand for shorter preparation times for rocket-propelled vehicles, traditional SINS initial alignment schemes are no longer sufficient. Summary of the Invention

[0006] The purpose of the present invention is to provide a fast alignment method based on a virtual gyroscope and an accelerometer, so as to significantly shorten the time for fine alignment and improve the alignment longitude.

[0007] In order to achieve the above tasks, the present invention adopts the following technical solutions:

[0008] A fast alignment method based on a virtual gyroscope and an accelerometer, comprising:

[0009] The strapdown inertial navigation system performs coarse alignment of the aircraft in the inertial coordinate system according to the bound initial position information of the aircraft, and constructs the initial attitude matrix of the strapdown inertial navigation system;

[0010] Constructing a mathematical model of a virtual gyroscope and a virtual accelerometer, and using the collected output values of the gyroscope and the accelerometer to obtain virtual output values of the gyroscope and the accelerometer through the mathematical model;

[0011] The strapdown inertial navigation system uses the virtual output values of the gyroscope and accelerometer to perform inertial navigation calculations, including the update of attitude matrix, velocity and position;

[0012] A Kalman filter is used for fine alignment, and the attitude matrix is fine-corrected based on the fine alignment result; the fine-corrected attitude matrix and the position and velocity are used for subsequent inertial navigation of the strapdown inertial navigation system.

[0013] Furthermore, the rough alignment in the aircraft inertial coordinate system is performed to construct an initial attitude matrix of the strapdown inertial navigation system, including:

[0014] The attitude matrix is obtained by the direction cosine matrix between the geocentric inertial coordinate system and the navigation coordinate system. , the direction cosine matrix between the solidification inertial coordinate system and the aircraft inertial coordinate system , the direction cosine matrix between the geocentric inertial coordinate system and the solidification inertial coordinate system Multiply them together to get:

[0015] for , decompose it into the direction cosine matrix between the earth coordinate system and the navigation coordinate system , the direction cosine matrix between the geocentric inertial coordinate system and the earth coordinate system The multiplication form is obtained based on the longitude, latitude and time interval of the strapdown inertial navigation system;

[0016] for , which is determined by solving the direction cosine matrix differential equation using the gyroscope output;

[0017] for , obtained by using the projection of the velocity vector in the geocentric inertial coordinate system and the solidification inertial coordinate system.

[0018] Furthermore, the , obtained by projecting the velocity vector in the geocentric inertial coordinate system and the solidification inertial coordinate system, including:

[0019] First determine the projection of gravitational acceleration in the geocentric inertial coordinate system , and then in the time period between the current time and the initial time Integrate to obtain the relationship between the aircraft speed and time in the geocentric inertial coordinate system ; Using the output value of the accelerometer in the aircraft inertial coordinate system , get the current accelerometer velocity value ; According to the method of orthogonalization of attitude matrix, at different moments 、 The relationship between aircraft speed and time 、 , and the corresponding accelerometer velocity value 、 Obtain .

[0020] Furthermore, the mathematical model of the virtual gyroscope and the virtual accelerometer is constructed, and the virtual output values of the gyroscope and the accelerometer are obtained by using the collected output values of the gyroscope and the accelerometer through the mathematical model, including:

[0021] First, based on the output values of the gyroscope and accelerometer collected over a period of time, the mean and standard deviation are calculated as follows:

[0022] , ;

[0023] , ;

[0024] Where, For the Axis gyroscope The output value at the moment, ; are the three axes of the gyroscope; , Indicates the total number of sampling moments; For the Axis accelerometer Output value at each moment; and Respectively The mean and standard deviation of the output values of the axis gyro; and Respectively The mean and standard deviation of the output values of the axis accelerometer;

[0025] Construct mathematical models of the virtual gyroscope and virtual accelerometer based on the mean and standard deviation, and generate virtual output values:

[0026] ;

[0027] ;

[0028] Where, , is the preset noise amplification factor; The standard deviation is Random noise; The standard deviation is Random noise; For the Axis gyroscope The virtual output value at a moment, For the Axis accelerometer The virtual output value at a moment.

[0029] Furthermore, the strapdown inertial navigation system uses the virtual output values of the gyroscope and accelerometer to perform inertial navigation solution, wherein the quaternion method is used for attitude matrix update calculation, firstly, based on the angular increments in the three directions within the navigation cycle output by the gyroscope, the equivalent angular increments in the three directions are determined; then, the attitude quaternion is updated and normalized using the equivalent angular increments, and the attitude matrix is updated using the normalized quaternion.

[0030] Furthermore, the strapdown inertial navigation system uses the virtual output values of the gyroscope and accelerometer to perform inertial navigation solution, wherein the speed update method is:

[0031] ;

[0032] ;

[0033] Where, For a navigation cycle Velocity increment of the strapdown inertial navigation system in the navigation coordinate system; The virtual output value generated by the mathematical model of the virtual accelerometer Projection in the navigation coordinate system; is the projection of the Earth's rotation angular rate in the navigation coordinate system; is the angular rate of rotation of the navigation coordinate system caused by the motion of the aircraft; is the speed of the strapdown inertial navigation system at time k-1; is the projection of the gravity acceleration vector in the navigation coordinate system; is the speed of the strapdown inertial navigation system at the current moment k.

[0034] Furthermore, the strapdown inertial navigation system uses the virtual output values of the gyroscope and accelerometer to perform inertial navigation solution, wherein the position update method is:

[0035] ;

[0036] ;

[0037] ;

[0038] Where, 、 are the latitudes of the strapdown inertial navigation system at the current time k and time k-1 respectively; 、 are the longitudes of the strapdown inertial navigation system at the current time k and time k-1 respectively; 、 are the altitudes of the strapdown inertial navigation system at the current time k and time k-1 respectively; is the navigation solution cycle; 、 、 are the north, east and celestial velocities of the strapdown inertial navigation system at the current time k in the navigation coordinate system; 、 、 are the north, east and celestial velocities of the strapdown inertial navigation system at time k-1 in the navigation coordinate system, 、 They represent the meridian radius and the meridian radius respectively.

[0039] Furthermore, the fine alignment using the Kalman filter and the fine correction of the attitude matrix based on the fine alignment result include:

[0040] According to the error equation of the strapdown inertial navigation system, the state equation and measurement equation of the fine alignment process are established;

[0041] Construct a recursive equation for the discrete Kalman filter and perform iterative Kalman filtering operations through the recursive equation until the specified time is reached to obtain the three misalignment angle estimates after precise alignment.

[0042] The attitude matrix of the strapdown inertial navigation system is corrected using the misalignment angle estimation value, and the corrected attitude matrix is used for inertial navigation.

[0043] Furthermore, the state equation and measurement equation of the fine alignment process are expressed as:

[0044] ;

[0045] ;

[0046] The dot above the parameter represents the first-order derivative of the parameter. is the state variable, is the measured variable, is the state transition matrix, and are system noise and measurement noise respectively; is the measurement matrix, It is a noise driven array;

[0047] ;

[0048] ;

[0049] in, It represents the estimated misalignment angles in the east, north and sky directions in the navigation coordinate system. Indicates the velocity error in the east, north and sky directions in the navigation coordinate system. Indicates the longitude error, latitude error and altitude error in the position error. is the gyro constant drift error in the x, y, and z directions in the airborne coordinate system, is the constant bias error of the accelerometer in the x, y, and z directions in the airborne coordinate system.

[0050] Furthermore, the attitude matrix of the strapdown inertial navigation system is corrected using the misalignment angle estimation value, which is expressed as:

[0051] ;

[0052] ;

[0053] in, To use the misalignment angle estimate Constructed skew-symmetric matrix; is the corrected posture matrix; is the identity matrix; is the current posture matrix.

[0054] A terminal device comprises a processor, a memory and a computer program stored in the memory; when the processor executes the computer program, the fast alignment method based on a virtual gyroscope and an accelerometer is implemented.

[0055] A computer-readable storage medium stores a computer program; when the computer program is executed by a processor, the rapid alignment method based on a virtual gyroscope and an accelerometer is implemented.

[0056] Compared with the prior art, the present invention has the following technical features:

[0057] 1. In view of the fact that traditional precise alignment schemes have high alignment accuracy but long alignment time, the present invention utilizes the powerful computing power of the onboard computer to virtualize the outputs of the gyroscope and accelerometer and perform direct calculations, thus saving the time required to collect data from the gyroscope and accelerometer. While retaining the advantage of high alignment accuracy of the precise alignment scheme, it overcomes the disadvantage of long alignment time.

[0058] 2. In the traditional Kalman filter equation, the noise characteristics of the gyroscope and accelerometer are usually pre-set, but the strapdown inertial navigation system is affected by the environment when it is actually used, and sometimes a large deviation is produced. In the present invention, the gyroscope and accelerometer data collected in real time are used to calculate the noise characteristics, thus avoiding the large drawback of parameter deviation. BRIEF DESCRIPTION OF THE DRAWINGS

[0059] Figure 1 Schematic diagram of the process of the present invention;

[0060] Figure 2 Schematic diagram of coarse alignment of an aircraft inertial navigation system according to one embodiment of the present invention;

[0061] Figure 3 This is a working principle diagram of a navigation solution according to an embodiment of the present invention;

[0062] Figure 4 is an east misalignment angle alignment curve in one embodiment of the present invention;

[0063] Figure 5 is a north misalignment angle alignment curve in one embodiment of the present invention;

[0064] Figure 6 This is an alignment curve for the celestial misalignment angle in one embodiment of the present invention. DETAILED DESCRIPTION

[0065] The present invention provides a fast alignment method based on virtual gyroscope and accelerometer, see the attached Figure 1 ,include:

[0066] Step 1: The strapdown inertial navigation system performs coarse alignment in the aircraft inertial coordinate system according to the bound initial position information of the aircraft, and constructs the initial attitude matrix of the strapdown inertial navigation system.

[0067] First, bind the initial position information of the aircraft to the strapdown inertial navigation system, including longitude ,latitude and height Then, perform coarse alignment in the inertial navigation system to obtain the initial attitude matrix. Coarse alignment is an alignment method that uses the slow change of the projection of gravity acceleration on the aircraft's inertial coordinate system to extract the north information, which is expressed as:

[0068] ;

[0069] in, is the aircraft inertial coordinate system, which adopts the right front upper coordinate system, and the coordinate origin is located at the center of mass of the strapdown inertial navigation system; n is the navigation coordinate system, which adopts the northeast celestial coordinate system ENU, and the coordinate origin is located at the center of mass of the strapdown inertial navigation system; i is the geocentric inertial coordinate system, and the coordinate origin is located at the center of mass of the reference earth model; ib0 represents the solidified inertial coordinate system that is solidified with the aircraft inertial coordinate system at the initial moment; is the direction cosine matrix between the aircraft inertial coordinate system and the navigation coordinate system, that is, the attitude matrix; is the direction cosine matrix between the geocentric inertial coordinate system and the navigation coordinate system; is the direction cosine matrix between the solidification inertial coordinate system and the aircraft inertial coordinate system; is the direction cosine matrix between the geocentric inertial coordinate system and the solidification inertial coordinate system.

[0070] (1-1) Strapdown Inertial Navigation System Longitude ,latitude and time interval To determine:

[0071] ;

[0072] in is the Earth's rotation angular velocity, is the direction cosine matrix between the earth coordinate system e and the navigation coordinate system n; is the direction cosine matrix between the geocentric inertial coordinate system and the earth coordinate system; is the precision parameter, is the latitude, specifically the longitude in the initial location information above ,latitude Perform the solution; and Represent the current time and the initial time respectively.

[0073] (1-2) The direction cosine matrix is determined by solving the differential equation of the gyroscope output:

[0074] ;

[0075] ;

[0076] In the formula express The first-order differential of for The identity matrix of is the output value of the gyroscope; is a skew-symmetric matrix of output values.

[0077] (1-3) The velocity vector is projected onto the two coordinate systems i and ib0 to obtain:

[0078] As the Earth rotates, the acceleration due to gravity Projection in the geocentric inertial coordinate system is a quantity that changes with time and can be expressed as:

[0079] ;

[0080] In the time period Integrating the above formula, we can get the relationship between the speed of the aircraft and time in the geocentric inertial coordinate system: :

[0081] ;

[0082] The output value of the accelerometer is expressed in the aircraft inertial coordinate system as :

[0083] ;

[0084] Where, is the accelerometer bias, is the projection of the gravitational acceleration in the aircraft inertial coordinate system.

[0085] The output value of the accelerometer in the solid inertial coordinate system is , then the velocity value of the accelerometer at time t is for:

[0086] ;

[0087] make , then:

[0088] ;

[0089] It can be seen from the above formula that the direction cosine matrix between the solidification inertial coordinate system and the geocentric inertial coordinate system can be obtained through the velocity vectors in different coordinate systems.

[0090] According to the method of orthogonalization of the attitude matrix, at different moments 、 The relationship between aircraft speed and time 、 , and the corresponding accelerometer velocity value 、 , the formula can be constructed as follows:

[0091] ;

[0092] Here, the superscript represents the transposition operation. At this point, the initial attitude matrix of the strapdown inertial navigation system can be obtained: ,like Figure 2 shown.

[0093] Step 2: constructing a mathematical model of a virtual gyroscope and a virtual accelerometer, and using the collected output values of the gyroscope and the accelerometer to obtain virtual output values of the gyroscope and the accelerometer through the mathematical model.

[0094] (2-1) First, based on the output values of the gyroscope and accelerometer collected over a period of time, calculate the mean and standard deviation respectively:

[0095] , ;

[0096] , ;

[0097] Where, For the Axis gyroscope The output value at the moment, ; are the three axes of the gyroscope; , Indicates the total number of sampling moments; For the Axis accelerometer Output value at each moment; and Respectively The mean and standard deviation of the output values of the axis gyro; and Respectively The mean and standard deviation of the output values of the axis accelerometer.

[0098] (2-2) Construct mathematical models of the virtual gyroscope and virtual accelerometer based on the mean and standard deviation, and generate virtual output values:

[0099] ;

[0100] ;

[0101] Where, , is the preset noise amplification factor; The standard deviation is Random noise; The standard deviation is Random noise; For the Axis gyroscope The virtual output value at a moment, For the Axis accelerometer The virtual output value at a moment.

[0102] Step 3: The strapdown inertial navigation system uses the virtual output values of the gyroscope and accelerometer to perform inertial navigation solution, including the update of attitude matrix, velocity and position.

[0103] (3-1) Use the classic quaternion method to update the attitude matrix.

[0104] a) Calculation of equivalent angular increment.

[0105] ;

[0106] Where, is the navigation solution cycle; 、 、 is the equivalent angular increment in the x, y, and z directions during the navigation solution cycle; 、 、 The angular increments in the x, y, and z directions within the navigation cycle output by the gyroscope are respectively, and the virtual output values generated by the mathematical model of the virtual gyroscope in step 2 are Integrate to obtain; is latitude; 、 are the north velocity and east velocity in the navigation coordinate system; is the altitude of the strapdown inertial navigation system; is the Earth's rotation angular rate; and are the radii of the Earth's meridian and meridian respectively, is the posture matrix.

[0107] b) Update the attitude quaternion using the equivalent angle increment and normalize it.

[0108] ;

[0109] ;

[0110] ;

[0111] ;

[0112] Where, are the attitude quaternions of the strapdown inertial navigation system at time k and time k+1 respectively; and are the components of the attitude quaternion before and after normalization respectively; is the equivalent angular increment 、 、 The model; The equivalent angular increment 、 、 Constructed matrix; for The identity matrix of .

[0113] c) Update calculation of the posture matrix.

[0114] ;

[0115] ;

[0116] Where, 、 、 、 are the attitude quaternions at time k+1 respectively The four components of is the posture matrix.

[0117] (3-2) Speed update.

[0118] ;

[0119] ;

[0120] Where, For a navigation cycle Velocity increment of the strapdown inertial navigation system in the navigation coordinate system; The virtual output value generated by the mathematical model of the virtual accelerometer in step 2 Projection in the navigation coordinate system; is the projection of the Earth's rotation angular rate in the navigation coordinate system; is the angular rate of rotation of the navigation coordinate system caused by the motion of the aircraft; is the speed of the strapdown inertial navigation system at time k-1; is the projection of the gravity acceleration vector in the navigation coordinate system; is the velocity of the strapdown inertial navigation system at the current time k, which is decomposed into the navigation coordinate system to obtain its three-dimensional velocity in three directions.

[0121] (3-3) Position update.

[0122] ;

[0123] ;

[0124] ;

[0125] Where, 、 are the latitudes of the strapdown inertial navigation system at the current time k and time k-1 respectively; 、 are the longitudes of the strapdown inertial navigation system at the current time k and time k-1 respectively; 、 are the altitudes of the strapdown inertial navigation system at the current time k and time k-1 respectively; is the navigation solution cycle; 、 、 are the north, east and celestial velocities of the strapdown inertial navigation system at the current time k in the navigation coordinate system; 、 、 are the north, east and celestial velocities of the strapdown inertial navigation system at time k-1 in the navigation coordinate system, 、 They represent the meridian radius and the meridian radius respectively. Figure 3 shown.

[0126] Step 4: Perform fine alignment using a Kalman filter, and perform fine correction on the attitude matrix using the fine alignment result; the finely corrected attitude matrix and the position and velocity are used for subsequent inertial navigation of the strapdown inertial navigation system.

[0127] (4-1) Based on the error equation of the strapdown inertial navigation system, the state equation and measurement equation of the fine alignment process are established:

[0128] ;

[0129] ;

[0130] The dot above the parameter represents the first-order derivative of the parameter. is the state variable, is the measured variable, is the state transition matrix, and are system noise and measurement noise respectively; is the measurement matrix, The noise driven array.

[0131] ;

[0132] ;

[0133] in, It represents the estimated misalignment angles in the east, north and sky directions in the navigation coordinate system. Indicates the velocity error in the east, north and sky directions in the navigation coordinate system. Indicates the longitude error, latitude error and altitude error in the position error. is the gyro constant drift error in the x, y, and z directions in the airborne coordinate system, is the constant bias error of the accelerometer in the x, y, and z directions in the airborne coordinate system.

[0134] for dimensional state transfer matrix, which can be written as:

[0135] ;

[0136] in, express The zero matrix of is the posture matrix.

[0137] ;

[0138] Where, 、 are the celestial and north-projected components of the Earth's rotation angular velocity vector respectively; is the radius of the Earth; is latitude; 、 It is the projection component of the three-axis specific force of the triaxial accelerometer in the east, north and sky directions.

[0139] (4-2) Construct the recursive equation of the discrete Kalman filter and perform iterative Kalman filter operation through the recursive equation until the operation reaches the specified time to obtain the three misalignment angle estimates after precise alignment .

[0140] (4-3) Using the estimated value of the misalignment angle Attitude matrix of strapdown inertial navigation system Correction is performed and the corrected attitude matrix is used for inertial navigation, thereby performing combined navigation of Kalman filtering and inertial navigation.

[0141] Pose Matrix The correction method is:

[0142] ;

[0143] ;

[0144] in, To use the misalignment angle estimate Constructed skew-symmetric matrix; is the corrected posture matrix; is the identity matrix; is the current posture matrix.

[0145] The above technical solution of the present invention can construct a virtual gyroscope and a virtual accelerometer on the basis of coarse alignment, and use the output values of the virtual gyroscope and the virtual accelerometer for fine alignment through the powerful computing power of the onboard computer, thereby greatly shortening the fine alignment time and improving the alignment accuracy.

[0146] In order to verify the effect of the present invention relative to the traditional initial alignment, a medium-precision strapdown inertial navigation device is used, whose gyro bias stability is about 0.03° / h and accelerometer bias stability is about 0.2mg; after power-on, the output values of the gyroscope and accelerometer are collected in real time at 5ms intervals for a total of 200s of data; in the traditional initial alignment scheme, the first 20s of data are used for coarse alignment of the inertial system, and the last 180s of data are used for fine alignment; in the scheme of the present invention, the first 20s of data are used for coarse inertial alignment and a virtual gyroscope and accelerometer model is constructed, and its output is used for an equivalent 180s of fine alignment. The fine alignment process curves of the two schemes are shown as follows. Figures 4 to 6 As shown in the figure, it can be seen that the alignment accuracy of the two schemes is comparable, and the actual operation time of the precise alignment of the present invention is about 8s, which has a significant effect in shortening the time.

[0147] It is worth noting that: 1) the calculation scheme of the present invention significantly shortens the fine alignment calculation time by shortening the navigation solution cycle, shortening the Kalman filter cycle, and improving the computing power of the onboard computer; in contrast, traditional initial alignment schemes are constrained by the sampling time, and the alignment time is basically fixed. 2) In actual applications, if the gyroscope and accelerometer undergo a significant startup process or are significantly disturbed during the alignment process, the short sampling time of this scheme may cause the mean value of the gyroscope and accelerometer to deviate from the true value, resulting in reduced alignment accuracy.

[0148] The above embodiments are only used to illustrate the technical solutions of the present application, rather than to limit them. Although the present application has been described in detail with reference to the aforementioned embodiments, those skilled in the art should understand that they can still modify the technical solutions described in the aforementioned embodiments, or make equivalent replacements for some of the technical features therein. These modifications or replacements do not deviate the essence of the corresponding technical solutions from the spirit and scope of the technical solutions of the various embodiments of the present application, and should all be included in the scope of protection of the present application.

Claims

1. A fast alignment method based on a virtual gyroscope and an accelerometer, characterized in that: include: The strapdown inertial navigation system performs coarse alignment of the aircraft in the inertial coordinate system according to the bound initial position information of the aircraft, and constructs the initial attitude matrix of the strapdown inertial navigation system; Constructing a mathematical model of a virtual gyroscope and a virtual accelerometer, using the collected output values of the gyroscope and the accelerometer, and obtaining virtual output values of the gyroscope and the accelerometer through the mathematical model, including: First, based on the output values of the gyroscope and accelerometer collected over a period of time, the mean and standard deviation are calculated as follows: ， ； ， ； Where, For the Axis gyroscope The output value at the moment, ; are the three axes of the gyroscope; , Indicates the total number of sampling moments; For the Axis accelerometer Output value at each moment; and Respectively The mean and standard deviation of the output values of the axis gyro; and Respectively The mean and standard deviation of the output values of the axis accelerometer; Construct mathematical models of the virtual gyroscope and virtual accelerometer based on the mean and standard deviation, and generate virtual output values: ； ； Where, , is the preset noise amplification factor; The standard deviation is Random noise; The standard deviation is Random noise; For the Axis gyroscope The virtual output value at a moment, For the Axis accelerometer Virtual output value at a moment; The strapdown inertial navigation system uses the virtual output values of the gyroscope and accelerometer to perform inertial navigation calculations, including the update of attitude matrix, velocity and position; Performing fine alignment using a Kalman filter and fine-correcting the attitude matrix based on the fine alignment result, including: According to the error equation of the strapdown inertial navigation system, the state equation and measurement equation of the fine alignment process are established; Construct a recursive equation for the discrete Kalman filter and perform iterative Kalman filtering operations through the recursive equation until the specified time is reached to obtain the three misalignment angle estimates after precise alignment. The attitude matrix of the strapdown inertial navigation system is corrected using the misalignment angle estimation value; The attitude matrix after fine correction and the position and velocity are used for subsequent inertial navigation of the strapdown inertial navigation system.

2. The rapid alignment method based on virtual gyroscope and accelerometer according to claim 1, characterized in that: The rough alignment in the aircraft inertial coordinate system is performed to construct the initial attitude matrix of the strapdown inertial navigation system, including: The attitude matrix is obtained by the direction cosine matrix between the geocentric inertial coordinate system and the navigation coordinate system. , the direction cosine matrix between the solidification inertial coordinate system and the aircraft inertial coordinate system , the direction cosine matrix between the geocentric inertial coordinate system and the solidification inertial coordinate system Multiply them together to get: for , decompose it into the direction cosine matrix between the earth coordinate system and the navigation coordinate system , the direction cosine matrix between the geocentric inertial coordinate system and the earth coordinate system The multiplication form is obtained based on the longitude, latitude and time interval of the strapdown inertial navigation system; for , determined by solving the direction cosine matrix differential equation using the gyroscope output; for , obtained by using the projection of the velocity vector in the geocentric inertial coordinate system and the solidification inertial coordinate system.

3. The rapid alignment method based on virtual gyroscope and accelerometer according to claim 2, characterized in that: The said , obtained by projecting the velocity vector in the geocentric inertial coordinate system and the solidification inertial coordinate system, including: First determine the projection of gravitational acceleration in the geocentric inertial coordinate system , and then in the time period between the current time and the initial time Integrate to obtain the relationship between the aircraft speed and time in the geocentric inertial coordinate system ; Using the output value of the accelerometer in the aircraft inertial coordinate system , get the current accelerometer velocity value ; According to the method of orthogonalization of attitude matrix, at different moments 、 The relationship between aircraft speed and time 、 , and the corresponding accelerometer velocity value 、 Obtain .

4. The rapid alignment method based on virtual gyroscope and accelerometer according to claim 1, characterized in that: The strapdown inertial navigation system uses virtual output values of a gyroscope and an accelerometer to perform inertial navigation solution, wherein a quaternion method is used when performing attitude matrix update calculation. First, based on the angular increments in three directions within a navigation cycle output by the gyroscope, equivalent angular increments in the three directions are determined; then, the attitude quaternion is updated and normalized using the equivalent angular increments, and the attitude matrix is updated using the normalized quaternion.

5. The rapid alignment method based on virtual gyroscope and accelerometer according to claim 1, characterized in that: The strapdown inertial navigation system uses the virtual output values of the gyroscope and accelerometer to perform inertial navigation solution, where the speed update method is: ； ； Where, For a navigation cycle Velocity increment of the strapdown inertial navigation system in the navigation coordinate system; The virtual output value generated by the mathematical model of the virtual accelerometer Projection in the navigation coordinate system; is the projection of the Earth's rotation angular rate in the navigation coordinate system; is the angular rate of rotation of the navigation coordinate system caused by the motion of the aircraft; is the speed of the strapdown inertial navigation system at time k-1; is the projection of the gravity acceleration vector in the navigation coordinate system; is the speed of the strapdown inertial navigation system at the current moment k.

6. The rapid alignment method based on virtual gyroscope and accelerometer according to claim 1, characterized in that: The strapdown inertial navigation system uses the virtual output values of the gyroscope and accelerometer to perform inertial navigation solution, where the position update method is: ； ； ； Where, 、 are the latitudes of the strapdown inertial navigation system at the current time k and time k-1 respectively; 、 are the longitudes of the strapdown inertial navigation system at the current time k and time k-1 respectively; 、 are the altitudes of the strapdown inertial navigation system at the current time k and time k-1 respectively; is the navigation solution cycle; 、 、 are the north, east and celestial velocities of the strapdown inertial navigation system at the current time k in the navigation coordinate system; 、 、 are the north, east and celestial velocities of the strapdown inertial navigation system at time k-1 in the navigation coordinate system, 、 They represent the meridian radius and the meridian radius respectively.

7. The rapid alignment method based on virtual gyroscope and accelerometer according to claim 1, characterized in that: The state equation and measurement equation of the fine alignment process are expressed as: ； ； The dot above the parameter represents the first-order derivative of the parameter. is the state variable, is the measured variable, is the state transition matrix, and are system noise and measurement noise respectively; is the measurement matrix, It is a noise driven array; ； ； in, It represents the estimated misalignment angles in the east, north and sky directions in the navigation coordinate system. Indicates the velocity error in the east, north and sky directions in the navigation coordinate system. Indicates the longitude error, latitude error and altitude error in the position error. is the gyro constant drift error in the x, y, and z directions in the airborne coordinate system, is the constant bias error of the accelerometer in the x, y, and z directions in the airborne coordinate system.

8. A terminal device comprising a processor, a memory, and a computer program stored in the memory; characterized in that: When the processor executes the computer program, the fast alignment method based on the virtual gyroscope and accelerometer according to any one of claims 1 to 7 is implemented.

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