Agile response method for machine system signal of rotation modulation inertial navigation system

Abstract

The invention belongs to the technical field of inertial navigation, and relates to a system signal agile response method of a rotary modulation inertial navigation system, which is suitable for scenes such as rapid starting or air restarting of an aircraft carrying rotary inertial navigation. By designing the method for rapidly outputting the angular rate and the acceleration data of the rotary inertial navigation, the traditional alignment process of the rotary inertial navigation can be canceled, the flight control data of the machine system can be provided for flight within several seconds after power-on initialization is completed, the signal response speed of the rotary inertial navigation machine system is rapidly increased, and the system reliability is improved. The method is suitable for practical scenes such as rapid takeoff or air restart of airplanes.

Description

Technical Field

[0001] The invention belongs to the technical field of inertial navigation, and relates to a method for a quick response of a rotation modulation inertial navigation system machine system signal. Background Art

[0002] The rotation modulation technology realizes the automatic cancellation of the gyro and accelerometer constant errors by introducing the alternating rotation of the rotating mechanism, thus ensuring the long-flight navigation accuracy. The additional angular velocity generated by the rotating mechanism is sensitive to the gyroscope, and the change in spatial position causes the accelerometer output to change, resulting in the inability of the gyroscope and accelerometer output to directly represent the angular velocity and acceleration information of the aircraft system. Usually, after the alignment process of several minutes or more than ten minutes is completed, the flight control signal of the aircraft system is obtained through coordinate conversion. Summary of the invention

[0003] The purpose of the present invention is to propose a method for agile response of a rotational modulation inertial navigation system machine system signal, which can greatly improve the response time of the rotational inertial navigation and output the data required for flight control.

[0004] Technical solution of the present invention: In order to achieve the above-mentioned invention purpose, a method for agile response of a rotationally modulated inertial navigation system to a machine system signal is proposed, an inertial measurement solidified coordinate system is established, and the attitude update matrix of the solidified system, IMU system, and machine system is calculated in real time by using the gyro output and the synchronized angle measurement value, and a conversion relationship between the inertial measurement coordinate system and the machine system is established, so as to output flight control data such as angular velocity and acceleration with the fastest response time.

[0005] The specific steps include:

[0006] Step 1: Periodically collect the output signals of the three-axis laser gyro and the three-axis accelerometer, set a low-pass filter, and obtain the outputs of the three-axis laser gyro and the three-axis accelerometer after jitter filtering;

[0007] Step 2: Perform accumulation processing on the gyro and accelerometer outputs for at least 10 beats to obtain the down-converted three-axis laser gyro orthogonal angular velocity and the three-axis accelerometer acceleration increment value;

[0008] Step 3: Record the IMU sensitive system as s system, and solidify the s system at the start of the calculation into the inertial solidification coordinate system s0 system, and use the real-time updated three-axis laser gyro orthogonal angular velocity increment to calculate the attitude conversion matrix between s system and s0 system at the current moment

[0009] Step 4: Use the rotating inertial navigation inner and outer frame angle measurement values, inner and outer frame installation error angles, and S- and S0-frame attitude conversion matrices synchronized with the gyro data The attitude conversion matrix between the computer system and the solidification coordinate system s0

[0010] Step 5: Use the attitude transformation matrix between the machine system and the solidification coordinate system s0 The computer system is relative to the three-axis attitude angle of the solidification coordinate system s0. The attitude angle is differentiated to obtain the angular velocity output of the machine system. The attitude transformation matrix Convert the real-time output of the accelerometer to the machine system output.

[0011] In a possible embodiment, in step 1, the original outputs of the three-axis laser gyro and the three-axis accelerometer are recorded as Gx, Gy, Gz, Ax, Ay, and Az, respectively, and are processed by a discretized low-pass filter:

[0012]

[0013] Wherein, Gx'(n) is the X gyro angular velocity increment value at the nth sampling moment after filtering, and Gx(nk) is the X gyro angular velocity increment value at the nkth sampling moment before filtering; k 、b k are the filter coefficients; Gx', Gy', Gz', Ax', Ay', and Az' are the X, Y, and Z axis angular velocities and acceleration increments after filtering, respectively.

[0014] In one possible embodiment, the three-axis laser gyro and the three-axis accelerometer output an oblique coordinate system.

[0015] In a possible embodiment, in step 2, it also includes converting the output of the three-axis gyroscope and the accelerometer from the oblique coordinate system to the orthogonal coordinate system using an oblique matrix:

[0016]

[0017] Among them, ωx, ωy, ωz, fx, fy, and fz are the three-axis angular velocity and acceleration increment values ​​of the orthogonal coordinate system respectively.

[0018] The conversion matrix coefficients for the three-axis gyro and three-axis plus oblique coordinate system;

[0019] Then, ωx, ωy, ωz, fx, fy, and fz are accumulated for at least 10 beats to obtain the incremental accumulated value of the three-axis angular velocity of the orthogonal system after frequency reduction ωx sum ,ωy sum ,ωz sum And the incremental cumulative value of the three-axis acceleration of the orthogonal system fx sum 、fy sum 、fz sum .

[0020] In a possible embodiment, in step 3, the attitude conversion matrix between the s system and the s0 system at the current moment is calculated using the real-time updated three-axis laser gyro orthogonal angular velocity increment value. The specific process includes:

[0021] Set the initial quaternion to is the quaternion of the attitude conversion between the s system and the s0 system at the initial moment;

[0022] In each subsequent calculation cycle, the current gyro angle increment ωx is used sum ,ωy sum ,ωz sum Real-time value, update quaternion

[0023] Where I is the three-dimensional unit matrix, ΔΘ=[ωx sum ωy sum ωz sum ] T , t k-1 is the kth calculation cycle, t k is the (k+1)th calculation cycle;

[0024] according to Calculate the attitude transformation matrix between the s system and the s0 system remember Then we have:

[0025]

[0026] Among them, q1, q2, q3, q4 are the quaternions of the current calculation period The four coefficients of .

[0027] In a possible embodiment, in step 4, the attitude conversion matrix between the machine system and the solidification coordinate system s0 is The specific calculation process includes:

[0028] Record the measured value α of the outer frame angle in the previous cycle and the current cycle Y (t k ), α Y (t k-1 ), α Z (t k ), α Z (t k-1 ), the time synchronization error between the angle measurement value and the gyro measurement value is compensated to obtain the inner and outer frame angle measurement value synchronized with the gyro; as shown in the following formula:

[0029]

[0030] Among them, αY_SYN , α Z_SYN is the angle measurement value after synchronization, Unsyn_T is the time that the angle measurement value leads the gyro value, and DT is the calculation period.

[0031] In a possible embodiment, the synchronized angle measurement value, the attitude conversion matrix between the s system and the s0 system in the current cycle, and the inner and outer frame installation error angle are used to calculate the conversion matrix from the b system to the s0 system:

[0032]

[0033] in, is the transformation matrix from the external frame coordinate system to the aircraft coordinate system, which is determined according to the pre-calibrated installation error angles δθ, δγ, and δψ from the external frame coordinate system to the aircraft coordinate system; δθ, δγ, and δψ are the installation error angles between the external frame system and the aircraft system's pitch axis, roll axis, and azimuth axis, respectively;

[0034]

[0035] δθ, δγ, and δψ are the installation error angles between the external frame moving coordinate system and the aircraft system’s pitch axis, roll axis, and azimuth axis, respectively;

[0036] is the transformation matrix from the static coordinate system of the outer frame to the dynamic coordinate system of the outer frame, according to the angle α of the outer frame rotation Z_SYN Sure:

[0037]

[0038] is the transformation matrix between the inner frame dynamic coordinate system and the outer frame static coordinate system, which is determined according to the calibrated inner and outer frame installation deflection angles δθ2 and δγ2;

[0039]

[0040] δθ2 and δγ2 are the installation error angles of the pitch and roll axes of the inner frame system and the outer frame, respectively;

[0041] is the transformation matrix from the inner frame static coordinate system to the inner frame dynamic coordinate system, according to the angle α of the inner frame rotation Y_SYN Sure:

[0042]

[0043] is the transformation matrix from the IMU coordinate system to the internal frame static coordinate system, which is determined according to the calibrated installation error angles δθ4, δγ4, and δψ4:

[0044]

[0045] δθ4, δγ4, and δψ4 are the installation error angles of the pitch axis, roll axis, and azimuth axis between the IMU coordinate system and the internal frame static coordinate system, respectively.

[0046] In one possible embodiment, using The three attitude angles θ′, γ′, ψ′ of the computer system b relative to the solidification coordinate system s0 are recorded at the current moment. The matrix is ​​as follows:

[0047]

[0048] Then the pitch angle, roll angle, and azimuth angle of the aircraft system b relative to the solidification system s0 are:

[0049] θ'=sin -1 (C 32 ),γ'=atan2(-C 31 ,C 33 ),θ'=atan2(C 12 ,C 22 )

[0050] The attitude angle difference is the attitude change rate:

[0051]

[0052] Using the attitude angle and attitude change rate, the machine system angular rate output is calculated as follows:

[0053]

[0054] in, are the angular velocities of the aircraft along the X, Y, and Z axes respectively.

[0055] Using the pose matrix calculated in real time Acceleration increment value fx after filtering sum 、fy sum 、fz sum Computer system acceleration:

[0056]

[0057] in, are the accelerations of the aircraft's X, Y, and Z axes, respectively, and DT is the calculation period.

[0058] Beneficial technical effects of the present invention:

[0059] The present invention designs a method for agile response of the machine system signals of a rotational modulation inertial navigation system, which does not require alignment and does not affect the rotational inertial navigation workflow. It utilizes the inertial measurement solidified coordinate system to establish a conversion matrix from the inertial measurement system to the machine system, and outputs flight control data at the fastest speed. It can support combat missions such as rapid aircraft deployment and is suitable for application scenarios such as rapid deployment or aerial restart. BRIEF DESCRIPTION OF THE DRAWINGS

[0060] Figure 1 A flow chart of a method according to a preferred embodiment of the present invention. DETAILED DESCRIPTION

[0061] In order to make the purpose, technical solution and advantages of the present invention more clear, the embodiments of the present invention are described in detail below. It should be noted that the embodiments and features in the embodiments of the present application can be combined with each other arbitrarily without conflict.

[0062] Embodiment 1:

[0063] like Figure 1 As shown, a method for agile response of a rotation modulation inertial navigation system machine signal comprises the following steps:

[0064] Step 1: Power on the system to complete the initialization of the external bus and internal data. Collect the 0.25ms three-axis laser gyro and three-axis accelerometer output signals in real time. Set a low-pass filter to obtain the three-axis gyro and accelerometer outputs after jitter filtering. The original outputs of the three-axis gyro and acceleration are recorded as Gx, Gy, Gz, Ax, Ay, and Az, respectively, and processed with the discretized low-pass filter.

[0065]

[0066] Wherein, Gx'(n) is the X-gyro output at the nth 0.25ms moment after filtering, and Gx(nk) is the X-gyro output at the nkth 2.5ms moment before filtering. k 、b k are the filter coefficients:

[0067] a0=1, a1=-6.0678, a2=15.8347, a3=-23.0316,

[0068] a4=20.1610, a5=-10.6191, a6=3.1157, a7=-0.3928

[0069] b0=8.0916, b1=-0.0002, b2=0.0004, b3=-0.0002,

[0070] b4=-0.0002, b5=0.0004, b6=-0.0002, b7=8.0916

[0071] Let Gx', Gy', Gz', Ax', Ay', and Az' be the outputs of the X, Y, and Z axis gyro and accelerometer after filtering, respectively.

[0072] Step 2: The outputs of the three-axis gyroscope and the three-axis accelerometer are in an oblique coordinate system. The oblique matrix is ​​used to convert the outputs of the three-axis gyroscope and the accelerometer from the oblique coordinate system to the orthogonal coordinate system.

[0073] Among them, ωx, ωy, ωz, fx, fy, and fz are the gyroscope and accelerometer outputs of the orthogonal coordinate system respectively.

[0074] The gyro and accelerometer outputs are accumulated for 10 times to obtain 2.5ms angular velocity and acceleration increment values.

[0075]

[0076] Among them, ωx sum ,ωy sum ,ωz sum 、fx sum 、fy sum 、fz sum The gyro and accelerometer increments are accumulated every 2.5ms. i ,ωy i ,ωz i 、fx i 、fy i 、fz i The gyro and accelerometer increments for every 0.25ms period of the i-th period within each 2.5ms.

[0077] Step 3: Record the IMU sensitive system as the s system, and solidify the s system at the start of the calculation into the s0 system fixed to the inertial space. Use the real-time updated three-axis gyro orthogonal output to calculate the attitude conversion matrix between the s system and the s0 system at the current moment.

[0078] At the start of the calculation, the initial quaternion is set to It is the quaternion of the attitude conversion between the s system and the s0 system at the initial moment.

[0079] In each subsequent 2.5ms calculation cycle, the current gyro angle increment ωx is used sum ,ωy sum ,ωz sum Real-time value, update quaternion

[0080] Where I is the three-dimensional unit matrix, ΔΘ=[ωx sum ωy sum ωz sum ] T , t k-1 is the kth 2.5ms calculation cycle, t k This is the (k+1)th 2.5ms calculation cycle.

[0081] according to Calculate the attitude transformation matrix between the s system and the s0 system remember Then we have:

[0082]

[0083] Step 4: Use the rotation angle measurement values ​​of the inner and outer frames of the rotating inertial navigation system synchronized with the gyro data, the inner and outer frame installation error angle computer system and the attitude conversion matrix of the solidified coordinate system S0 system

[0084] Record the measured values ​​α of the inner and outer frame angles in the previous cycle and the current cycle Y (t k ), α Y (t k-1 ), α Z (t k ), α Z (t k-1 ), the time synchronization error between the angle measurement value and the gyro measurement value is compensated to obtain the inner and outer frame angle measurement value synchronized with the gyro. As shown in the following formula:

[0085]

[0086] Among them, α Y_SYN , α Z_SYN is the angle measurement value after synchronization, Unsyn_T is the time that the angle measurement value leads the gyro value, and DT is 0.0025s.

[0087] The transformation matrix from the b system (aircraft coordinate system) to the s0 system is calculated using the synchronized angle measurement value, the attitude transformation matrix of the s system and the s0 system in the current cycle, and the installation error angle of the inner and outer frames:

[0088]

[0089] in, is the transformation matrix from the external frame coordinate system to the aircraft coordinate system, which is determined based on the pre-calibrated installation error angles δθ, δγ, and δψ from the external frame coordinate system to the aircraft coordinate system. δθ, δγ, and δψ are the installation error angles between the external frame system and the aircraft system's pitch axis, roll axis, and azimuth axis, respectively.

[0090]

[0091] δθ, δγ, and δψ are the installation error angles between the external frame dynamic coordinate system and the aircraft system's pitch axis, roll axis, and azimuth axis, respectively.

[0092] is the transformation matrix from the static coordinate system of the outer frame to the dynamic coordinate system of the outer frame, according to the angle α of the outer frame rotation Z_SYN Sure:

[0093]

[0094] It is the transformation matrix between the inner frame dynamic coordinate system and the outer frame static coordinate system, which is determined according to the calibrated inner and outer frame installation deflection angles δθ2 and δγ2.

[0095]

[0096] δθ2 and δγ2 are the installation error angles of the pitch and roll axes of the inner frame system and the outer frame, respectively.

[0097] is the transformation matrix from the inner frame static coordinate system to the inner frame dynamic coordinate system, according to the angle α of the inner frame rotation Y_SYN Sure:

[0098]

[0099] is the transformation matrix from the IMU coordinate system to the internal frame static coordinate system, which is determined according to the calibrated installation error angles δθ4, δγ4, and δψ4:

[0100]

[0101] δθ4, δγ4, δψ4 are the installation error angles of the pitch axis, roll axis, and azimuth axis between the IMU coordinate system and the internal frame static coordinate system respectively.

[0102] Step 5: Use the posture transformation matrix The computer system's three-axis attitude angle relative to the solidification coordinate system s0 is differentiated to obtain the angular velocity output of the machine system. Using the attitude transformation matrix Convert the real-time output of the accelerometer to the machine system output. The matrix is ​​as follows

[0103]

[0104] Then the pitch angle, roll angle, and azimuth angle of the aircraft system b relative to the solidification system s0 are:

[0105] θ'=sin -1 (C32 ),γ'=atan2(-C 31 ,C 33 ),θ'=atan2(C 12 ,C 22 )

[0106] The attitude angle difference is the attitude change rate:

[0107]

[0108] Using the attitude angle and attitude change rate, the machine system angular rate output is calculated as follows:

[0109]

[0110] in, are the angular velocities of the aircraft along the X, Y, and Z axes respectively.

[0111] Using the pose matrix calculated in real time The filtered accelerometer outputs fx, fy, fz are the computer system accelerations:

[0112]

[0113] in, They are the accelerations of the aircraft on the X, Y, and Z axes. DT is 0.0025s.

[0114] Embodiment 2:

[0115] A method for agile response of a rotation modulation inertial navigation system machine system signal comprises the following steps:

[0116] Step 1: Power on the system to complete the initialization of the external bus and internal data. Collect the 0.5ms three-axis fiber optic gyroscope and three-axis accelerometer output signals in real time. Set a low-pass filter to obtain the three-axis gyroscope and accelerometer outputs after jitter filtering. The original outputs of the three-axis gyroscope and acceleration are recorded as Gx, Gy, Gz, Ax, Ay, and Az, respectively, and processed with a discretized low-pass filter.

[0117]

[0118] Wherein, Gx'(n) is the X-gyro output at the nth 0.5ms moment after filtering, and Gx(nk) is the X-gyro output at the nkth 5ms moment before filtering. k 、b k are the filter coefficients:

[0119] a0=1, a1=0, a2=0, a3=0,

[0120] a4=0, a5=0, a6=0, a7=0

[0121] b0=0, b1=0, b2=0, b3=0,

[0122] b4=0, b5=0, b6=0, b7=0

[0123] Let Gx', Gy', Gz', Ax', Ay', and Az' be the outputs of the X, Y, and Z axis gyro and accelerometer after filtering, respectively.

[0124] Step 2: The outputs of the three-axis gyroscope and the three-axis accelerometer are in an oblique coordinate system. The oblique matrix is ​​used to convert the outputs of the three-axis gyroscope and the accelerometer from the oblique coordinate system to the orthogonal coordinate system.

[0125] Among them, ωx, ωy, ωz, fx, fy, and fz are the gyroscope and accelerometer outputs of the orthogonal coordinate system respectively.

[0126] The gyro and accelerometer outputs are accumulated for 10 times to obtain 5 ms angular velocity and acceleration increment values.

[0127]

[0128] Among them, ωx sum ,ωy sum ,ωz sum 、fx sum 、fy sum 、fz sum The gyro and accelerometer increments are accumulated every 5ms. i ,ωy i ,ωz i 、fx i 、fy i 、fz i The gyro and accelerometer increments for the i-th 0.5ms period within each 5ms.

[0129] Step 3: Record the IMU sensitive system as the s system, and solidify the s system at the start of the calculation into the s0 system fixed to the inertial space. Use the real-time updated three-axis gyro orthogonal output to calculate the attitude conversion matrix between the s system and the s0 system at the current moment.

[0130] At the start of the calculation, the initial quaternion is set to It is the quaternion of the attitude conversion between the s system and the s0 system at the initial moment.

[0131] In each subsequent 5ms calculation cycle, the current gyro angle increment ωx is used sum ,ωy sum ,ωz sumReal-time value, update quaternion

[0132] Where I is the three-dimensional unit matrix, ΔΘ=[ωx sum ωy sum ωz sum ] T , t k-1 is the kth 2.5ms calculation cycle, t k This is the (k+1)th 5ms calculation cycle.

[0133] according to Calculate the attitude transformation matrix between the s system and the s0 system remember Then we have:

[0134]

[0135] Step 4: Use the rotation angle measurement values ​​of the inner and outer frames of the rotating inertial navigation system synchronized with the gyro data, the inner and outer frame installation error angle computer system and the attitude conversion matrix of the solidified coordinate system S0 system

[0136] Record the measured values ​​α of the inner and outer frame angles in the previous cycle and the current cycle Y (t k ), α Y (t k-1 ), α Z (t k ), α Z (t k-1 ), the time synchronization error between the angle measurement value and the gyro measurement value is compensated to obtain the inner and outer frame angle measurement value synchronized with the gyro. As shown in the following formula:

[0137]

[0138] Among them, α Y_SYN , α Z_SYN is the angle measurement value after synchronization, Unsyn_T is the time that the angle measurement value leads the gyro value, and DT is 0.0025s.

[0139] The transformation matrix from the b system (aircraft coordinate system) to the s0 system is calculated using the synchronized angle measurement value, the attitude transformation matrix of the s system and the s0 system in the current cycle, and the installation error angle of the inner and outer frames:

[0140]

[0141] in, is the transformation matrix from the external frame coordinate system to the aircraft coordinate system, which is determined based on the pre-calibrated installation error angles δθ, δγ, and δψ from the external frame coordinate system to the aircraft coordinate system. δθ, δγ, and δψ are the installation error angles between the external frame system and the aircraft system's pitch axis, roll axis, and azimuth axis, respectively.

[0142]

[0143] δθ, δγ, and δψ are the installation error angles between the external frame dynamic coordinate system and the aircraft system's pitch axis, roll axis, and azimuth axis, respectively.

[0144] is the transformation matrix from the static coordinate system of the outer frame to the dynamic coordinate system of the outer frame, according to the angle α of the outer frame rotation Z_SYN Sure:

[0145]

[0146] It is the transformation matrix between the inner frame dynamic coordinate system and the outer frame static coordinate system, which is determined according to the calibrated inner and outer frame installation deflection angles δθ2 and δγ2.

[0147]

[0148] δθ2 and δγ2 are the installation error angles of the pitch and roll axes of the inner frame system and the outer frame, respectively.

[0149] is the transformation matrix from the inner frame static coordinate system to the inner frame dynamic coordinate system, according to the angle α of the inner frame rotation Y_SYN Sure:

[0150]

[0151] is the transformation matrix from the IMU coordinate system to the internal frame static coordinate system, which is determined according to the calibrated installation error angles δθ4, δγ4, and δψ4:

[0152]

[0153] δθ4, δγ4, δψ4 are the installation error angles of the pitch axis, roll axis, and azimuth axis between the IMU coordinate system and the internal frame static coordinate system respectively.

[0154] Step 5: Use the posture transformation matrix The computer system's three-axis attitude angle relative to the solidification coordinate system s0 is differentiated to obtain the angular velocity output of the machine system. Using the attitude transformation matrix Convert the real-time output of the accelerometer to the machine system output. The matrix is ​​as follows

[0155]

[0156] Then the pitch angle, roll angle, and azimuth angle of the aircraft system b relative to the solidification system s0 are:

[0157] θ'=sin -1 (C 32 ),γ'=atan2(-C 31 ,C 33 ),θ'=atan2(C 12 ,C 22 )

[0158] The attitude angle difference is the attitude change rate:

[0159]

[0160] Using the attitude angle and attitude change rate, the machine system angular rate output is calculated as follows:

[0161]

[0162] in, are the angular velocities of the aircraft along the X, Y, and Z axes respectively.

[0163] Using the pose matrix calculated in real time The filtered accelerometer outputs fx, fy, fz are the computer system accelerations:

[0164]

[0165] in, They are the accelerations of the aircraft on the X, Y, and Z axes. DT is 0.005s.

[0166] Although the embodiments disclosed in the present invention are as above, the contents described are only embodiments adopted to facilitate understanding of the present invention and are not intended to limit the present invention. Any technician in the field to which the present invention belongs can make any modifications and changes in the form and details of implementation without departing from the spirit and scope disclosed in the present invention, but the patent protection scope of the present invention shall still be subject to the scope defined in the attached claims.

Claims

1. A method for agile response of a rotation modulation inertial navigation system system signal, characterized in that: Step 1: Periodically collect the output signals of the three-axis laser gyro and the three-axis accelerometer, and obtain the outputs of the three-axis laser gyro and the three-axis accelerometer after jitter filtering by setting a low-pass filter; Step 2: Perform at least 10 beats of accumulation processing on the gyro and accelerometer outputs to obtain the three-axis laser gyro orthogonal angular velocity and the three-axis accelerometer acceleration increment after frequency reduction; Step 3: Record the IMU sensitive system as the s system, and solidify the s system at the start of the calculation as the inertial solidification coordinate system s0 system, and use the real-time updated three-axis laser gyro orthogonal angular velocity increment to calculate the attitude conversion matrix between the s system and the s0 system at the current moment Step 4: Use the rotating inertial navigation inner and outer frame angle measurement values, inner and outer frame installation error angles, and S- and S0-frame attitude conversion matrices synchronized with the gyro data The attitude conversion matrix between the computer system and the solidification coordinate system s0 Step 5: Use the attitude transformation matrix between the machine system and the solidification coordinate system s0 The computer system is relative to the three-axis attitude angle of the solidification coordinate system s0. The attitude angle is differentiated to obtain the angular velocity output of the machine system. The attitude transformation matrix Convert the real-time output of the accelerometer to the machine system output.

2. The method for agile response of a rotation modulation inertial navigation system system signal according to claim 1, characterized in that: In step 1, the original outputs of the three-axis laser gyro and the three-axis accelerometer are recorded as Gx, Gy, Gz, Ax, Ay, and Az, respectively, and are processed by a discretized low-pass filter: Wherein, Gx'(n) is the X gyro angular velocity increment value at the nth sampling moment after filtering, and Gx(nk) is the X gyro angular velocity increment value at the nkth sampling moment before filtering; k 、b k are the filter coefficients; Gx', Gy', Gz', Ax', Ay', and Az' are the X, Y, and Z axis angular velocities and acceleration increments after filtering, respectively.

3. The method for agile response of a rotation modulation inertial navigation system system signal according to claim 2, characterized in that: The output of the three-axis laser gyro and the three-axis accelerometer is an oblique coordinate system.

4. The method for agile response of a rotation modulation inertial navigation system system signal according to claim 3, characterized in that: It also includes using the tilt matrix to convert the outputs of the three-axis laser gyro and the three-axis accelerometer from the tilt coordinate system to the orthogonal coordinate system. The specific process includes: Among them, ωx, ωy, ωz, fx, fy, and fz are the three-axis angular velocity and acceleration increment values ​​of the orthogonal coordinate system respectively. The conversion matrix coefficients for the three-axis gyro and three-axis plus oblique coordinate system; Then, ωx, ωy, ωz, fx, fy, and fz are accumulated for at least 10 beats to obtain the incremental accumulated value of the three-axis angular velocity of the orthogonal system after frequency reduction ωx sum ,ωy sum ,ωz sum And the incremental cumulative value of the three-axis acceleration of the orthogonal system fx sum 、fy sum 、fz sum .

5. The method for agile response of a rotation modulation inertial navigation system system signal according to claim 4, characterized in that: In step 3, the attitude conversion matrix between the s system and the s0 system at the current moment is calculated using the real-time updated three-axis laser gyro orthogonal angular velocity increment value The specific process includes: Set the initial quaternion to is the quaternion of the attitude conversion between the s system and the s0 system at the initial moment; In each subsequent calculation cycle, the current gyro angle increment ωx is used sum ,ωy sum ,ωz sum Real-time value, update quaternion Where I is the three-dimensional unit matrix, ΔΘ=[ωx sum ωy sum ωz sum ] T , t k-1 is the kth calculation cycle, t k is the (k+1)th calculation cycle; according to Calculate the attitude transformation matrix between the s system and the s0 system remember Then we have: Among them, q1, q2, q3, q4 are the quaternions of the current calculation period The four coefficients of .

6. The method for agile response of a rotation modulation inertial navigation system system signal according to claim 5, characterized in that: In step 4, the attitude conversion matrix between the machine system and the solidification coordinate system s0 is The specific calculation process includes: Record the measured value α of the outer frame angle in the previous cycle and the current cycle Y (t k ), α Y (t k-1 ), α Z (t k ), α Z (t k-1 ), the time synchronization error between the angle measurement value and the gyro measurement value is compensated to obtain the inner and outer frame angle measurement value synchronized with the gyro; as shown in the following formula: Among them, α Y_SYN , α Z_SYN is the angle measurement value after synchronization, Unsyn_T is the time that the angle measurement value leads the gyro value, and DT is the calculation period.

7. The method for agile response of a rotation modulation inertial navigation system system signal according to claim 6, characterized in that: Using the synchronized angle measurement value, the attitude conversion matrix between the current cycle s system and s0 system, and the inner and outer frame installation error angle, the conversion matrix from the computer system b system to the s0 system is: in, is the transformation matrix from the external frame coordinate system to the aircraft coordinate system, which is determined according to the pre-calibrated installation error angles δθ, δγ, and δψ from the external frame coordinate system to the aircraft coordinate system; δθ, δγ, and δψ are the installation error angles between the external frame system and the aircraft system's pitch axis, roll axis, and azimuth axis, respectively; δθ, δγ, and δψ are the installation error angles between the external frame moving coordinate system and the aircraft system pitch axis, roll axis, and azimuth axis, respectively; is the transformation matrix from the static coordinate system of the outer frame to the dynamic coordinate system of the outer frame, according to the angle α of the outer frame rotation Z_SYN Sure: is the transformation matrix between the inner frame dynamic coordinate system and the outer frame static coordinate system, which is determined according to the calibrated inner and outer frame installation deflection angles δθ2 and δγ2; δθ2 and δγ2 are the installation error angles of the pitch and roll axes of the inner frame system and the outer frame, respectively; is the transformation matrix from the static coordinate system of the inner frame to the dynamic coordinate system of the inner frame, according to the angle α of the inner frame rotation Y_SYN Sure: is the transformation matrix from the IMU coordinate system to the internal frame static coordinate system, which is determined according to the calibrated installation error angles δθ4, δγ4, and δψ4: δθ4, δγ4, and δψ4 are the installation error angles of the pitch axis, roll axis, and azimuth axis between the IMU coordinate system and the internal frame static coordinate system, respectively.

8. The method for agile response of a rotation modulation inertial navigation system system signal according to claim 7, characterized in that: use The three attitude angles θ′, γ′, ψ′ of the computer system b relative to the solidification coordinate system s0 are recorded at the current moment. The matrix is ​​as follows: Then the pitch angle, roll angle, and azimuth angle of the aircraft system b relative to the solidification system s0 are: θ'=sin -1 (C 32 )、γ'=atan2(-C 31 ,C 33 )、θ'=atan2(C 12 ,C 22 ) The attitude angle difference is the attitude change rate: Using the attitude angle and attitude change rate, the machine system angular rate output is calculated as follows: in, are the angular velocities of the aircraft along the X, Y, and Z axes respectively. Using the pose matrix calculated in real time Acceleration increment value fx after filtering sum 、fy sum 、fz sum Computer system acceleration: in, are the accelerations of the aircraft's X, Y, and Z axes, respectively, and DT is the calculation period.

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