A SINS / GNSS airborne alignment method for aircraft during high overload periods

Abstract

This application relates to a SINS / GNSS airborne alignment method for aircraft during high-G phases, comprising: step S1, calculating the first position information of the aircraft at each measurement moment during the high-G phase flight based on SINS; step S2, outputting the second position information of the aircraft at the corresponding measurement moment based on GNSS; step S3, constructing an observation measurement based on the first and second position information, and performing measurement residual verification; if the measurement residual verification fails, proceed to step S4, discarding the observation measurement as bad observation information and returning to step S1; if the measurement residual verification passes, proceed to step S5, correcting the attitude error of SINS using the observation measurement, obtaining the quaternion of the projectile attitude at the current moment, and completing the airborne alignment. This application provides a SINS / GNSS airborne alignment method for aircraft during high-G phases, complementing the advantages of SINS and GNSS, with low requirements for gyroscope accuracy, high system fault tolerance, and fast and accurate alignment speed.

Description

Technical Field

[0001] This invention relates to the field of aerial alignment technology, and more specifically to a SINS / GNSS aerial alignment method for aircraft during high-G phases. Background Technology

[0002] Currently, two types of navigation systems are widely used on aircraft: Strapdown Inertial Navigation System (SINS) and Global Navigation Satellite System (GNSS). SINS is a navigation technology based on Newton's laws, using inertial devices to measure and calculate real-time information about the aircraft, thereby obtaining information such as attitude, position, and velocity. As a purely autonomous navigation system, SINS does not rely on external information and does not radiate any energy, exhibiting good dynamic performance and thus enjoying widespread application. However, due to the inherent errors of the inertial devices and their integration operations, errors accumulate over time, causing navigation accuracy to gradually diverge. GNSS, on the other hand, is a satellite-based radio navigation technology with ultra-high precision, all-weather, and all-time capabilities. As a non-purely autonomous navigation system, GNSS requires receiving satellite signals, which are susceptible to blockage or interference, and has a low data refresh rate.

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

[0004] The purpose of this application is to address the technical problems in related technologies where the use of SINS in ground self-alignment methods results in high requirements for gyroscope accuracy and low fault tolerance, and the use of GNSS is not suitable for the high overload range of aircraft and navigation accuracy is easily affected.

[0005] This application provides a SINS / GNSS airborne alignment method for aircraft during high overload periods, including the following steps:

[0006] Step S1: Calculate the first position information of the aircraft at each measurement moment during the high G-force flight based on SINS. The first position information includes latitude L. SINS Longitude λ SINS Height h SINS ;

[0007] Step S2: Output the second position information of the aircraft at the corresponding measurement time based on GNSS, the second position information including latitude L GNSS Longitude λ GNSS Height h GNSS ;

[0008] Step S3: Construct an observation measurement based on the first and second location information, and perform measurement residual verification;

[0009] If the measurement residual verification fails, proceed to step S4, discard the observation as bad observation information, and return to step S1;

[0010] If the measurement residual verification is passed, proceed to step S5, correct the attitude error of the SINS using the observation, obtain the projectile attitude quaternion at the current moment, and complete the aerial alignment.

[0011] In one implementation, step S1, calculating the first position information of the aircraft at each measurement moment during the high-G flight phase based on SINS, includes:

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

[0013] Step S12: Collect the SINS data of the aircraft during the high overload flight process, perform pure inertial navigation calculation and Kalman filter time update to obtain the first position information at each measurement moment.

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

[0015] In one implementation, step S12, collecting SINS data of the aircraft during high-G flight and performing pure inertial navigation calculation and Kalman filter time update, includes:

[0016] Step S121: Collect SINS data to perform pure inertial navigation attitude update, velocity update and position update calculations. The SINS data includes the apparent velocity increment and angle increment output by the missile-borne inertial combination device.

[0017] Step S122: Perform Kalman filter time update.

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

[0019]

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

[0021] —These represent the eastward, northward, and celestial speeds before the recursive update; —These represent the recursively updated eastward, northward, and celestial velocities, respectively; t h L is the sampling period of the inertial navigation system. k-1 ,λ k-1 ,h k-1 —These represent the latitude, longitude, and altitude before the recursive update; L k ,λ k ,h k —These are the recursively updated geographical latitude, longitude, and altitude, which serve as the primary location information.

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

[0023] Step S21: Based on GNSS, obtain the PDOP value and projectile overload change rate output by the GNSS receiver at the corresponding measurement time.

[0024] Step S22: Determine whether the PDOP value is ≤ a set threshold and the projectile overload change rate. Is it ≤ the threshold value?

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

[0026] If neither is true, the location information at the measurement time is discarded as bad observation information, and the process returns to step S1.

[0027] In one implementation, the threshold value for the PDOP is set to ≤6.

[0028] In one embodiment, step S3, constructing an observation measurement based on the first location information and the second location information, and performing measurement residual verification includes:

[0029] Step S31: Construct the observation Z based on the first location information and the second location information. k , expressed as:

[0030]

[0031] Step S32: Construct the observation matrix H k And matrix A,

[0032] Observation matrix H k Expressed as:

[0033] 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 verification:

[0037] Taking the square root of the diagonal elements of matrix A yields vector A. diagsqrt ;

[0038] Vector Z k -H k X k / k-1 The vector |Z is obtained by taking the absolute value of each element in the vector |Z|. k -H k X k / k-1 |, where X k / k-1 —Predicted state value;

[0039] Determine the vector |Z k -H k X k / k-1 Are all elements in vector A less than or equal to vector A? diagsqrt Three times the corresponding element;

[0040] If so, it is considered to have passed the measurement residual verification;

[0041] If not, it is considered that the measurement residual verification has failed.

[0042] In one implementation, step S5, correcting the attitude error of the SINS using the observation to obtain the projectile attitude quaternion at the current moment, includes:

[0043] Step S51: Observe the Z k The data is fed into a Kalman filter for Kalman filtering measurement updates to obtain the estimated state X. k ;

[0044] Step S52: Through the estimated state X k The pure inertial navigation attitude quaternion q at the current moment SINS Perform output correction to obtain the projectile attitude quaternion q at the current moment.

[0045] In one implementation, step S52 involves estimating the state X. k The pure inertial navigation attitude quaternion q at the current moment SINS Perform output correction to obtain the projectile attitude quaternion q at the current moment, including:

[0046] Step S521: Calculate the corrected 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| represents the modulus of Φ;

[0049] Step S522: Adjust the quaternion q′ of the pure inertial navigation attitude at the current moment. SINS Perform output correction to obtain the projectile attitude quaternion q at the current moment.

[0050] The quaternion q of the pure inertial navigation at the current moment SINS Expressed as:

[0051] The projectile attitude quaternion q at the current moment is expressed as:

[0052] The beneficial effects of the technical solutions provided in this application include:

[0053] This application provides a SINS / GNSS airborne alignment method for aircraft during high-G phases. It constructs an observation measurement based on first position information provided by SINS and second position information provided by GNSS, selecting position error terms while discarding velocity error terms, which are more sensitive to high-dynamic environments. Furthermore, it eliminates undesirable observation information through measurement residual verification, avoiding the adverse effects of the high-dynamic environment on airborne alignment during high-G phases. This method complements the advantages of SINS and GNSS, fully utilizing the strong observability of SINS heading errors during high-G phases, which has low requirements for gyroscope accuracy and high system fault tolerance. It also utilizes GNSS data, which is not typically applicable to high-G phases, enabling alignment to be completed during high-G phases of aircraft flight, resulting in fast and high-precision alignment. Attached Figure Description

[0054] To more clearly illustrate the technical solutions in the embodiments of this application, the accompanying drawings used in the description of the embodiments will be briefly introduced below. Obviously, the accompanying drawings described below are only some embodiments of this application. For those skilled in the art, other drawings can be obtained based on these drawings without creative effort.

[0055] Figure 1 This is a flowchart of a SINS / GNSS air alignment method for aircraft during high overload periods, according to an embodiment of the present invention. Detailed Implementation

[0056] To enable those skilled in the art to better understand the present application, the technical solutions in the embodiments of the present application will be clearly and completely described below with reference to the accompanying drawings. Obviously, the described embodiments are only some embodiments of the present application, and not all embodiments. Based on the embodiments in the present application, all other embodiments obtained by those of ordinary skill in the art without creative effort are within the scope of protection of the present application.

[0057] First, let's explain some of the proper nouns:

[0058] High overload period: The period during which the aircraft's engines operate.

[0059] PDOP (Position Dilution of Precision) is a measure of the accuracy of a GNSS positioning system. The PDOP value reflects the impact of satellite geometry on positioning accuracy; a smaller PDOP value indicates higher positioning accuracy, and vice versa.

[0060] The overload change rate of a projectile is an important parameter describing how quickly the overload of an aircraft changes over time during flight. It is the amount of change in the overload of the projectile per unit time, and it intuitively reflects the magnitude of the overload change every second.

[0061] like Figure 1 As shown, Figure 1 This is a flowchart of a SINS / GNSS air alignment method for aircraft during high overload periods, according to an embodiment of the present invention.

[0062] This embodiment provides a SINS / GNSS air-to-air alignment method for aircraft during high-G phases, including the following steps:

[0063] Step S1: Calculate the first position information of the aircraft at each measurement moment during the high G-force flight based on SINS. The first position information includes latitude L. SINS Longitude λ SINS Height h SINS ;

[0064] Step S2: Based on GNSS output, the second position information of the aircraft at the corresponding measurement time is output. The second position information includes latitude L. GNSS Longitude λ GNSS Height h GNSS ;

[0065] Step S3: Construct observations based on the first and second location information, and perform measurement residual verification;

[0066] If the measurement residual verification fails, proceed to step S4, discard the observation as bad observation information, and return to step S1;

[0067] If the measurement residual verification is passed, proceed to step S5, correct the attitude error of SINS using the observation, obtain the projectile attitude quaternion at the current moment, and complete the air alignment.

[0068] This embodiment provides a SINS / GNSS airborne alignment method for aircraft during high-G phases. It constructs an observation measurement based on first position information provided by SINS and second position information provided by GNSS, selecting position error terms while discarding velocity error terms, which are more sensitive to high-dynamic environments. Furthermore, it eliminates undesirable observation information through measurement residual verification, avoiding the adverse effects of the high-dynamic environment during high-G phases on airborne alignment. This method complements the advantages of SINS and GNSS, fully utilizing the strong observability of SINS heading errors during high-G phases, its low requirement for gyroscope accuracy, and its high system fault tolerance. It also utilizes GNSS data, which is not typically applicable to high-G phases, enabling alignment to be completed during these phases with high speed and accuracy. This embodiment is applicable to scenarios requiring airborne alignment, such as aircraft, launch vehicles, missiles, and rockets.

[0069] The following is a detailed explanation and elaboration of each step.

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

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

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

[0073] Specifically, the ground coarse alignment result q0 of the aircraft is obtained as the initial attitude information of the missile body, and the initial velocity v0 and initial position p0 of the missile body are obtained at the same time. The above information is used as the initial values ​​for the pure inertial navigation of the aircraft body.

[0074] Step S12: Collect SINS data of the aircraft during the high overload period of flight, perform pure inertial navigation calculation and Kalman filter time update to obtain the first position information at each measurement moment.

[0075] In one embodiment, step S12, collecting SINS data during the high-G flight process of the aircraft and performing pure inertial navigation calculation and Kalman filter time update, includes:

[0076] Step S121: Collect SINS data to perform pure inertial navigation attitude update, velocity update and position update calculation. The SINS data includes 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 navigation system at a certain frequency (such as a sampling period of 5ms or 10ms).

[0078] Step S122: Perform Kalman filter time update.

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

[0080]

[0081] Where: Δθ x ,Δθ y ,Δθ z —These represent the angular increments output by the onboard inertial navigation system at the current moment, in radians; q0(t) k-1 ),q1(t k-1 ),q2(t k-1 ),q3(t k-1 — These are the attitude quaternions before the recursive update; q0(t) k ),q1(t k ),q2(t k ),q3(t k — These are the attitude quaternions after recursive updates.

[0082] In one embodiment, the pure inertial navigation velocity is updated, and the calculation formula is as follows:

[0083]

[0084]

[0085] Where: t h The sampling period of the inertial navigation system is expressed in seconds; ΔV x ,ΔV y ,ΔV z —These represent the apparent velocity increments output by the onboard inertial navigation system at the current moment, in meters per second; —These represent the eastward, northward, and celestial velocities before the recursive update, all in meters per second; —These are the recursively updated eastward, northward, and celestial velocities, all in meters per second; —These represent the drag accelerations in the north, sky, and east directions, respectively, in m / s². 2 ; —These represent the Coriolis accelerations in the north, sky, and east directions, respectively, in m / s². 2 g k —Current gravitational acceleration, in m / s² 2 ;

[0086] The direction cosine matrix represents the distance from the projectile's own system to the navigation system.

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

[0088]

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

[0090] —These represent the eastward, northward, and celestial speeds before the recursive update; —These represent the recursively updated eastward, northward, and celestial velocities, respectively; t h L is the sampling period of the inertial navigation system. k-1 ,λ k-1 ,h k-1 —These represent the geographic latitude, longitude, and altitude before the recursive update, in radians, radians, and meters, respectively; L k ,λ k ,h k —These are the recursively updated geographical latitude, longitude, and altitude, in radians, radians, and meters, respectively, serving as the primary 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] Where X is the filter state vector; φ x ,φ y ,φ z —These represent the attitude misalignment angles, in rad; δv E ,δv N ,δv U —These represent the velocity errors in the navigation system, in m / s; δL, δλ, δh —These represent the position errors, in rad, rad, and m, respectively; ε x ,ε y ,ε z —These represent the zero bias of the x, y, and z gyroscopes, in rad / s; δK gx ,δK gy ,δK gz —These represent the x, y, and z gyroscope scaling errors, in ppm; Δ x ,Δ y ,Δ z —Add zero offset to x, y, and z respectively, in m / s. 2 ;δK ax ,δK ay ,δK az —Add scale error to x, y, and z respectively, in ppm.

[0095] The equation for the filter time update is:

[0096]

[0097] Wherein, the state transition 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 from the previous step; A k / k-1 —The state transition matrix calculated based on the current satellite measurement pseudorange, pseudorange rate, satellite position and velocity, and the current position and velocity of the inertial navigation system; —A k / k-1 The device; Q k —Process noise variance matrix; X k / k-1 —Predicted state 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; —The projection of the rotational angular rate of the carrying system relative to the inertial frame onto the carrying system; (f n ×) — Vector f b The antisymmetric matrix projected in the navigation frame; (v n ×)—vector v n An antisymmetric matrix.

[0101] The above scheme utilizes Kalman filtering (KF) technology during the high-G flight phase of the aircraft. By modeling the pure inertial navigation error through linear system state equations, it fully leverages the strong observability of the pure inertial navigation heading error during the high-G phase of the aircraft and the 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 GNSS, includes:

[0103] Step S21: Based on GNSS, obtain the PDOP value and projectile overload change rate output by the GNSS receiver at the corresponding measurement time.

[0104] Step S22: Determine whether the PDOP value is ≤ a set threshold and the projectile overload change rate. Is it ≤ the threshold value?

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

[0106] If neither is true, the location information at the measurement time is discarded as bad observation information, and the process returns to step S1.

[0107] In one embodiment, the threshold value for the PDOP value is set to ≤6, and the threshold value for the overload change rate of the projectile is determined based on the maximum overload change rate that the GNSS satellite navigation receiver can accommodate.

[0108] The above scheme uses the PDOP value to correspond to the satellite configuration. By using the satellite configuration and the overload change rate of the projectile as indicators, when the PDOP value and the overload change rate of the projectile do not meet the set conditions, they are discarded as bad observation information to avoid their adverse effects on the system and improve the system's fault tolerance.

[0109] In one embodiment, step S3, constructing an observation measurement based on the first location information and the second location information, and performing measurement residual verification includes:

[0110] Step S31: Construct the observation Z based on the first location information and the second location information. k , expressed as:

[0111]

[0112] Step S32: Construct the observation matrix H k And matrix A,

[0113] Observation matrix H k Expressed as:

[0114] 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 verification:

[0118] Taking the square root of the diagonal elements of matrix A yields vector A. diagsqrt ;

[0119] Vector Z k -H k X k / k-1 The vector |Z is obtained by taking the absolute value of each element in the vector |Z|. k -H k X k / k-1 |, where X k / k-1 —Predicted state value;

[0120] Judgment vector |Z k -H k X k / k-1 Are all elements in vector A less than or equal to vector A? diagsqrt Three times the corresponding element;

[0121] If so, it is considered to have passed the measurement residual verification;

[0122] If not, it is considered that the measurement residual verification has failed.

[0123] The above scheme selects the position error term and discards the velocity error term, which is more sensitive to the high dynamic environment, when constructing the observation. The measurement residual is used as an indicator to eliminate bad observation information and avoid the adverse effects of the high dynamic environment of the aircraft's high overload period on the air alignment.

[0124] In one embodiment, step S5, correcting the attitude error of the SINS using the observations and obtaining the projectile attitude quaternion at the current moment, includes:

[0125] Step S51: Observe the Z k The data is fed into a Kalman filter for Kalman filtering measurement updates to obtain the estimated state X. k .

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

[0127]

[0128] Step S52: Estimate state X k The pure inertial navigation attitude quaternion q at the current moment SINS Perform output correction to obtain the projectile attitude quaternion q at the current moment.

[0129] In one embodiment, step S52 involves estimating state X. k The pure inertial navigation attitude quaternion q at the current moment SINS Perform output correction to obtain the projectile attitude quaternion q at the current moment, including:

[0130] Step S521: Calculate the corrected 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| represents the modulus of Φ;

[0133] Step S522: Adjust the quaternion q′ of the pure inertial navigation attitude at the current moment. SINS Perform output correction to obtain the projectile attitude quaternion q at the current moment.

[0134] The quaternion q of the pure inertial navigation at the current moment SINS Expressed as:

[0135] The projectile attitude quaternion q at the current moment is expressed as:

[0136] By using the above method, the online optimal estimation and correction of pure inertial navigation attitude error can be performed on the measurement after removing bad observation information, so as to obtain high-precision attitude information of the aircraft.

[0137] In the description of this application, it should be noted that the terms "upper," "lower," etc., indicating the orientation or positional relationship are based on the orientation or positional relationship shown in the accompanying drawings, and are only for the convenience of describing this application and simplifying the description, and do not indicate or imply that the device or element referred to must have a specific orientation, or be constructed and operated in a specific orientation, and therefore should not be construed as a limitation of this application. Unless otherwise expressly specified and limited, the terms "installed," "connected," and "linked" should be interpreted broadly. For example, they can refer to a fixed connection, a detachable connection, or an integral connection; they can refer to a mechanical connection or an electrical connection; they can refer to a direct connection or an indirect connection through an intermediate medium; they can refer to the internal communication between two elements. For those skilled in the art, the specific meaning of the above terms in this 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 used merely to distinguish one entity or operation from another, and do not necessarily require or imply any such actual relationship or order between these entities or operations. Furthermore, the terms "comprising," "including," or any other variations thereof are intended to cover non-exclusive inclusion, such that a process, method, article, or apparatus that comprises a list of elements includes not only those elements but also other elements not expressly listed, or elements inherent to such a process, method, article, or apparatus. Without further limitations, an element defined by the phrase "comprising one..." does not exclude the presence of other identical elements in the process, method, article, or apparatus that includes the element.

[0139] In the description of the embodiments of this application, unless otherwise stated, " / " means "or". For example, A / B can mean A or B. The "and / or" in the text is merely a description of the relationship between related objects, indicating that there can be three relationships. For example, A and / or B can mean: A exists alone, A and B exist simultaneously, and B exists alone. In addition, in the description of the embodiments of this application, "multiple" means two or more.

[0140] The above are merely specific embodiments of this application, enabling those skilled in the art to understand or implement this application. Various modifications to these embodiments will be readily apparent to those skilled in the art, and the general principles defined herein may be implemented in other embodiments without departing from the spirit or scope of this application. Therefore, this application is not to be limited to the embodiments shown herein, but is to be accorded the widest scope consistent with the principles and novel features claimed herein.

Claims

1. A SINS / GNSS air-to-air alignment method for aircraft during high-G phases, characterized in that, Includes the following steps: Step S1: Calculate the first position information of the aircraft at each measurement moment during the high G-force flight based on SINS. The first position information includes latitude. ,longitude ,high ; Step S2: Output the second position information of the aircraft at the corresponding measurement time based on GNSS, the second position information including latitude. ,longitude ,high ; Step S3: Construct an observation measurement based on the first and second location information, and perform measurement residual verification; If the measurement residual verification fails, proceed to step S4, discard the observation as bad observation information, and return to step S1; If the measurement residual verification is passed, proceed to step S5, correct the attitude error of the SINS using the observation, obtain the projectile attitude quaternion at the current moment, and complete the air alignment. Step S3, constructing the observation measurement based on the first and second location information and performing measurement residual verification, includes: Step S31: Construct an observation measurement based on the first location information and the second location information. , expressed as: ； Step S32: Construct the observation matrix sum matrix , Observation matrix Expressed as: ; matrix Expressed as: ; in, —State one-step prediction variance; Represents the observation noise matrix; Step S33: Perform measurement residual verification: matrix Taking the square root of the diagonal elements yields the vector. ; vector The vector is obtained by taking the absolute value of each element. ,in —Predicted state value; Judgment Vector Does every element in the vector not exceed the value of the vector? Three times the corresponding element; If so, it is considered to have passed the measurement residual verification; If not, it is considered that the measurement residual verification has failed.

2. The SINS / GNSS air-to-air alignment method for aircraft during high-G phases as described in claim 1, characterized in that, Step S1, calculating the first position information of the aircraft at each measurement moment during the high G-force flight based on SINS, includes: Step S11: Obtain the initial attitude, initial velocity, and initial position of the aircraft as initial navigation values ​​for SINS; Step S12: Collect the SINS data of the aircraft during the high overload flight process, perform pure inertial navigation calculation and Kalman filter time update to obtain the first position information at each measurement moment.

3. The SINS / GNSS air-to-air alignment method for aircraft during high-G phases as described in claim 2, characterized in that, The ground alignment result of the aircraft is used as the initial attitude of the aircraft.

4. The SINS / GNSS air-to-air alignment method for aircraft during high-G phases as described in claim 2, characterized in that, Step S12, collecting SINS data of the aircraft during high-G flight and performing pure inertial navigation calculation and Kalman filter time update, includes: Step S121: Collect SINS data to perform pure inertial navigation attitude update, velocity update and position update calculations. The SINS data includes the apparent velocity increment and angle increment output by the missile-borne inertial combination device. Step S122: Perform Kalman filter time update.

5. The SINS / GNSS air-to-air alignment method for aircraft during high-G phases as described in claim 4, characterized in that, The formula for calculating the position update of pure inertial navigation is as follows: ； in, Represents the principal radius of curvature of the Earth's geoid; Represents the radius of the principal curvature of the Earth's meridian; —These represent the eastward, northward, and celestial speeds before the recursive update; —These are the recursively updated eastward, northward, and celestial speeds, respectively; The sampling period of the inertial combination device; —These represent the geographic latitude, longitude, and altitude before the recursive update; —These are the recursively updated geographical latitude, longitude, and altitude, which serve as the primary location information.

6. The SINS / GNSS air-to-air alignment method for aircraft during high-G phases as described in claim 1, characterized in that, Step S2, outputting the second position information of the aircraft at the corresponding measurement time based on GNSS, includes: Step S21: Based on GNSS, obtain the PDOP value and projectile overload change rate output by the GNSS receiver at the corresponding measurement time. : Step S22: Determine whether the PDOP value is ≤ a set threshold and the projectile overload change rate. Is it ≤ the threshold value? If both are true, then output the position information at that measurement time as the second position information; If neither is true, the location information at the measurement time is discarded as bad observation information, and the process returns to step S1.

7. The SINS / GNSS air-to-air alignment method for aircraft during high-G phases as described in claim 6, characterized in that, The threshold value for PDOP is set to ≤6.

8. The SINS / GNSS air-to-air alignment method for aircraft during high-G phases as described in claim 1, characterized in that, Step S5, correcting the attitude error of the SINS using the observations, and obtaining the projectile attitude quaternion at the current moment includes: Step S51: Observe the measurement The data is fed into a Kalman filter for Kalman filtering measurement updates to obtain the estimated state. ; Step S52: Based on the estimated state The pure inertial navigation attitude quaternion at the current moment Perform output correction to obtain the projectile attitude quaternion at the current moment. .

9. The SINS / GNSS air-to-air alignment method for aircraft during high-G phases as described in claim 8, characterized in that, Step S52, through the estimated state The pure inertial navigation attitude quaternion at the current moment Perform output correction to obtain the projectile attitude quaternion at the current moment. include: Step S521: Calculate the corrected quaternion : The calculation formula is: ; in, The misalignment angle is expressed as the estimated state. The first three dimensions: ; for Modulus; Step S522: Correct the quaternion The pure inertial navigation attitude quaternion at the current moment Perform output correction to obtain the projectile attitude quaternion at the current moment. , Attention quaternion of pure inertial navigation at the current moment Expressed as: ; The quaternion of the projectile's attitude at the current moment Expressed as: .

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