A method and system for estimating projectile attitude error

Abstract

A method and system for estimating projectile attitude error includes: calculating the projectile's ground heading angle deviation YawErrGnd and ground pitch angle deviation PitErrGnd; and estimating the projectile's attitude error based on its real-time position information (Lon M ,Lat M Alt M ), projectile pitch angle P, projectile heading angle Y, projectile roll angle R, and calibration point position information (Lon T Lat T Alt T The theoretical frame azimuth angle (SeekerYawNav) and elevation angle (SeekerPitNav) of the seeker are calculated in reverse. These angles are then subtracted from the actual frame azimuth angle (SeekerYaw) and elevation angle (SeekerPit) of the seeker to obtain the missile's flight azimuth deviation (YawErrReal) and elevation deviation (PitErrReal), respectively. Kalman filtering is then performed using the missile's ground heading angle deviation (YawErrGnd) and ground elevation angle deviation (PitErrGnd), as well as the flight azimuth deviation (YawErrReal) and flight elevation angle deviation (PitErrReal), to obtain estimated values ​​for the missile's azimuth deviation (YawErrCom) and elevation deviation (PitErrCom). This invention utilizes the frame angle information of the missile-borne frame laser seeker to reduce the missile's navigation attitude error and improve navigation accuracy.

Description

Technical Field

[0001] This invention belongs to the field of precision guidance technology, and in particular relates to a method and system for estimating the attitude error of a projectile. Background Technology

[0002] With the rapid development of satellite navigation technology, it is appearing in precision-guided weapons and equipment across multiple fields with its low cost and high efficiency. Consequently, countermeasures against satellite navigation are becoming increasingly widespread, ranging from jamming to deception. As the instability of satellite navigation increases, there are also higher demands on missile-borne pure inertial navigation systems, especially for navigation accuracy requirements under long-endurance pure inertial conditions, which is causing a rapid increase in overall missile costs. Methods to improve the accuracy of pure inertial navigation mainly focus on improving initial attitude accuracy. Currently, the estimation and compensation of attitude errors are often hampered by issues such as low maneuverability of the carrier platform, low structural precision, and wing deformation during flight. How to utilize the frame angle information of the frame laser seeker to estimate and compensate for missile attitude errors is a worthy research direction. The missile-borne frame laser seeker can track a target point in both azimuth and elevation channels. Based on the positional relationship between the seeker and the target point, it outputs the frame azimuth and elevation angle information of the target point in real time, featuring high accuracy and low latency. It has already been applied to the guidance modification of various types of munitions both domestically and internationally. Summary of the Invention

[0003] This invention proposes a method and system for estimating projectile attitude error. When the projectile enters tracking mode, the azimuth and elevation angle information of the onboard laser-guided head is used to compensate for the projectile's attitude error. This allows for more accurate measurement of the projectile's attitude information and solves the problem of low navigation attitude accuracy in existing frame-guided laser weapons.

[0004] This invention proposes a method for estimating projectile attitude error, comprising:

[0005] Based on the pitch and heading angle information of the carrier aircraft and the pitch and heading angle information of the missile mounted on the carrier aircraft, calculate the ground heading angle deviation YawErrGn and the ground pitch angle deviation PitErrGn of the missile.

[0006] Based on the real-time position information of the missile during flight (Lon M ,Lat M Alt M The missile's pitch angle P, heading angle Y, roll angle R, actual frame azimuth angle SeekerYaw and actual frame pitch angle SeekerPit of the seeker, and ground calibration point location information (Lon T ,Lat T Alt T), calculate the flight azimuth deviation YawErrReal between the actual frame azimuth angle SeekerYaw and the theoretical frame azimuth angle SeekerYawNav, and the flight pitch deviation PitErrReal between the actual frame pitch angle SeekerPit and the theoretical frame pitch angle SeekerPitNav;

[0007] The ground heading angle deviation YawErrGnd and the ground pitch angle deviation PitErrGnd of the missile body are used as the output of the state equation of the Kalman filter. The flight azimuth angle deviation YawErrReal and the flight pitch angle deviation PitErrReal are used as the observation information of the Kalman filter equation to obtain the estimated values ​​of the missile body azimuth angle error YawErrCom and the estimated values ​​of the missile body pitch angle deviation PitErrCom.

[0008] Furthermore, the pitch and heading information of the carrier aircraft is provided by the carrier aircraft's navigation system, while the pitch and heading information of the missile mounted on the carrier aircraft is measured using a GNSS-based dual-antenna direction finding device.

[0009] Furthermore, the method for obtaining the actual frame azimuth angle (SeekerYaw) and actual frame pitch angle (SeekerPit) information of the seeker head is as follows: a laser illuminator is set within a specified distance range of the ground calibration point, and the angle between the line connecting the laser illuminator and the calibration point and the aircraft flight path is within a specified range. When the horizontal distance OffsetDis between the projectile and the ground calibration point is less than or equal to a specified threshold, the laser illuminator illuminates the ground calibration point, the seeker head enters the tracking mode, and the seeker head outputs the actual frame azimuth angle (SeekerYaw) and actual frame pitch angle (SeekerPit) information in real time.

[0010] Furthermore, after the seeker enters tracking mode, the following calculations are made: the flight azimuth deviation YawErrReal between the actual frame azimuth angle SeekerYaw and the theoretical frame azimuth angle SeekerYawNav, and the flight pitch deviation PitErrReal between the actual frame pitch angle SeekerPit and the theoretical frame pitch angle SeekerPitNav. Specifically, these calculations include:

[0011] The three-directional distance between the missile and the ground calibration point [X] n ,Y n Z n ] T From Geography to Missile System [X] b ,Y b Z b ]T :

[0012] Calculate the three-way distance between the missile and the ground calibration point:

[0013]

[0014] Among them, Lon M Lat M Alt M For the position information of the missile assembly navigation, Lon T Lat T Alt T For calibration point location information, R M R is the radius of the Earth's meridian. N Let X be the radius of the Earth's circumpolar orbit; then calculate [X]. b ,Y b Z b ] T ,as follows:

[0015]

[0016] in This is the attitude transfer matrix from the geographic system to the missile system, in the following form:

[0017]

[0018] Where P represents the projectile's pitch angle, Y represents the projectile's heading angle, and R represents the projectile's roll angle;

[0019] The formulas for calculating the azimuth angle (SeekerYawNav) and pitch angle (SeekerPitNav) of the theoretical frame of the seeker are as follows:

[0020]

[0021] SeekerPitNav=π / 2-arctan2(X b / Z b )

[0022] The azimuth deviation YawErrReal and pitch deviation PitErrReal are calculated as follows:

[0023] YawErrReal=SeekerYaw-SeekerYawNav

[0024] PitErr Real=SeekerPit-SeekerPitNav.

[0025] Furthermore, the method further includes compensating the estimated values ​​of the missile body azimuth angle error YawErrCom and the estimated values ​​of the missile body pitch angle deviation PitErrCom into the corresponding missile body attitude information.

[0026] This invention proposes a projectile attitude error estimation system, comprising:

[0027] The ground attitude deviation calculation module calculates the ground heading angle deviation YawErrGnd and the ground pitch angle deviation PitErrGnd of the missile based on the pitch angle and heading angle information of the carrier aircraft and the pitch angle and heading angle information of the missile mounted on the carrier aircraft.

[0028] The flight attitude deviation calculation module calculates the deviation based on the real-time position information (Lon) of the missile during flight. M ,Lat M Alt M The missile's pitch angle P, heading angle Y, roll angle R, actual frame azimuth angle SeekerYaw and actual frame pitch angle SeekerPit of the seeker, and ground calibration point location information (Lon T ,Lat T Alt T ), calculate the flight azimuth deviation YawErrReal between the actual frame azimuth angle SeekerYaw and the theoretical frame azimuth angle SeekerYawNav, and the flight pitch deviation PitErrReal between the actual frame pitch angle SeekerPit and the theoretical frame pitch angle SeekerPitNav;

[0029] The attitude deviation estimation module uses the ground heading angle deviation YawErrGnd and the ground pitch angle deviation PitErrGnd of the missile body as the output of the state equation of the Kalman filter, and uses the flight azimuth angle deviation YawErrReal and the flight pitch angle deviation PitErrReal of the flight as the observation information of the Kalman filter equation to perform Kalman filtering, and obtains the estimated values ​​of the missile body azimuth angle deviation YawErrCom and the estimated values ​​of the missile body pitch angle deviation PitErrCom.

[0030] Furthermore, the pitch and heading information of the carrier aircraft is provided by the carrier aircraft's navigation system, while the pitch and heading information of the missile mounted on the carrier aircraft is measured using a GNSS-based dual-antenna direction finding device.

[0031] Furthermore, the method for obtaining the actual frame azimuth angle (SeekerYaw) and actual frame pitch angle (SeekerPit) information of the seeker head is as follows: a laser illuminator is set within a specified distance range of the ground calibration point, and the angle between the line connecting the laser illuminator and the calibration point and the aircraft flight path is within a specified range. When the horizontal distance OffsetDis between the projectile and the ground calibration point is less than or equal to a specified threshold, the laser illuminator illuminates the ground calibration point, the seeker head enters the tracking mode, and the seeker head outputs the actual frame azimuth angle (SeekerYaw) and actual frame pitch angle (SeekerPit) information in real time.

[0032] Furthermore, after the seeker enters tracking mode, the following calculations are made: the flight azimuth deviation YawErrReal between the actual frame azimuth angle SeekerYaw and the theoretical frame azimuth angle SeekerYawNav, and the flight pitch deviation PitErrReal between the actual frame pitch angle SeekerPit and the theoretical frame pitch angle SeekerPitNav. Specifically, these calculations include:

[0033] The three-directional distance between the missile and the ground calibration point [X] n ,Y n Z n ] T From Geography to Missile System [X] b ,Y b Z b ] T :

[0034] Calculate the three-way distance between the missile and the ground calibration point:

[0035]

[0036] Among them, Lon M Lat M Alt M For the position information of the missile assembly navigation, Lon T Lat T Alt T For calibration point location information, R M R is the radius of the Earth's meridian. N Let X be the radius of the Earth's circumpolar orbit; then calculate [X]. b ,Y b Z b ] T ,as follows:

[0037]

[0038] in This is the attitude transfer matrix from the geographic system to the missile system, in the following form:

[0039]

[0040] Where P represents the projectile's pitch angle, Y represents the projectile's heading angle, and R represents the projectile's roll angle;

[0041] The formulas for calculating the azimuth angle (SeekerYawNav) and pitch angle (SeekerPitNav) of the theoretical frame of the seeker are as follows:

[0042]

[0043] SeekerPitNav=π / 2-arctan2(X b / Z b )

[0044] The azimuth deviation YawErrReal and pitch deviation PitErrReal are calculated as follows:

[0045] YawErrReal=SeekerYaw-SeekerYawNav

[0046] PitErr Real=SeekerPit-SeekerPitNav.

[0047] Furthermore, the system further includes an error compensation module that compensates the estimated values ​​of the missile body azimuth angle error YawErrCom and the estimated values ​​of the missile body pitch angle deviation PitErrCom into the missile body's attitude information.

[0048] This invention aims to reduce the navigation attitude error of the missile body by utilizing the frame angle information of the missile-borne frame laser seeker, thereby improving the navigation accuracy of the missile body. This is beneficial for pure inertial navigation laser terminal guidance weapons to quickly and accurately acquire targets and improve the effectiveness of laser-guided weapon munitions. Attached Figure Description

[0049] Figure 1 This is a flowchart of the steps of a projectile attitude error estimation method based on a missile-borne frame laser seeker in an embodiment of the present invention;

[0050] Figure 2 This refers to the position and angle relationship information between the missile and the calibration point in this embodiment of the invention;

[0051] Figure 3 This is a schematic diagram of the flight path of the scheme in the embodiment of the present invention;

[0052] Figure 4 This invention provides measurement and filtering deviation information. Detailed Implementation

[0053] The specific implementation methods are described in detail below with reference to the accompanying drawings. This invention proposes a method and system for estimating projectile attitude error.

[0054] The overall steps of the projectile attitude error estimation method of the present invention include:

[0055] S1. Initialization Steps: Pre-set the ground calibration points and calibration flight path according to the seeker's performance parameters. See [link / reference] Figure 3 This facilitates the seeker to quickly and accurately capture the laser signal at the calibration point during flight.

[0056] The attitude error information of the projectile's heading angle and pitch angle was collected on the ground. The average value of multiple collections was taken to obtain the ground attitude deviation information YawErrGnd and PitErrGnd of the projectile.

[0057] In the missile attitude error estimation method based on the missile-borne frame laser seeker, the coordinates of the ground calibration points need to be obtained in advance. The selection of the calibration flight path needs to be calculated based on the seeker's frame azimuth range [LYawLim, RYawLim], frame elevation range [LPitLim, HPitLim], and the seeker's effective recognition distance EffDis.

[0058] WGS-84 coordinates of ground calibration points (Lon T ,Lat T Alt T Data can be collected in advance on the ground or using a high-precision airborne pod.

[0059] The calibration line is generally required to be directly above the calibration point, and the line height (HairLine) should meet the following formula.

[0060] HairLine≤EffDis×sin(LPitLim)···(1)

[0061] The horizontal distance between the starting point of the flight path and the calibration point, StarDis, satisfies the following formula.

[0062] StarDis≥EffDis×cos(LPitLim)···(2)

[0063] The horizontal offset OffsetDis of the flight path satisfies the following formula.

[0064] OffsetDis≤EffDis×atan(min[LYawLim,RYawLim])···(3)

[0065] S2. Ground Projectile Attitude Deviation Calculation Steps: Based on the pitch and yaw angle information of the carrier aircraft and the pitch and yaw angle information of the projectile mounted on the carrier aircraft, calculate the ground yaw angle deviation YawErrGnd and the ground pitch angle deviation PitErrGnd of the projectile. In specific implementation, a GNSS-based dual-antenna lateral device can be used to measure the pitch and yaw angle information of the projectile when it is mounted on the carrier aircraft, and obtain the pitch and yaw angle information of the carrier aircraft at the same moment. Using these two information, the ground yaw angle deviation YawErrGnd and the ground pitch angle deviation PitErrGnd of the projectile can be calculated. The above angle information can be used as the state equation output of the Kalman filter.

[0066] S3. Calculation steps for missile attitude deviation: After the seeker enters tracking mode, based on the missile's real-time position information (Lon... M ,Lat M Alt M ), missile pitch angle P, missile heading angle Y, missile roll angle R, and ground calibration point location information (Lon T ,Lat T Alt T The theoretical frame azimuth angle SeekerYawNav and theoretical frame pitch angle SeekerPitNav of the seeker are calculated in reverse. The difference between the above angles and the actual frame azimuth angle SeekerYaw and frame elevation angle SeekerPit output by the seeker is calculated to calculate the flight azimuth angle deviation YawErrReal and the flight pitch angle deviation PitErrReal. The above angle information is used as the observation information for Kalman filtering.

[0067] Specifically, a laser illuminator is pre-installed on the ground. The distance between the laser illuminator and the ground calibration point is no more than 2km, and the angle between the laser illuminator and the flight path is between 30° and 60°. Once the horizontal distance OffsetDis between the missile and the ground calibration point meets the design requirements, the laser illuminator continuously illuminates the ground calibration point. After the seeker acquires the target and enters the tracking mode, it will output the frame azimuth angle SeekerYaw and the frame pitch angle SeekerPit in real time.

[0068] When the horizontal distance OffsetDis between the missile body and the ground calibration point is greater than a specified threshold, the missile enters cruise mode. Once the seeker enters tracking mode, it utilizes the missile body's integrated navigation position information (Lon...). M ,Lat M Alt M ), calibration point location information (Lon) T ,Lat T Alt TThe projectile attitude information (heading angle Y, pitch angle P, roll angle R) and seeker frame angle information (SeekerYaw and SeekerPit) are used to calculate the projectile attitude error information YawErrReal and PitErrReal in real time.

[0069] The three-directional distance between the bullet and the calibration point [X] n ,Y n Z n ] T From Geography to Missile System [X] b ,Y b Z b ] T The calculation formula is as follows:

[0070] Calculate the distance between the projectile and the calibration point in three directions:

[0071]

[0072] Among them, R M R is the radius of the Earth's meridian. N Let X be the radius of the Earth's circumpolar orbit. Then calculate [X]. b ,Y b Z b ] T ,as follows:

[0073]

[0074] in This is the attitude transfer matrix from the geographic system to the missile system, in the following form:

[0075]

[0076] Where P represents the projectile's pitch angle, Y represents the projectile's heading angle, and R represents the projectile's roll angle.

[0077] At this point, the azimuth angle SeekerYawNav and the elevation angle SeekerPitNav of the theoretical frame of the seeker can be calculated based on the positional relationship of the missile system. The formulas are as follows:

[0078]

[0079] SeekerPitNav=Pi / 2-atan2(X b / Z b (8)

[0080] The theoretical azimuth and pitch angles of the seeker head calculated in reverse differ from the actual azimuth and pitch angles of the seeker head frame. These are the YawErrReal and PitErrReal that we require. The calculation formulas are as follows:

[0081] YawErrReal=SeekerYaw-SeekerYawNav···(9)

[0082] PitErr Real=SeekerPit-SeekerPitNav···(10).

[0083] S4. Error Estimation Step: Kalman filtering is performed on the ground heading angle deviation YawErrGnd and ground pitch angle deviation PitErrGn, as well as the launch azimuth angle deviation YawErrReal and launch pitch angle deviation PitErrReal, to obtain the estimated values ​​of the launch azimuth angle error YawErrCom and the estimated value of the launch pitch angle error PitErrCom. Furthermore, the estimated values ​​of the launch attitude deviations can be used to compensate for the launch attitude information in real time.

[0084] The present invention also proposes a projectile attitude error estimation system, including: a ground projectile attitude deviation calculation module, a flight projectile attitude deviation calculation module, and a projectile attitude error estimation module.

[0085] The ground-based projectile attitude deviation calculation module calculates the ground heading angle deviation YawErrGnd and the ground pitch angle deviation PitErrGnd of the projectile based on the pitch and heading angle information of the carrier aircraft and the pitch and heading angle information of the projectile mounted on the carrier aircraft. The pitch and heading angle information of the carrier aircraft and the pitch and heading angle information of the projectile mounted on the carrier aircraft are obtained by measuring them separately using a GNSS-based dual-antenna direction finding device.

[0086] The missile body attitude deviation calculation module is attached, based on the missile body's real-time position information (London). M ,Lat M Alt M The missile's pitch angle P, heading angle Y, roll angle R, seeker's frame azimuth angle SeekerYaw and pitch angle SeekerPit, and ground calibration point location information (Lon T ,Lat T Alt T ), calculate the flight azimuth deviation YawErrReal between the theoretical frame azimuth angle and the actual frame azimuth angle of the seeker, and the flight pitch deviation PitErrReal between the theoretical pitch angle and the actual frame pitch angle of the seeker.

[0087] The method for obtaining the azimuth angle (SeekerYaw) and elevation angle (SeekerPit) information of the frame laser seeker is as follows: A laser illuminator is set up within a specified distance range of the ground calibration point. The angle between the line connecting the laser illuminator and the calibration point and the aircraft flight path is within a specified range. When the horizontal distance OffsetDis between the projectile and the ground calibration point is less than a specified threshold, the laser illuminator illuminates the ground calibration point, and the seeker enters the tracking mode. The seeker outputs the frame azimuth angle (SeekerYaw) and elevation angle (SeekerPit) information of the calibration point in real time.

[0088] The three-directional distances of the "missile-calibration point" [X] n ,Y n Z n ] T From Geography to Missile System [X] b ,Y b Z b ] T :

[0089] Calculate the three-directional distances between the missile and the calibration point:

[0090]

[0091] Among them, Lon M Lat M Alt M For the position information of the missile assembly navigation, Lon T Lat T Alt T For calibration point location information, R M R is the radius of the Earth's meridian. N Let X be the radius of the Earth's circumpolar orbit; then calculate [X]. b ,Y b Z b ] T ,as follows:

[0092]

[0093] in This is the attitude transfer matrix from the geographic system to the missile system, in the following form:

[0094]

[0095] Where P represents the projectile's pitch angle, Y represents the projectile's heading angle, and R represents the projectile's roll angle;

[0096] The formulas for calculating the seeker frame azimuth angle (SeekerYawNav) and seeker frame pitch angle (SeekerPitNav) are as follows:

[0097]

[0098] SeekerPitNav=π / 2-arctan2(X b / Z b )

[0099] The azimuth deviation YawErrReal and pitch deviation PitErrReal are calculated as follows:

[0100] YawErrReal=SeekerYaw-SeekerYawNav

[0101] PitErrReal=SeekerPit-SeekerPitNav.

[0102] The error estimation module performs Kalman filtering on the ground heading angle deviation YawErrGnd and ground pitch angle deviation PitErrGnd of the missile body, as well as the flight azimuth angle deviation YawErrReal and flight pitch angle deviation PitErrReal, to obtain the estimated values ​​of the missile body azimuth angle error YawErrCom and the estimated values ​​of the missile body pitch angle deviation PitErrCom.

[0103] The missile attitude error estimation system of the present invention further includes an error compensation module, which compensates the estimated value of the missile azimuth angle error YawErrCom and the estimated value of the missile pitch angle deviation PitErrCom into the missile attitude information.

[0104] Example

[0105] To make the objectives, technical solutions, and advantages of the present invention clearer, the embodiments disclosed in the present invention will be described in further detail below with reference to the accompanying drawings.

[0106] like Figure 1 In this embodiment, the projectile attitude error estimation method based on the missile-borne frame laser seeker includes:

[0107] Step 101: Pre-set the ground calibration point and calibration route according to the seeker performance parameters.

[0108] In this embodiment, the position coordinates of the ground calibration points can be acquired from the ground and uploaded to the missile's onboard navigation system in real time via a data link. Both the calibration point coordinates and the aircraft's position information use position information in the WGS-84 coordinate system. Other methods can also be used to obtain the coordinates, provided that the overall accuracy error of the calibration point coordinates is less than 50 meters.

[0109] The flight path altitude should generally not exceed 1 / 2 of the Hairline; the horizontal offset distance should generally not exceed 2 / 3 of the OffsetDis.

[0110] Step 102: Ground-based acquisition of projectile attitude deviation information YawErrGnd and PitErrGnd.

[0111] In this embodiment, a GNSS-based dual-antenna lateral device is used on the ground to measure the pitch and heading angles of the missile when it is mounted on the carrier aircraft, and to obtain the pitch and heading angles of the carrier aircraft at the same time. The heading angle deviation YawErrGnd and pitch angle deviation PitErrGnd of the missile are calculated using the two. Generally, no less than 7 data collections are required, and the mean and standard deviation of the measurement results are statistically analyzed.

[0112] Step 103, Ground illumination calibration point, calculate the missile attitude error information YawErrReal and PitErrReal based on the navigation information and seeker frame angle information during the missile's flight process.

[0113] In this embodiment, the location of the laser illumination point should be rationally planned. The location of the ground laser illumination point should comprehensively consider the actual weather and the power of the illuminator. Generally, the distance between the illuminator and the calibration point should not exceed 2km, and the angle between the line connecting the illuminator and the calibration flight path should be between 30° and 60°. Command control is used to control the laser illuminator to illuminate the calibration point. To facilitate target acquisition by the missile-borne frame laser seeker, a target with a specific reflection coefficient can be set up at the calibration point. During flight, the seeker frame angle information and navigation information are acquired at fixed intervals using TiCom. Generally, in this embodiment, TiCom is taken as 200ms. Navigation position and angle information are expressed in radians.

[0114] Step 104 uses a Kalman filter to filter the attitude error information described above. The filtering step size is 1 second, and the attitude angle error information from the above steps is filtered in real time to obtain the estimated values ​​of the projectile attitude error, YawErrCom and PitErrCom.

[0115] Step 105: Obtain the real-time attitude error information YawErrCom and PitErrCom after filtering.

[0116] Step 106 compensates the filtered and updated YawErrCom and PitErrCom to the projectile attitude information. The specific compensation strategy needs to be considered comprehensively based on the time and location; generally, dynamic constraints are applied to the attitude error information YawErrCom and PitErrCom. Figure 4 As can be seen, the error value after filtering is more accurate and changes more smoothly.

[0117] As the system implementation is in contrast to the method implementation, it is described in a simpler way. For relevant details, please refer to the description in the method implementation section.

[0118] Although the present invention has been disclosed above with reference to preferred embodiments, it is not intended to limit the present invention. Any person skilled in the art can make possible changes and modifications to the technical solutions of the present invention by utilizing the methods and techniques disclosed above without departing from the spirit and scope of the present invention. Therefore, any simple modifications, equivalent changes and alterations made to the above embodiments based on the technical essence of the present invention without departing from the content of the technical solutions of the present invention shall fall within the protection scope of the technical solutions of the present invention.

[0119] The contents not described in detail in this specification are common knowledge to those skilled in the art.

Claims

1. A method for estimating projectile attitude error, characterized in that, include: Based on the pitch and heading angle information of the carrier aircraft and the pitch and heading angle information of the missile mounted on the carrier aircraft, calculate the ground heading angle deviation YawErrGnd and the ground pitch angle deviation PitErrGnd of the missile. Based on the real-time position information of the missile during flight (Lon M ,Lat M Alt M The missile's pitch angle P, heading angle Y, roll angle R, actual frame azimuth angle SeekerYaw and actual frame pitch angle SeekerPit information, and ground calibration point location information (Lon T ,Lat T Alt T ), calculate the flight azimuth deviation YawErrReal between the actual frame azimuth angle Seeker Yaw and the theoretical frame azimuth angle Seeker YawNav, and the flight pitch deviation PitErrReal between the actual frame pitch angle SeekerPit and the theoretical frame pitch angle Seeker PitNav; The ground heading angle deviation YawErrGnd and the ground pitch angle deviation PitErrGnd of the projectile are used as the output of the state equation of the Kalman filter. The flight azimuth angle deviation YawErr Real and the flight pitch angle deviation PitErr Real are used as the observation information of the Kalman filter equation to obtain the estimated value of the projectile azimuth angle error YawErrCom and the estimated value of the projectile pitch angle deviation PitErrCom. The method for obtaining the actual frame azimuth angle (Seeker Yaw) and actual frame pitch angle (Seeker Pit) information of the seeker head is as follows: a laser illuminator is set within a specified distance range of the ground calibration point, and the angle between the line connecting the laser illuminator and the calibration point and the aircraft flight path is within a specified range. When the horizontal distance OffsetDis between the projectile and the ground calibration point is less than or equal to a specified threshold, the laser illuminator illuminates the ground calibration point, the seeker head enters the tracking mode, and the seeker head outputs the actual frame azimuth angle (Seeker Yaw) and actual frame pitch angle (Seeker Pit) information in real time. Once the seeker enters tracking mode, the following calculations are made: the flight azimuth deviation YawErrReal between the actual frame azimuth angle Seeker Yaw and the theoretical frame azimuth angle Seeker YawNav, and the flight pitch deviation PitErrReal between the actual frame pitch angle Seeker Pit and the theoretical frame pitch angle Seeker PitNav. Specifically, these calculations include: The three-directional distance between the missile and the ground calibration point [X] n ,Y n Z n ] T From Geography to Missile System [X] b ,Y b Z b ] T : Calculate the three-way distance between the missile and the ground calibration point: Among them, Lon M Lat M Alt M For the position information of the missile assembly navigation, Lon T Lat T Alt T For calibration point location information, R M R is the radius of the Earth's meridian. N Let X be the radius of the Earth's circumpolar orbit; then calculate [X]. b ,Y b Z b ] T ,as follows: in This is the attitude transfer matrix from the geographic system to the missile system, in the following form: Where P represents the projectile's pitch angle, Y represents the projectile's heading angle, and R represents the projectile's roll angle; The theoretical frame azimuth angle Seeker YawNav and the theoretical frame pitch angle SeekerPitNav of the seeker are calculated using the following formulas: SeekerPitNav=π / 2-arctan2(X b / Z b ) The azimuth deviation (YawErr Real) and pitch deviation (PitErr Real) are calculated as follows: YawErrReal=SeekerYaw-SeekerYawNav PitErr Real=SeekerPit-SeekerPitNav.

2. The projectile attitude error estimation method according to claim 1, characterized in that, The aircraft's pitch and heading angle information is provided by the aircraft's navigation system, while the pitch and heading angle information of the missile mounted on the aircraft is measured using a GNSS-based dual-antenna direction finding device.

3. The projectile attitude error estimation method according to claim 1, characterized in that, The method further includes compensating the estimated values ​​of the missile body azimuth error YawErrCom and the estimated values ​​of the missile body pitch deviation PitErrCom into the corresponding missile body attitude information.

4. A projectile attitude error estimation system, characterized in that, include: The ground attitude deviation calculation module calculates the ground heading angle deviation YawErrGnd and the ground pitch angle deviation PitErrGnd of the missile based on the pitch angle and heading angle information of the carrier aircraft and the pitch angle and heading angle information of the missile mounted on the carrier aircraft. The flight attitude deviation calculation module calculates the deviation based on the real-time position information (Lon) of the missile during flight. M ,Lat M Alt M The missile's pitch angle P, heading angle Y, roll angle R, actual frame azimuth angle SeekerYaw and actual frame pitch angle SeekerPit information, and ground calibration point location information (Lon T ,Lat T Alt T ), calculate the flight azimuth deviation YawErrReal between the actual frame azimuth angle See ker Yaw and the theoretical frame azimuth angle See ker YawNav, and the flight pitch deviation PitErrReal between the actual frame pitch angle See ker Pit and the theoretical frame pitch angle See ker PitNav; The attitude deviation estimation module uses the ground heading angle deviation YawErrGnd and the ground pitch angle deviation PitErrGnd of the missile body as the output of the state equation of the Kalman filter, and uses the flight azimuth angle deviation YawErr Real and the flight pitch angle deviation PitErr Real as the observation information of the Kalman filter equation to perform Kalman filtering, and obtains the estimated value of the missile body azimuth angle deviation YawErrCom and the estimated value of the missile body pitch angle deviation PitErrCom. The method for obtaining the actual frame azimuth angle (Seeker Yaw) and actual frame pitch angle (Seeker Pit) information of the seeker head is as follows: a laser illuminator is set within a specified distance range of the ground calibration point, and the angle between the line connecting the laser illuminator and the calibration point and the aircraft flight path is within a specified range. When the horizontal distance OffsetDis between the projectile and the ground calibration point is less than or equal to a specified threshold, the laser illuminator illuminates the ground calibration point, the seeker head enters the tracking mode, and the seeker head outputs the actual frame azimuth angle (Seeker Yaw) and actual frame pitch angle (Seeker Pit) information in real time. Once the seeker enters tracking mode, the following calculations are made: the flight azimuth deviation YawErrReal between the actual frame azimuth angle Seeker Yaw and the theoretical frame azimuth angle Seeker YawNav, and the flight pitch deviation PitErrReal between the actual frame pitch angle Seeker Pit and the theoretical frame pitch angle Seeker PitNav. Specifically, these calculations include: The three-directional distance between the missile and the ground calibration point [X] n ,Y n Z n ] T From Geography to Missile System [X] b ,Y b Z b ] T : Calculate the three-way distance between the missile and the ground calibration point: Among them, Lon M Lat M Alt M For the position information of the missile assembly navigation, Lon T Lat T Alt T For calibration point location information, R M R is the radius of the Earth's meridian. N Let X be the radius of the Earth's circumpolar orbit; then calculate [X]. b ,Y b Z b ] T ,as follows: in This is the attitude transfer matrix from the geographic system to the missile system, in the following form: Where P represents the projectile's pitch angle, Y represents the projectile's heading angle, and R represents the projectile's roll angle; The theoretical frame azimuth angle Seeker YawNav and the theoretical frame pitch angle SeekerPitNav of the seeker are calculated using the following formulas: See ker PitNav=π / 2-arctan2(X b / WITH b ) The azimuth deviation (YawErr Real) and pitch deviation (PitErr Real) are calculated as follows: YawErr Real=See ker Yaw-See ker YawNav PitErr Real=See ker Pit-See ker PitNav.

5. The projectile attitude error estimation system according to claim 4, characterized in that, The aircraft's pitch and heading angle information is provided by the aircraft's navigation system, while the pitch and heading angle information of the missile mounted on the aircraft is measured using a GNSS-based dual-antenna direction finding device.

6. The projectile attitude error estimation system according to claim 4, characterized in that, The system further includes an error compensation module that compensates the estimated values ​​of the missile body azimuth error YawErrCom and the estimated values ​​of the missile body pitch angle deviation PitErrCom into the missile body's attitude information.

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