A method of horizontal aiming and attitude correction in parallel with erecting

Abstract

This invention proposes a method for horizontal aiming and attitude correction in parallel with aircraft erection, belonging to the field of aircraft navigation and positioning technology. The method includes the following steps: S1. When the aircraft is horizontal, the onboard inertial navigation system (INS) performs coarse alignment; simultaneously, the pitch, roll, and tilt angles of the INS are calculated using the output of the accelerometer; S2. At the end of the coarse alignment, the photoelectric collimator prism is collimated and the measurement results are output; the coarse azimuth A of the INS is calculated. c Initialization vector A 0 S3. After establishing the attitude matrix using the aircraft's inertial navigation system's coarse azimuth, pitch, and roll data, immediately switch to attitude tracking, while simultaneously updating vector A in real time. i S4. After the aircraft's inertial navigation system (INS) enters attitude tracking, the vehicle-mounted INS fine alignment and aircraft erection are carried out in parallel. S5. After erection, the vehicle-mounted INS fine alignment is completed before the aircraft ignition, and the attitude correction of the aircraft's INS is performed. This invention solves the problem of the long total time for aiming and erection during the ground preparation phase of vehicle-mounted aircraft in the prior art.

Description

Technical Field

[0001] This invention belongs to the field of aircraft navigation and positioning technology, specifically relating to a method for horizontal aiming and attitude correction in parallel with take-off and erection, which is particularly suitable for vehicle-mounted aircraft. Background Technology

[0002] Many aircraft, especially vehicle-mounted aircraft, require reduced ground preparation time. For example, when adding new satellites to a satellite constellation urgently needs to be done, a launch vehicle should be launched as quickly as possible. Initial alignment (i.e., aiming) is a necessary step in pre-launch preparation and typically takes several minutes. To shorten ground preparation time, the initial alignment time for some aircraft is forced to be reduced, which introduces significant azimuth errors. Auxiliary equipment (such as global navigation satellite systems, star sensors) or other auxiliary information can be used to correct the flight navigation errors caused by these azimuth errors. However, these auxiliary devices increase the complexity and cost of the aircraft. Furthermore, global navigation satellite systems are unavailable in some situations. Therefore, reducing ground preparation time without compromising the azimuth and attitude accuracy of the aircraft's inertial navigation system is a challenge.

[0003] There are many methods for initial alignment of vehicle-mounted aircraft. These methods can be broadly categorized into three types: moving base alignment, vertical aiming / self-alignment, and horizontal aiming / self-alignment. Both moving base-based alignment and moving base self-alignment require vehicle maneuvering. Furthermore, changes in the mounting angles between the aircraft's inertial navigation system (INS) and onboard equipment such as the odometer and laser Doppler velocimeter during vehicle maneuvering reduce alignment accuracy.

[0004] Vertical aiming methods are widely used due to their high accuracy. However, the deployment, erection, and dismantling of aiming equipment such as electro-optical theodolites require significant time. While vertical self-alignment of the inertial navigation system (INS) avoids manual operation, it still requires several minutes of self-alignment after the aircraft has been erected. Thanks to its advantages of automated operation and rapid implementation, horizontal aiming or self-alignment methods have become increasingly popular in recent years. In traditional horizontal aiming methods, the vehicle-mounted INS performs self-alignment, using the self-alignment results and the output of the electro-optical collimator to calculate the azimuth of the aircraft's INS. The INS then enters attitude tracking and performs erection. If a horizontal self-alignment method is used, this self-alignment also needs to be performed before the aircraft is erected.

[0005] In summary, both vertical aiming / self-aligning methods and horizontal aiming / self-aligning methods in the existing technology have the drawback that the aircraft erection cannot be carried out in parallel with aiming, which consumes valuable ground preparation time and needs to be improved. Summary of the Invention

[0006] This invention provides a method for horizontal aiming and attitude correction that is performed in parallel with the erection process, in order to solve the problem in the prior art that the erection of vehicle-mounted aircraft cannot be carried out in parallel with aiming, and that the total time for aiming and erection during the ground preparation phase of the aircraft is long.

[0007] The objective of this invention is achieved through the following technical solution:

[0008] A method for horizontal aiming and attitude correction in parallel with erection includes the following steps:

[0009] S1. When the aircraft is horizontal, the vehicle-mounted inertial navigation system performs coarse alignment; at the same time, the pitch, roll and tilt angles of the aircraft's inertial navigation system are calculated using the output of the accelerometer.

[0010] S2. At the end of the coarse alignment, the photoelectric collimator prism is aligned and the measurement results are output. Using the coarse alignment azimuth, pitch and roll of the vehicle-mounted inertial navigation system (INS), the collimator measurement results, the tilt angle of the aircraft INS, and the pre-calibrated parameters, the coarse azimuth A of the aircraft INS is calculated. c Initialization vector A 0 Vectors D(0) and B(0) are initialized as needed;

[0011] A 0 It can be obtained through the following formula:

[0012] A 0 =I 3×3

[0013] I 3×3 It is a 3×3 identity matrix;

[0014] S3. After establishing the attitude matrix using the aircraft's inertial navigation system (INS) with approximate azimuth, pitch, and roll, the INS immediately switches to attitude tracking, while simultaneously updating vector A in real time. i Whether vectors D(T) and B(T) are updated in real time depends on the requirements;

[0015] Where B(T) is used to update D(T+Δt) at the next time step;

[0016] T = iΔt;

[0017] Δt is the sampling interval of the aircraft's inertial navigation system;

[0018] S4. After the aircraft inertial navigation system switches to attitude tracking, the vehicle-mounted inertial navigation system fine alignment and aircraft erection are carried out in parallel.

[0019] S5. After the erection is completed, the vehicle-mounted inertial navigation system is precisely aligned and the attitude of the inertial navigation system is corrected before the aircraft is ignited.

[0020] As a further optimization, the attitude correction steps for the aircraft's inertial navigation system are performed as follows:

[0021] S5.1 After the erection is completed, the vehicle-mounted inertial navigation system continues to perform precise alignment until the aircraft ignites;

[0022] The initial inertial navigation system (INS) azimuth correction is calculated using the following formula:

[0023] φ U (0) = A c -A r

[0024] In the formula, A r The precise azimuth of the aircraft's inertial navigation system at the end of the coarse alignment is calculated using the azimuth, pitch, and roll of the vehicle-mounted inertial navigation system during reverse fine alignment, as well as the measurement results of the collimator, the tilt angle of the aircraft's inertial navigation system, and the pre-calibrated parameters.

[0025] Let φ E (0) and φ N (0) is 0, and the vector φ(0) can be written as: φ(0)=[0 0 φ U (0)] T ;

[0026] S5.2, φ(0) obtained through the above steps, and the real-time updated vectors D(T) and A i The vector φ(T) is estimated, and the attitude matrix is ​​corrected according to the following formula.

[0027]

[0028] in, Calculated using the following formula:

[0029]

[0030] pass Calculate the azimuth, pitch, and roll angles of the aircraft's inertial navigation system at the current moment, and complete the attitude correction.

[0031] As a further optimization, in the horizontal aiming and attitude correction method that runs parallel to the erection, it is necessary to estimate and compensate for the attitude tracking error caused by the gyroscope zero bias. The gyroscope zero bias vector ε b Estimate using the following formula:

[0032]

[0033] in, The average gyroscope output of the aircraft's inertial navigation system before erection;

[0034] The initial attitude matrix is ​​calculated using the initial azimuth, pitch, and roll of the aircraft's inertial navigation system.

[0035] This is the vector of Earth's rotational angular velocity;

[0036] ω ie L represents the Earth's rotational angular velocity, and L represents the latitude.

[0037] As a further optimization, the initialization vectors D(0) and B(0) are obtained by the following formula:

[0038] D(0)=0 3×3

[0039]

[0040] Among them, 0 3×3 It is a 3×3 zero matrix;

[0041] The initial attitude matrix is ​​calculated using the initial azimuth, pitch, and roll of the aircraft's inertial navigation system.

[0042] As a further optimization, after the aircraft's inertial navigation system switches to attitude tracking, D(T) is updated in real time using the following formula:

[0043] D(T)=AD(T-Δt)+B(T-Δt)

[0044] B(T) is updated in real time using the following formula:

[0045] As a further optimization, the vector φ(T) is estimated using the following formula:

[0046] φ(T)=A i φ(0)+D(T)ε b

[0047] A i Updated in real time using the following formula:

[0048] A i =A·A i-1

[0049] Among them, A i-1 Updated at the previous moment.

[0050] As a further optimization, the horizontal aiming and attitude correction method, which operates in parallel with the erection process, can be optimized when there is no need to estimate and compensate for the attitude tracking error caused by the gyroscope zero bias vector ε. b Since the vector is zero, the average gyroscope output of the aircraft's inertial navigation system does not need to be saved during the coarse alignment of the vehicle-mounted inertial navigation system, and there is no need to initialize or update B(T) and D(T); the vector φ(T) is estimated by the following formula:

[0051] φ(T)=A i φ(0)

[0052] A i Updated in real time using the following formula: A i =A·A i-1

[0053] Among them, A i-1 Updated at the previous moment.

[0054] Based on the same technical concept, this invention also proposes a horizontal aiming and attitude correction method that runs parallel to erection, comprising the following steps:

[0055] S1. When the aircraft is horizontal, the vehicle-mounted inertial navigation system performs coarse alignment; at the same time, the pitch, roll and tilt angles of the aircraft's inertial navigation system are calculated using the output of the accelerometer.

[0056] S2. At the end of the coarse alignment, the photoelectric collimator prism is aligned and the measurement results are output. Using the coarse alignment azimuth, pitch and roll of the vehicle-mounted inertial navigation system (INS), the collimator measurement results, the tilt angle of the aircraft INS, and the pre-calibrated parameters, the coarse azimuth A of the aircraft INS is calculated. c ;

[0057] S3. After establishing the attitude matrix through the aircraft's inertial navigation system's rough azimuth, pitch, and roll, the aircraft's inertial navigation system immediately switches to attitude tracking.

[0058] S4. After the aircraft inertial navigation system switches to attitude tracking, the vehicle-mounted inertial navigation system fine alignment and aircraft erection are carried out in parallel.

[0059] S5. After the erection is completed, while the vehicle-mounted inertial group continues to perform fine alignment, it uses the acceleration output for an additional tens of seconds to calculate the initial attitude value, replaces the inaccurate current attitude of the aircraft inertial group, and completes the current orientation and attitude correction of the aircraft inertial group.

[0060] During the initial attitude calculation, the initial inertial navigation system azimuth correction φ U (0) is calculated using the following formula;

[0061] φ U (0)=A′ c -A′ r

[0062] In the formula, A′ c The position is the position from the coarse alignment of the vehicle-mounted inertial navigation system to the end of the fine alignment.

[0063] A′ r To ensure the precise alignment of the vehicle-mounted inertial navigation system at the end of the alignment process;

[0064] Once the current orientation and attitude of the aircraft's inertial navigation system are corrected, no further fine alignment of the onboard inertial navigation system will be performed before the aircraft ignites.

[0065] Based on the same technical concept, this invention also proposes a horizontal aiming and attitude correction method that runs parallel to erection, comprising the following steps:

[0066] S1. When the aircraft is horizontal, the vehicle-mounted inertial navigation system performs self-alignment. At the same time, the output of the accelerometer is used to calculate the pitch, roll and tilt angles of the aircraft's inertial navigation system. The photoelectric collimator straightens the collimator prism and outputs the measurement results.

[0067] S2. After the pitch, roll and tilt calculations of the aircraft inertial navigation system are completed, the original data of the aircraft inertial navigation system is saved. The vehicle-mounted inertial navigation system continues to self-align, and the aircraft begins to stand up.

[0068] S3. After erection, the initial orientation of the aircraft's inertial navigation system is calculated using the vehicle-mounted inertial navigation system self-alignment results, the tilt angle of the aircraft's inertial navigation system in horizontal state, the measurement results of the electro-optical collimator, and the pre-calibrated parameters.

[0069] S4. Establish the initial attitude matrix using the initial azimuth, pitch, and roll of the aircraft inertial navigation system (INS); then use the saved raw data of the aircraft INS to perform attitude tracking and calculate the final azimuth and attitude of the aircraft INS.

[0070] The beneficial technical effects achieved by this invention are:

[0071] By performing horizontal aiming and attitude correction simultaneously with the erection, the technical solution proposed in this invention, compared with the traditional horizontal aiming method, can significantly shorten the total time of aiming and erection without reducing the final orientation and attitude accuracy of the inertial group during the ground preparation phase of the aircraft. It solves the problem of the long total time of inertial group aiming and erection during the ground preparation phase of the aircraft due to the inability to perform the erection of the vehicle-mounted aircraft in parallel with aiming. It has outstanding substantive features and significant progress. Attached Figure Description

[0072] Figure 1 This is a schematic diagram of the layout of a vehicle-mounted directional aiming device according to one specific embodiment of the present invention;

[0073] Figure 2 This is a flowchart of one specific embodiment of the present invention. Detailed Implementation

[0074] The technical solution of the present invention will be further described in detail below with reference to the accompanying drawings and specific embodiments. Obviously, the described embodiments are only a part of the embodiments of the present invention, and not all of them. All other embodiments obtained by those skilled in the art based on the embodiments of the present invention without creative effort are within the scope of protection claimed by the present invention.

[0075] A method for horizontal aiming and attitude correction in parallel with erection employs a real-time error estimation model. The real-time error estimation model for azimuth and attitude during inertial navigation system attitude tracking established in this invention is as follows:

[0076] When the strapdown inertial navigation system is stationary or moving nearby, the attitude error equation of the strapdown inertial navigation system can be expressed as:

[0077]

[0078] Where, φ=[φ E φ N φ U ] T The attitude error of the navigation coordinate system is calculated relative to the navigation coordinate system.

[0079] φ U This refers to the azimuth error;

[0080] φ E ,φ N This refers to attitude error;

[0081] This is the vector of Earth's rotational angular velocity;

[0082] ω ie ω represents the Earth's rotational angular velocity, and L represents the latitude.

[0083] ε n =[ε E ε N ε U ] T This is the zero bias vector of the gyroscope in the navigation coordinate system.

[0084] Formula (1) can also be written as:

[0085]

[0086] in, The attitude matrix;

[0087] ε b =[ε x ε y ε z ] T This is the zero-bias vector of the gyroscope in the carrier coordinate system;

[0088]

[0089] Formula (2) can be expressed as the following linear discrete equation

[0090]

[0091] Where, φ(t+Δt)=[φE (t+Δt) φ N (t+Δt) φ U (t+Δt)] T The attitude error of the navigation coordinate system relative to the calculated navigation coordinate system at time t+Δt;

[0092] φ(t)=[φ E (t) φ N (t) φ U (t)] T Let t be the attitude error of the navigation coordinate system relative to the calculated navigation coordinate system;

[0093] Let be the attitude matrix at time t;

[0094] I 3×3 It is a 3×3 identity matrix;

[0095] Δt is the sampling interval of the aircraft's inertial navigation system.

[0096] definition:

[0097]

[0098] as well as:

[0099]

[0100] Then formula (4) can be expressed as:

[0101] φ(t+Δt)=Aφ(t)+B(t)ε b (7)

[0102] The following formula is further obtained:

[0103] φ(T)=A i φ(t)+D(T)ε b (8)

[0104] in:

[0105] T=t+iΔt (9)

[0106] In formula (8):

[0107] A i =A·A i-1 (10)

[0108] Where A i-1 Updated at the previous moment.

[0109] In formula (8), matrix D(T) can be written as:

[0110] D(T)=AD(T-Δt)+B(T-Δt) (11)

[0111] in:

[0112] D(t=0) 3×3 (12)

[0113] When t = 0, formula (8) can be written as:

[0114] φ(T)=A i φ(0)+D(T)ε b (13)

[0115] in:

[0116] T=iΔt (14)

[0117] At this point, formulas (12), (6), and (10) are respectively:

[0118] D(0)=0 3×3 (15)

[0119]

[0120] A 0 =I 3×3 (17)

[0121] in The initial attitude matrix can be calculated from the initial azimuth, pitch, and roll of the aircraft's inertial navigation system.

[0122] 0 3×3 It is a 3×3 zero matrix.

[0123] Equation (13) is the real-time error estimation model. If there is no need to estimate and compensate for the attitude tracking error caused by the gyroscope's zero bias, Equation (13) can be simplified to:

[0124] φ(T)=A i φ(0) (18)

[0125] like Figure 1 As shown, a specific embodiment of a horizontal aiming and attitude correction method parallel to erection is implemented using the aforementioned real-time error estimation model. A prism in the vehicle-mounted aircraft is fixedly connected to the aircraft's inertial navigation system (INS), which is entirely installed inside the aircraft. A glass panel is installed on the shell opposite the prism. When the aircraft is horizontal, an optoelectronic collimator fixed to the vehicle-mounted INS collimates the prism, and the parallel light emitted from the collimator passes through the glass. The installation error angle between the prism and the aircraft INS, as well as the three-dimensional installation error angle between the collimator and the vehicle-mounted INS, are pre-calibrated. Figure 2 As shown, in this specific embodiment, the horizontal aiming and attitude correction method, which runs parallel to erection, is performed according to the following steps:

[0126] S1. When the aircraft is horizontal, the vehicle-mounted inertial navigation system performs coarse alignment; at the same time, the pitch, roll and tilt angles of the aircraft's inertial navigation system are calculated using the output of the accelerometer.

[0127] During the coarse alignment of the vehicle-mounted inertial navigation system, the average gyroscope output of the aircraft's inertial navigation system is saved (if there is no need to estimate and compensate for the attitude tracking error caused by the gyroscope zero bias, then there is no need to save the average gyroscope output).

[0128] S2. At the end of the coarse alignment, the photoelectric collimator straightens the collimator prism and outputs the measurement results.

[0129] Using the coarse alignment azimuth, pitch and roll measurements from the vehicle-mounted inertial navigation system (INS), the collimator measurements, the aircraft INS tilt angle, and pre-calibrated parameters, the coarse azimuth A of the aircraft INS can be calculated. c .

[0130] Initialize vectors D(0), B(0), and A using formulas (15)-(17). 0 .

[0131] S3. After establishing the attitude matrix through the azimuth, pitch and roll of the aircraft inertial navigation system, the aircraft inertial navigation system immediately switches to attitude tracking and updates D(T) in real time through formula (11).

[0132] After the aircraft inertial navigation system switches to attitude tracking, B(T) is updated in real time using formula (6). B(T) will be used to update D(T+Δt) at the next moment.

[0133] A i The formula (10) is used for real-time updates.

[0134] S4. After the aircraft inertial navigation system switches to attitude tracking, the vehicle-mounted inertial navigation system fine alignment and aircraft erection are carried out in parallel.

[0135] S5. After the erection is completed, the vehicle-mounted inertial navigation system continues to perform precise alignment until the aircraft ignites.

[0136] S5.1, Precise bearing of the aircraft's inertial navigation system at the end of coarse alignment A r It can be calculated from the azimuth, pitch, and roll of the vehicle-mounted inertial navigation system through reverse precision alignment, as well as some previously obtained parameters (the collimator's measurement results, the aircraft's inertial navigation system's tilt angle, and pre-calibrated parameters).

[0137] φ is calculated using the following formula. U (0):

[0138] φ U (0) = A c -A r (19)

[0139] φU (0) can also be obtained through φ U (0)=A′ c -A′ r We obtain, where A′ c A′ represents the azimuth of the vehicle-mounted inertial navigation system from the point of coarse alignment to the point of final fine alignment. r To ensure the precise alignment of the vehicle-mounted inertial navigation system at the end of the alignment process;

[0140] Considering φ E (0) and φ N (0) is unpredictable, so we set it to 0, and then the vector φ(0) can be written as:

[0141] φ(0)=[0 0 φ U (0)] T (20)

[0142] gyroscope zero bias vector ε b The following formula can be used to estimate: (If there is no need to estimate and compensate for the attitude tracking error caused by the gyroscope zero bias, then there is no need to calculate the gyroscope zero bias vector ε.) b )

[0143]

[0144] in, The average gyroscope output of the aircraft's inertial navigation system before erection;

[0145] The initial attitude matrix can be obtained through azimuth A. r The pitch and roll are calculated.

[0146] S5.2, φ(0) and ε obtained above b And the vectors D(T) and A updated in real time i φ(T) can be estimated using formula (13) or (18). Then the attitude matrix is ​​corrected using the following formula.

[0147]

[0148] in, Calculated using the following formula

[0149]

[0150] pass Calculate the azimuth, pitch, and roll angles of the aircraft's inertial navigation system at the current moment, and complete the attitude correction.

[0151] When the aircraft is level, the onboard inertial navigation system (INS) performs coarse alignment. Simultaneously, the pitch, roll, and tilt angles of the INS are calculated using the accelerometer output. At the end of the coarse alignment, the electro-optical collimator prism is aligned and the measurement results are output. Using the coarse alignment azimuth, pitch, and roll of the onboard INS, the collimator measurement results, the tilt angle of the INS, and pre-calibrated parameters, the coarse azimuth A of the INS can be calculated. c After establishing the attitude matrix using the azimuth, pitch, and roll parameters of the aircraft's inertial navigation system (INS), the INS immediately transitions to attitude tracking, simultaneously performing fine alignment of the vehicle-mounted INS and aircraft erection. The precise azimuth A of the aircraft's INS at the end of coarse alignment is... r The azimuth, pitch, and roll of the vehicle-mounted inertial navigation system (INS) can be calculated using the precisely aligned azimuth, pitch, and roll parameters, as well as previously defined parameters (collimator measurements, INS tilt angle, and pre-calibrated parameters). The current azimuth correction for the INS is then calculated using the following formula:

[0152] φ U (T)=A c -A r (twenty four)

[0153] After the aircraft takes off, the initial attitude value is calculated using the acceleration output over an additional several tens of seconds, replacing the inaccurate current attitude of the aircraft's inertial navigation system (INS) (this inaccurate current attitude is caused by the initial orientation error of the INS when it enters attitude tracking). At this point, the current orientation and attitude of the aircraft's INS have been corrected.

[0154] The beneficial technical effects achieved by this specific embodiment are as follows:

[0155] 1. A model for estimating real-time azimuth and attitude errors during inertial navigation system (INS) attitude tracking was established. The influence of initial azimuth and attitude errors of the aircraft's INS and gyroscope zero bias on real-time azimuth and attitude errors during attitude tracking was revealed, making post-compensation of initial azimuth and gyroscope zero bias of the aircraft's INS possible.

[0156] 2. Based on the real-time error estimation model, an inertial navigation system attitude tracking error estimation and correction process was designed, which enabled the parallel execution of aircraft erection and vehicle-mounted inertial navigation system precision alignment, significantly shortening the total aiming and erection time.

[0157] 3. A simple horizontal aiming and attitude correction method that runs parallel to the aircraft's erection is proposed. After the aircraft is erected, the initial attitude value is calculated using the acceleration output during the extra tens of seconds, replacing the inaccurate current attitude of the aircraft's inertial navigation system, thus eliminating the matrix calculation in the real-time error estimation model.

[0158] 4. A parallel aiming method based on aircraft inertial navigation system (INS) data storage and computation is also proposed. This method obtains the final azimuth and attitude of the aircraft INS by offline calculation of the aircraft INS attitude tracking results after the vehicle-mounted INS self-alignment is completed, thus achieving parallel operation of vehicle-mounted INS self-alignment and erection. Compared with existing technologies, this method can shorten the total aiming and erection time. However, compared with the other two schemes, this method has two drawbacks: first, it requires a large amount of stored aircraft INS data, and the longer the erection time, the larger the amount of data stored; second, it requires additional time to calculate the final azimuth and attitude of the aircraft INS, and is not real-time.

[0159] In summary, based on the same technical concept, this specific embodiment proposes three methods for horizontal aiming and attitude correction that are performed in parallel with erection. These methods involve establishing an error model and performing matrix calculations, performing initial value calculations, and employing offline calculations, respectively, to perform horizontal aiming and attitude correction simultaneously with erection. Compared with traditional horizontal aiming methods, the proposed technical solutions can significantly shorten the total time for aiming and erection without reducing the final azimuth and attitude accuracy of the inertial navigation system (INS) during the aircraft's ground preparation phase. This solves the problem of the long total time for INS aiming and erection during the ground preparation phase of the aircraft, which is caused by the inability to perform vehicle-mounted aircraft erection in parallel with aiming in the existing technology. These solutions have outstanding substantive features and significant progress.

Claims

1. A method of horizontal aiming and attitude correction in parallel with erecting, characterized in that, It comprises the following steps: S1, when the aircraft is horizontal, the vehicle-mounted inertial unit is coarsely aligned; meanwhile, the output of the accelerometer is used to calculate the pitch, roll and inclination of the aircraft inertial unit; S2, at the end of coarse alignment, photoelectric collimator collimation prism and output measurement results; using the coarse alignment azimuth, pitch and roll of the vehicle-mounted inertial unit, the measurement results of the collimator, the inclination of the aircraft inertial unit and the pre-calibrated parameters, the coarse azimuth A of the aircraft inertial unit is calculated c , initialization vector A 0 , vector D(0), B(0) is determined whether to initialize as needed; A 0 By the following equation: A 0 = I 3×3 I 3×3 is a 3x3 identity matrix; S3, after establishing the attitude matrix through the coarse azimuth, pitch and roll of the aircraft inertial unit, the aircraft inertial unit immediately enters the attitude tracking, and the vector A is updated in real time i The vectors D(T) and B(T) are determined according to the need whether to update in real time. Wherein, B(T) is used to update D(T+Δt) of the next time; T=iΔt; Δt is the sampling interval of the aircraft inertial unit; S4, after the aircraft inertial unit is switched into attitude tracking, the vehicle-mounted inertial unit is precisely aligned and the aircraft is erected in parallel; S5, after the erection is completed, the vehicle-mounted inertial unit is precisely aligned and the attitude of the aircraft inertial unit is corrected before the aircraft is ignited.

2. The method of claim 1, wherein, The step of correcting the attitude of the aircraft inertial unit is performed according to the following method: S5.1, after the erection is completed, the vehicle-mounted inertial unit is precisely aligned and the attitude is corrected before the aircraft is ignited. The initial azimuth correction amount of the aircraft inertial unit is calculated according to the following formula: φ U (0) = A c -A r In the formula, A r is the accurate azimuth of the aircraft inertial unit at the end of coarse alignment, which is calculated by the azimuth, pitch and roll of the reverse accurate alignment of the vehicle-mounted inertial unit, the measurement results of the collimator, the inclination of the aircraft inertial unit and the pre-marked parameters; Let φ E (0) and φ N (0) be 0, the vector φ(0) is written as: φ(0) = [0 0 φ U (0)] T ; S5.2, the vector φ(0) obtained in the above steps and the real-time updated vectors D(T) and A i , the vector φ(T) is estimated, and the attitude matrix is corrected according to the following formula wherein calculated by the formula: By The azimuth, pitch and roll angles of the current time of the aircraft's inertial measurement unit are calculated to complete the attitude correction. The vector φ(T) is estimated according to the following formula: φ(T) = A i φ(0) + D(T)ε b where ε b represents the gyro zero offset vector.

3. The method of claim 2, wherein, In the horizontal aiming and attitude correction method in parallel with the erecting, the attitude tracking error caused by the gyro zero bias vector ε b is estimated by the following formula: wherein, is the average output of the gyro of the vertical forebody inertial unit; is the initial attitude matrix, which is calculated from the initial yaw, pitch and roll of the aircraft's inertial measurement unit (IMU); is the earth rotation angular velocity vector; ω ie is the earth rotation angular velocity and L is the latitude.

4. The method according to any one of claims 1 to 3, characterized in that, The initial vectors D(0) and B(0) are obtained according to the following formula: D(0)＝0 3×3 wherein 0 3×3 is a 3x3 zero matrix; is the initial attitude matrix, which is calculated from the initial yaw, pitch and roll of the aircraft's inertial measurement unit (IMU).

5. The method of claim 4, wherein, After the aircraft inertial unit is switched into attitude tracking, D(T) is updated in real time according to the following formula: D(T)=AD(T-Δt)+B(T-Δt) B(T) is updated in real time by the following equation:

6. The method of claim 5, wherein, The A i is updated in real time by the equation: A i = A · A i-1 wherein A i-1 Updated at the previous time.

7. The method of claim 2, wherein, The horizontal aiming and attitude correction method in parallel with erecting does not need to estimate and compensate the attitude tracking error caused by gyro zero bias b The gyro average output of the aircraft inertial unit during the rough alignment of the vehicle-mounted inertial unit is not needed to be saved, and B(T), D(T) is not needed to be initialized and updated; the vector φ(T) is estimated according to the following formula: φ(T) = A i φ(0) A i is updated in real time by the equation: A i = A · A i-1 wherein A i-1 Updated at the previous time.

8. A method of horizontal aiming and attitude correction in parallel with erecting, characterized in that, It comprises the following steps: S1, when the aircraft is horizontal, the vehicle-mounted inertia unit is coarsely aligned; meanwhile, the output of the accelerometer is used to calculate the inclination, roll and pitch of the aircraft inertial unit; S2, at the end of the coarse alignment, the photoelectric collimator collimates the prism and outputs the measurement result; using the coarse alignment azimuth, pitch and roll of the vehicle-mounted inertial unit, the measurement result of the collimator, the inclination of the aircraft inertial unit and the pre-labeled parameters, the coarse azimuth A of the aircraft inertial unit is calculated c ; S3, after the attitude matrix is established through the coarse azimuth, pitch and roll of the aircraft inertial unit, the aircraft inertial unit is immediately switched into attitude tracking; S4, after the aircraft inertial unit is switched into attitude tracking, the vehicle-mounted inertia unit is precisely aligned and the aircraft is erected in parallel; S5, after the erection is complete, the vehicle-mounted inertial unit is precisely aligned and the attitude is corrected by using the output of the additional tens of seconds of acceleration to replace the inaccurate current attitude of the aircraft inertial unit, so as to complete the current azimuth and attitude correction of the aircraft inertial unit; During the attitude initial value calculation process, the initial aircraft inertial measurement unit azimuth correction amount φ U (0) is calculated by the following formula: φ U (0) = A' c - A' r In the formula, A' c is the azimuth of the vehicle-mounted inertial unit from the coarse alignment to the end of the fine alignment. A′ r Azimuth of the end time of the fine alignment of the vehicle-mounted inertial unit After the current azimuth and attitude correction of the aircraft inertial unit is completed, the vehicle-mounted inertial unit is no longer precisely aligned before the aircraft is ignited.

9. A method of horizontal aiming and attitude correction in parallel with erecting, characterized in that, It comprises the following steps: S1, when the aircraft is horizontal, the vehicle-mounted inertio unit is self-aligned, and the output of the accelerometer is used to calculate the pitch, roll and inclination of the aircraft inertio unit, and the collimation prism of the photoelectric collimator is collimated and the measurement result is output; S2, after the pitch, roll and inclination of the aircraft inertio unit are calculated, the original data of the aircraft inertio unit is saved, the vehicle-mounted inertio unit continues to be self-aligned, and the aircraft starts to be erected; S3, after the erection is completed, the initial azimuth of the aircraft inertio unit in the horizontal state is calculated through the self-aligned result of the vehicle-mounted inertio unit, the inclination of the aircraft inertio unit in the horizontal state, the measurement result of the photoelectric collimator and the pre-marked parameters; S4, the initial attitude matrix is established through the initial azimuth, pitch and roll of the aircraft inertio unit, and then the saved original data of the aircraft inertio unit is used for attitude tracking to calculate the final azimuth and attitude of the aircraft inertio unit.

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