SINS / GNSS (strapdown inertial navigation system / global navigation satellite system) air alignment method for large overload section of aircraft

Abstract

The invention relates to an SINS / GNSS air alignment method for a large overload section of an aircraft. The method comprises the steps that S1, first position information of the aircraft at each measurement moment in the flight process of the large overload section is calculated based on an SINS; s2, outputting second position information of the aircraft at a corresponding measurement moment based on the GNSS; s3, constructing an observed quantity according to the first position information and the second position information, and carrying out measurement residual error verification; if the observation quantity does not pass the measurement residual error verification, entering a step S4, taking the observation quantity as bad observation information to be abandoned, and returning to the step S1; if the measurement residual error verification is passed, entering the step S5, correcting the attitude error of the SINS through the observed quantity, and obtaining the missile attitude quaternion at the current moment so as to complete aerial alignment. According to the SINS / GNSS air alignment method for the large overload section of the aircraft, the advantages of the SINS and the GNSS are complemented, the requirement for the precision of the gyroscope is low, the system error-tolerant rate is high, the alignment speed is high, and the precision is high.

Description

Technical Field

[0001] The invention relates to the technical field of aerial alignment, and in particular to a SINS / GNSS aerial alignment method for a large overload section of an aircraft. Background Art

[0002] At present, there are two widely used navigation systems on aircraft, one is the strapdown inertial navigation system (SINS), and the other is the global navigation satellite system (GNSS). The strapdown inertial navigation system is a navigation technology based on Newton's theorem, which uses inertial devices to measure the real-time information of the carrier and perform calculations to obtain the carrier's attitude, position, speed and other information. As a purely autonomous navigation system, SINS does not rely on external information and does not radiate any energy to the outside. It has good dynamic performance and is widely used. However, due to the errors of the inertial devices themselves and their integral operations, the errors will accumulate over time, and the navigation accuracy will gradually diverge. The global satellite navigation system is a satellite-based radio navigation technology with the characteristics of ultra-high precision, all-time, and all-weather. As a non-pure autonomous navigation system, GNSS needs to receive satellite signals, and the signals are easily blocked or interfered with, and the data refresh rate is low.

[0003] In the related technology, if the traditional ground self-alignment method uses SINS, a high-precision gyroscope is required to complete high-precision alignment, so the gyroscope has high requirements on accuracy and low fault tolerance. If GNSS is used, the high dynamic environment of the aircraft's high overload period (i.e., the working period of the aircraft engine) can easily have a significant impact on the navigation accuracy of GNSS, so GNSS is generally not suitable for use during this period. Summary of the invention

[0004] The present application aims to solve the technical problems that the ground self-alignment method in the related art uses SINS, which leads to high requirements for gyroscope accuracy and low fault tolerance, and the use of GNSS is not suitable for the large overload section of the aircraft and the navigation accuracy is easily affected.

[0005] The present application embodiment provides a SINS / GNSS aerial alignment method for a large overload section of an aircraft, comprising the following steps:

[0006] Step S1: Calculate the first position information of the aircraft at each measurement time during the high-overload flight based on SINS, wherein the first position information includes the latitude L SINS , longitude λ SINS , height h SINS ;

[0007] Step S2: output second position information of the aircraft at the corresponding measurement time based on GNSS, wherein the second position information includes latitude L GNSS , longitude λ GNSS , height h GNSS ;

[0008] Step S3, constructing an observation quantity according to the first position information and the second position information, and performing measurement residual verification;

[0009] If the measurement residual check fails, the process proceeds to step S4, where the observation is discarded as bad observation information, and the process returns to step S1;

[0010] If the measurement residual check is passed, the process proceeds to step S5, where the attitude error of the SINS is corrected using the observed value, and the missile body attitude quaternion at the current moment is obtained to complete the air alignment.

[0011] In one embodiment, the step S1, calculating the first position information of the aircraft at each measurement time during the high-overload flight process based on SINS, comprises:

[0012] Step S11, obtaining the initial attitude, initial velocity and initial position of the aircraft as the initial navigation values ​​of SINS;

[0013] Step S12: collecting SINS data of the aircraft during the flight of the large overload section, performing pure inertial navigation solution and Kalman filter time update, and obtaining the first position information at each measurement moment.

[0014] In one implementation, the ground alignment result of the aircraft is used as the initial posture of the aircraft.

[0015] In one embodiment, the step S12, collecting SINS data of the aircraft during the high overload flight to perform pure inertial navigation solution and Kalman filter time update, includes:

[0016] Step S121, collecting SINS data to perform pure inertial navigation attitude update, velocity update and position update solution, wherein the SINS data includes the apparent velocity increment and angle increment output by the missile-borne inertial combination device;

[0017] Step S122: Update the Kalman filter time.

[0018] In one implementation, the pure inertial navigation position update solution formula is:

[0019]

[0020] in, Represents the radius of principal curvature of the Earth's circumplex; represents the principal radius of curvature of the earth's meridian;

[0021] —respectively the eastward, northward and celestial velocities before recursive update; —respectively the eastward, northward and celestial velocities after recursive update; t h is the sampling period of the inertial combination device; L k-1 ,λ k-1 ,h k-1 —respectively the geographic latitude, longitude and altitude before recursive update; L k ,λ k ,h k —respectively, the geographical latitude, longitude and altitude after recursive update, as the first location information.

[0022] In one embodiment, the step S2 of outputting the second position information of the aircraft at the corresponding measurement time based on GNSS includes:

[0023] Step S21: Obtain the PDOP value and the missile overload change rate output by the GNSS receiver at the corresponding measurement time based on the GNSS.

[0024] Step S22: Determine whether the PDOP value is less than or equal to the set threshold value and the missile overload change rate. Is it ≤ threshold value?

[0025] If both are true, then the position information at the measurement time is output as the second position information;

[0026] If no, the position information at the measurement time is discarded as bad observation information, and the process returns to step S1.

[0027] In one embodiment, the set threshold value of the PDOP value is ≤6.

[0028] In one implementation, the step S3, constructing an observation quantity according to the first position information and the second position information, and performing measurement residual verification includes:

[0029] Step S31: construct an observation quantity Z according to the first position information and the second position information k , expressed as:

[0030]

[0031] Step S32: construct the observation matrix H k and the matrix A,

[0032] The observation matrix H k Expressed as:

[0033] The matrix A is expressed as:

[0034] Among them, P k / k-1 —state one-step prediction variance;

[0035] R k represents the observation noise matrix;

[0036] Step S33: perform measurement residual check:

[0037] Take the square root of the diagonal elements of matrix A and get vector A diagsqrt ;

[0038] The vector Z k -H k X k / k-1 Take the absolute value of each element in to get the vector |Z k -H k X k / k-1 |, where X k / k-1 —state prediction value;

[0039] Judgment vector |Z k -H k X k / k-1 Is any element in | not greater than vector A? diagsqrt 3 times the corresponding element in;

[0040] If yes, it is considered to have passed the measurement residual check;

[0041] If not, it is considered to have failed the measurement residual calibration.

[0042] In one embodiment, the step S5, correcting the attitude error of the SINS by using the observation quantity to obtain the quaternion of the missile body attitude at the current moment, comprises:

[0043] Step S51: Observation quantity Z k Send it to the Kalman filter and perform Kalman filter measurement update to obtain the estimated state X k ;

[0044] Step S52: By estimating the state X k The pure inertial navigation attitude quaternion q at the current moment SINS Output correction is performed to obtain the projectile attitude quaternion q at the current moment.

[0045] In one embodiment, the step S52, by estimating the state X k The pure inertial navigation attitude quaternion q at the current moment SINS Output correction is performed to obtain the current missile body attitude quaternion q, which includes:

[0046] Step S521, calculate the correction quaternion q′:

[0047] The calculation formula is:

[0048] Where Φ is the misalignment angle, expressed as the estimated state X k The first three dimensions: |Φ| is the modulus of Φ;

[0049] Step S522: Correct the quaternion q′ to the current pure inertial navigation attitude quaternion q SINS Perform output correction to obtain the current missile attitude quaternion q.

[0050] The current pure inertial navigation attitude quaternion q SINS Expressed as:

[0051] The current missile attitude quaternion q is expressed as:

[0052] The beneficial effects brought by the technical solution provided in the embodiments of the present application include:

[0053] The present application provides a SINS / GNSS aerial alignment method for a large overload section of an aircraft. An observation quantity is constructed according to first position information provided by SINS and second position information provided by GNSS, a position error term is selected and a speed error term which is more sensitive to a high dynamic environment is discarded, and then bad observation information is eliminated through measurement residual verification, so as to avoid the adverse effect of the high dynamic environment of the large overload section of the aircraft on the aerial alignment. The advantages of SINS and GNSS are complemented, and the strong observability of the heading error of SINS in the large overload section of the aircraft is fully utilized, the gyro accuracy requirement is low, and the system fault tolerance rate is high; GNSS data which is not applicable to the large overload section of the aircraft is utilized, and alignment can be completed in the large overload section of the aircraft flight, and the alignment speed is fast and the accuracy is high. BRIEF DESCRIPTION OF THE DRAWINGS

[0054] In order to more clearly illustrate the technical solutions in the embodiments of the present application, the drawings required for use in the description of the embodiments will be briefly introduced below. Obviously, the drawings described below are only some embodiments of the present application. For ordinary technicians in this field, other drawings can be obtained based on these drawings without creative work.

[0055] Figure 1 The present invention is a flowchart of a SINS / GNSS aerial alignment method for a high-G segment of an aircraft according to an embodiment of the present invention. DETAILED DESCRIPTION

[0056] In order to enable those skilled in the art to better understand the solution of the present application, the technical solution in the embodiments of the present application will be clearly and completely described below in conjunction with the drawings in the embodiments of the present application. Obviously, the described embodiments are only part of the embodiments of the present application, not all of the embodiments. Based on the embodiments in the present application, all other embodiments obtained by ordinary technicians in this field without creative work are within the scope of protection of this application.

[0057] First, some proper nouns are explained:

[0058] High overload period: the period when the aircraft's engine is working.

[0059] PDOP (Position Dilution of Precision) refers to the position precision factor, which is used to measure the accuracy of the GNSS positioning system. The PDOP value reflects the impact of satellite geometric distribution on positioning accuracy. The smaller the value, the higher the positioning accuracy; conversely, the larger the value, the lower the positioning accuracy.

[0060] The projectile overload change rate is an important parameter that describes how fast the overload of an aircraft changes over time during flight. It is the change in the projectile overload per unit time, which intuitively reflects the magnitude of the overload change in each second.

[0061] like Figure 1 As shown, Figure 1 The present invention is a flowchart of a SINS / GNSS aerial alignment method for a high-G segment of an aircraft according to an embodiment of the present invention.

[0062] This embodiment provides a SINS / GNSS air alignment method for an aircraft in a high-G segment, comprising the following steps:

[0063] Step S1: Calculate the first position information of the aircraft at each measurement time during the high-overload flight based on SINS. The first position information includes the latitude L SINS , longitude λ SINS , height h SINS ;

[0064] Step S2: Output the second position information of the aircraft at the corresponding measurement time based on GNSS, the second position information includes the latitude L GNSS , longitude λ GNSS , height h GNSS ;

[0065] Step S3, constructing an observation quantity according to the first position information and the second position information, and performing measurement residual verification;

[0066] If the measurement residual check fails, the process proceeds to step S4, where the observation is discarded as bad observation information, and the process returns to step S1;

[0067] If the measurement residual check passes, the process proceeds to step S5, where the SINS attitude error is corrected using the observed value, and the missile body attitude quaternion at the current moment is obtained to complete the air alignment.

[0068] The present embodiment provides a SINS / GNSS aerial alignment method for the high overload section of an aircraft, constructs an observation quantity based on the first position information provided by SINS and the second position information provided by GNSS, selects the position error term and discards the speed error term that is more sensitive to the high dynamic environment, and then eliminates the bad observation information through the measurement residual check, so as to avoid the adverse effect of the high dynamic environment of the high overload section of the aircraft on the aerial alignment, and complements the advantages of SINS and GNSS, making full use of the strong observability of the heading error of SINS in the high overload section of the aircraft, with low requirements for gyro accuracy and high system fault tolerance; utilizes the GNSS data that is not applicable to the high overload section of the aircraft, and can complete the alignment in the high overload section of the aircraft flight, with fast alignment speed and high accuracy. The present embodiment is applicable to scenes such as aircraft, launch vehicles, missiles, rockets, etc. that need to be aligned in the air.

[0069] Each step is described and explained in detail below.

[0070] In one embodiment, step S1, calculating the first position information of the aircraft at each measurement time during the high-overload flight process based on SINS, includes:

[0071] Step S11, obtaining the initial attitude, initial velocity and initial position of the aircraft as the initial navigation values ​​of SINS.

[0072] In one embodiment, the ground alignment result of the aircraft is used as the initial attitude of the aircraft.

[0073] Specifically, the rough ground alignment result q0 of the aircraft is obtained as the initial attitude information of the projectile, and the initial velocity v0 and initial position p0 of the projectile are obtained at the same time, and the above information is used as the initial value of the pure inertial navigation of the aircraft.

[0074] Step S12: collecting SINS data of the aircraft during the high-overload flight, performing pure inertial navigation solution and Kalman filter time update, and obtaining the first position information at each measurement moment.

[0075] In one embodiment, step S12, collecting SINS data of the aircraft during the high-g segment flight process to perform pure inertial navigation solution and Kalman filter time update, includes:

[0076] Step S121, collecting SINS data to perform pure inertial navigation attitude update, velocity update and position update solution, the SINS data including the apparent velocity increment and angle increment output by the missile-borne inertial combination device;

[0077] Specifically, the missile-borne navigation computer receives the apparent velocity increment and angle increment information output by the missile-borne inertial combination device at a certain frequency (such as a sampling period of 5ms or 10ms).

[0078] Step S122: Update the Kalman filter time.

[0079] In one embodiment, the quaternion method is used to perform pure inertial navigation attitude update:

[0080]

[0081] Where: Δθ x ,Δθ y ,Δθ z —are the angle increments output by the missile-borne inertial combination device at the current moment, in radians; q0(t k-1 ),q1(t k-1 ),q2(t k-1 ),q3(t k-1 )—are the attitude quaternions before recursive update; q0(t k ),q1(t k ),q2(t k ),q3(t k )—are the attitude quaternions after recursive update.

[0082] In one embodiment, pure inertial navigation speed update is performed, and the calculation formula is:

[0083]

[0084]

[0085] Where: t h is the sampling period of the inertial combination device, in seconds; ΔV x ,ΔV y ,ΔV z —are the apparent velocity increments output by the missile-borne inertial combination device at the current moment, in meters per second; —respectively the eastward, northward and celestial velocities before recursive update, all in meters per second; —respectively the eastward, northward and celestial velocities after recursive update, all in meters per second; — are the accelerations in the north, sky and east directions, in m / s 2 ; —Coriolis acceleration in the north, celestial and east directions, in m / s 2 ; g k —Current gravitational acceleration, in m / s 2 ;

[0086] Represents the direction cosine matrix from the missile's own system to the navigation system.

[0087] In one embodiment, the pure inertial navigation position update solution formula is:

[0088]

[0089] in, Represents the radius of principal curvature of the Earth's circumplex; represents the principal radius of curvature of the earth's meridian;

[0090] —respectively the eastward, northward and celestial velocities before recursive update; —respectively the eastward, northward and celestial velocities after recursive update; t h is the sampling period of the inertial combination device; L k-1 ,λ k-1 ,h k-1 —respectively the geographic latitude, longitude and altitude before recursive update, in radians, radians and meters respectively; L k ,λ k ,h k —respectively represent the recursively updated geographic latitude, longitude, and altitude, in radians, radians, and meters, as the first location information.

[0091] In one embodiment, step S122, performing Kalman filter time update includes:

[0092] The following states are selected as the system states of the filter:

[0093] X=[φ x φ y φ z δv E δv N δv U δL δλ δh ε x ε y ε z δK gx δK gy δK gz Δ x Δ y Δ z δK ax δK ay δKaz ];

[0094] Among them, X—filter state vector; φ x ,φ y ,φ z — respectively, attitude misalignment angle, in rad; δv E ,δv N ,δv U — respectively represent the velocity error in the navigation system, in m / s; δL, δλ, δh — respectively represent the position error, in rad, rad, m; ε x ,ε y ,ε z —x, y, z gyro bias, in rad / s; δK gx ,δK gy ,δK gz —respectively x, y, z gyro scale errors, in ppm; Δ x ,Δ y ,Δ z —x, y, z plus table zero deviation, unit is m / s 2 ; δK ax ,δK ay ,δK az —x, y, z plus scale error respectively, in ppm.

[0095] The equation for filter time update is:

[0096]

[0097] Among them, the state transfer matrix A k / k-1 The calculation is as follows: A k / k-1 =I 21×21 +FΔT

[0098]

[0099]

[0100] Among them, X k-1 —The state estimate of the previous step; A k / k-1 —The state transfer matrix calculated based on the current satellite measurement pseudorange, pseudorange rate, satellite position and velocity, and the current position and velocity of inertial navigation; —A k / k-1 Device; Q k —process noise variance matrix; X k / k-1 —state prediction value; P k / k-1 —state one-step prediction variance; P k-1 —The variance matrix of the previous state estimate; -vector The antisymmetric matrix of ; —The projection of the rotation angular velocity of the carrier system relative to the inertial system on the carrier system; (f n ×)—vector f b The antisymmetric matrix projected in the navigation system; (v n ×)—vector v n The antisymmetric matrix of .

[0101] Through the above scheme, Kalman filter (KF) technology is used during the high-overload flight of the aircraft, and the pure inertial navigation error is modeled through the linear system state equation, which makes full use of the characteristics of strong observability of the pure inertial navigation heading error in the high-overload section of the aircraft and fast convergence of the Kalman filter.

[0102] In one embodiment, step S2, outputting the second position information of the aircraft at the corresponding measurement time based on the GNSS, comprises:

[0103] Step S21: Obtain the PDOP value and the missile overload change rate output by the GNSS receiver at the corresponding measurement time based on the GNSS.

[0104] Step S22: Determine whether the PDOP value is less than or equal to the set threshold value and the missile overload change rate. Is it ≤ threshold value?

[0105] If both are true, then the position information at the measurement time is output as the second position information;

[0106] If no, the position information at the measurement time is discarded as bad observation information, and the process returns to step S1.

[0107] In one embodiment, the setting threshold of the PDOP value is ≤6, and the threshold value of the missile overload change rate is determined according to the maximum overload change rate that the GNSS satellite navigation receiver can adapt to.

[0108] Through the above scheme, the PDOP value corresponds to the satellite configuration. The satellite configuration and the missile overload change rate are used as indicators. When the PDOP value and the missile overload change rate do not meet the set conditions, they are eliminated as bad observation information to avoid their adverse effects on the system and improve the system fault tolerance.

[0109] In one embodiment, step S3, constructing an observation quantity according to the first position information and the second position information, and performing measurement residual verification includes:

[0110] Step S31: construct an observation quantity Z according to the first position information and the second position information k , expressed as:

[0111]

[0112] Step S32: construct the observation matrix H k and the matrix A,

[0113] The observation matrix H k Expressed as:

[0114] The matrix A is expressed as:

[0115] Among them, P k / k-1 —state one-step prediction variance;

[0116] R k represents the observation noise matrix;

[0117] Step S33: perform measurement residual check:

[0118] Take the square root of the diagonal elements of matrix A and get vector A diagsqrt ;

[0119] The vector Z k -H k X k / k-1 Take the absolute value of each element in to get the vector |Z k -H k X k / k-1 |, where X k / k-1 —state prediction value;

[0120] Judgment vector |Z k -H k X k / k-1 Is any element in | not greater than vector A? diagsqrt 3 times the corresponding element in;

[0121] If yes, it is considered to have passed the measurement residual check;

[0122] If not, it is considered to have failed the measurement residual calibration.

[0123] Through the above scheme, when constructing the observation quantity, the position error term is selected and the velocity error term which is more sensitive to the high dynamic environment is discarded. The measurement residual is used as an indicator to eliminate the bad observation information and avoid the adverse effect of the high dynamic environment of the large overload section of the aircraft on the air alignment.

[0124] In one embodiment, step S5, correcting the attitude error of the SINS by using the observation quantity to obtain the quaternion of the missile body attitude at the current moment includes:

[0125] Step S51: Observation quantity Z k Send it to the Kalman filter and perform Kalman filter measurement update to obtain the estimated state X k .

[0126] Specifically, the Kalman filter measurement update includes:

[0127]

[0128] Step S52: By estimating the state X k The pure inertial navigation attitude quaternion q at the current moment SINS Output correction is performed to obtain the projectile attitude quaternion q at the current moment.

[0129] In one embodiment, step S52, by estimating the state X k The pure inertial navigation attitude quaternion q at the current moment SINS Output correction is performed to obtain the current missile body attitude quaternion q, which includes:

[0130] Step S521, calculate the correction quaternion q′:

[0131] The calculation formula is:

[0132] Where Φ is the misalignment angle, expressed as the estimated state X k The first three dimensions: |Φ| is the modulus of Φ;

[0133] Step S522: Correct the quaternion q′ to the current pure inertial navigation attitude quaternion q SINS Perform output correction to obtain the current missile attitude quaternion q.

[0134] The current pure inertial navigation attitude quaternion q SINS Expressed as:

[0135] The current missile attitude quaternion q is expressed as:

[0136] Through the above scheme, the pure inertial navigation attitude error is optimally estimated and corrected online by measuring after eliminating bad observation information, so as to obtain high-precision attitude information of the aircraft.

[0137] In the description of the present application, it should be noted that the terms "upper", "lower", etc. indicate orientations or positional relationships based on the orientations or positional relationships shown in the accompanying drawings, and are only for the convenience of describing the present application and simplifying the description, rather than indicating or implying that the device or element referred to must have a specific orientation, be constructed and operated in a specific orientation, and therefore cannot be understood as a limitation on the present application. Unless otherwise clearly specified and limited, the terms "installed", "connected", and "connected" should be understood in a broad sense, for example, it can be a fixed connection, a detachable connection, or an integral connection; it can be a mechanical connection or an electrical connection; it can be a direct connection, or it can be an indirect connection through an intermediate medium, or it can be a connection between the two elements. For those of ordinary skill in the art, the specific meanings of the above terms in the present application can be understood according to the specific circumstances.

[0138] It should be noted that, in this application, relational terms such as "first" and "second" are only used to distinguish one entity or operation from another entity or operation, and do not necessarily require or imply any such actual relationship or order between these entities or operations. Moreover, the terms "include", "comprise" or any other variants thereof are intended to cover non-exclusive inclusion, so that a process, method, article or device including a series of elements includes not only those elements, but also other elements not explicitly listed, or also includes elements inherent to such process, method, article or device. In the absence of further restrictions, the elements defined by the sentence "comprise a ..." do not exclude the presence of other identical elements in the process, method, article or device including the elements.

[0139] In the description of the embodiments of the present application, unless otherwise specified, “ / ” means or, for example, A / B can mean A or B; the “and / or” in the text is merely a description of the association relationship of associated objects, indicating that three relationships may exist, for example, A and / or B can mean: A exists alone, A and B exist at the same time, and B exists alone. In addition, in the description of the embodiments of the present application, “multiple” refers to two or more than two.

[0140] The above is only a specific implementation of the present application, so that those skilled in the art can understand or implement the present application. Various modifications to these embodiments will be apparent to those skilled in the art, and the general principles defined herein can be implemented in other embodiments without departing from the spirit or scope of the present application. Therefore, the present application will not be limited to the embodiments shown herein, but will conform to the widest range consistent with the principles and novel features applied for herein.

Claims

1. A SINS / GNSS air alignment method for aircraft high-G segment, characterized in that: The following steps are involved: Step S1: Calculate the first position information of the aircraft at each measurement time during the high-overload flight based on SINS, wherein the first position information includes the latitude L SINS , longitude λ SINS , height h SINS ; Step S2: output second position information of the aircraft at the corresponding measurement time based on GNSS, wherein the second position information includes latitude L GNSS , longitude λ GNSS , height h GNSS ; Step S3, constructing an observation quantity according to the first position information and the second position information, and performing measurement residual verification; If the measurement residual check fails, the process proceeds to step S4, where the observation is discarded as bad observation information, and the process returns to step S1; If the measurement residual check is passed, the process proceeds to step S5, where the attitude error of the SINS is corrected using the observed value, and the missile body attitude quaternion at the current moment is obtained to complete the air alignment.

2. A SINS / GNSS aerial alignment method for a large overload section of an aircraft as claimed in claim 1, characterized in that: The step S1, calculating the first position information of the aircraft at each measurement moment during the high overload flight process based on SINS, comprises: Step S11, obtaining the initial attitude, initial velocity and initial position of the aircraft as the initial navigation values ​​of SINS; Step S12: collecting SINS data of the aircraft during the flight of the large overload section, performing pure inertial navigation solution and Kalman filter time update, and obtaining the first position information at each measurement moment.

3. A SINS / GNSS aerial alignment method for aircraft high-G segment as claimed in claim 2, characterized in that: The ground alignment result of the aircraft is used as the initial attitude of the aircraft.

4. A SINS / GNSS aerial alignment method for aircraft high-G segment as claimed in claim 2, characterized in that: The step S12, collecting the SINS data of the aircraft during the high overload flight to perform pure inertial navigation solution and Kalman filter time update, comprises: Step S121, collecting SINS data to perform pure inertial navigation attitude update, velocity update and position update solution, wherein the SINS data includes the apparent velocity increment and angle increment output by the missile-borne inertial combination device; Step S122: Update the Kalman filter time.

5. A SINS / GNSS aerial alignment method for aircraft high-G segment as claimed in claim 4, characterized in that: The pure inertial navigation position update solution formula is: in, Represents the radius of principal curvature of the Earth's circumplex; represents the principal radius of curvature of the earth's meridian; —respectively the eastward, northward and celestial velocities before recursive update; —respectively the eastward, northward and celestial velocities after recursive update; t h is the sampling period of the inertial combination device; L k-1 ,λ k-1 ,h k-1 —respectively the geographic latitude, longitude and altitude before recursive update; L k ,λ k ,h k —respectively, the geographical latitude, longitude and altitude after recursive update, as the first location information.

6. A SINS / GNSS aerial alignment method for aircraft high-G segment as claimed in claim 1, characterized in that: The step S2, outputting the second position information of the aircraft at the corresponding measurement time based on GNSS, comprises: Step S21: Obtain the PDOP value and the missile overload change rate output by the GNSS receiver at the corresponding measurement time based on the GNSS. Step S22: Determine whether the PDOP value is less than or equal to the set threshold value and the missile overload change rate. Is it ≤ threshold value? If both are true, then the position information at the measurement time is output as the second position information; If no, the position information at the measurement time is discarded as bad observation information, and the process returns to step S1.

7. A SINS / GNSS aerial alignment method for aircraft high-G segment as claimed in claim 6, characterized in that: The setting threshold of the PDOP value is ≤6.

8. A SINS / GNSS aerial alignment method for aircraft high-G segment as claimed in claim 1, characterized in that: The step S3, constructing an observation quantity according to the first position information and the second position information, and performing measurement residual verification comprises: Step S31: construct an observation quantity Z according to the first position information and the second position information k , expressed as: Step S32: construct the observation matrix H k and the matrix A, The observation matrix H k Expressed as: The matrix A is expressed as: Among them, P k / k-1 —state one-step prediction variance; R k represents the observation noise matrix; Step S33: perform measurement residual check: Take the square root of the diagonal elements of matrix A and get vector A diagsqrt ; The vector Z k -H k X k / k-1 Take the absolute value of each element in to get the vector |Z k -H k X k / k-1 |, where X k / k-1 —state prediction value; Judgment vector |Z k -H k X k / k-1 Is any element in | not greater than vector A? diagsqrt 3 times the corresponding element in; If yes, it is considered to have passed the measurement residual check; If not, it is considered to have failed the measurement residual calibration.

9. A SINS / GNSS aerial alignment method for aircraft high-G segment as claimed in claim 8, characterized in that: The step S5, correcting the attitude error of the SINS by using the observation quantity to obtain the quaternion of the missile body attitude at the current moment, comprises: Step S51: Observation quantity Z k Send it to the Kalman filter and perform Kalman filter measurement update to obtain the estimated state X k ; Step S52: By estimating the state X k The pure inertial navigation attitude quaternion q at the current moment SINS Output correction is performed to obtain the projectile attitude quaternion q at the current moment.

10. A SINS / GNSS air alignment method for aircraft high-G segment as claimed in claim 9, characterized in that: In step S52, the estimated state X k The pure inertial navigation attitude quaternion q at the current moment SINS Output correction is performed to obtain the current missile body attitude quaternion q, which includes: Step S521, calculate the correction quaternion q′: The calculation formula is: Where Φ is the misalignment angle, expressed as the estimated state X k The first three dimensions: |Φ| is the modulus of Φ; Step S522: Correct the quaternion q′ to the current pure inertial navigation attitude quaternion q SINS Perform output correction to obtain the current missile attitude quaternion q. The current pure inertial navigation attitude quaternion q SINS Expressed as: The current missile attitude quaternion q is expressed as:

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