CN115453600A - Flight state identification method and system based on airborne double-antenna GNSS and MINS combined navigation - Google Patents

Abstract

The invention relates to the technical field of navigation, in particular to a flight state identification method based on airborne double-antenna GNSS and MINS combined navigation

Judging that the body is in a constant-speed cruising state or a ground static state; when the temperature is higher than the set temperature

Computing the output of a horizontal two-axis accelerometer

To make a specific flight status determination. The method can judge the flight state of the aircraft body according to the resolving speed of the GNSS data, and judge whether to correct the roll angle based on different conditions after judging the flight state of the aircraft. The invention also discloses a flight state identification system based on the airborne double-antenna GNSS and MINS combined navigation.

Description

Flight status identification method based on airborne dual-antenna GNSS and MINS integrated navigation and its system

technical field

The invention relates to the technical field of navigation, in particular to a flight state identification method and system based on integrated navigation of airborne dual-antenna GNSS and MINS.

Background technique

The basic principle of integrated navigation is to use information fusion technology to combine various navigation systems such as radio, satellite, astronomy, terrain and scene matching through signal processing methods such as optimal estimation and digital filtering. In order to give full play to the advantages of various navigation technologies, it can achieve higher navigation accuracy and reliability than any single navigation method.

Compared with inertial navigation, GNSS has the advantages of low cost, high navigation accuracy, and the error does not accumulate over time. The navigation information output by the GNSS navigation system is used as the observation of the system state, and the state of the system (position, speed, etc.) ) and the error are optimally estimated to realize the calibration and error compensation of the inertial navigation system. The advantages of inertial navigation system, such as autonomy, real-time, and continuous, can make up for the shortcomings of GNSS that are susceptible to interference and poor reliability in dynamic environments.

In different landform environments, the body will face various environmental interferences, and there may be situations where satellite signals fail. At this time, the integrated navigation system loses the update of satellite measurement signals, and the attitude may gradually drift, resulting in error angles. Gradually increasing will affect the attitude accuracy of the navigation system. To this end, it is necessary to identify the state of the body, and to use which method of error compensation based on different states.

Contents of the invention

One of the purposes of the present invention is to provide a flight state identification method based on the integrated navigation of airborne dual-antenna GNSS and MINS, which can judge the state of the airframe according to the calculation speed of GNSS data, and after judging the state of the airframe, based on different Status, to determine whether to correct the roll angle.

A flight status identification method based on airborne dual-antenna GNSS and MINS integrated navigation, including the following:

时，判定机体处于地面准备态，此时采用GNSS数据，包括航向、俯仰角、速度和位置进行误差补偿修正，同时利用加速度计进行横滚角的修正；当

时，判定机体为运动状态，此时使用GNSS数据，包括航向、俯仰角、速度、位置进行误差补偿修正，横滚角不进行误差补偿修正，其中，v_ε为速度阈值参数，依据GNSS的速度量测噪声进行取值。Analyze the data sent by GNSS to judge the validity of the data. If the GNSS data is valid, the calculation speed of the GNSS data will be used as the basis for judging the flight status of the aircraft.

, it is determined that the airframe is in the ground preparation state. At this time, GNSS data, including heading, pitch angle, speed and position, are used for error compensation and correction, and the accelerometer is used to correct the roll angle; when

, it is determined that the body is in motion. At this time, GNSS data is used, including heading, pitch angle, speed, and position for error compensation correction, and roll angle is not for error compensation correction. Among them, v _ε is the speed threshold parameter, according to the GNSS speed Measure the noise to take the value.

时，判定机体处于地面准备态，此时采用GNSS数据，包括航向、俯仰角、速度和位置进行误差补偿修正，同时利用加速度计进行横滚角的修正，由于GNSS只能提供航向和俯仰角，在进行飞行姿态识别并确认当前状态可以使用加速度计进行误差补偿时，则可使用加速度计获得横滚角量测值；当

时，判定机体为运动状态，此时使用GNSS数据，包括航向、俯仰角、速度、位置进行误差补偿修正，横滚角不进行误差补偿修正，将横滚角量测误差设为零。The beneficial effect of the present invention is that: the present invention can judge the flight state of the airframe according to the calculation speed of the GNSS data when the GNSS data is valid, specifically when

At this time, it is determined that the airframe is in the ground preparation state. At this time, GNSS data, including heading, pitch angle, speed and position, are used for error compensation and correction. At the same time, the accelerometer is used to correct the roll angle. Since GNSS can only provide heading and pitch angle, When performing flight attitude recognition and confirming that the current state can use the accelerometer for error compensation, the accelerometer can be used to obtain the roll angle measurement value; when

, it is determined that the body is in motion. At this time, the GNSS data, including heading, pitch angle, speed, and position, are used for error compensation correction. The roll angle does not perform error compensation correction, and the roll angle measurement error is set to zero.

The present invention is based on the flight state identification method of airborne dual-antenna GNSS and MINS integrated navigation, also includes:

When the GNSS data is invalid, use the accelerometer to identify the flight state, and calculate the accelerometer output at time k according to the following formula

时，判定机体处于稳定态，此时的机体为匀速巡航状态或者地面静止状态，此阶段使用加速度计进行姿态误差补偿修正，其中，

为阈值参数，其取值视导航系统所在环境的具体噪声水平设定；当

时，计算水平双轴加速度计的输出

来进行具体飞行状态判定，判定方法如下：when

When , it is determined that the body is in a stable state. At this time, the body is in a constant speed cruising state or a static state on the ground. At this stage, the accelerometer is used to correct the attitude error compensation. Among them,

is the threshold parameter, and its value depends on the specific noise level of the environment where the navigation system is located; when

When , calculate the output of the horizontal two-axis accelerometer

To determine the specific flight status, the determination method is as follows:

时，与陀螺仪数据进行结合判定，机体处于转弯状态或者盘旋状态，其中，

为阈值参数，依据水平加速度计的噪声具体取值，when

When , combined with the gyroscope data to determine, the body is in a turning state or a hovering state, wherein,

is the threshold parameter, according to the specific value of the noise of the horizontal accelerometer,

时，机体的机动状态大，处于起飞或者降落阶段。

When , the maneuvering state of the airframe is large, and it is in the stage of take-off or landing.

The beneficial effect is that: through the method, when the GNSS signal is valid, the accelerometer and the gyroscope can also be used to effectively identify the flight state of the airframe.

The present invention is based on the flight state identification method of airborne dual-antenna GNSS and MINS integrated navigation, also includes:

In the static alignment phase of the airframe, the following formulas are used to obtain the local reference gravity acceleration

以便用于与加速度计输出

作比较，从而识别机体当前状态。The beneficial effect is that the local reference gravitational acceleration is calculated in advance in the static alignment phase of the airframe

for use with the accelerometer output

For comparison, so as to identify the current state of the body.

The present invention is based on the flight state identification method of airborne dual-antenna GNSS and MINS integrated navigation, and the validity judgment method of GNSS data is:

The calculation error of GNSS data is judged by three precision factors, which are PDOP, VDOP and HDOP respectively, PDOP ² =HDOP ² +VDOP ² , where PDOP represents the square sum of standard deviations between longitude, latitude and height. The root sign, VDOP means the square root of the standard deviation between longitude and latitude, and HDOP means the standard deviation of the height. When any precision factor is greater than 3, it is judged that the GNSS data is invalid.

The beneficial effect is that the dual-antenna GNSS is based on the single-antenna GNSS, with the antenna 1 as the reference station and the antenna 2 as the terminal station, the coordinates of the two antennas are obtained through the differential positioning method and the baseline vector between the two antennas is obtained, and then the heading is completed. With the measurement of pitch angle, the calculation error of GNSS data can be judged by the above three precision factors. When any precision factor is greater than 3, the accuracy of GNSS data does not meet the fusion requirements of integrated navigation. At this time, GNSS data will be invalid and unavailable .

The present invention is based on the flight state identification method of the airborne dual-antenna GNSS and MINS integrated navigation, which also includes predicting the duration of the GNSS data invalid phase under the current navigation route, and if the predicted duration is greater than the set threshold, then sending out the first prompt message.

The beneficial effect is that: if the duration of the GNSS data invalid stage under the current navigation route is predicted to be greater than the set threshold, it means that under the current navigation route, the body will not receive GNSS signals for a long time or the calculated GNSS data will be invalid. In this case, only the data collected by the MINS navigation system is used for attitude calculation and error compensation. Because the MINS navigation system has insufficient error accumulation over time, there may be a problem of large errors, resulting in a decrease in navigation accuracy. Therefore, The first prompt message needs to be issued.

The flight state identification method based on the airborne dual-antenna GNSS and MINS integrated navigation of the present invention also includes calculating the duration of the GNSS data invalid phase, and when the calculated duration is greater than the set threshold, a second prompt message is issued.

The beneficial effect is: if the calculated duration of the GNSS data invalid phase under the current navigation route is greater than the set threshold, it means that under the current navigation route, the body has not been able to receive GNSS signals for a long time or the calculated GNSS data is invalid Not available. In this case, only the data collected by the MINS navigation system is used for attitude calculation and error compensation. Because the MINS navigation system has insufficient error accumulation over time, there may be a problem of large errors, resulting in reduced navigation accuracy. Therefore, it is necessary to issue a second prompt message.

时，判定机体处于地面准备态，此时采用GNSS数据，包括航向、俯仰角、速度和位置进行误差补偿修正，同时利用加速度计进行横滚角的修正；当

时，判定机体为运动状态，此时使用GNSS数据，包括航向、俯仰角、速度、位置进行误差补偿修正，横滚角不进行误差补偿修正，其中，v_ε为速度阈值参数，依据GNSS的速度量测噪声进行取值。Another object of the present invention is to provide a flight status identification system based on airborne dual-antenna GNSS and MINS integrated navigation, including a GNSS data validity analysis module, which is used to analyze the data sent by GNSS and judge the validity of the data; The flight state recognition module is used to judge the flight state of the body according to the calculation speed of the GNSS data when the GNSS data is valid.

, it is determined that the airframe is in the ground preparation state. At this time, GNSS data, including heading, pitch angle, speed and position, are used for error compensation and correction, and the accelerometer is used to correct the roll angle; when

, it is determined that the body is in motion. At this time, GNSS data is used, including heading, pitch angle, speed, and position for error compensation correction, and roll angle is not for error compensation correction. Among them, v _ε is the speed threshold parameter, according to the GNSS speed Measure the noise to take the value.

Through this system, the flight state of the airframe can be judged according to the calculation speed of the GNSS data, and after the flight state of the aircraft is judged, it can be judged whether to correct the roll angle based on different situations.

The present invention is based on the flight state recognition system of airborne dual-antenna GNSS and MINS integrated navigation, and the flight state recognition module is also used for when the GNSS data is invalid, using the accelerometer to carry out flight state recognition, according to the following formula to obtain the k moment accelerometer output

时，判定机体处于稳定态，此时的机体为匀速巡航状态或者地面静止状态，此阶段使用加速度计进行姿态误差补偿修正，其中，

为阈值参数，其取值视导航系统所在环境的具体噪声水平设定；当

时，计算水平双轴加速度计的输出

来进行具体飞行状态判定，判定方法如下：when

When , it is determined that the body is in a stable state. At this time, the body is in a constant speed cruising state or a stationary state on the ground. At this stage, the accelerometer is used to correct the attitude error compensation. Among them,

is the threshold parameter, and its value depends on the specific noise level of the environment where the navigation system is located; when

When , calculate the output of the horizontal two-axis accelerometer

To determine the specific flight status, the determination method is as follows:

时，与陀螺仪数据进行结合判定，机体处于转弯状态或者盘旋状态，其中，

为阈值参数，依据水平加速度计的噪声具体取值，when

When , combined with the gyroscope data to determine, the body is in a turning state or a hovering state, wherein,

is the threshold parameter, according to the specific value of the noise of the horizontal accelerometer,

时，机体的机动状态大，处于起飞或者降落阶段。

When , the maneuvering state of the airframe is large, and it is in the stage of take-off or landing.

Through the flight state identification module, the accelerometer and gyroscope can also be used to effectively identify the flight state of the airframe when the GNSS signal is valid.

Description of drawings

Fig. 1 is the flow chart of the flight state identification method based on the integrated navigation of airborne dual antenna GNSS and MINS in the present invention.

detailed description

The preferred embodiments of the present invention will be described below in conjunction with the accompanying drawings. It should be understood that the preferred embodiments described below are only used to explain the present invention and do not limit the protection scope of the present invention.

The terms "first", "second" and the like in the specification, claims, and embodiments of the present application are used to distinguish similar objects, rather than to describe a specific order or sequence.

The present invention is described in further detail below by preferred specific embodiments:

Embodiment one

As shown in accompanying drawing 1: present embodiment discloses a kind of flight status identification method based on airborne dual-antenna GNSS and MINS integrated navigation, comprising the following contents:

时，判定机体处于地面准备态，此时采用GNSS数据，包括航向、俯仰角、速度和位置进行误差补偿修正，同时利用加速度计进行横滚角的修正；当

时，判定机体为运动状态，此时使用GNSS数据，包括航向、俯仰角、速度、位置进行误差补偿修正，横滚角不进行误差补偿修正，其中，v_ε为速度阈值参数，依据GNSS的速度量测噪声进行取值。Analyze the data sent by GNSS to judge the validity of the data. If the GNSS data is valid, the calculation speed of the GNSS data will be used as the basis for judging the flight status of the airframe.

, it is determined that the airframe is in the ground preparation state. At this time, the GNSS data, including heading, pitch angle, speed and position, are used for error compensation and correction, and the accelerometer is used to correct the roll angle; when

, it is determined that the body is in motion. At this time, GNSS data is used, including heading, pitch angle, speed, and position for error compensation correction, and roll angle is not for error compensation correction. Among them, v _ε is the speed threshold parameter, which is based on the speed Measure the noise to take the value.

In this embodiment, the validity judgment method of GNSS data is: the solution error of GNSS data is judged by three precision factors, which are respectively PDOP, VDOP and HDOP, and PDOP ² =HDOP ² +VDOP ² , wherein, PDOP represents It is the square root of the standard deviation between longitude, latitude and height, VDOP means the square root of the standard deviation between longitude and latitude, HDOP means the standard deviation of height, when any precision factor is greater than 3 , it is judged that the GNSS data is invalid.

In this embodiment, the discrete Kalman filter is used for the error compensation, and the indirect method is used for the selection of state quantities. The state equation and the measurement equation are respectively used to reflect the state characteristics of the integrated navigation and the relationship between the measurement information and the state.

The state equation of integrated navigation is as follows:

和加速度计相关漂移

作为状态向量(共15维)，如下：Select MINS attitude misalignment angle [φ _E φ _N φ _U ] ^T , velocity error in northeast direction [δv _E δv _N δv _U ] ^T , high latitude and longitude position error [δL δλ δh] ^T , gyroscope related drift

and accelerometer related drift

As a state vector (a total of 15 dimensions), as follows:

The state transition matrix of the system at time t is:

In the formula, F _I (t) is the error matrix of the inertial navigation system, which can be expressed as the following form:

In the formula, α _s =diag(1/τ _sx 1/τ _sy 1/τ _sz )(s=g,a), 1/τsi(s=g,a; i=x,y,z) is Marko time-dependent constant.

Choose the random white noise of the gyroscope [w _gεx w _gεy w _gεz ] ^T , the random white noise of the accelerometer [w _aεx w _aεy w _aεz ] ^T , the first-order Markov driving white noise of the gyroscope [η _gx η _gy η _g ] _z ^T and the first-order Markov driven white noise [η _ax η _ay η _az ] ^T of the accelerometer as the noise of the system, so the system noise matrix is:

W(t)＝[w _gεx w _gεy w _gεz w _aεx w _aεy w _aεz η _gx η _gy η _gz η _ax η _ay η _az ]

The covariance matrix P corresponding to the noise matrix is:

The system noise allocation matrix is:

The designed dual-antenna GNSS and MINS integrated navigation system uses nine-dimensional data of position, velocity and attitude data as observation values, and the measurement equation is as follows:

Z(t) _9×1 ＝H(t) _9×15 X(t) _15×1 +R(t) _9×1

为INS解算姿态，

为GNSS和磁力计提供的姿态角。in

Calculate attitude for INS,

Attitude angle provided for GNSS and magnetometer.

H(t)＝[H _Φ (t) _3×15 H _v (t) _3×15 H _p (t) _3×15 ] ^T ,

in,

H _Φ ＝[I _3×3 0 _3×12 ]

H _v ＝[0 _3×3 I _3×3 0 _3×9 ]

H _p ＝[0 _3×6 I _3×3 0 _3×6 ]

Measurement noise matrix R(t)＝[R _Φ (t) R _v (t) R _p (t)] ^T , where R _Φ (t) is the white noise of GNSS and magnetometer, R _v (t) and R _p (t) is the white noise of the velocity and position measurements of the GNSS receiver, respectively.

In this embodiment, when the GNSS data is invalid, the accelerometer is used to identify the flight state, and the accelerometer output at time k is obtained according to the following formula

In the static alignment phase of the airframe, the following formulas are used to obtain the local reference gravity acceleration

时，判定机体处于稳定态，此时的机体为匀速巡航状态或者地面静止状态，此阶段使用加速度计进行姿态误差补偿修正，其中，

为阈值参数，其取值视导航系统所在环境的具体噪声水平设定；当

时，计算水平双轴加速度计的输出

来进行具体飞行状态判定，判定方法如下：when

When , it is determined that the body is in a stable state. At this time, the body is in a constant speed cruising state or a static state on the ground. At this stage, the accelerometer is used to correct the attitude error compensation. Among them,

is the threshold parameter, and its value depends on the specific noise level of the environment where the navigation system is located; when

When , calculate the output of the horizontal two-axis accelerometer

To determine the specific flight status, the determination method is as follows:

时，与陀螺仪数据进行结合判定，机体处于转弯状态或者盘旋状态，其中，

为阈值参数，依据水平加速度计的噪声具体取值，when

When , combined with the gyroscope data to determine, the body is in a turning state or a hovering state, wherein,

is the threshold parameter, according to the specific value of the noise of the horizontal accelerometer,

时，机体的机动状态大，处于起飞或者降落阶段。

When , the maneuvering state of the airframe is large, and it is in the stage of take-off or landing.

The flight state identification method based on the airborne dual-antenna GNSS and MINS integrated navigation of this embodiment also includes predicting the duration of the GNSS data invalid phase under the current navigation route, and if the predicted duration is greater than the set threshold, a first prompt is issued information.

In this embodiment, it also includes calculating the duration of the GNSS data invalid phase, and when the calculated duration is greater than the set threshold, a second prompt message is issued.

时，判定机体处于地面准备态，此时采用GNSS数据，包括航向、俯仰角、速度和位置进行误差补偿修正，同时利用加速度计进行横滚角的修正；当

时，判定机体为运动状态，此时使用GNSS数据，包括航向、俯仰角、速度、位置进行误差补偿修正，横滚角不进行误差补偿修正，其中，v_ε为速度阈值参数，依据GNSS的速度量测噪声进行取值。This embodiment also discloses a flight state identification system based on airborne dual-antenna GNSS and MINS integrated navigation, including a GNSS data validity analysis module, which is used to analyze the data sent by GNSS and judge the validity of the data; flight state identification The module is used to judge the flight state of the body according to the calculation speed of the GNSS data when the GNSS data is valid.

, it is determined that the airframe is in the ground preparation state. At this time, GNSS data, including heading, pitch angle, speed and position, are used for error compensation and correction, and the accelerometer is used to correct the roll angle; when

, it is determined that the body is in motion. At this time, GNSS data is used, including heading, pitch angle, speed, and position for error compensation correction, and roll angle is not for error compensation correction. Among them, v _ε is the speed threshold parameter, according to the GNSS speed Measure the noise to take the value.

In the present embodiment, the flight state identification module is also used to use the accelerometer to identify the flight state when the GNSS data is invalid, and obtain the accelerometer output at k time according to the following formula

时，判定机体处于稳定态，此时的机体为匀速巡航状态或者地面静止状态，此阶段使用加速度计进行姿态误差补偿修正，其中，

为阈值参数，其取值视导航系统所在环境的具体噪声水平设定；当

时，计算水平双轴加速度计的输出

来进行具体飞行状态判定，判定方法如下：when

When , it is determined that the body is in a stable state. At this time, the body is in a constant speed cruising state or a stationary state on the ground. At this stage, the accelerometer is used to correct the attitude error compensation. Among them,

is the threshold parameter, and its value depends on the specific noise level of the environment where the navigation system is located; when

When , calculate the output of the horizontal two-axis accelerometer

To determine the specific flight status, the determination method is as follows:

时，与陀螺仪数据进行结合判定，机体处于转弯状态或者盘旋状态，其中，

为阈值参数，依据水平加速度计的噪声具体取值，when

When , combined with the gyroscope data to determine, the body is in a turning state or a hovering state, wherein,

is the threshold parameter, according to the specific value of the noise of the horizontal accelerometer,

时，机体的机动状态大，处于起飞或者降落阶段。

When , the maneuvering state of the airframe is large, and it is in the stage of take-off or landing.

Embodiment two

The difference between this embodiment and Embodiment 1 is that the flight status identification system of the airborne dual-antenna GNSS and MINS integrated navigation also includes a GNSS signal analysis module, and the GNSS signal analysis module is used to judge whether the current GNSS signal is available, if If not available, the current GNSS signal is not available to the ground control terminal. Specifically, the GNSS signal analysis module is used to determine whether the current GNSS signal is available, including that there is currently no GNSS signal, or the accuracy of the current GNSS signal does not meet the requirements.

之差，判断GNSS信号的精度，即判断GNSS信号是否可用。其中D为机体两个天线之间的实际安装距离，

为根据两天线的位置坐标，计算得到的两天线的定位距离。Compared with the first embodiment of judging whether the GNSS signal is available by the precision factor, in this embodiment, the GNSS signal analysis module calculates D and

The difference is to judge the accuracy of the GNSS signal, that is, to judge whether the GNSS signal is available. Where D is the actual installation distance between the two antennas of the airframe,

is the positioning distance of the two antennas calculated according to the position coordinates of the two antennas.

The ground control terminal of this embodiment, after receiving that the current GNSS signal is unavailable, retrieves the historical navigation data of the current airframe location, and judges whether the current airborne dual antenna module of the current airframe fails according to the historical navigation data, and if so, reports to The GNSS signal analysis module sends a fault signal.

In this embodiment, the historical navigation data includes the historical airframe receiving GNSS signals, the initial attitude information of the historical airframe, the airframe attitude information after error compensation, and the calculated speed and position information. Further, the historical navigation data It is multiple sets of associated stored navigation data closest to the navigation time of the current airframe, and the ground control terminal is used to comprehensively analyze the multiple sets of associated stored navigation data to determine whether the current airborne dual-antenna module is faulty.

The ground control terminal judges whether the airborne dual-antenna module of the current airframe is faulty according to the historical navigation data. The specific analysis method is: the historical airframe receiving the GNSS signal refers to whether the historical airframe has received the GNSS signal. If the historical airframe has not received the GNSS signal signal, the ground control terminal can directly judge that the airborne dual-antenna module of the current airframe has not failed; if the probability that the historical airframe has not received the GNSS signal reaches more than 60%, then further analyze the remaining historical airframes that have received the GNSS signal. Initial attitude information, airframe attitude information after error compensation, and calculated speed and position information, if the analysis of the remaining historical airframes that received GNSS signals did not use the second navigation data received by GNSS signals for error compensation accounted for more than 60% , it is judged that the airborne dual-antenna module of the current airframe is not faulty; if the proportion of the historical airframe receiving the GNSS signal reaches more than 80%, or the proportion of the historical airframe receiving the GNSS signal is between 60-80%, and the received The historical body of the GNSS signal uses the second navigation data received by the GNSS signal to perform error compensation. If the error is within the allowable range and the ratio reaches more than 90%, it is judged that the dual antenna module onboard the current body is faulty.

The preferred embodiment of the present application has been described in detail above in conjunction with the accompanying drawings. The typical known structures and common knowledge technologies in the preferred embodiment are not described here too much. Those of ordinary skill in the art can learn from the inspiration given by this embodiment. , in combination with one's own ability to perfect and implement the technical solution of the present invention, some typical known structures, known methods or common knowledge technologies should not become obstacles for those of ordinary skill in the art to implement the present application.

The scope of protection required by this application shall be based on the content of the claims, and the content of the invention, specific implementation methods, and the contents recorded in the drawings of the specification are used to interpret the claims.

Within the scope of the technical conception of the present application, some modifications can also be made to the specific embodiments of the present application, and the specific embodiments after these modifications should also be regarded as within the protection scope of the present application.

Claims (8)

1. The flight state identification method based on airborne dual-antenna GNSS and MINS integrated navigation, is characterized in that, comprises the following content: Analyze the data sent by GNSS to judge the validity of the data. If the GNSS data is valid, the calculation speed of the GNSS data will be used as the basis for judging the flight status of the aircraft.

, it is determined that the airframe is in the ground preparation state. At this time, GNSS data, including heading, pitch angle, speed and position, are used for error compensation and correction, and the accelerometer is used to correct the roll angle; when

, it is determined that the body is in motion. At this time, GNSS data is used, including heading, pitch angle, speed, and position for error compensation correction, and roll angle is not for error compensation correction. Among them, v _ε is the speed threshold parameter, according to the GNSS speed Measure the noise to take the value. 2. the flight state identification method based on airborne dual-antenna GNSS and MINS integrated navigation according to claim 1, is characterized in that, also comprises: When the GNSS data is invalid, use the accelerometer to identify the flight state, and calculate the accelerometer output at time k according to the following formula

when

When , it is determined that the body is in a stable state. At this time, the body is in a constant speed cruising state or a static state on the ground. At this stage, the accelerometer is used to correct the attitude error compensation. Among them,

is the threshold parameter, and its value depends on the specific noise level of the environment where the navigation system is located; when

When , calculate the output of the horizontal two-axis accelerometer

To determine the specific flight status, the determination method is as follows: when

When , combined with the gyroscope data to determine, the body is in a turning state or a hovering state, wherein,

is the threshold parameter, which is based on the specific value of the noise of the horizontal accelerometer;

When , the maneuvering state of the airframe is large, and it is in the stage of take-off or landing. 3. the flight state identification method based on airborne dual-antenna GNSS and MINS integrated navigation according to claim 2, is characterized in that, also comprises: In the static alignment phase of the airframe, the following formulas are used to obtain the local reference gravity acceleration

4. the flight state identification method based on airborne dual-antenna GNSS and MINS integrated navigation according to claim 3, is characterized in that, the validity judgment method of GNSS data is: The calculation error of GNSS data is judged by three precision factors, which are PDOP, VDOP and HDOP respectively, PDOP ² =HDOP ² +VDOP ² , where PDOP represents the square sum of standard deviations between longitude, latitude and height. The root sign, VDOP means the square root of the standard deviation between longitude and latitude, and HDOP means the standard deviation of the height. When any precision factor is greater than 3, it is judged that the GNSS data is invalid. 5. the flight state identification method based on airborne dual-antenna GNSS and MINS integrated navigation according to claim 2, is characterized in that, also comprises the duration of the GNSS data invalid stage under the current navigation route is predicted, if the predicted duration is greater than If the threshold is set, the first prompt message will be issued. 6. The flight state identification method based on airborne dual-antenna GNSS and MINS integrated navigation according to claim 2, further comprising calculating the duration of the GNSS data invalid stage, when the calculated duration is greater than the set threshold , a second prompt message is issued. 7. the flight state recognition system based on airborne dual antenna GNSS and MINS integrated navigation according to claim 6, is characterized in that, comprises GNSS data validity analysis module, is used for analyzing the data that GNSS sends, judges that data is valid The flight state identification module is used to judge the flight state of the airframe according to the calculation speed of the GNSS data when the GNSS data is valid.

, it is determined that the airframe is in the ground preparation state. At this time, GNSS data, including heading, pitch angle, speed and position, are used for error compensation and correction, and the accelerometer is used to correct the roll angle; when

, it is determined that the body is in motion. At this time, GNSS data is used, including heading, pitch angle, speed, and position for error compensation correction, and roll angle is not for error compensation correction. Among them, v _ε is the speed threshold parameter, according to the GNSS speed Measure the noise to take the value. 8. The flight state identification system based on airborne dual antenna GNSS and MINS integrated navigation according to claim 7, wherein the flight state identification module is also used to utilize the accelerometer to fly when the GNSS data is invalid State recognition, according to the following formula to obtain the accelerometer output at time k

when

When , it is determined that the body is in a stable state. At this time, the body is in a constant speed cruising state or a stationary state on the ground. At this stage, the accelerometer is used to correct the attitude error compensation. Among them,

is the threshold parameter, and its value depends on the specific noise level of the environment where the navigation system is located; when

When , calculate the output of the horizontal two-axis accelerometer

To determine the specific flight status, the determination method is as follows: when

When , combined with the gyroscope data to determine, the body is in a turning state or a hovering state, wherein,

is the threshold parameter, which is based on the specific value of the noise of the horizontal accelerometer;

When , the maneuvering state of the airframe is large, and it is in the stage of take-off or landing.

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