A pseudo-linear strapdown inertial navigation solution updating method

Abstract

The application discloses a pseudo-linear strapdown inertial navigation solution updating method, which comprises the following steps: S1, acquiring strapdown inertial navigation related data; S2, updating and solving the strapdown inertial navigation related data by using a pseudo-linear strapdown inertial navigation data updating method to obtain an updating result of the strapdown inertial navigation; and S3, improving navigation precision according to the updating result of the strapdown inertial navigation. The application adopts a navigation parameter vector model to uniformly solve navigation, and has high solution precision.

Description

TECHNICAL FIELD

[0001] The present application relates to the field of strapdown inertial navigation technology, and particularly relates to a pseudo-linear strapdown inertial navigation solution updating method. BACKGROUND

[0002] Traditional strapdown inertial navigation solution methods are all attitude solutions using attitude matrix, quaternion, Rodrigues parameter or Euler angle, and then converting the accelerometer output in the carrier coordinate system to the navigation coordinate system according to the corresponding coordinate conversion calculation rules. Since the above attitude parameters are not defined in the linear space, the velocity update equation is nonlinear, which is not convenient for directly using the strapdown inertial navigation system model for error analysis, filtering and other operations. SUMMARY

[0003] The present application aims to provide a pseudo-linear strapdown inertial navigation solution updating method, which uses a navigation parameter vector model for unified navigation solution, and has high solution accuracy.

[0004] The technical solution of the present application to solve the above technical problems is as follows:

[0005] The present application provides a pseudo-linear strapdown inertial navigation solution updating method, which comprises:

[0006] S1: obtaining strapdown inertial related data;

[0007] S2: updating and solving the strapdown inertial related data using a pseudo-linear strapdown inertial data updating method to obtain the updating result of the strapdown inertial navigation;

[0008] S3: improving the navigation accuracy according to the updating result of the strapdown inertial navigation.

[0009] Optionally, the strapdown inertial related data comprises attitude differential equation, velocity differential equation and position differential equation.

[0010] The pseudo-linear strapdown inertial data updating method comprises:

[0011] S21: pseudo-linearizing the attitude differential equation to obtain a pseudo-linear attitude update equation;

[0012] S22: pseudo-linearizing the velocity differential equation to obtain a pseudo-linear velocity update equation;

[0013] S23: pseudo-linearizing the position differential equation to obtain a pseudo-linear position update equation;

[0014] S24: obtaining a navigation parameter vector differential equation according to the pseudo-linear attitude update equation, the pseudo-linear velocity update equation and the pseudo-linear position update equation.

[0015] S25: performing an updating operation on the navigation parameter vector differential equation to obtain an updated navigation parameter vector;

[0016] S26: obtaining an updating result of the strapdown inertial navigation system according to the updated navigation parameter vector.

[0017] Optionally, in the step S21, the pseudo-linear attitude updating equation is:

[0018]

[0019] wherein A 11 represents an attitude quaternion conversion system matrix and is an output angular velocity of a gyroscope, is a navigation coordinate system rotation angular velocity, functions M L (·) and M R (·) respectively represent a matrix form of a left multiplication quaternion and a matrix form of a right multiplication quaternion in quaternion multiplication, and

[0020] represents an attitude quaternion, i represents an inertial coordinate system, b represents a carrier coordinate system, n represents a navigation coordinate system, q0, q1, q2 and q3 respectively represent four elements of the quaternion.

[0021] Optionally, the step S22 comprises:

[0022] S221: obtaining the velocity differential equation of the strapdown inertial navigation system;

[0023] S222: performing a transformation on a velocity vector in the velocity differential equation to obtain a velocity quaternion;

[0024] S223: obtaining the pseudo-linear velocity updating equation according to the velocity quaternion and the attitude quaternion.

[0025] Optionally, in the step S22, the pseudo-linear velocity updating equation is:

[0026]

[0027] wherein A 21 and A 22 respectively represent a conversion system matrix from the attitude quaternion to the velocity quaternion and a velocity quaternion conversion system matrix, and represents an output angular velocity of a gyroscope, and function ML (·) and M R (·) represent the matrix form of the left-multiplying quaternion and the matrix form of the right-multiplying quaternion in quaternion multiplication respectively and represents the specific force of the accelerometer output, represents the sum of the centripetal acceleration and the Coriolis acceleration and and respectively represent the earth rotation angular velocity and the rotation angular velocity of the navigation system relative to the earth system caused by the carrier motion, v n represents the velocity vector in the navigation system, g n represents the gravity acceleration vector, represents the multiplication operation of the quaternion, represents the rotation speed of the navigation system relative to the inertial system, represents the attitude quaternion, μ n represents the velocity quaternion and b represents the carrier coordinate system and n represents the navigation coordinate system.

[0028] Optionally, the step S23 comprises:

[0029] S231: obtaining a position vector according to the latitude, the longitude and the height of the carrier;

[0030] S232: obtaining a constant matrix in an update period according to the position differential equation and the velocity vector;

[0031] S233: obtaining a conversion system matrix of the attitude quaternion to the position quaternion according to the constant matrix, the position vector and the rotation angular velocity of the navigation coordinate system;

[0032] S234: obtaining a conversion system matrix of the position quaternion according to the gyro output angular velocity;

[0033] S235: obtaining the pseudo-linear position update equation according to the position vector, the attitude quaternion, the velocity quaternion, the conversion system matrix of the attitude quaternion to the position quaternion and the conversion system matrix of the position quaternion.

[0034] Optionally, in the step S23, the pseudo-linear position update equation is:

[0035]

[0036] wherein, A 31 and A 33 respectively represent the conversion system matrix of the attitude quaternion to the position quaternion and the conversion system matrix of the position quaternion and Function M L One of (·) and M R (·) represents the matrix form of the left quaternion in quaternion multiplication, the other represents the matrix form of the right quaternion and represents the matrix C R is the inverse matrix of R and μ n represents the velocity quaternion and v n represents the velocity vector in the navigation system, represents the differential of the position vector, represents the attitude quaternion, represents the multiplication operation of the quaternion, represents the rotation angular velocity of the navigation coordinate system, represents the gyro output angular velocity, p represents the new position quaternion and p represents the position vector and p = [L λ h] τ , L represents the latitude, λ represents the longitude, h represents the height, R M represents the meridian principal curvature radius, R N represents the principal curvature radius of the prime vertical.

[0037] Optionally, the step S25 comprises:

[0038] S251: obtaining an initial value of a navigation parameter vector;

[0039] S252: obtaining a gyro output angular increment sequence and an accelerometer output velocity increment sequence in an update period;

[0040] S253: obtaining an angular velocity function and a specific force function according to the angular increment sequence and the velocity increment sequence;

[0041] S254: obtaining a system matrix function according to the angular velocity function and the specific force function;

[0042] S255: obtaining a state transition matrix according to the system matrix function and a state transition matrix in the update period;

[0043] S256: obtaining an updated navigation parameter vector according to the state transition matrix and the initial value of the navigation parameter vector.

[0044] Optionally, in the step S26, the updated navigation parameter vector N(T) is:

[0045] N(T) = Φ(T, 0)N(0)

[0046] Wherein, Φ(T, 0) represents a state transition matrix, and N(0) represents an initial value of a navigation parameter vector.

[0047] Optionally, the step S26 comprises:

[0048] S261: converting the updated navigation parameter vector into a posture quaternion, a velocity vector and a position vector;

[0049] S262: obtaining an update result of the SINS according to the posture quaternion, the velocity vector and the position vector.

[0050] The present application has the following beneficial effects:

[0051] The pseudo-linear SINS update method provided by the present application can solve linear system equations to realize SINS update, and has high algorithm accuracy in large maneuvering conditions, such as navigation calculation of a rotating projectile with a large roll angular velocity. Compared with traditional SINS algorithms, the pseudo-linear model-based algorithm has the advantages of high accuracy, simple structure and linearization. BRIEF DESCRIPTION OF DRAWINGS

[0052] Figure 1 A flowchart of the pseudo-linear SINS update method provided by the present application. DETAILED DESCRIPTION

[0053] The principles and features of the present application are described below in combination with the accompanying drawings, and the examples are only used to explain the present application and not to limit the scope of the present application.

[0054] EMBODIMENT

[0055] The technical solution of the present application to solve the above technical problems is as follows:

[0056] The present application provides a pseudo-linear SINS update method, as shown in Figure 1 The pseudo-linear SINS update method comprises the following steps:

[0057] S1: obtaining SINS related data;

[0058] Optionally, the SINS related data comprises a posture differential equation, a velocity differential equation and a position differential equation.

[0059] S2: performing update calculation processing on the SINS related data by using a pseudo-linear SINS data update method to obtain an update result of the SINS;

[0060] Here, the pseudo-linear SINS data update method comprises:

[0061] S21: pseudo-linearizing the attitude differential equation to obtain a pseudo-linear attitude update equation;

[0062] Alternatively, in step S21, the pseudo-linear attitude update equation is:

[0063]

[0064] wherein A 11 represents a system matrix of attitude quaternion conversion system and is a gyro output angular velocity, is a navigation coordinate system rotation angular velocity, functions M L (·) and M R (·) respectively represent a matrix form of left multiplication quaternion and a matrix form of right multiplication quaternion in quaternion multiplication and represents an attitude quaternion, i represents an inertial coordinate system, b represents a carrier coordinate system, n represents a navigation coordinate system, q0, q1, q2 and q3 respectively represent four elements of a quaternion.

[0065] Specifically, in a strapdown inertial navigation algorithm, an attitude quaternion differential equation is

[0066]

[0067] In the formula: is an attitude quaternion, is an angular velocity of a carrier coordinate system relative to a navigation coordinate system. Since cannot be directly measured by a gyro, it needs to be expanded,

[0068] further written as:

[0069]

[0070] In the formula: is a gyro output angular velocity, is a navigation coordinate system rotation angular velocity, functions M L (·) and M R (·) respectively represent a matrix form of left multiplication quaternion and a matrix form of right multiplication quaternion in quaternion multiplication and A 11 is used to represent a system matrix in the model, that is:

[0071]

[0072] The attitude update of the navigation solution can be expressed as:

[0073]

[0074] S22: performing pseudo-linearization on the velocity differential equation to obtain a pseudo-linear velocity update equation;

[0075] Optionally, the step S22 comprises:

[0076] S221: obtaining the velocity differential equation of the SINS;

[0077] S222: transforming the velocity vector in the velocity differential equation to obtain a velocity quaternion;

[0078] S223: obtaining the pseudo-linear velocity update equation according to the velocity quaternion and an attitude quaternion.

[0079] Optionally, in the step S22, the pseudo-linear velocity update equation is:

[0080]

[0081] wherein A 21 and A 22 respectively represent a conversion system matrix from an attitude quaternion to a velocity quaternion and a velocity quaternion conversion system matrix, and denotes an output angular velocity of a gyroscope, and functions M L (·) and M R (·) respectively represent a matrix form of a left multiplication quaternion in quaternion multiplication and a matrix form of a right multiplication quaternion, and denotes an output specific force of an accelerometer, denotes a sum of centripetal acceleration and Coriolis acceleration, and and respectively represent an earth rotation angular velocity and a rotation angular velocity of a navigation system relative to an earth system caused by a carrier motion, v n denotes a velocity vector in a navigation system, g n denotes a gravitational acceleration vector, denotes a multiplication operation of a quaternion, denotes a rotation velocity of the navigation system relative to an inertial system, denotes an attitude quaternion, μ n denotes a velocity quaternion, and b denotes a carrier coordinate system, and n denotes a navigation coordinate system.

[0082] Specifically, the force ratio equation is one of the core components of strapdown inertial navigation system (SINS) calculations; solving the force ratio equation allows for velocity updates. The force ratio equation in the navigation frame of a strapdown inertial navigation system is as follows:

[0083]

[0084] In the formula: v n It is a velocity vector. It is the attitude matrix. and These are the Earth's rotational angular velocity and the rotational angular velocity of the navigation system relative to the Earth system caused by the carrier's motion, g. n It is the gravitational acceleration vector.

[0085] exist By using quaternions to transform the ratio in the b-system to the n-system, we obtain:

[0086]

[0087] As can be seen from the first term on the right side of the above equation, the velocity update equation is nonlinear when using quaternions for navigation calculation.

[0088] To obtain a linear velocity update equation, the velocity vector v n Perform the transformation and define a new velocity vector—the velocity quaternion.

[0089]

[0090] Where: μ n For velocity quaternions, For v n The zero scalar quaternion will be used consistently throughout the text. This represents the zero scalar quaternion of a three-dimensional vector. Clearly, according to... The velocity vector v can also be solved. n .

[0091] right Differentiate both sides

[0092]

[0093] Will Write in quaternion form

[0094]

[0095] In the formula: Represents the sum of centripetal acceleration and Coriolis acceleration, and Indicates centripetal acceleration. This represents Coriolis acceleration.

[0096] Will and Substitute Get

[0097]

[0098] In the formula And Still nonlinear, but the speed change is small in an update cycle, and the values of the three angular velocities are small, so in the update cycle And Can be considered as constants, so

[0099] It is converted into a pseudo-linear equation.

[0100] After rearrangement:

[0101]

[0102] S23: Perform pseudo-linearization processing on the position differential equation to obtain a pseudo-linear position update equation;

[0103] Alternatively, the step S23 comprises:

[0104] S231: Obtain a position vector according to the latitude, longitude and altitude of the carrier;

[0105] S232: Obtain a constant matrix in an update cycle according to the position differential equation and the velocity vector;

[0106] S233: Obtain a conversion system matrix from attitude quaternion to position quaternion according to the constant matrix, the position vector and the rotation angular velocity of the navigation coordinate system;

[0107] S234: Obtain a position quaternion conversion system matrix according to the angular velocity output by the gyroscope;

[0108] S235: Obtain the pseudo-linear position update equation according to the position vector, the attitude quaternion, the velocity quaternion, the conversion system matrix from attitude quaternion to position quaternion and the position quaternion conversion system matrix.

[0109] Alternatively, in the step S23, the pseudo-linear position update equation Is:

[0110]

[0111] Wherein, A 31 And A 33 Respectively represent the conversion system matrix from attitude quaternion to position quaternion and the position quaternion conversion system matrix, and ​ Function M L (·) and M R (·) represent the matrix form of the left quaternion in quaternion multiplication, the other represents the matrix form of the right quaternion and represents the inverse matrix of matrix C R and μ n represents the velocity quaternion and v n represents the velocity vector in the navigation system, represents the differential of the position vector, represents the attitude quaternion, represents the multiplication operation of the quaternion, represents the rotation angular velocity of the navigation coordinate system, represents the gyroscope output angular velocity, ρ represents the new position quaternion and p represents the position vector and p = [L λ h] τ , L represents the latitude, λ represents the longitude, h represents the height, R M represents the meridian principal radius of curvature, R N represents the prime vertical principal radius of curvature.

[0112] Specifically, the latitude, longitude and height of the carrier in the geographic system are usually calculated as the position update result of the inertial navigation solution, and the position vector is defined as

[0113] p = [L λ h] T

[0114] In the formula, L is the latitude, λ is the longitude, and h represents the height. The traditional position update equation is

[0115]

[0116] In the formula,

[0117]

[0118] R M is the meridian radius of the earth, R N is the prime vertical radius.

[0119] Substituting the formula into , we can obtain

[0120]

[0121] Obviously, is nonlinear. By defining a new position quaternion The position update equation is converted into a pseudo-linear form, and C R In the update period, the change is small and can be regarded as a constant, and the definition of p is Taking the derivative of both sides, we get

[0122]

[0123] The above formula is rewritten in matrix form as

[0124]

[0125] In the formula,

[0126]

[0127]

[0128] So far, by introducing a new position quaternion p, the traditional position update equation has also been converted into a pseudo-linear equation about attitude quaternion, velocity quaternion and position quaternion.

[0129] S24: Obtain a navigation parameter vector differential equation according to the pseudo-linear attitude update equation, the pseudo-linear velocity update equation and the pseudo-linear position update equation;

[0130] S25: Perform an update operation on the navigation parameter vector differential equation to obtain an updated navigation parameter vector;

[0131] Optionally, the step S25 comprises:

[0132] S251: Obtain a navigation parameter vector initial value;

[0133] Here, the navigation parameter vector initial value can be set according to the initial attitude, velocity and position of the carrier.

[0134] Suppose the measured initial attitude quaternion of the carrier is The initial velocity vector is v n (0), and the initial position vector is p(0) = [L λ h] T The initial navigation parameter vector is:

[0135]

[0136] In the formula,

[0137] S252: Obtain a gyro output angular increment sequence and an accelerometer output velocity increment sequence in an update period;

[0138] In the update period [0, T], n times of sampling are performed, and the sampling interval is Ts , to obtain the angle increment vector sequence Δθ1, Δθ2, …, Δθ n and the velocity increment vector sequence Δv1, Δv2, …, Δv n .

[0139] S253: According to the angle increment sequence and the velocity increment sequence, an angular velocity function and a specific force function are obtained;

[0140] The angular velocity function is an n-1 order function The specific force function is an n-1 order function The relationship between the angular velocity function coefficient and the specific force function coefficient and the angle increment, the velocity increment, respectively, is

[0141]

[0142] Wherein:

[0143] T s is the sampling interval of the gyroscope and the accelerometer, so as to obtain the coefficient of the angular velocity function and the specific force function , respectively

[0144]

[0145] S254: According to the angular velocity function and the specific force function, a system matrix function is obtained;

[0146] The angular velocity function and the specific force function are substituted into the system matrix function F(t) to obtain

[0147]

[0148] Wherein:

[0149] S255: According to the system matrix function and the state transition matrix in the update period, a state transition matrix is obtained;

[0150] According to , the derivatives of F(t) are substituted to calculate the derivatives of Φ(t, 0) at t=0 Φ (i) (0, 0), i=0, 1, 2, …, m;

[0151] According to , the state update matrix Φ(T, 0) in the update period [0, T] is calculated.

[0152] S256: Based on the state transition matrix and the initial values ​​of the navigation parameter vector, the updated navigation parameter vector is obtained.

[0153] Alternatively, the updated navigation parameter vector N(T) is:

[0154] N(T)=Φ(T,0)N(0)

[0155] Where Φ(T, 0) represents the state transition matrix, and N(0) represents the initial value of the navigation parameter vector.

[0156] S26: Based on the updated navigation parameter vector, obtain the updated result of the strapdown inertial navigation system.

[0157] Alternatively, step S26 may include:

[0158] S261: Convert the updated navigation parameter vector into attitude quaternions, velocity vector, and position vector;

[0159] S262: Based on the attitude quaternion, velocity vector, and position vector, obtain the update result of the strapdown inertial navigation system.

[0160] The updated navigation parameter vector N(T) = Φ(T, 0)N(0) is converted into attitude quaternions in the geographic system. velocity vector v n and position vector p, i.e. Due to the updated navigation parameter vector It is a 12x1 column vector, therefore the attitude quaternion It is equal to the first four elements of N(T); velocity Where μ n The four middle elements of N(T); positions The new positional quaternion ρ is equal to the last four elements of N(T).

[0161] This allows us to obtain the solution output from the strapdown inertial navigation system.

[0162] S3: Improve navigation accuracy based on the updated results of the strapdown inertial navigation system.

[0163] The present invention has the following beneficial effects:

[0164] The pseudo-linear strapdown inertial navigation system (SINS) solution update method provided by this invention can solve linear system equations to achieve SINS solution updates, and it exhibits high algorithm accuracy under high maneuverability conditions, such as navigation solutions for spinning projectiles with large roll velocities. Compared with traditional SINS algorithms, the algorithm based on the pseudo-linear model has advantages such as high accuracy, simple structure, and linearization.

[0165] The above merely describes preferred embodiments of the present application, and is not used to limit the present application, and any modification, equivalent replacement, improvement, etc. made within the spirit and principle of the present application shall be included in the protection scope of the present application.

Claims

1. A pseudo-linear strapdown inertial navigation system (SINS) solution update method, characterized in that, The pseudo-linear strapdown inertial navigation system solution update method includes: S1: Obtain relevant data from strapdown inertial navigation; S2: The strapdown inertial navigation data is updated and solved using a pseudo-linear strapdown inertial navigation data update method to obtain the updated results of the strapdown inertial navigation. The differential equation for strapdown inertial navigation is: In step S2, the pseudo-linear strapdown inertial navigation data update method includes: S21: Perform pseudo-linearization on the attitude differential equation to obtain the pseudo-linear attitude update equation; The pseudolinear attitude update equation for: in, A 11 Represents the attitude quaternion transformation system matrix and , The gyroscope outputs angular velocity. For the rotational angular velocity of the navigation coordinate system, the function M L (·)and M R (·) represent the matrix forms of left-multiplication and right-multiplication of a quaternion, respectively. , , Represents attitude quaternions, i Indicates an inertial coordinate system. b Indicates the carrier coordinate system. n Indicates the navigation coordinate system. , , and Each of the four elements represents a quaternion; S22: Perform pseudo-linearization on the velocity differential equation to obtain a pseudo-linear velocity update equation; The pseudo-linear velocity update equation for: in, A 21 and A 22 Let these represent the transformation system matrices from attitude quaternions to velocity quaternions and the transformation system matrix from velocity quaternions, respectively. , , The function represents the angular velocity output by the gyroscope. M L (·)and M R (·) represent the matrix forms of left-multiplication and right-multiplication of a quaternion, respectively. , , This indicates the specific force output by the accelerometer. Represents the sum of centripetal acceleration and Coriolis acceleration, and , and These represent the Earth's rotational angular velocity and the rotational angular velocity of the navigation system relative to the Earth system caused by the carrier's motion, respectively. This represents the velocity vector in the navigation frame. Represents the gravitational acceleration vector, "Represents the multiplication operation of quaternions, This represents the rotational speed of the navigation frame relative to the inertial frame. Represents attitude quaternions, Represents a velocity quaternion and , b Indicates the carrier coordinate system. n Indicates the navigation coordinate system; S23: Perform pseudo-linearization on the position differential equation to obtain the pseudo-linear position update equation; The pseudo-linear position update equation for: in, A 31 and A 33 Let these represent the transformation system matrix from attitude quaternions to position quaternions and the transformation system matrix from position quaternions, respectively. , ,function M L (·)and M R In the parentheses (·), one represents the matrix form of left-multiplying a quaternion, and the other represents the matrix form of right-multiplying a quaternion. , , Representation matrix The inverse matrix and , Represents a velocity quaternion and , This represents the velocity vector in the navigation frame. Represents the differential of the position vector. Represents attitude quaternions, "Represents the multiplication operation of quaternions, Indicates the angular velocity of rotation of the navigation coordinate system. This indicates the angular velocity output by the gyroscope. Describe the new positional quaternion and , Represents a position vector and , L Indicates latitude, λ Indicates longitude. h Indicates altitude, This represents the radius of the principal curvature of the meridian. Indicates the principal radius of curvature of the circle; S24: Based on the pseudolinear attitude update equation, the pseudolinear velocity update equation, and the pseudolinear position update equation, the navigation parameter vector differential equation is obtained; S25: Perform an update operation on the differential equation of the navigation parameter vector to obtain the updated navigation parameter vector; S26: Based on the updated navigation parameter vector, obtain the updated result of the strapdown inertial navigation system; S3: Improve navigation accuracy based on the updated results of the strapdown inertial navigation system.

2. The pseudo-linear strapdown inertial navigation system solution update method according to claim 1, characterized in that, Step S22 includes: S221: Obtain the velocity differential equation of the strapdown inertial navigation system; S222: Transform the velocity vector in the velocity differential equation to obtain the velocity quaternion; S223: Based on the velocity quaternion and attitude quaternion, the pseudo-linear velocity update equation is obtained.

3. The pseudo-linear strapdown inertial navigation system solution update method according to claim 1, characterized in that, Step S23 includes: S231: Obtain the position vector based on the latitude, longitude, and altitude of the carrier; S232: Based on the position differential equation and velocity vector, obtain the constant matrix within the update cycle; S233: Based on the constant matrix, the position vector, and the rotation angular velocity of the navigation coordinate system, obtain the transformation system matrix from attitude quaternion to position quaternion; S234: Obtain the position quaternion transformation system matrix based on the angular velocity output by the gyroscope; S235: Based on the position vector, the attitude quaternion, the velocity quaternion, the transformation system matrix from the attitude quaternion to the position quaternion, and the position quaternion transformation system matrix, the pseudo-linear position update equation is obtained.

4. The pseudo-linear strapdown inertial navigation system solution update method according to claim 1, characterized in that, Step S25 includes: S251: Obtain the initial values ​​of the navigation parameter vector; S252: Obtain the gyroscope output angle increment sequence and accelerometer output velocity increment sequence within the update cycle; S253: Based on the angular increment sequence and the velocity increment sequence, obtain the angular velocity function and the specific force function; S254: Based on the angular velocity function and the specific force function, obtain the system matrix function; S255: Obtain the state transition matrix based on the system matrix function and the state transition matrix within the update period; S256: Based on the state transition matrix and the initial values ​​of the navigation parameter vector, the updated navigation parameter vector is obtained.

5. The pseudo-linear strapdown inertial navigation system solution update method according to claim 1 or 4, characterized in that, In step S26, the updated navigation parameter vector for: in, Represents the state transition matrix. This represents the initial value of the navigation parameter vector.

6. The pseudo-linear strapdown inertial navigation system solution update method according to claim 2, characterized in that, Step S26 includes: S261: Convert the updated navigation parameter vector into attitude quaternions, velocity vector, and position vector; S262: Based on the attitude quaternion, velocity vector, and position vector, obtain the update result of the strapdown inertial navigation system.

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