Method and device for calibrating attitude of laser inertial measurement unit in erecting process of launcher

Abstract

The invention provides a calibration method and device for the attitude of a laser inertial measurement unit in the erecting process of a launching rack. The method comprises the following steps: acquiring attitude angle data obtained through an experiment, wherein the attitude angle data comprises attitude angles output by an inertial measurement unit at different pitch angles in the erection process of a launcher; based on a rotation transformation matrix among a navigation coordinate system, an inertial measurement unit carrier coordinate system and a launcher carrier coordinate system, solving the relationship between the installation error between the inertial measurement unit and the launcher and the attitude angle; and calculating the installation error based on the relationship and the attitude angle data to realize calibration of the attitude of the inertial measurement unit in the erection process of the launcher. The problems that in the erecting process of the launcher, the installation error between the inertial measurement unit and the azimuth angle of the launcher is not fixed, and the azimuth angle output by the inertial measurement unit changes along with erecting of the launcher are solved, and the precision of the azimuth angle output by the inertial measurement unit in the erecting process is improved.

Description

Technical Field

[0001] The invention belongs to the technical field of laser inertial group calibration, and in particular relates to a method and a device for calibrating the attitude of a laser inertial group during the erection process of a launcher. Background Art

[0002] The positioning and orientation system includes a laser inertial group (hereinafter referred to as the inertial group), an odometer, a satellite antenna, an altimeter, a display controller, etc. The inertial group is installed on the launcher of the launch vehicle and is an important part of the positioning and orientation system. Figure 2 It is a schematic diagram of the coordinate system of the inertial group carrier relative to the coordinate system of the launcher carrier. The inertial group should be able to accurately measure the azimuth value of the launcher, but there is an installation error between the inertial group and the launcher (azimuth error μ x , pitch angle error μ y , Roll angle error μ z ), in order to improve the attitude measurement accuracy of the inertial group, the installation error should be corrected and compensated.

[0003] The IMU is usually installed on the chassis of the launch vehicle. The difference between the azimuth (also called heading angle) of the laser strapdown IMU system and the azimuth of the system on the launch vehicle chassis caused by the installation error is a fixed value. In normal cases, the IMU calibration only needs to use a theodolite to measure the difference between the azimuth of the standard point on the launch vehicle chassis and the azimuth output by the laser strapdown IMU as the installation error angle. When the IMU is installed on the launcher, during the erection of the launcher, the difference between the azimuth of the IMU and the launcher is not fixed. The azimuth output by the IMU changes with the erection of the launcher, which will cause the error of the above calibration method to increase or even fail. Summary of the invention

[0004] In order to solve the above problems existing in the prior art, the present invention provides a method and device for calibrating the attitude of the laser inertial group during the erection of the launcher, which can make the azimuth angle output by the inertial group during the erection of the launcher accurately represent the heading of the launch vehicle launcher system.

[0005] In order to achieve the above object, the present invention adopts the following technical solutions.

[0006] In a first aspect, the present invention provides a method for calibrating the attitude of a laser inertial group during the erection of a launcher, comprising the following steps:

[0007] Acquire the attitude angle data obtained through experiments, including the attitude angle output by the inertial group at different pitch angles during the launcher erection process;

[0008] Based on the rotation transformation matrix between the navigation coordinate system, the IMU carrier coordinate system and the launcher carrier coordinate system, solving the relationship between the installation error between the IMU and the launcher and the attitude angle;

[0009] The installation error is calculated based on the relationship and the attitude angle data to achieve calibration of the inertial group attitude during the erection of the launcher.

[0010] Furthermore, the method for solving the relationship between the installation error between the inertial group and the launcher and the attitude angle includes:

[0011] Define n as the navigation coordinate system, b as the inertial group carrier coordinate system, and B as the launcher carrier coordinate system; the rotation transformation matrix from n to B is for:

[0012]

[0013] In the formula, θ and γ are the azimuth, pitch, and roll angles required to rotate from the n system to the B system, respectively;

[0014] The installation error between the inertial group and the launcher is set as the pitch angle installation error μ x , Roll angle installation error μ y 、Azimuth installation error μ z ;μ x , μ y , μ z Both are close to 0. After approximation with formula (1), the rotation transformation matrix from system b to system B is obtained:

[0015]

[0016] During the erection of the launcher, the roll angle γ is approximately 0. Substituting γ = 0 into equation (1) yields the simplified Then we get the rotation transformation matrix from B system to n system

[0017]

[0018] Based on matrix multiplication, the rotation transformation matrix from b system to n system is obtained

[0019]

[0020] In the formula, for Elements corresponding to rows and columns;

[0021] Set the azimuth, pitch, and roll angles required to rotate from the n system to the b system to be θ', γ', when the launcher pitch angle θ = 0, substitute θ = 0 into equation (4) and combine with equation (1) to obtain:

[0022]

[0023] Thus we get:

[0024] μ x = -θ' (7)

[0025] μ y = -γ'; (8)

[0026] When the launcher pitch angle θ is not 0, we get:

[0027]

[0028] From (9) and (10), we can get: z =γ' / tanθ';

[0029] Due to the error in the attitude output by the inertial group, the pitch angle and roll angle of the launcher after erection are set to θ1′ and γ1′, and the pitch angle and roll angle of the launcher when the pitch angle is 0 are set to θ0′ and γ0′, μ z The calculation formula is:

[0030]

[0031] The installation error between the inertial group and the launch pad can be calculated according to equations (7), (8) and (11).

[0032] Furthermore, the method of obtaining attitude angle data through experiments includes:

[0033] S1. After the launch vehicle stops at the test site, it is turned off and stationary, and the vehicle body is leveled so that the pitch angle of the launcher is 0;

[0034] S2, the IMU starts to seek north and obtains the pitch angle θ' and the roll angle γ';

[0035] S3, inertial group shutdown;

[0036] S4, the inertial group starts to seek north, and obtains the pitch angle θ0′ and the roll angle γ0′; the launcher is erected to a pitch angle of α, and the pitch angle θ1′ and the roll angle γ1′ are obtained;

[0037] S5, inertial group shutdown.

[0038] Furthermore, steps S2 and S4 are repeated to measure N sets of data: θ i ′、γ i ′ are θ' and γ' measured for the i-th time; θ0′ i ,γ0′ i are θ0′, γ0′, θ1′ measured for the i-th time respectively i ,γ1′ i are θ1′ measured at the i-th time i ,γ1′ i, i=1,2,......,N.

[0039] Furthermore, the method for calculating the installation error based on the relationship and the attitude angle data includes:

[0040] Calculate the pitch angle installation error and roll angle installation error as follows:

[0041]

[0042] In the formula, μ x , μ y They are the pitch angle installation error and the roll angle installation error respectively;

[0043] Calculate the azimuth installation error as follows:

[0044]

[0045] In the formula, μ z is the azimuth installation error.

[0046] In a second aspect, the present invention provides a device for calibrating the attitude of a laser inertial group during the erection of a launcher, comprising:

[0047] A data acquisition module is used to obtain the attitude angle data obtained through experiments, including the attitude angle output by the inertial group at different pitch angles during the erection process of the launcher;

[0048] A relationship solving module, used for solving the relationship between the installation error between the inertial group and the launcher and the attitude angle based on the rotation transformation matrix between the navigation coordinate system, the inertial group carrier coordinate system and the launcher carrier coordinate system;

[0049] The inertial group calibration module is used to calculate the installation error based on the relationship and the attitude angle data, so as to realize the calibration of the inertial group attitude during the erection process of the launcher.

[0050] Furthermore, the method for solving the relationship between the installation error between the inertial group and the launcher and the attitude angle includes:

[0051] Define n as the navigation coordinate system, b as the inertial group carrier coordinate system, and B as the launcher carrier coordinate system; the rotation transformation matrix from n to B is for:

[0052]

[0053] In the formula, θ and γ are the azimuth, pitch, and roll angles required to rotate from the n system to the B system, respectively;

[0054] The installation error between the inertial group and the launcher is set as the pitch angle installation error μ x, Roll angle installation error μ y 、Azimuth installation error μ z ;μ x , μ y , μ z Both are close to 0. After approximation with formula (1), the rotation transformation matrix from system b to system B is obtained:

[0055]

[0056] During the erection of the launcher, the roll angle γ is approximately 0. Substituting γ = 0 into equation (1) yields the simplified Then we get the rotation transformation matrix from B system to n system

[0057]

[0058] Based on matrix multiplication, the rotation transformation matrix from b system to n system is obtained

[0059]

[0060] In the formula, for Elements corresponding to rows and columns;

[0061] Set the azimuth, pitch, and roll angles required to rotate from the n system to the b system to be θ', γ', when the launcher pitch angle θ = 0, substitute θ = 0 into equation (4) and combine with equation (1) to obtain:

[0062]

[0063] Thus we get:

[0064] μ x = -θ' (7)

[0065] μ y = -γ'; (8)

[0066] When the launcher pitch angle θ is not 0, we get:

[0067]

[0068] From (9) and (10), we can get: z =γ' / tanθ';

[0069] Due to the error in the attitude output by the inertial group, the pitch angle and roll angle of the launcher after erection are set to θ1′ and γ1′, and the pitch angle and roll angle of the launcher when the pitch angle is 0 are set to θ0′ and γ0′, μ zThe calculation formula is:

[0070]

[0071] The installation error between the inertial group and the launch pad can be calculated according to equations (7), (8) and (11).

[0072] Furthermore, the method of obtaining attitude angle data through experiments includes:

[0073] S1. After the launch vehicle stops at the test site, it is turned off and stationary, and the vehicle body is leveled so that the pitch angle of the launcher is 0;

[0074] S2, the IMU starts to seek north and obtains the pitch angle θ' and the roll angle γ';

[0075] S3, inertial group shutdown;

[0076] S4, the inertial group starts to seek north, and obtains the pitch angle θ′0 and the roll angle γ′0; the launcher is erected to a pitch angle of α, and the pitch angle θ′1 and the roll angle γ′1 are obtained;

[0077] S5, inertial group shutdown.

[0078] Furthermore, steps S2 and S4 are repeated to measure N sets of data: θ′ i , γ′ i are θ′ and γ′ measured for the i-th time respectively; θ′ 0i , γ′ 0i are θ′0, γ′0, θ′ measured for the i-th time respectively. 1i , γ′ 1i are θ′ measured at the i-th time 1i , γ′ 1i , i=1,2,......,N.

[0079] Furthermore, the method for calculating the installation error based on the relationship and the attitude angle data includes:

[0080] Calculate the pitch angle installation error and roll angle installation error as follows:

[0081]

[0082] In the formula, μ x , μ y They are the pitch angle installation error and the roll angle installation error respectively;

[0083] Calculate the azimuth installation error as follows:

[0084]

[0085] In the formula, μ z is the azimuth installation error.

[0086] Compared with the prior art, the present invention has the following beneficial effects.

[0087] The present invention obtains the attitude angle data output by the inertial group at different pitch angles during the erection of the launcher through experiments, solves the relationship between the installation error between the inertial group and the launcher and the attitude angle based on the navigation coordinate system, the inertial group carrier coordinate system and the launcher carrier coordinate system, calculates the installation error based on the relationship and the attitude angle data, and realizes the calibration of the inertial group attitude during the erection of the launcher. The present invention solves the problems that the installation error between the inertial group and the azimuth angle of the launcher is not fixed during the erection of the launcher, and the azimuth angle output by the inertial group changes with the erection of the launcher, thereby improving the azimuth angle accuracy of the inertial group output during the erection process. BRIEF DESCRIPTION OF THE DRAWINGS

[0088] Figure 1 The present invention is a flowchart of a method for calibrating the attitude of a laser inertial group during the erection of a launcher according to an embodiment of the present invention.

[0089] Figure 2 This is a schematic diagram of the IMU carrier coordinate system relative to the launch pad carrier coordinate system.

[0090] Figure 3 The figure is a schematic diagram of the installation of the IMU on the launch pad, where θ is the vertical angle of the launch pad, the X-axis, Y-axis, and Z-axis are the pitch axis, roll axis, and azimuth axis of the launch pad carrier coordinate system, and the x-axis, y-axis, and z-axis are the pitch axis, roll axis, and azimuth axis of the IMU carrier coordinate system.

[0091] Figure 4 The present invention is a block diagram of a device for calibrating the attitude of a laser inertial group during the erection of a launcher according to an embodiment of the present invention. DETAILED DESCRIPTION

[0092] In order to make the purpose, technical solution and advantages of the present invention clearer and more understandable, the present invention is further described below in conjunction with the accompanying drawings and specific implementation methods. Obviously, the described embodiments are only part of the embodiments of the present invention, not all of the embodiments. Based on the embodiments of the present invention, all other embodiments obtained by ordinary technicians in this field without making creative work are within the scope of protection of the present invention.

[0093] Figure 1 The present invention is a flowchart of a method for calibrating the attitude of a laser inertial group during the erection of a launcher, which includes the following steps:

[0094] Step 101, acquiring attitude angle data obtained through experiments, including attitude angles output by the inertial group at different pitch angles during the erection process of the launcher;

[0095] Step 102, based on the rotation transformation matrix between the navigation coordinate system, the inertial group carrier coordinate system and the launcher carrier coordinate system, solve the relationship between the installation error between the inertial group and the launcher and the attitude angle;

[0096] Step 103, calculating the installation error based on the relationship and the attitude angle data, so as to calibrate the attitude of the inertial group during the erection process of the launcher.

[0097] As mentioned above, there is an installation error (μ x , μ y , μ z ),like Figure 2 As shown. In order to improve the attitude measurement accuracy of the inertial group, it is necessary to correct and compensate for the installation error. For the inertial group installed on the chassis of the launch vehicle, the difference between the azimuth of the inertial group carrier system and the azimuth of the launch vehicle chassis carrier system caused by the installation error is a fixed value. In this case, the inertial group attitude calibration only needs to use a theodolite to measure the difference between the azimuth of the standard point of the launch vehicle chassis and the azimuth output by the inertial group as the installation error angle. However, when the inertial group is installed on the launcher, during the erection of the launcher, the difference between the azimuth of the inertial group and the launcher is not fixed, and the azimuth output by the inertial group changes with the erection of the launcher, which increases the error of the above calibration method or even makes it invalid. To this end, this embodiment proposes a method for calibrating the attitude of the laser inertial group during the erection of the launcher, which can effectively eliminate or weaken the influence of the non-fixed difference between the azimuth of the inertial group and the launcher during the erection of the launcher.

[0098] In this embodiment, step 101 is mainly used to obtain attitude angle data obtained through experiments. The attitude angle data includes the attitude angles (azimuth, pitch and roll) output by the inertial group at different pitch angles during the erection of the launcher, for example, the attitude angle output by the inertial group when the launcher is in a horizontal state, that is, when the pitch angle is 0°, and the attitude angle output by the inertial group when the pitch angle of the launcher is 60°.

[0099] In this embodiment, step 102 is mainly used to solve the relationship between the installation error between the inertial group and the launcher and the attitude angle output by the inertial group during the erection of the launcher. This embodiment involves three coordinate systems, namely the navigation coordinate system, the inertial group carrier coordinate system and the launcher carrier coordinate system. This embodiment is based on the rotation transformation matrix between the above three different coordinate systems, after some approximate processing, to solve the relationship between the installation error and the attitude angle output by the inertial group during the erection of the launcher, that is, the formula for calculating the installation error angle based on the attitude angle.

[0100] In this embodiment, step 103 is mainly used to calculate the installation error between the inertial group and the launcher. Step 102 has obtained a formula for calculating the installation error angle according to the attitude angle. Substituting the attitude angle data obtained through experiments in step 101 into the calculation formula of the installation error angle, the installation error can be obtained, thereby realizing the calibration of the inertial group attitude during the erection process of the launcher.

[0101] As an optional embodiment, a method for solving the relationship between the installation error between the inertial group and the launcher and the attitude angle includes:

[0102] Define n as the navigation coordinate system, b as the inertial group carrier coordinate system, and B as the launcher carrier coordinate system; the rotation transformation matrix from n to B is for:

[0103]

[0104] In the formula, θ and γ are the azimuth, pitch, and roll angles required to rotate from the n system to the B system, respectively;

[0105] The installation error between the inertial group and the launcher is set as the pitch angle installation error μ x , Roll angle installation error μ y 、Azimuth installation error μ z ;μ x , μ y , μ z Both are close to 0. After approximation with formula (1), the rotation transformation matrix from system b to system B is obtained:

[0106]

[0107] During the erection of the launcher, the roll angle γ is approximately 0. Substituting γ = 0 into equation (1) yields the simplified Then we get the rotation transformation matrix from B system to n system

[0108]

[0109] Based on matrix multiplication, the rotation transformation matrix from b system to n system is obtained

[0110]

[0111] In the formula, for Elements corresponding to rows and columns;

[0112] Set the azimuth, pitch, and roll angles required to rotate from the n system to the b system to be θ', γ', when the launcher pitch angle θ = 0, substitute θ = 0 into equation (4) and combine with equation (1) to obtain:

[0113]

[0114] Thus we get:

[0115] μ x = -θ' (7)

[0116] μ y = -γ'; (8)

[0117] When the launcher pitch angle θ is not 0, we get:

[0118]

[0119] From (9) and (10), we can get: z =γ' / tanθ';

[0120] Due to the error in the attitude output by the inertial group, the pitch angle and roll angle of the launcher after erection are set to θ1′ and γ1′, and the pitch angle and roll angle of the launcher when the pitch angle is 0 are set to θ0′ and γ0′, μ z The calculation formula is:

[0121]

[0122] The installation error between the inertial group and the launch pad can be calculated according to equations (7), (8) and (11).

[0123] This embodiment provides a technical solution for solving the relationship between the installation error and the attitude angle output by the inertial group at different pitch angles during the erection of the launcher. First, three coordinate systems are defined: the n system is defined as the navigation coordinate system, the b system is the inertial group carrier coordinate system, and the B system is the launcher carrier coordinate system, and the rotation transformation matrix from the n system to the B system is listed. As shown in formula (1). Formula (1) is actually the expression of the rotation transformation matrix between any two coordinate systems, except that the three angles θ, γ, Corresponding to the two corresponding coordinate systems, formula (1) is used many times in this embodiment. Then the installation error between the inertial group and the launcher is set as the pitch angle installation error μ x , Roll angle installation error μ y 、Azimuth installation error μ z . Use μ x , μ y , μ z Replace θ, γ, The rotation transformation matrix from b system to B system can be obtained Because μ x , μy , μ z are all small angles, that is, close to 0. After approximation, the simplified expression shown in formula (2) is obtained. The approximation method includes: x , μ y , μ z The cosine of is approximately 1, and will contain more than two μ x , μ y , μ z The product of the sine factors is approximately 0, such as sinμ y sinμ z = 0. Since the roll angle γ during the erection of the launcher is approximately 0, substituting γ = 0 into equation (1) yields the simplified Find the transpose to get the rotation transformation matrix from B system to n system As shown in formula (3). Get the rotation transformation matrix from b system to n system As shown in formula (4). In order to avoid the matrix in formula (4) being too long, the expressions of some unused elements in the matrix are expressed by the row and column numbers of the elements, such as express Next, set the azimuth, pitch, and roll angles required to rotate from the b system to the n system to be θ', γ', combined with formula (1) (using θ' and γ' are replaced by θ, γ) to obtain The expression is:

[0124]

[0125] From the above formula, we can know that the output azimuth of the inertial group is affected by the installation error μ x , μ z and the azimuth of the launcher and the influence of the pitch angle θ.

[0126] Then, by substituting θ=0 into equation (4), we can solve the launcher pitch angle θ=0 Combining (1), we can get the expressions of θ' and γ' as (5) and (6). According to (5) and (6), we can get μ x , μ y The expressions of θ' and γ' are as follows (7) and (8). Then, according to equation (4), we can get the expressions of θ' and γ' when θ is not 0 by approximation, as shown in (9) and (10), thus obtaining μ z =γ' / tanθ'. Due to the error in the output attitude of the inertial group, μ z=γ' / tanθ' The calculation error is large. Therefore, this embodiment eliminates the calculation error to a certain extent by calculating the angle difference, that is, using the difference θ1' and γ1' after the launcher is erected at a certain angle and θ0' and γ0' when the launcher pitch angle is 0, θ1'-θ0', γ1'-γ0' to solve μ z , as shown in formula (11). Formula (11) is μ z =γ' / tanθ' is obtained by approximate processing. Thus, the installation error μ is obtained x , μ y , μ z The calculation formulas (7), (8) and (11) are used.

[0127] As an optional embodiment, the method of obtaining attitude angle data through experiments includes:

[0128] S1. After the launch vehicle stops at the test site, it is turned off and stationary, and the vehicle body is leveled so that the pitch angle of the launcher is 0;

[0129] S2, the IMU starts to seek north and obtains the pitch angle θ' and the roll angle γ';

[0130] S3, inertial group shutdown;

[0131] S4, the inertial group starts to seek north, and obtains the pitch angle θ0′ and the roll angle γ0′; the launcher is erected to a pitch angle of α, and the pitch angle θ1′ and the roll angle γ1′ are obtained;

[0132] S5, inertial group shutdown.

[0133] This embodiment provides a technical solution for obtaining attitude angle data through experiments. This embodiment obtains attitude angle data by executing steps S1 to S5. Step S1 is the experimental preparation. After the launch vehicle is parked at the test site, it is turned off and stationary; then the vehicle body is leveled to make the pitch angle of the launcher 0. Step S2 is used to obtain the calculation μ x , μ y The required attitude angle data is obtained by turning on the inertial group to find the north, and obtaining the pitch angle θ' and roll angle γ' output by the inertial group when the pitch angle of the launcher is 0. Step S3 is to turn off the inertial group after obtaining a set of experimental data. S4 is used to obtain the calculation μ z The required attitude angle data. The same is true when the inertial group is turned on to find north. First, the pitch angle θ0′ and the roll angle γ0′ of the launcher are measured when the pitch angle is 0. Then the launcher is erected to a pitch angle of α (such as 60°) to obtain the pitch angle θ′1 and the roll angle γ′1. Step S5 is to shut down the inertial group after obtaining another set of experimental data.

[0134] As an optional embodiment, steps S2 and S4 are repeated to measure N groups of data: θ′ i , γ′ i are θ′ and γ′ measured for the i-th time respectively; θ′0i , γ′ 0i are θ′0, γ′0, θ′ measured for the i-th time respectively. 1i , γ′ 1i are θ′ measured at the i-th time 1i , γ′ 1i , i=1,2,......,N.

[0135] In order to improve the accuracy of attitude calibration, this embodiment performs multiple (N) repeated measurements on each attitude angle, such as N = 5, and then calculates the installation error angle μ of the inertial group relative to the launcher by taking the average value. x , μ y , μ z .

[0136] As an optional embodiment, the method for calculating the installation error based on the relationship and the attitude angle data includes:

[0137] Calculate the pitch angle installation error and roll angle installation error as follows:

[0138]

[0139] In the formula, μ x , μ y They are the pitch angle installation error and the roll angle installation error respectively;

[0140] Calculate the azimuth installation error as follows:

[0141]

[0142] In the formula, μ z is the azimuth installation error.

[0143] This embodiment provides a specific formula for calculating the installation error based on experimental data. x , μ y The calculation formulas (12) and (13) of are obtained based on formulas (7) and (8), that is, the average of the calculation results obtained by formulas (7) and (8). Similarly, μ z The calculation formula (14) is obtained based on formula (11), that is, the average of the calculation results obtained by formula (11) is obtained.

[0144] Figure 4 The present invention is a schematic diagram of a device for calibrating the attitude of a laser inertial group during the erection of a launcher according to an embodiment of the present invention, wherein the device comprises:

[0145] The data acquisition module 11 is used to acquire the attitude angle data obtained through experiments, including the attitude angle output by the inertial group at different pitch angles during the launcher erection process;

[0146] A relationship solving module 12 is used to solve the relationship between the installation error between the inertial group and the launcher and the attitude angle based on the rotation transformation matrix between the navigation coordinate system, the inertial group carrier coordinate system and the launcher carrier coordinate system;

[0147] The inertial group calibration module 13 is used to calculate the installation error based on the relationship and the attitude angle data, so as to realize the calibration of the inertial group attitude during the erection process of the launcher.

[0148] The device of this embodiment can be used to perform Figure 1 The technical solution of the method embodiment shown has similar implementation principles and technical effects, which will not be described in detail here. The same is true for the following embodiments, which will not be described in detail.

[0149] As an optional embodiment, a method for solving the relationship between the installation error between the inertial group and the launcher and the attitude angle includes:

[0150] Define n as the navigation coordinate system, b as the inertial group carrier coordinate system, and B as the launcher carrier coordinate system; the rotation transformation matrix from n to B is for:

[0151]

[0152] In the formula, θ and γ are the azimuth, pitch, and roll angles required to rotate from the n system to the B system, respectively;

[0153] The installation error between the inertial group and the launcher is set as the pitch angle installation error μ x , Roll angle installation error μ y 、Azimuth installation error μ z ;μ x , μ y , μ z Both are close to 0. After approximation with formula (1), the rotation transformation matrix from system b to system B is obtained:

[0154]

[0155] During the erection of the launcher, the roll angle γ is approximately 0. Substituting γ = 0 into equation (1) yields the simplified Then we get the rotation transformation matrix from B system to n system

[0156]

[0157] Based on matrix multiplication, the rotation transformation matrix from b system to n system is obtained

[0158]

[0159] In the formula, for Elements corresponding to rows and columns;

[0160] Set the azimuth, pitch, and roll angles required to rotate from the n system to the b system to be θ', γ', when the launcher pitch angle θ = 0, substitute θ = 0 into equation (4) and combine with equation (1) to obtain:

[0161]

[0162] Thus we get:

[0163] μ x = -θ' (7)

[0164] μ y = -γ'; (8)

[0165] When the launcher pitch angle θ is not 0, we get:

[0166]

[0167] From (9) and (10), we can get: z =γ' / tanθ';

[0168] Due to the error in the attitude output by the inertial group, the pitch angle and roll angle of the launcher after erection are set to θ′1 and γ′1, and the pitch angle and roll angle of the launcher when the pitch angle is 0 are set to θ′0 and γ′0, μ z The calculation formula is:

[0169]

[0170] The installation error between the inertial group and the launch pad can be calculated according to equations (7), (8) and (11).

[0171] As an optional embodiment, the method of obtaining attitude angle data through experiments includes:

[0172] S1. After the launch vehicle stops at the test site, it is turned off and stationary, and the vehicle body is leveled so that the pitch angle of the launcher is 0;

[0173] S2, the IMU starts to seek north and obtains the pitch angle θ' and the roll angle γ';

[0174] S3, inertial group shutdown;

[0175] S4, the inertial group starts to seek north, and obtains the pitch angle θ0′ and the roll angle γ0′; the launcher is erected to a pitch angle of α, and the pitch angle θ1′ and the roll angle γ1′ are obtained;

[0176] S5, inertial group shutdown.

[0177] As an optional embodiment, steps S2 and S4 are repeated to measure N groups of data: θ i ′、γ i ′ are θ′ and γ′ measured for the i-th time; θ′ 0i , γ′ 0i are θ′0, γ′0, θ′ measured for the i-th time respectively. 1i , γ′ 1i are θ′ measured at the i-th time 1i , γ′ 1i , i=1,2,......,N.

[0178] As an optional embodiment, the method for calculating the installation error based on the relationship and the attitude angle data includes:

[0179] Calculate the pitch angle installation error and roll angle installation error as follows:

[0180]

[0181] In the formula, μ x , μ y They are the pitch angle installation error and the roll angle installation error respectively;

[0182] Calculate the azimuth installation error as follows:

[0183]

[0184] In the formula, μ z is the azimuth installation error.

[0185] The above is only a specific embodiment of the present invention, but the protection scope of the present invention is not limited thereto. Any changes or substitutions that can be easily thought of by a person skilled in the art within the technical scope disclosed by the present invention should be included in the protection scope of the present invention. Therefore, the protection scope of the present invention shall be subject to the protection scope of the claims.

Claims

1. A method for calibrating the attitude of a laser inertial group during the erection of a launcher, characterized in that: The following steps are involved: Acquire the attitude angle data obtained through experiments, including the attitude angle output by the inertial group at different pitch angles during the launcher erection process; Based on the rotation transformation matrix between the navigation coordinate system, the IMU carrier coordinate system and the launcher carrier coordinate system, solving the relationship between the installation error between the IMU and the launcher and the attitude angle; The installation error is calculated based on the relationship and the attitude angle data to achieve calibration of the inertial group attitude during the erection of the launcher.

2. The method for calibrating the attitude of the laser inertial group during the erection of the launcher according to claim 1, characterized in that: The method for solving the relationship between the installation error between the inertial group and the launcher and the attitude angle includes: Define n as the navigation coordinate system, b as the inertial group carrier coordinate system, and B as the launcher carrier coordinate system; the rotation transformation matrix from n to B is for: In the formula, θ and γ are the azimuth, pitch, and roll angles required to rotate from the n system to the B system, respectively; The installation error between the inertial group and the launcher is set as the pitch angle installation error μ x , Roll angle installation error μ y 、Azimuth installation error μ z ;μ x , μ y , μ z Both are close to 0. After approximation with formula (1), the rotation transformation matrix from system b to system B is obtained: During the erection of the launcher, the roll angle γ is approximately 0. Substituting γ = 0 into equation (1) yields the simplified Then we get the rotation transformation matrix from B system to n system Based on matrix multiplication, the rotation transformation matrix from b system to n system is obtained In the formula, for Elements corresponding to rows and columns; Set the azimuth, pitch, and roll angles required to rotate from the n system to the b system to be θ', γ', when the launcher pitch angle θ = 0, substitute θ = 0 into equation (4) and combine with equation (1) to obtain: Thus we get: m x =-θ' (7) m y =-γ'; (8) When the launcher pitch angle θ is not 0, we get: From (9) and (10), we can get: z =γ' / tanθ'; Due to the error in the attitude output by the inertial group, the pitch angle and roll angle of the launcher after erection are set to θ1′ and γ1′, and the pitch angle and roll angle of the launcher when the pitch angle is 0 are set to θ0′ and γ0′, μ z The calculation formula is: The installation error between the inertial group and the launch pad can be calculated according to equations (7), (8) and (11).

3. The method for calibrating the attitude of the laser inertial group during the erection of the launcher according to claim 1, characterized in that: Methods for obtaining attitude angle data through experiments include: S1. After the launch vehicle stops at the test site, it is turned off and stationary, and the vehicle body is leveled so that the pitch angle of the launcher is 0; S2, the IMU starts to seek north and obtains the pitch angle θ' and the roll angle γ'; S3, inertial group shutdown; S4, the inertial group starts to seek north, and obtains the pitch angle θ0′ and the roll angle γ0′; the launcher is erected to a pitch angle of α, and the pitch angle θ1′ and the roll angle γ1′ are obtained; S5, inertial group shutdown.

4. The method for calibrating the attitude of the laser inertial group during the erection of the launcher according to claim 3 is characterized in that: Steps S2 and S4 are repeated to measure N sets of data: θ i ′、γ i ′ are θ' and γ' measured for the i-th time; θ' 0i , γ′ 0i are θ′0, γ′0, θ′ measured for the i-th time respectively. 1i , γ′ 1i are θ′ measured at the i-th time 1i , γ′ 1i , i=1,2,......,N.

5. The method for calibrating the attitude of the laser inertial group during the erection of the launcher according to claim 4, characterized in that: The method for calculating the installation error based on the relationship and the attitude angle data includes: Calculate the pitch angle installation error and roll angle installation error as follows: In the formula, μ x , μ y They are the pitch angle installation error and the roll angle installation error respectively; Calculate the azimuth installation error as follows: In the formula, μ z is the azimuth installation error.

6. A device for calibrating the attitude of a laser inertial group during the erection of a launcher, characterized in that: include: A data acquisition module is used to obtain the attitude angle data obtained through experiments, including the attitude angle output by the inertial group at different pitch angles during the erection process of the launcher; A relationship solving module, used for solving the relationship between the installation error between the inertial group and the launcher and the attitude angle based on the rotation transformation matrix between the navigation coordinate system, the inertial group carrier coordinate system and the launcher carrier coordinate system; The inertial group calibration module is used to calculate the installation error based on the relationship and the attitude angle data, so as to realize the calibration of the inertial group attitude during the erection process of the launcher.

7. The device for calibrating the attitude of the laser inertial group during the erection of the launcher according to claim 6, characterized in that: The method for solving the relationship between the installation error between the inertial group and the launcher and the attitude angle includes: Define n as the navigation coordinate system, b as the inertial group carrier coordinate system, and B as the launcher carrier coordinate system; the rotation transformation matrix from n to B is for: In the formula, θ and γ are the azimuth, pitch, and roll angles required to rotate from the n system to the B system, respectively; The installation error between the inertial group and the launcher is set as the pitch angle installation error μ x , Roll angle installation error μ y 、Azimuth installation error μ z ;μ x , μ y , μ z Both are close to 0. After approximation with formula (1), the rotation transformation matrix from system b to system B is obtained: During the erection of the launcher, the roll angle γ is approximately 0. Substituting γ = 0 into equation (1) yields the simplified Then we get the rotation transformation matrix from B system to n system Based on matrix multiplication, the rotation transformation matrix from b system to n system is obtained In the formula, for Elements corresponding to rows and columns; Set the azimuth, pitch, and roll angles required to rotate from the n system to the b system to be θ', γ', when the launcher pitch angle θ = 0, substitute θ = 0 into equation (4) and combine with equation (1) to obtain: Thus we get: m x =-θ' (7) m y =-γ'; (8) When the launcher pitch angle θ is not 0, we get: From (9) and (10), we can get: z =γ' / tanθ'; Due to the error in the attitude output by the inertial group, the pitch angle and roll angle of the launcher after erection are set to θ1′ and γ1′, and the pitch angle and roll angle of the launcher when the pitch angle is 0 are set to θ0′ and γ0′, μ z The calculation formula is: The installation error between the inertial group and the launch pad can be calculated according to equations (7), (8) and (11).

8. The device for calibrating the attitude of the laser inertial group during the erection of the launcher according to claim 6, characterized in that: Methods for obtaining attitude angle data through experiments include: S1. After the launch vehicle stops at the test site, it is turned off and stationary, and the vehicle body is leveled so that the pitch angle of the launcher is 0; S2, the IMU starts to seek north and obtains the pitch angle θ' and the roll angle γ'; S3, inertial group shutdown; S4, the inertial group starts to seek north, and obtains the pitch angle θ0′ and the roll angle γ0′; the launcher is erected to a pitch angle of α, and the pitch angle θ1′ and the roll angle γ1′ are obtained; S5, inertial group shutdown.

9. The device for calibrating the attitude of the laser inertial group during the erection of the launcher according to claim 8, characterized in that: Steps S2 and S4 are repeated to measure N sets of data: θ i ′、γ i ′ are θ' and γ' measured for the i-th time; θ0′ i ,γ0′ i are θ0′, γ0′, θ1′ measured for the i-th time respectively i ,γ1′ i are θ1′ measured at the i-th time i ,γ1′ i , i=1,2,......,N.

10. The device for calibrating the attitude of the laser inertial group during the erection of the launcher according to claim 9, characterized in that: The method for calculating the installation error based on the relationship and the attitude angle data includes: Calculate the pitch angle installation error and roll angle installation error as follows: In the formula, μ x , μ y They are the pitch angle installation error and the roll angle installation error respectively; Calculate the azimuth installation error as follows: In the formula, μ z is the azimuth installation error.

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