Method for updating strapdown inertial navigation solutions based on launch-centered earth-fixed frame

Abstract

The present invention relates to a method for updating strapdown inertial navigation solutions based on a launch-centered earth-fixed (LCEF) frame (g frame). The present invention uses the g frame as a navigation reference frame of a medium-to-short-range surface-to-surface missile. This is beneficial to establish a relative relationship between the missile and the ground so as to keep the same missile parameters required by a missile control and guidance system. The calculation of a navigation algorithm in the g frame is moderate, which is suitable for an embedded system.

Description

TECHNICAL FIELD

The present invention belongs to the field of flight vehicle navigation, guidance and control, and relates to a reference frame for a flight vehicle navigation algorithm, in particular to a method for updating strapdown inertial navigation solutions based on a launch-centered earth-fixed (LCEF) frame.

BACKGROUND

A suitable navigation reference frame is conducive to the design and application of the flight vehicle navigation and guidance system. The local level (LL) frame describes the attitude and position of the carrier in the near-space motion, and the launch-centered inertial (LCI) frame (a frame) describes the attitude and position of the ballistic missile and the carrier rocket. The LCEF frame can be used to establish a motion equation of the rocket relative to the ground, and is suitable for medium-to-short-range surface-to-surface missiles. It is possible to choose the LCEF frame as the navigation coordinate system to quickly resolve the attitude and position of the missile relative to the launch point, which is beneficial to the data interaction with the flight control and guidance system. The methods for updating the strapdown inertial navigation solutions based on the LL/LCI frame are mature, but there is no perfect method for updating the strapdown inertial navigation solutions based on the LCEF frame.

SUMMARY

The present invention provides a method for updating strapdown inertial navigation solutions in a launch-centered earth-fixed (LCEF) frame, which fills in the gap of the prior art. The present invention uses the LCEF frame as a navigation reference frame of a medium-to-short-range surface-to-surface missile (SSM). This is beneficial to establish a relative relationship between the missile and the ground so as to keep the same missile parameters required by a missile control and guidance system. The calculation of a navigation algorithm in the LCEF frame is moderate, which is suitable for an embedded system.

A method for updating strapdown inertial navigation solutions based on an LCEF frame (g frame) includes the following steps:

step 1, using a surface-to-surface missile (SSM) as a body, and establishing a body frame (b frame), which points front, top, and right;

step 2: expressing a navigation equation in the g frame as:

$\left[\begin{array}{c}{\stackrel{.}{P}}^g\\ {\stackrel{.}{V}}^g\\ {\stackrel{.}{R}}_b^g\end{array}\right]=\left[\begin{array}{c}{V}^g\\ R_b^g{f}^b-2{\Omega }_{\mathrm{ag}}^g{V}^g+g^g\\ R_b^g\left({\Omega }_{\mathrm{ab}}^b-{\Omega }_{\mathrm{ag}}^b\right)\end{array}\right]$

where, P^g, V^g and R_b^g are position, velocity and attitude matrices of the body in the g frame, and the corresponding equations are position, velocity and attitude navigation equations; f^b is a measured value of an accelerometer triad; g^g is a gravity of the body in the g frame; Ω_ab^b is an anti-symmetric matrix corresponding to a measured value ω_ab^b of a gyroscope triad; Ω_ag^b is an anti-symmetric matrix corresponding to a rotational angular velocity ω_ag^b of the g frame relative to an a frame;

step 3: performing attitude update on the SSM, including the following sub-steps:

step 3-1: resolving the attitude navigation equation in step 2 by a quaternion method:q_b(m)^g(m)=q_g(m-1)^g(m)q_b(m-1)^g(m-1)q_b(m)^g(m-1)

where, q_b(m-1)^g(m-1) is a transformation quaternion from the b frame to the g frame at a t_m-1 moment, that is, an attitude quaternion at t_m-1; q_b(m)^g(m) is an attitude quaternion at t_m; q_g(m-1)^g(m) is a transformation quaternion of the g frame from t_m-1 to t_m; q_b(m)^b(m-1) is a transformation quaternion calculated from an angular increment from t_m-1 to t_m;

step 3-2: calculating with an equivalent rotation vector method, where

an equivalent rotation vector ζm of the g frame from t_m-1 to t_m is expressed as:

${ϛ}_m={\int }_{t_{m-1}}^{t_m}{\omega }_{\mathrm{ag}}^g\left(t\right)\mathrm{dt}\approx {\omega }_{\mathrm{ag}}^{g}T$

an equivalent rotation vector Φ_m of the body frame relative to the inertial frame from t_m-1 to t_m is expressed as:

${\Phi }_m={\int }_{t_{m-1}}^{t_m}{\omega }_{\mathrm{ab}}^b\left(t\right)dt$

step 3-3: applying the angular increment measured by the gyroscope triad to an actual project to calculate:

$\begin{array}{cc}{\Phi }_m=\Delta {\theta }_1+\Delta {\theta }_2+\frac{2}{3}\left(\Delta {\theta }_1\times \Delta {\theta }_2\right)& \left(0.1\right)\end{array}$

where

$\Delta {\theta }_1={\int }_{t_{m-1}}^{t_{m-1}+\frac{T}{2}}{\omega }_{\mathrm{ab}}^b\left(\tau \right)d\tau ,\Delta {\theta }_2={\int }_{t_{m-1}+\frac{T}{2}}^{t_m}{\omega }_{\mathrm{ab}}^b\left(\tau \right)d\tau ;$ a rotation vector Φ and a corresponding quaternion q(h) have the following calculation relationship:

$\begin{array}{cc}q\left(h\right)=\cos \frac{\Phi }{2}+\frac{\Phi }{\Phi }\sin \frac{\Phi }{2}& \left(0.2\right)\end{array}$

thus quaternions q_g(m-1)^g(m) and q_b(m)^b(m-1) corresponding to ζ_m and Φ_m in are obtained, and substituted into q_b(m)^g(m)=q_g(m-1)^g(m)q_b(m-1)^g(m-1)q_b(m)^b(m-1) to complete attitude update;

step 4: performing velocity update on the SSM, including the following sub-steps:

step 4-1: integrating the velocity navigation equation in step 2 over a period [t_m-1, t_m]:V_m^g=V_m-1^g+ΔV_sf(m)^g+ΔV_cor/g(m)^g

where, V_m^g and V_m-1^g are the velocity of the SSM in the g frame at t_m and t_m-1; ΔV_sf(m)^g and ΔV_cor/g(m)^g are a specific force velocity increment and a velocity increment of a harmful acceleration of the navigation coordinate system in the period T=t_m−t_m-1;

Step 4-2: substituting ΔV_cor/g(m)^g with a value at t_m-1/2=(t_m-1+t_m)/2:ΔV_cor/g(m)^g≈[−2ω_ag(m-1/2)^g×V_m-1/2^g+g_m-1/2^g]T;

performing matrix chain multiplication and decomposition on ΔV_sf(m)^g:ΔV_sf(m)^g=R_b(m-1)^g(m-1)ΔV+R_b(m-1)^g(m-1)(ΔV_rot(m)+ΔV_scul(m))−(ΔV′_rot(m)+ΔV′_scul(m))

where R_b(m-1)^g(m-1) is an attitude matrix at t_m-1; ΔV is an integral of a measured value of the accelerometer triad in T; ΔV_rot(m) and ΔV′_rot(m) are rotation error compensation quantities of velocity; ΔV_scul(m) and ΔV′_scul(m) are scull error compensation quantities;

step 4-3: expressing ΔV_sf(m)^g by an angular velocity increment and a velocity increment at t_m-1/2 and t_m, and setting:

$\{\begin{array}{c}\Delta {\theta }_1={\int }_{t_{m-1}}^{t_{m-1}+T/2}{\omega }_{\mathrm{ab}}^b\left(\tau \right)d\tau \\ \Delta {\theta }_2={\int }_{t_{m-1}+T/2}^{t_m}{\omega }_{\mathrm{ab}}^b\left(\tau \right)d\tau \\ \Delta V_1={\int }_{t_{m-1}}^{t_{m-1}+T/2}{f}^b\left(\tau \right)d\tau \\ \Delta V_2={\int }_{t_{m-1}+T/2}^{t_m}{f}^b\left(\tau \right)d\tau \\ \Delta \theta ={\int }_{t_{m-1}}^{t_m}{\omega }_{\mathrm{ab}}^b\left(\tau \right)d\tau \\ \Delta V={\int }_{t_{m-1}}^{t_m}{f}^b\left(\tau \right)d\tau \end{array}\{\begin{array}{c}{\mathrm{\Delta \theta }}_1^{\prime }={\int }_{t_{m-1}}^{t_{m-1}+T/2}{\omega }_{\mathrm{ag}}^g\left(\tau \right)d\tau \approx \frac{T}{2}{\omega }_{\mathrm{ag}}^g\\ {\mathrm{\Delta \theta }}_2^{\prime }={\int }_{t_{m-1}+T/2}^{t_m}{\omega }_{\mathrm{ag}}^g\left(\tau \right)d\tau \approx \frac{T}{2}{\omega }_{\mathrm{ag}}^g\\ \Delta V_1^{\prime }={\int }_{t_{m-1}}^{t_{m-1}+T/2}{R}_{b\left(m-1\right)}^{g\left(m-1\right)}{f}^b\left(\tau \right)d\tau =R_{b\left(m-1\right)}^{g\left(m-1\right)}\Delta V_1\\ \Delta V_2^{\prime }={\int }_{t_{m-1}+T/2}^{t_m}{R}_{b\left(m-1\right)}^{g\left(m-1\right)}{f}^b\left(\tau \right)d\tau =R_{b\left(m-1\right)}^{g\langle m-1)}\Delta V_2\\ {\mathrm{\Delta \theta }}^{\prime }={\int }_{t_{m-1}}^{t_m}{\omega }_{\mathrm{ag}}^g\left(\tau \right)\mathrm{d\tau }={\mathrm{\Delta \theta }}_1^{\prime }+{\mathrm{\Delta \theta }}_2^{\prime }\\ \Delta V^{\prime }={\int }_{t_{m-1}}^{t_m}{R}_{b\left(m-1\right)}^{g\left(m-1\right)}{f}^b\left(\tau \right)d\tau =R_{b\left(m-1\right)}^{g\left(m-1\right)}\Delta V\end{array}$

where, a two-subsample velocity rotation error compensation algorithm is as follows:

$\Delta V_{\mathrm{rot}\left(m\right)}=\frac{1}{2}\Delta \theta \times \mathrm{\Delta V}$ $\Delta V_{\mathrm{rot}\left(m\right)}^{\prime }=\frac{1}{2}A{\theta }^{\prime }\times \Delta V^{\prime }$

a two-subsample velocity scull error compensation algorithm is as follows:

$\Delta V_{\mathrm{scul}\left(m\right)}=\frac{2}{3}\left(\Delta {\theta }_1\times \Delta V_2+\Delta V_1\times \Delta {\theta }_2\right)$ $\Delta V_{\mathrm{scul}\left(m\right)}^{\prime }=\frac{2}{3}\left({\mathrm{\Delta \theta }}_1^{\prime }\times \Delta V_2^{\prime }+\Delta V_1^{\prime }\times {\mathrm{\Delta \theta }}_2^{\prime }\right)$ $\mathrm{and}$ $\Delta V_{\mathrm{rot}\left(m\right)}^{\prime }+\Delta V_{\mathrm{scul}\left(m\right)}^{\prime }\approx \frac{T}{6}{\omega }_{\mathrm{ag}}^g\times \left[R_{b\left(m-1\right)}^{g\left(m-1\right)}\left(\Delta V_1+5\Delta V_2\right)\right]$

where, the complete algorithm of the specific force velocity increment in the g frame is as follows:

$\Delta V_{\mathrm{sf}\left(m\right)}^g=R_{b\left(m-1\right)}^{g\left(m-1\right)}\Delta V-\frac{T}{6}{\omega }_{\mathrm{ag}}^g\times \left[R_{b\left(m-1\right)}^{g\left(m-1\right)}\left(\Delta V_1+5\Delta V_2\right)\right]+R_{b\left(m-1\right)}^{g\left(m-1\right)}\left(\Delta V_{\mathrm{rot}\left(m\right)}+\Delta V_{\mathrm{scul}\left(m\right)}\right)$

step 5: performing position update on the SSM, including the following sub-steps:

step 5-1: obtaining the following equation from the velocity navigation equation in step 2:{dot over (P)}^g(t)=V^g(t)

• • where

$V^g\left(t\right)=V_{m-1}^g+\Delta V_{\mathrm{sf}}^g\left(t\right)+\Delta V_{\mathrm{cor}/g}^g\frac{t-t_{m-1}}{T}\left(t_{m-1}\le t\le t_m\right);$

step 5-2: integrating both sides of V^g(t) in step 5-1 in [t_m-1, t_m] to obtain:

$P_m^g=P_{m-1}^g+\left[V_{m-1}^g+\frac{1}{2}\Delta V_{\mathrm{cor}/g\left(m\right)}^g\right]T+\Delta P_{\mathrm{sf}\left(m\right)}^g$

where, P_m^g and P_m-1^g are positions of the SSM in the g frame at t_m and t_m-1;ΔP_sf(m)^g=R_b(m-1)^g(m-1)(S_Δvm+ΔP_rotm+ΔP_scrlm)−(ΔP′_rotm+ΔP′_sclm)and

$S_{\Delta \mathrm{vm}}=T\left[\frac{5}{6}\Delta V_1+\frac{1}{6}\Delta V_2\right]$

indicates a secondary integral increment of a specific force;

$\Delta P_{\mathrm{rotm}}=T\left(\Delta {\theta }_1\times \left[\frac{5}{18}\Delta V_1+\frac{1}{6}\Delta V_2\right]+\Delta {\theta }_2\times \left[\frac{1}{6}\Delta V_1+\frac{1}{18}\Delta V_2\right]\right)$ $\Delta P_{\mathrm{rotm}}^{\prime }=\tau \left({\mathrm{\Delta \theta }}_1^{\prime }\times \left[\frac{5}{18}\Delta V_1^{\prime }+\frac{1}{6}\Delta V_2^{\prime }\right]+{\mathrm{\Delta \theta }}_2^{\prime }\times \left[\frac{1}{6}\Delta V_1^{\prime }+\frac{1}{18}\Delta V_2^{\prime }\right]\right)$

indicates a rotation effect compensation quantity in position calculation;

$\Delta P_{\mathrm{scrlm}}=T\left({\mathrm{\Delta \theta }}_1\times \left[\frac{11}{90}\Delta V_1+\frac{1}{10}\Delta V_2\right]+\Delta {\theta }_2\times \left[\frac{1}{90}\Delta V_2-\frac{7}{30}\Delta V_1\right]\right)$ $\Delta P_{\mathrm{scrlm}}^{\prime }=T\left(\Delta {\theta }_1^{\prime }\times \left[\frac{11}{90}\Delta V_1^{\prime }+\frac{1}{10}\Delta V_2^{\prime }\right]+\Delta {\theta }_2^{\prime }\times \left[\frac{1}{90}\Delta V_2^{\prime }-\frac{7}{30}\Delta V_1^{\prime }\right]\right)$

indicates a scroll effect compensation quantity in position calculation;

step 5-3: finally obtaining the complete algorithm of the ΔP_sf(m)^g position increment in the g frame:

$\Delta P_{\mathrm{sf}\left(m\right)}^g=R_{b\left(m-1\right)}^{g\left(m-1\right)}\left(S_{\Delta \mathrm{vm}}+\Delta P_{\mathrm{rotm}}+\Delta P_{\mathrm{scrlm}}\right)-\frac{T^2}{6}{\omega }_{\mathrm{ag}}^g\times \left(R_{b\left(m-1\right)}^{g\left(m-1\right)}\left[\Delta V_1+\Delta V_2\right]\right)$

according to the result in step 5-2 and

$\Delta P_{\mathrm{rotm}}^{\prime }+\Delta P_{\mathrm{scrlm}}^{\prime }\approx \frac{T^2}{6}{\omega }_{\mathrm{ag}}^g\times \left(R_{b\left(m-1\right)}^{g\left(m-1\right)}\left[\Delta V_1+\Delta V_2\right]\right)$

thus completing the updates of the strapdown inertial navigation solutions in terms of attitude, velocity and position of the SSM in the g frame.

The present invention has the following beneficial effects.

In the present invention, the attitude and position of the SSM in the g frame calculated by the method are consistent with the data of the flight control and guidance system. This method uses an angular increment and a velocity increment to complete the update of the strapdown inertial navigation solutions. It takes into account the linear and angular vibrations of the SSM in flight, and compensates for the calculation of a cone error, a scull error and a scroll error. In this way, the method reduces the errors of data between the SSM and the system, and improves the hit rate of the SSM.

BRIEF DESCRIPTION OF DRAWINGS

FIG. 1: Mechanization of strapdown inertial navigation in an LCEF frame.

DETAILED DESCRIPTION

Referring to FIG. 1, coordinate systems in the present invention are defined as follows:

1) Earth-centered inertial (ECI) frame (i frame).

2) Earth-centered earth-fixed (ECEF) frame (e frame).

3) Body frame (b frame), which is a forward-up-right body frame. A surface-to-surface missile (SSM) serves as a body in the present invention.

4) Launch-centered earth-fixed (LCEF) frame (g frame). An origin of coordinates o is at a launch point; ox is in a horizontal plane of the launch point, and points to a launch aiming direction; oy is along a vertical direction of the launch point; oxyz is a right-handed coordinate system. The g frame is fixed on the earth and rotates with the earth.

5) Launch-centered inertial (LCI) frame (a frame), which coincides with the g frame at the moment of launching, and then remains unchanged in an inertial space.

The method for updating strapdown inertial navigation solutions of the SSM based on the g frame includes attitude update, velocity update and position update.

Step 1: Calculate a corresponding rotation vector according to a gyroscope triad angular increment of the SSM in an update period T, a rotational angular velocity ω_ag^g of the g frame (g frame) relative to the LCI frame (a frame) and the update period T; calculate a corresponding quaternion according to a relationship between the rotation vector and the quaternion; finally, complete attitude update by an quaternion recursion equation.

Step 2: Calculate a velocity increment of a harmful acceleration according to a velocity V_m-1^g of the SSM at a previous moment, the rotational angular velocity ω_ag^g of the g frame relative to the a frame and a gravity g^g of the SSM; calculate a specific force velocity increment according to the gyroscope triad angular increment and an accelerometer triad velocity increment of the SSM in the update period; add to the velocity V_m-1^g of the previous moment to complete velocity update.

Step 3: Calculate an integral of the velocity V_m-1^g of the previous moment and the velocity increment of the harmful acceleration obtained in step 2 with T; calculate an integral of the specific force velocity increment based on the angular increment and velocity increment in the update period; finally, sum the integral and a position P_m-1^g of the missile at the previous moment to complete position update.

The algorithm of the present invention is described in detail below.

A navigation equation in the g frame is expressed as:

$\begin{array}{cc}\left[\begin{array}{c}{\stackrel{.}{P}}^g\\ {\stackrel{.}{V}}^g\\ {\stackrel{.}{R}}_b^g\end{array}\right]=\left[\begin{array}{c}{V}^g\\ R_b^g{f}^b-2{\Omega }_{\mathrm{ag}}^g{V}^g+g^g\\ R_b^g\left({\Omega }_{\mathrm{ab}}^b-{\Omega }_{\mathrm{ag}}^b\right)\end{array}\right]& \left(1.1\right)\end{array}$

In the equation, P^g, V^g and R_b^g are position, velocity and attitude matrices of the SSM in the g frame, and the corresponding equations are position, velocity and attitude navigation equations; f^b is a measured value of an accelerometer triad; g^g is a gravity of the SSM in the g frame; ω_ab^b is an anti-symmetric matrix corresponding to a measured value ω_ab^b of a gyroscope triad; and ω_ag^b is an anti-symmetric matrix corresponding to the rotational angular velocity ω_ag^b of the g frame relative to the a frame.

(1) Attitude Update

The attitude navigation equation is resolved using a quaternion method; a numerical recursion of an attitude update quaternion equation is performed:q_b(m)^g(m)=q_g(m-1)^g(m)q_b(m-1)^g(m-1)q_b(m)^b(m-1)  (1.2)

In the equation, q_b(m-1)^g(m-1) is a transformation quaternion from the b frame to the g frame at a t_m-1 moment, that is, an attitude quaternion at t_m-1; q_b(m)^g(m) is an attitude quaternion at a t_m moment; q_g(m-1)^g(m) is a transformation quaternion of the g frame from t_m-1 to t_m; q_b(m)^b(m-1) is a transformation quaternion calculated from the angular increment from t_m-1 to t_m.

Because there is a non-commutative error of rotation when the rigid body makes limited rotation, an equivalent rotation vector method is used. When the angular increment is used to calculate an equivalent vector, this non-commutative error is appropriately compensated, and the rotation vector algorithm is used to obtain an optimized algorithm in a cone motion environment. The calculation method is as follows:

(1) The calculation of q_g(m-1)^g(m) uses the rotational angular velocity ω_ag^g(t) (t_m-1≤t≤t_m); t_m−t_m-1=T, and the equivalent rotation vector of the g frame from t_m-1 to t_m is ζ_m:

$\begin{array}{cc}{ϛ}_m={\int }_{t_{m-1}}^{t_m}{\omega }_{\mathrm{ag}}^g\left(t\right)\mathrm{dt}\approx {\omega }_{\mathrm{ag}}^{g}T& \left(1.3\right)\end{array}$

(2) The calculation of q_b(m)^b(m-1) uses the rotational angular velocity ω_ab^b(t), which is the angular velocity measured by the gyroscope triad; the equivalent rotation vector of the body frame relative to the inertial frame from the t_m-1 moment to t_m is Φ_m:

$\begin{array}{cc}{\Phi }_m={\int }_{t_{m-1}}^{t_m}{\omega }_{\mathrm{ab}}^b\left(t\right)dt& \left(1.4\right)\end{array}$

In an actual project, the angular increment measured by the gyroscope triad is used to calculate:Φ_m=Δθ₁+Δθ₂+⅔(Δθ₁×Δθ₂)  (1.5)

In the equation,

$\Delta {\theta }_1={\int }_{t_{m-1}}^{t_{m-1}+\frac{T}{2}}{\omega }_{\mathrm{ab}}^b\left(\tau \right)d\tau ,\Delta {\theta }_2={\int }_{t_{m-1}+\frac{T}{2}}^{t_m}{\omega }_{\mathrm{ab}}^b\left(\tau \right)d\tau .$ The rotation vector Φ and the corresponding quaternion q(h) have the following calculation relationship:

$\begin{array}{cc}q\left(h\right)=\cos \frac{\Phi }{2}+\frac{\Phi }{\Phi }\sin \frac{\Phi }{2}& \left(1.6\right)\end{array}$

The quaternions q_g(m-1)^g(m) and q_b(m)^b(m-1) corresponding to ζ_m and Φ_m are calculated respectively, and are substituted into (1.2) to complete attitude update.

(2) Velocity Update

The velocity navigation equation is integrated over a period [t_m-1, t_m] to obtain the following recursive formV_m^g=V_m-1^g+ΔV_sf(m)^g+ΔV_cor/g(m)^g  (1.7)

In the equation, V_m^g and V_m-1^g are the velocity of the SSM in the g frame at t_m and t_m-1, and

$\begin{array}{cc}\Delta V_{\mathrm{sf}\left(m\right)}^g={\int }_{t_{m-1}}^{t_m}{R}_b^g\left(t\right)f^b\left(t\right)\mathrm{dt}& \left(1.8\right)\\ \Delta V_{\mathrm{cor}/g\left(m\right)}^g={\int }_{t_{m-1}}^{t_m}\left[-2{\omega }_{\mathrm{ag}}^g\left(t\right)\times V^g\left(t\right)+g^g\left(t\right)\right]\mathrm{dt}& \left(1.9\right)\end{array}$

ΔV_sf(m)^g and ΔV_cor/g(m)^g are respectively the specific force velocity increment and the velocity increment of the harmful acceleration of the navigation coordinate system in period T=t_m−t_m-1.

The integrand of ΔV_cor/g(m)^g is a slow time-varying variable, and the value at t_m-1/2=(t_m-1−t_m) is used to approximate equation (1.9) as:ΔV_cor/g(m)^g=[−2ω_ag(m-1/2)^g×V_m-1/2^g+g_g(m-1/2)^g]T  (1.10)

The specific force velocity increment ΔV_sf(m)^g is calculated by matrix chain multiplication and decomposition and approximate expansion:ΔV_sf(m)^g=R_b(m-1)^g(m-1)ΔV+R_b(m-1)^g(m-1)(ΔV_rot(m)+ΔV_scul(m))−(ΔV′_rot(m)+ΔV′_scul(m))  (1.11)

R_b(m-1)^g(m-1) is the attitude matrix at t_m-1; ΔV is the integral of the measured value of the accelerometer triad in T; and ΔV_rot(m) and ΔV′_rot(m) are rotation error compensation quantities of velocity, which are caused due to the spatial rotation change of a specific force direction during the resolving period; ΔV_scul(m) and ΔV′_scul(m) are scull error compensation quantities.

ΔV_scul(m) is expressed by the angular velocity increment and the velocity increment at t_m-1/2 and t_m:

$\{\begin{array}{c}\Delta {\theta }_1={\int }_{t_{m-1}}^{t_{m-1}+T/2}{\omega }_{\mathrm{ab}}^b\left(\tau \right)d\tau \\ \Delta {\theta }_2={\int }_{t_{m-1}+T/2}^{t_m}{\omega }_{\mathrm{ab}}^b\left(\tau \right)d\tau \\ \Delta V_1={\int }_{t_{m-1}}^{t_{m-1}+T/2}{f}^b\left(\tau \right)d\tau \\ \Delta V_2={\int }_{t_{m-1}+T/2}^{t_m}{f}^b\left(\tau \right)d\tau \\ \Delta \theta ={\int }_{t_{m-1}}^{t_m}{\omega }_{\mathrm{ab}}^b\left(\tau \right)d\tau \\ \Delta V={\int }_{t_{m-1}}^{t_m}{f}^b\left(\tau \right)d\tau \end{array}\{\begin{array}{c}{\mathrm{\Delta \theta }}_1^{\prime }={\int }_{t_{m-1}}^{t_{m-1}+T/2}{\omega }_{\mathrm{ag}}^g\left(\tau \right)d\tau \approx \frac{T}{2}{\omega }_{\mathrm{ag}}^g\\ {\mathrm{\Delta \theta }}_2^{\prime }={\int }_{t_{m-1}+T/2}^{t_m}{\omega }_{\mathrm{ag}}^g\left(\tau \right)d\tau \approx \frac{T}{2}{\omega }_{\mathrm{ag}}^g\\ \Delta V_1^{\prime }={\int }_{t_{m-1}}^{t_{m-1}+T/2}{R}_{b\left(m-1\right)}^{g\left(m-1\right)}{f}^b\left(\tau \right)d\tau =R_{b\left(m-1\right)}^{g\left(m-1\right)}\Delta V_1\\ \Delta V_2^{\prime }={\int }_{t_{m-1}+T/2}^{t_m}{R}_{b\left(m-1\right)}^{g\left(m-1\right)}{f}^b\left(\tau \right)d\tau =R_{b\left(m-1\right)}^{g\left(m-1\right)}\Delta V_2\\ {\mathrm{\Delta \theta }}^{\prime }={\int }_{t_{m-1}}^{t_m}{\omega }_{\mathrm{ag}}^g\left(\tau \right)d\tau ={\mathrm{\Delta \theta }}_1^{\prime }+{\mathrm{\Delta \theta }}_2^{\prime }\\ \Delta V^{\prime }={\int }_{t_{m-1}}^{t_m}{R}_{b\left(m-1\right)}^{g\left(m-1\right)}{f}^b\left(\tau \right)d\tau =R_{b\left(m-1\right)}^{g\left(m-1\right)}\Delta V\end{array}$

A two-subsample velocity rotation error compensation algorithm is:

$\begin{array}{cc}\Delta V_{\mathrm{rot}\left(m\right)}=\frac{1}{2}\Delta \theta \times \mathrm{\Delta V}& \left(1.12\right)\\ \Delta V_{\mathrm{rot}\left(m\right)}^{\prime }=\frac{1}{2}A{\theta }^{\prime }\times \Delta V^{\prime }& \left(1.13\right)\end{array}$

A two-subsample scull error compensation algorithm is:

$\begin{array}{cc}\Delta V_{\mathrm{scul}\left(m\right)}=\frac{2}{3}\left(\Delta {\theta }_1\times \Delta V_2+\Delta V_1\times \Delta {\theta }_2\right)& \left(1.14\right)\end{array}$ $\begin{array}{cc}\Delta V_{\mathrm{scul}\left(m\right)}^{\prime }=\frac{2}{3}\left(\Delta {\theta }_1^{\prime }\times \Delta V_2^{\prime }+\Delta V_1^{\prime }\times \Delta {\theta }_2^{\prime }\right)& \left(1.15\right)\end{array}$ $\begin{array}{cc}\mathrm{and}& \end{array}$ $\begin{array}{cc}\Delta V_{\mathrm{rot}\left(m\right)}^{\prime }+\Delta V_{\mathrm{scul}\left(m\right)}^{\prime }\approx \frac{T}{6}{\omega }_{\mathrm{ag}}^g\times \left[R_{b\left(m-1\right)}^{g\left(m-1\right)}\left(\Delta V_1+5\Delta V_2\right)\right]& \left(1.16\right)\end{array}$

Thus, the complete algorithm of the specific force velocity increment in the g frame is obtained, that is, Equation (1.11) can be expressed as:

$\begin{array}{cc}\Delta V_{\mathrm{sf}\left(m\right)}^g=R_{b\left(m-1\right)}^{g\left(m-1\right)}\Delta V-\frac{T}{6}{\omega }_{\mathrm{ag}}^g\times \left[R_{b\left(m-1\right)}^{g\left(m-1\right)}\left(\Delta V_1+5\Delta V_2\right)\right]+R_{b\left(m-1\right)}^{g\left(m-1\right)}\left(\Delta V_{\mathrm{rot}\left(m\right)}+\Delta V_{\mathrm{scul}\left(m\right)}\right)& \left(1.17\right)\end{array}$

(3) Position Update

According to the position navigation equation, the following equation is obtained:{dot over (P)}^g(t)=V^g(t)  (1.10)

Due to the short time period [t_m-1,t_m], the velocity compensation term of the harmful acceleration changes very slowly during this time period, and its integral can be approximated as a linear function of time. According to Equation (1.7), both sides of (1.21) are integrated in [t_m-1,t_m]:

$\begin{array}{cc}{V}^g\left(t\right)=V_{m-1}^g+\Delta V_{\mathrm{sf}}^g\left(t\right)+\Delta V_{\mathrm{cor}/g}^g\frac{t-t_{m-1}}{T}\left(t_{m-1}\le t\le t_m\right)& \left(1.21\right)\end{array}$

to obtain:

$\begin{array}{cc}{P}_m^g=P_{m-1}^g+\left[V_{m-1}^g+\frac{1}{2}\Delta V_{\mathrm{cor}/g\left(m\right)}^g\right]T+\Delta P_{\mathrm{sf}\left(m\right)}^g& \left(1.22\right)\end{array}$

In the equation, P_m^g and P_m-1^g are the position of the SSM in the g frame at t_m and t_m-1.ΔP_sf(m)^g=R_b(m-1)^g(m-1)(S_Δvm+ΔP_rotm+ΔP_scrlm)−(ΔP′_rotm+ΔP′_scrlm)  (1.23)

And:

$\begin{array}{cc}{S}_{\Delta \mathrm{vm}}=T\left[\frac{5}{6}\Delta V_1+\frac{1}{6}\Delta V_2\right]& \left(1.24\right)\end{array}$

indicates a secondary integral increment of the specific force.

$\begin{array}{cc}\Delta P_{\mathrm{rotm}}=T\left(\Delta {\theta }_1\times \left[\frac{5}{18}\Delta V_1+\frac{1}{6}\Delta V_2\right]+\Delta {\theta }_2\times \left[\frac{1}{6}\Delta V_1+\frac{1}{18}\Delta V_2\right]\right)& \left(1.25\right)\\ \Delta P_{\mathrm{rotm}}^{\prime }=\tau \left({\mathrm{\Delta \theta }}_1^{\prime }\times \left[\frac{5}{18}\Delta V_1^{\prime }+\frac{1}{6}\Delta V_2^{\prime }\right]+{\mathrm{\Delta \theta }}_2^{\prime }\times \left[\frac{1}{6}\Delta V_1^{\prime }+\frac{1}{18}\Delta V_2^{\prime }\right]\right)& \left(1.26\right)\end{array}$

indicates a rotation effect compensation quantity in position calculation.

$\begin{array}{cc}\Delta P_{\mathrm{scrlm}}=T\left({\mathrm{\Delta \theta }}_1\times \left[\frac{11}{90}\Delta V_1+\frac{1}{10}\Delta V_2\right]+\Delta {\theta }_2\times \left[\frac{1}{90}\Delta V_2-\frac{7}{30}\Delta V_1\right]\right)& \left(1.27\right)\\ \Delta P_{\mathrm{scrlm}}^{\prime }=T\left(\Delta {\theta }_1^{\prime }\times \left[\frac{11}{90}\Delta V_1^{\prime }+\frac{1}{10}\Delta V_2^{\prime }\right]+\Delta {\theta }_2^{\prime }\times \left[\frac{1}{90}\Delta V_2^{\prime }-\frac{7}{30}\Delta V_1^{\prime }\right]\right)& \left(1.28\right)\end{array}$

indicates a scroll effect compensation quantity in position calculation. And

$\begin{array}{cc}\Delta P_{\mathrm{rotm}}^{\prime }+\Delta P_{\mathrm{scrlm}}^{\prime }\approx \frac{T^2}{6}{\omega }_{\mathrm{ag}}^g\times \left(R_{b\left(m-1\right)}^{g\left(m-1\right)}\left[\Delta V_1+\Delta V_2\right]\right)& \left(1.29\right)\end{array}$

Thus, the complete algorithm of the ΔP_sf(m)^g position increment in the g frame is obtained:

$\begin{array}{cc}\Delta P_{\mathrm{sf}\left(m\right)}^g=R_{b\left(m-1\right)}^{g\left(m-1\right)}\left(S_{\Delta \mathrm{vm}}+\Delta P_{\mathrm{rotm}}+\Delta P_{s\mathrm{crlm}}\right)-\frac{T^2}{6}{\omega }_{\mathrm{ag}}^g\times \left(R_{b\left(m-1\right)}^{g\left(m-1\right)}\left[\Delta V_1+\Delta V_2\right]\right).& \left(1.3\right)\end{array}$

Claims

1. A method for updating strapdown inertial navigation solutions using a launch-centered earth-fixed (LCEF) frame (g frame), comprising the following steps: step 1: using a surface-to-surface missile (SSM) as a body to establish a relative relationship between the SSM and a ground to keep identical missile parameters required by a missile control and guidance system, and establishing a body frame (b frame), which points front, top, and right;step 2: expressing a navigation equation in the g frame as: