Common-mode fault monitoring method of inertial reference system

Abstract

The invention provides a common-mode fault monitoring method of an inertial reference system, and belongs to the technical field of aviation navigation. According to the method, the detection of the common-mode fault of the redundant inertial navigation system is realized, and the transmission of an internal information link of the inertial navigation system is divided into three parts, namely a sensing link, a resolving link and a transmission link, according to the sensing link, angular velocity information of a direct-mode angular velocity sensor is introduced to be compared with angular velocity information output by a three-redundancy inertial navigation system for monitoring, and on the basis of ensuring that the angular velocity information of a gyroscope is correct, position information input by a satellite receiver or position information calculated by vacuum speed recursion is further utilized; the attitude can be directly calculated through integration, meanwhile, inertial attitude calculation can be conducted on output information of the gyroscope and the accelerometer, and fault detection can be achieved by comparing the difference of the attitudes calculated through the two methods with a set detection threshold; the resolving link realizes detection by adding an independent FPGA (Field Programmable Gate Array) and an attitude resolving algorithm different from that in a CPU (Central Processing Unit); the transmission link realizes rewinding detection of output attitude information by adding an FPGA (Field Programmable Gate Array) chip or a protocol chip for monitoring. The method avoids using a heterogeneous inertial navigation system to solve the problem of common-mode fault of a redundancy inertial navigation system, can save cost, and improves the safety of the system.

Description

Technical Field

[0001] The invention belongs to the technical field of civil aviation navigation, and in particular relates to a common mode fault monitoring method for an inertial reference system. Background Art

[0002] The Inertial Reference System (IRS) is an autonomous navigation system that does not rely on external information and can work independently around the clock and in all weather conditions, providing the aircraft with the necessary navigation information such as attitude, heading, speed, and position.

[0003] At present, fault monitoring of inertial reference systems is usually achieved by adding redundant configurations, including system-level redundancy and sensor-level redundancy. System-level redundancy is achieved by using two or more inertial reference systems and performing fault monitoring by comparing the consistency of data output; sensor redundancy is achieved by using non-orthogonal configurations of gyroscopes and accelerometers or increasing the number of sensors for fault monitoring.

[0004] The above two types of redundant configuration schemes have improved the system monitoring capability when the inertial reference system fails to a certain extent. However, when a common mode failure occurs in the inertial reference system, such methods cannot realize the monitoring of the common mode failure. Summary of the invention

[0005] The purpose of the present invention is to solve the common mode fault monitoring of the inertial reference system. The present invention proposes a common mode fault monitoring method for the inertial reference system. The common mode causes are analyzed from the inside of the system, and the information transmission links where the common mode fault of the system may occur are identified, namely, the "sensing link", the "solving link" and the "transmission link". A fault monitoring scheme is designed for the three links, including the "sensing link" using a direct mode angular velocity sensor (DMRS) to monitor the angular velocity information of the gyroscope, and then using the correct gyro information to monitor the output information of the accelerometer; the "solving link" uses a non-similar attitude solution method to achieve attitude monitoring; and the "transmission link" uses an additional non-similar channel to achieve wraparound monitoring of the output attitude information. Finally, the common mode fault monitoring of the inertial reference system is achieved through the monitoring results of the three links.

[0006] The technical solution of the present invention is as follows: In order to achieve the above-mentioned invention object, a common-mode fault monitoring method of an inertial reference system is proposed, which is applied to a three-redundant inertial reference system. The three-redundant inertial reference system has the same redundancy composition, including a chassis, an inertial measurement unit (IMU), and electronic circuit components including a CPU, an FPGA, etc., wherein the IMU senses the angular velocity and acceleration information of the aircraft through a gyroscope and an accelerometer respectively to form a sensing link; the CPU receives the angular velocity and acceleration information to perform attitude calculation to form a calculation link; the I / O partition outputs the navigation information to the aircraft platform and other equipment through an external interface to form a transmission link.

[0007] The method comprises the following steps:

[0008] S1, first the gyro senses the carrier movement and outputs the gyro pulse number, and then the angular velocity signal ω of the carrier relative to the navigation system is obtained through calculation. gyro , the IRS sends the angular velocity signal to the flight control computer (FCM) for voting. At the same time, the FCM introduces the angular velocity information ω output by the DMRS DMRS , using the angular velocity information ω after voting DMRS Monitor the angular velocity information to determine whether a common mode failure occurs in the gyro in the IRS. If it is determined that a common mode failure occurs in the gyro, an alarm is issued; otherwise, the process proceeds to step S2;

[0009] In a possible embodiment, in step S1, the voted ω DMRS The specific process of monitoring the angular velocity information and determining whether a common mode failure occurs in the gyro in the IRS includes: obtaining the maximum time T for the angular velocity information to be out of sync based on the out-of-sync time of the IRS and DMRS angular velocity information when receiving the information in the FCM, the out-of-sync time caused by the low-pass filtering delay of the IRS and DMRS angular velocities, and the out-of-sync time caused by the sampling, calculation, and output of the IRS and DMRS angular velocities, and then obtaining the angular velocity error caused by the out-of-sync time, plus the maximum error of the angular velocity outputs of the two, and setting the monitoring threshold accordingly. If the difference between the two exceeds the monitoring threshold, it is determined that a common mode failure occurs in the gyro.

[0010] S2, according to the angular velocity information ω after voting DMRS and the position information obtained by the satellite receiver, or the IRS monitoring attitude information obtained by true airspeed calculation;

[0011] In a possible embodiment, according to the angular velocity information ω after voting DMRS The specific process of calculating the IRS monitoring attitude information using the position information obtained by the satellite receiver includes:

[0012] IRS obtains the carrier's position information from the satellite receiver, including longitude λ GNSS , Latitude L GNSS , height H GNSS , combined with the initial attitude information of the carrier during the alignment process, including the pitch angle θ, the roll angle γ, and the heading angle ψ, the initial attitude transfer matrix C is obtained b n , the transformation matrix of navigation system n relative to earth system e

[0013]

[0014] According to the differential equation:

[0015]

[0016] Implementing the Matrix of updates; among them, is the transformation matrix of the machine system b relative to the inertial system i; is the angular velocity output by the gyroscope, i.e. ω gyro , and then according to

[0017]

[0018] Directly calculate the satellite monitoring attitude information φ gyro , where ω ie is the Earth's rotation angular velocity, The calculation relies on the position information provided by satellites.

[0019] In one possible embodiment, considering the scenario where the satellite navigation position information may be discontinuous, the satellite position information is lost at this time, and the atmospheric data system is used to output information such as true vacuum speed, and the true vacuum speed is approximated as the "ground speed". The position is calculated recursively through the true vacuum speed. When the satellite fails, the position is provided by the satellite instead, the earth curvature error is compensated, and the integral calculation of the angular velocity information output by the gyroscope is maintained.

[0020] According to the angular velocity information ω after voting DMRS The specific process of calculating the true airspeed to obtain the IRS monitoring attitude information includes:

[0021] Assume that the atmospheric vacuum speed is The angle of attack is α, the sideslip angle is β, and the airspeed output by the atmospheric data system is a scalar. It is necessary to use the angle of attack and sideslip information to convert to the b system. The formula is:

[0022]

[0023] After converting the speed to the navigation coordinate system, the speed can be integrated to obtain the current position longitude λ Air , Latitude L Air , height H Air , the conversion formula is as follows:

[0024]

[0025] Note: is the attitude matrix of the previous shot. After the position is updated, the attitude matrix will be updated. The navigation coordinate system velocity information is integrated to obtain the current position information, and the attitude matrix is ​​updated to obtain the true airspeed monitoring attitude information φ Air .

[0026] S3, using the angular velocity information output by the gyroscope and the acceleration information output by the accelerometer Get the inertial solution posture φ;

[0027] The basic process is divided into posture update

[0028]

[0029] Speed ​​Update

[0030]

[0031] Location Updates

[0032]

[0033] in, and The position information used is derived from the gyroscope and accelerometer; is the rotation of navigation system n relative to inertial system i, including the angular velocity of the Earth's rotation and the IRS moves near the Earth's surface. The navigation system rotates due to the curvature of the Earth's surface.

[0034]

[0035] In the formula, v N , v E are the north velocity and east velocity in the geographic system.

[0036] S4, setting the monitoring threshold. If the difference between the monitored attitude information and the inertial solution attitude exceeds the monitoring threshold, it is considered that the accelerometer may have a common mode failure, and the system will send an alarm message to the FCM.

[0037] In one possible embodiment, when satellite data is available, satellite monitoring attitude information φ is used preferentially. gyro Monitor the inertial solution attitude φ; when satellite information is unavailable due to interruptions, use true airspeed to monitor attitude information φ Air Monitor the inertial solution attitude φ.

[0038] In a possible embodiment, after the monitoring of the "sensing link" is completed, the common mode fault of the "solution link" is monitored.

[0039] The attitude solution algorithm is implemented in the navigation computer, so an independent FPGA is added to implement an attitude solution algorithm that is not similar to that in S5. The attitude is calculated to monitor the attitude results in the "solution link". The complementary filtering algorithm is considered. According to S3, the attitude φ can be obtained by directly integrating the gyro angular velocity. gyro During uniform motion, the three-axis acceleration is converted to the Earth-fixed coordinate system. The attitude angle φ measured by the accelerometer can be obtained based on the carrier's combined acceleration as gravity acceleration.acc ,according to

[0040] φ OF =k*φ gyro +(1-k)φ acc

[0041] Get the complementary filtering attitude result φ OF , where k is the complementary filter weight coefficient.

[0042] The attitude monitoring threshold is determined based on the error requirement of the attitude information φ output by the IRS and the output error of the attitude calculated by the complementary filtering algorithm used for monitoring. If the difference between the two exceeds the monitoring threshold, it is considered that a common mode failure occurs in the attitude solution link, and the system provides an alarm message to the FCM.

[0043] In a possible embodiment, after the monitoring of the sensing link and the solution link is completed, the common mode fault of the transmission link is monitored.

[0044] After monitoring the "solution link", the IRS will process the solution results of the navigation computer through the FPGA chip and then output them to the FCM through the ARINC429 bus. In order to avoid common-mode failure of the "transmission link" and cause output errors, a monitoring FPGA chip or protocol chip is added to realize the wraparound detection of the output attitude information. The solution results of the navigation computer are input into the two FPGA chips respectively, latched, and then compared for data consistency. If the data of the two are inconsistent, the system will alarm and prohibit the ARINC429 bus output through hardware means.

[0045] The advantages and effects of the present invention may be:

[0046] (1) The problem of being unable to monitor when a common mode failure occurs in the redundant inertial reference system is solved.

[0047] (2) Common-mode failures can be located from the system level to the sensor level, accurately identifying which link the common-mode failure occurs at. BRIEF DESCRIPTION OF THE DRAWINGS

[0048] In order to more clearly explain the technical solution implemented by the present invention, the following will be a simple explanation of the drawings needed in the description of the present invention. Obviously, the drawings described below are only some embodiments of the present invention. For those skilled in the art, other drawings can be obtained based on these drawings without creative work.

[0049] Figure 1 It is a flow chart of a common mode fault monitoring method for an inertial reference system according to the present invention;

[0050] Figure 2It is a principle structure diagram of a common mode fault monitoring method for an inertial reference system according to the present invention;

[0051] Figure 3 It is a "sensing link" monitoring principle structure diagram of a common mode fault monitoring method for an inertial reference system of the present invention;

[0052] Figure 4 It is a structural diagram of the monitoring principle of a "solution link" of a common mode fault monitoring method for an inertial reference system according to the present invention. DETAILED DESCRIPTION

[0053] In order to make the purpose, technical solution and advantages of the embodiments of the present invention clearer, the technical solution in the embodiments of the present invention will be clearly and completely described below in conjunction with the drawings in the embodiments of the present invention. Obviously, the described embodiments are part of the embodiments of the present invention, not all of the embodiments. Based on the embodiments of the present invention, all other embodiments obtained by ordinary technicians in this field without making creative work are within the scope of protection of the present invention.

[0054] The features of various aspects of the embodiments of the present invention will be described in detail below. In the detailed description below, many specific details are proposed to fully understand the present invention. However, it is obvious to those skilled in the art that the present invention can also be implemented without these specific details. The following description of the embodiments is only for a better understanding of the present invention by illustrating examples of the present invention. The present invention is not limited to any specific settings and methods provided below, but covers all product structures, any improvements, replacements, etc. of the methods covered without departing from the spirit of the present invention. In the various drawings and the following description, known structures and technologies are not shown to avoid unnecessary ambiguity of the present invention.

[0055] It should be noted that, in the absence of conflict, the embodiments of the present invention and the features in the embodiments can be combined with each other, and the embodiments can refer to and quote each other. The present invention will be described in detail below with reference to the accompanying drawings and in combination with the embodiments.

[0056] The method may include the following steps:

[0057] like Figure 1 As shown, a common mode fault monitoring method for an inertial reference system is as follows: Figure 2As shown, it is applied to a three-degree-of-freedom inertial reference system. The three-degree-of-freedom inertial reference system has the same redundancy components, including a chassis, an inertial measurement unit (IMU), and electronic circuit components including a CPU, an FPGA, etc., wherein the IMU senses the angular velocity and acceleration information of the aircraft through a gyroscope and an accelerometer, respectively, to form a sensing link; the CPU receives the angular velocity and acceleration information for attitude calculation, to form a calculation link; the I / O partition outputs the navigation information to the aircraft platform and other equipment through an external interface, to form a transmission link.

[0058] The specific steps include:

[0059] Step 1: Figure 3 As shown in the figure, for the monitoring of the "sensing link", the gyro first senses the carrier movement and outputs the gyro pulse number. After calculation, the angular velocity signal ω of the carrier relative to the navigation system is obtained. gyro , the IRS will send the angular velocity signal to the flight control computer (FCM) for voting. At the same time, the FCM will introduce the angular velocity information ω output by the DMRS DMRS , after voting DMRS Will be used to monitor DMRS Information is used to determine whether a common mode failure occurs in the gyro in the IRS.

[0060] Step 2: According to the asynchronous time when the IRS and DMRS angular velocity information is received in the FCM, the asynchronous time caused by the low-pass filter delay of the IRS and DMRS angular velocity, and the asynchronous time caused by the sampling, calculation, and output of the IRS and DMRS angular velocity, the maximum asynchronous time of the angular velocity information of the two is obtained as T, and then the angular velocity error caused by the asynchronous time is obtained. Add the maximum error of the angular velocity output of the two, and the monitoring threshold can be set accordingly. If the difference between the two exceeds the monitoring threshold, it is considered that the gyro has a common mode fault and the system alarms;

[0061] Step 3: IRS obtains the carrier's location information from the satellite receiver, including longitude λ GNSS , Latitude L GNSS , height H GNSS , combined with the initial attitude information of the carrier during the alignment process, including the pitch angle θ, the roll angle γ, and the heading angle ψ, the initial attitude transfer matrix can be obtained The transformation matrix of navigation system n relative to earth system e

[0062]

[0063] According to the differential equation:

[0064]

[0065] The matrix can be realized Of which, is the transformation matrix of the machine system b relative to the inertial system i. is the angular velocity output by the gyroscope, i.e. ω gyro , and then according to

[0066]

[0067] Directly calculate the posture φ gyro , where ω ie is the Earth's rotation angular velocity, The calculation relies on the position information provided by satellites.

[0068] Step 4: Consider the scenario where satellite navigation position information may be discontinuous. In this case, the satellite position information is lost. The atmospheric data system outputs information such as true vacuum speed, and the true vacuum speed is approximated as the "ground speed". The position is calculated recursively through the true vacuum speed. When the satellite fails, the position is provided instead of the satellite, the earth curvature error is compensated, and the integral calculation of the angular velocity information output by the gyroscope is maintained. Assume that the atmospheric vacuum speed is The angle of attack is α, the sideslip angle is β, and the airspeed output by the atmospheric data system is a scalar. It is necessary to use the angle of attack and sideslip information to convert to the b system. The formula is:

[0069]

[0070] After converting the speed to the navigation coordinate system, the speed can be integrated to obtain the current position longitude λ Air , Latitude L Air , height H Air , the conversion formula is as follows:

[0071]

[0072] Note: is the attitude matrix of the previous shot. After the position is updated, the attitude matrix will be updated. The velocity information of the navigation coordinate system is integrated to obtain the current position information, and the attitude matrix is ​​updated to obtain the attitude φ Air .

[0073] Step 5: Inertial attitude solution φ uses the angular velocity information output by the gyro and the acceleration information output by the accelerometer The basic process is divided into posture update

[0074]

[0075] Speed ​​Update

[0076]

[0077] Location Updates

[0078]

[0079] in, and The position information used is derived from a combination of the gyroscope and the accelerometer. is the rotation of navigation system n relative to inertial system i, including the angular velocity of the Earth's rotation and the IRS moves near the Earth's surface. The navigation system rotates due to the curvature of the Earth's surface.

[0080]

[0081] In the formula, v N , v E are the north velocity and east velocity in the geographic system.

[0082] Step 6: When satellite data is available, use φ first gyro Monitor φ, set the monitoring threshold according to the angular velocity integrated attitude error and the attitude accuracy of the IRS navigation mode. If the difference between the two exceeds the monitoring threshold, it is considered that the accelerometer may have a common mode failure, and the system will send an alarm message to FCM for judgment; similarly, when the satellite information is unavailable due to interruption, use φ Air Monitor posture φ.

[0083] Step 7: After completing the monitoring of the "sensing link", Figure 4 As shown, the common mode fault of the "solution link" is monitored. The attitude solution algorithm is implemented in the navigation computer, so an independent FPGA is added to implement a non-similar attitude solution algorithm in S5. The attitude is calculated to monitor the attitude results in the "solution link". The complementary filtering algorithm is considered. According to S3, the attitude φ can be obtained by directly integrating the gyro angular velocity. gyro During uniform motion, the three-axis acceleration is converted to the Earth-fixed coordinate system. The attitude angle φ measured by the accelerometer can be obtained based on the carrier's combined acceleration as gravity acceleration. acc ,according to

[0084] φ OF =k*φ gyro +(1-k)φ acc

[0085] Get the complementary filtering attitude result φ OF , where k is the complementary filter weight coefficient.

[0086] Step 8: Determine the attitude monitoring threshold based on the error requirement of the attitude information φ output by the IRS and the output error of the attitude calculated by the complementary filtering algorithm used for monitoring. If the difference between the two exceeds the monitoring threshold, it is considered that a common mode failure occurs in the attitude solution link, and the system provides an alarm message to the FCM.

[0087] Step 9: After the "solution link" monitoring, the IRS will process the solution results of the navigation computer through the FPGA chip and output them to the FCM through the ARINC429 bus. In order to avoid common mode failure of the "transmission link" and cause output errors, a monitoring FPGA chip or protocol chip is added to realize the wraparound detection of the output attitude information. The solution results of the navigation computer are input into the two FPGA chips respectively, latched, and then compared for data consistency. If the data of the two are inconsistent, the system will alarm and prohibit the ARINC429 bus output through hardware means.

[0088] Finally, it should be noted that the above embodiments are only used to illustrate the technical solutions of the present invention, but the protection scope of the present invention is not limited thereto. Any technician familiar with the field can easily think of various equivalent modifications or substitutions within the technical scope disclosed by the present invention, and these modifications or substitutions should all be covered within the protection scope of the present invention.

Claims

1. A common mode fault monitoring method for an inertial reference system, characterized in that: The following steps are involved: S1, first the gyro senses the carrier movement and outputs the gyro pulse number, and then the angular velocity signal ω of the carrier relative to the navigation system is obtained through calculation. gyro , IRS sends the angular velocity signal to FCM for voting. At the same time, FCM introduces the angular velocity information ω output by DMRS DMRS , using the angular velocity information ω after voting DMRS Monitor angular velocity information to determine whether a common mode failure occurs in the gyro in the IRS. If a common mode failure occurs in the gyro, an alarm is issued. Otherwise, go to step S2; S2, according to the angular velocity information ω after voting DMRS And the position information obtained by the satellite receiver, or the true airspeed calculation to obtain the IRS monitoring attitude information; S3, using the angular velocity information output by the gyroscope and the acceleration information output by the accelerometer Get the inertial solution attitude φ; S4, set the monitoring threshold. If the difference between the monitoring attitude information and the inertial solution attitude exceeds the monitoring threshold, it is considered that the accelerometer may have a common mode failure, and the system will send an alarm message to the FCM.

2. The common mode fault monitoring method of an inertial reference system according to claim 1, characterized in that: In step S1, the voted ω DMRS The specific process of monitoring the angular velocity information and determining whether a common mode failure occurs in the gyro in the IRS includes: obtaining the maximum time T for the angular velocity information to be out of sync based on the out-of-sync time of the IRS and DMRS angular velocity information when receiving the information in the FCM, the out-of-sync time caused by the low-pass filtering delay of the IRS and DMRS angular velocities, and the out-of-sync time caused by the sampling, calculation, and output of the IRS and DMRS angular velocities, and then obtaining the angular velocity error caused by the out-of-sync time, plus the maximum error of the angular velocity outputs of the two, and setting the monitoring threshold accordingly. If the difference between the two exceeds the monitoring threshold, it is determined that a common mode failure occurs in the gyro.

3. The common mode fault monitoring method of an inertial reference system according to claim 1, characterized in that: In step S2, according to the angular velocity information ω after voting DMRS The specific process of calculating the IRS monitoring attitude information using the position information obtained by the satellite receiver includes: IRS obtains the carrier's position information from the satellite receiver, including longitude λ GNSS , Latitude L GNSS , height H GNSS , combined with the initial attitude information of the carrier during the alignment process, including the pitch angle θ, the roll angle γ, and the heading angle ψ, the initial attitude transfer matrix is ​​obtained The transformation matrix of navigation system n relative to earth system e According to the differential equation: Implementing the Matrix of updates; among them, is the transformation matrix of the machine system b relative to the inertial system i; is the angular velocity output by the gyroscope, i.e. ω gyro , and then according to Directly calculate the satellite monitoring attitude information φ gyro , where ω ie is the Earth's rotation angular velocity, The calculation relies on the position information provided by satellites.

4. The common mode fault monitoring method of an inertial reference system according to claim 1, characterized in that: In step S2, according to the angular velocity information ω after voting DMRS The specific process of calculating the true airspeed to obtain the IRS monitoring attitude information includes: Assume that the atmospheric vacuum speed is The angle of attack is α, the sideslip angle is β, and the airspeed output by the atmospheric data system is a scalar. It is necessary to use the angle of attack and sideslip information to convert to the b system. The formula is: After converting the speed to the navigation coordinate system, the speed can be integrated to obtain the current position longitude λ Air , Latitude L Air , height H Air , the conversion formula is as follows: Note: The above formula is the attitude matrix of the previous shot. After the position is updated, the attitude matrix will be updated. The navigation coordinate system velocity information is integrated to obtain the current position information, and the attitude matrix is ​​updated to obtain the true airspeed monitoring attitude information φ Air .

5. The common mode fault monitoring method of an inertial reference system according to claim 1, characterized in that: In step S3, the basic process is divided into posture update Speed ​​Update Location Updates in, and The position information used is derived from the gyroscope and accelerometer; is the rotation of navigation system n relative to inertial system i, including the angular velocity of the Earth's rotation and the IRS moves near the Earth's surface. The navigation system rotates due to the curvature of the Earth's surface. In the formula, v N , v E are the north velocity and east velocity in the geographic system.

6. The common mode fault monitoring method of an inertial reference system according to claim 1, characterized in that: In step S3, when satellite data is available, the satellite monitoring attitude information φ is used preferentially. gyro Monitor the inertial solution attitude φ; when satellite information is unavailable due to interruptions, use true airspeed to monitor attitude information φ Air Monitor the inertial solution attitude φ.

7. The common mode fault monitoring method of an inertial reference system according to claim 1, characterized in that: It also includes solution link monitoring. By adding an independent FPGA, a non-similar attitude solution algorithm is implemented in S5. The attitude is calculated to monitor the attitude results in the solution link. The complementary filtering algorithm is used. According to S3, the attitude φ can be obtained by directly integrating the gyro angular velocity. gyro During uniform motion, the three-axis acceleration is converted to the Earth-fixed coordinate system. The attitude angle φ measured by the accelerometer can be obtained based on the carrier's combined acceleration as gravity acceleration. acc ,according to f OF =k*φ gyro +(1-k)φ acc Get the complementary filtering attitude result φ OF , where k is the complementary filter weight coefficient; The attitude monitoring threshold is determined based on the error requirement of the attitude information φ output by the IRS and the output error of the attitude calculated by the complementary filtering algorithm used for monitoring. If the difference between the two exceeds the monitoring threshold, it is considered that a common mode failure occurs in the attitude solution link, and the system provides an alarm message to the FCM.

8. A common mode fault monitoring method for an inertial reference system according to claim 7, characterized in that: It also includes transmission link monitoring. After the solution link monitoring, the IRS will process the solution results of the navigation computer through the FPGA chip and output them to the FCM through the ARINC429 bus. An FPGA chip or protocol chip for monitoring is added to realize the wraparound detection of the output attitude information. The solution results of the navigation computer are input into two FPGA chips respectively, latched, and then the data consistency is compared. If the data of the two are inconsistent, the system will alarm and prohibit the ARINC429 bus output through hardware means.

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