A common-mode fault monitoring method for inertial reference systems

Abstract

This invention proposes a common-mode fault monitoring method for inertial reference systems, belonging to the field of aviation navigation technology. This method detects common-mode faults in redundant inertial navigation systems by dividing the internal information link transmission of the inertial navigation system into three parts: a "sensing link," a "calculation link," and a "transmission link." The "sensing link" monitors the common-mode faults by comparing the angular velocity information from a direct-mode angular velocity sensor with the angular velocity information output by the triple-redundant inertial navigation system. While ensuring the accuracy of the gyroscope angular velocity information, it further utilizes position information input from a satellite receiver or position information calculated via vacuum velocity recursion to directly integrate and calculate the attitude. Simultaneously, the output information from the gyroscope and accelerometer can be used for inertial attitude calculation. The difference between the attitude calculated by the two methods is compared with a set detection threshold to achieve fault detection. The "calculation link" uses a separate FPGA to implement a different attitude calculation algorithm than the CPU for detection. The "transmission link" uses a monitoring FPGA chip or protocol chip to achieve loopback detection of the output attitude information. This method avoids the need for heterogeneous inertial navigation systems to address common-mode faults in redundant inertial navigation systems, saving costs and improving system safety.

Description

Technical Field

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

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

[0003] Currently, fault monitoring of inertial reference systems is typically achieved by adding redundant configurations, including system-level redundancy and sensor-level redundancy. System-level redundancy involves using two or more inertial reference systems and monitoring faults by comparing the consistency of their data outputs; sensor redundancy involves using non-orthogonal configurations of gyroscopes and accelerometers or increasing the number of sensors for fault monitoring.

[0004] Both of the above-mentioned redundancy configuration schemes improve the system monitoring capability when the inertial reference system fails to some extent. However, when the inertial reference system experiences a common-mode failure, these methods cannot monitor the common-mode failure. Summary of the Invention

[0005] The purpose of this invention is to address the common-mode fault monitoring problem in inertial reference systems. This invention proposes a common-mode fault monitoring method for inertial reference systems. It analyzes the causes of common-mode faults within the system and identifies the information transmission links that may lead to common-mode faults: the "sensing link," the "resolver link," and the "transmission link." Fault monitoring schemes are designed for these three links. The "sensing link" uses a direct-mode angular velocity sensor (DMRS) to monitor the angular velocity information of the gyroscope, and then uses the correct gyroscope information to monitor the output information of the accelerometer. The "resolver link" uses a dissimilar attitude calculation method to achieve attitude monitoring. The "transmission link" uses the addition of dissimilar channels to achieve loop-around monitoring of the output attitude information. Finally, the common-mode fault monitoring of the inertial reference system is achieved through the monitoring results of these three links.

[0006] The technical solution of this invention: To achieve the above-mentioned objective, a common-mode fault monitoring method for an inertial reference system is proposed, which is applied to a triple-redundant inertial reference system. Each redundancy in the triple-redundant inertial reference system has the same components, including a chassis, an inertial measurement unit (IMU), and electronic circuit components including a CPU, FPGA, etc. The IMU senses the angular velocity and acceleration information of the aircraft through gyroscopes and accelerometers, respectively, forming a sensing link; the CPU receives the angular velocity and acceleration information and performs attitude calculation, forming a calculation link; the I / O partition outputs navigation information to the aircraft platform and other devices through external interfaces, forming a transmission link.

[0007] The method includes the following steps:

[0008] S1, firstly, the gyroscope senses the motion of the carrier and outputs the number of gyroscope pulses. After calculation, the angular velocity signal ω of the carrier relative to the navigation system is obtained. gyro The IRS sends the angular velocity signal to the Flight Control Computer (FCM) for voting. Simultaneously, the FCM incorporates the angular velocity information ω output from the DMRS. DMRS Using the angular velocity information ω after voting DMRS Monitor angular velocity information to determine if a common-mode fault has occurred in the gyroscope in the IRS. If a common-mode fault is detected, an alarm is issued; otherwise, proceed to step S2.

[0009] In one possible embodiment, in step S1, the ω after voting is adopted. DMRS The specific process of monitoring angular velocity information and determining whether the gyroscope in the IRS has a common-mode fault includes: obtaining the maximum time of angular velocity information asynchrony between the IRS and DMRS in the FCM, the time of asynchrony caused by the low-pass filtering delay of the angular velocities of the IRS and DMRS, and the time of asynchrony caused by the sampling, calculation, and output of the angular velocities of the IRS and DMRS, and then obtaining the angular velocity error caused by the time of asynchrony between the two. Adding the maximum error of the angular velocity output of the two, a monitoring threshold is set accordingly. If the difference between the two exceeds the monitoring threshold, it is determined that the gyroscope has a common-mode fault.

[0010] S2, based on the angular velocity information ω after the vote DMRS And the position information obtained from the satellite receiver, or the monitoring attitude information of the IRS obtained from the vacuum velocity calculation;

[0011] In one possible embodiment, based on the angular velocity information ω after voting DMRS The specific process of calculating the IRS monitoring attitude information from the position information obtained by the satellite receiver includes:

[0012] IRS obtains the carrier's location information, including longitude λ, from the satellite receiver. GNSS Latitude L GNSS Height H GNSS By combining the initial attitude information of the carrier during the alignment process, including pitch angle θ, roll angle γ, and yaw angle ψ, the initial attitude transfer matrix C is obtained. b n The transformation matrix between the navigation system n and the Earth system e.

[0013]

[0014] According to the differential equation:

[0015]

[0016] Implement matrix The update; among them, Let be the transformation matrix of the machine system b relative to the inertial system i; where The angular velocity output by the gyroscope is ω. gyro And then according to

[0017]

[0018] The satellite monitoring attitude information φ is obtained directly through calculation. gyro , where ω ie This is the Earth's rotational angular velocity. The calculations rely on location information provided by satellites.

[0019] In one possible embodiment, considering scenarios where satellite navigation position information may be discontinuous and satellite position information is lost, information such as vacuum velocity is output from the atmospheric data system. Vacuum velocity is approximated as "Ground velocity," and position is calculated recursively using vacuum velocity. In the event of satellite failure, this can replace the satellite in providing position, compensate for Earth curvature errors, and maintain the integral calculation of angular velocity information output by the gyroscope.

[0020] Based on the angular velocity information ω after the vote DMRS The specific process of calculating the vacuum velocity to obtain the monitoring attitude information of the IRS includes:

[0021] Let the atmospheric vacuum velocity be With angle of attack α and sideslip angle β, the airspeed output by the atmospheric data system is a scalar, which needs to be converted to the b-series using the angle of attack and sideslip angle information. The formula is:

[0022]

[0023] After converting the velocity to the navigation coordinate system, the velocity 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: The above formula The attitude matrix is ​​the one from the previous frame; it will be updated after the position is updated. The current position information is obtained by integrating the velocity information from the navigation coordinate system, and the attitude matrix is ​​updated accordingly, resulting in the vacuum velocity monitoring attitude information φ. Air .

[0026] S3 uses the angular velocity information output by the gyroscope and the acceleration information output by the accelerometer. The inertial attitude φ is obtained;

[0027] Its basic process is divided into pose update

[0028]

[0029] Speed ​​Update

[0030]

[0031] Location update

[0032]

[0033] in, and The position information used is calculated based on the gyroscope and accelerometer. The rotation of the navigation frame n relative to the inertial frame i, including the Earth's rotational angular velocity. And the IRS moves near the Earth's surface due to the rotation of the navigation system caused by the curvature of the Earth's surface.

[0034]

[0035] In the formula, v N v E These represent the northward and eastward velocities under the geographical system.

[0036] S4. Set a monitoring threshold. If the difference between the monitored attitude information and the inertial calculated attitude exceeds the monitoring threshold, it is assumed that the accelerometer may have a common-mode fault, and the system will send an alarm message to the FCM.

[0037] In one possible embodiment, satellite monitoring attitude information φ is preferentially used when satellite data is available. gyro The inertial attitude φ is monitored; when satellite information is unavailable due to intermittent delays, the vacuum velocity is used to monitor the attitude information φ. Air Monitor inertial calculation of attitude φ.

[0038] In one possible embodiment, after monitoring the "sensing link" is completed, monitoring is performed on common-mode faults in the "computation link".

[0039] The attitude calculation algorithm is implemented in the navigation computer, so a separate FPGA is added to implement an attitude calculation algorithm that is dissimilar to that in S5. The attitude calculation implements monitoring of the attitude results in the "resolution link." A complementary filtering algorithm is considered. Based on the direct integration of gyro angular velocity in S3, the attitude φ can be obtained. gyro During uniform motion, by converting the triaxial acceleration to a geocentric coordinate system, the attitude angle φ measured by the accelerometer can be obtained by considering the resultant acceleration of the carrier as gravitational acceleration.acc ,according to

[0040] φ OF =k*φ gyro +(1-k)φ acc

[0041] Obtain the complementary filter attitude result φ OF In the formula, 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 fault has occurred in the attitude calculation link, and the system provides alarm information to the FCM.

[0043] In one possible embodiment, after monitoring the sensing link and the computation link is completed, common-mode faults in the transmission link are monitored.

[0044] After monitoring the "resolver link", the IRS will process the solution results of the navigation computer through the FPGA chip and output them to the FCM via the ARINC429 bus. In order to avoid common-mode failure of the "transmission link" and thus output errors, a monitoring FPGA chip or protocol chip is added to realize the loopback 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 two data are inconsistent, the system alarms and disables the ARINC429 bus output through hardware means.

[0045] The advantages and effects of this invention can be:

[0046] (1) It solves the problem of not being able to monitor when a common-mode fault occurs in a redundant inertial reference system.

[0047] (2) Common mode faults can be located from the system level to the sensor level, accurately identifying which part has a common mode fault. Attached Figure Description

[0048] To more clearly illustrate the technical solutions implemented in this invention, a simple explanation of the accompanying drawings used in the description of this invention will be provided below. Obviously, the drawings described below are merely some embodiments of this invention. For those skilled in the art, other drawings can be obtained based on these drawings without any creative effort.

[0049] Figure 1 This is a flowchart illustrating a common-mode fault monitoring method for an inertial reference system according to the present invention.

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

[0051] Figure 3 This is a schematic diagram of the "sensor link" monitoring principle structure of a common-mode fault monitoring method for an inertial reference system according to the present invention.

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

[0053] To make the objectives, technical solutions, and advantages of the embodiments of the present invention clearer, the technical solutions of the embodiments of the present invention will be clearly and completely described below with reference to the accompanying drawings. Obviously, the described embodiments are only some, not all, of the embodiments of the present invention. Based on the embodiments of the present invention, all other embodiments obtained by those skilled in the art without creative effort are within the scope of protection of the present invention.

[0054] The features of various aspects of the embodiments of the present invention will now be described in detail. In the following detailed description, numerous specific details are set forth to provide a thorough understanding of the invention. However, it will be apparent to those skilled in the art that the invention can also be practiced without these specific details. The following description of the embodiments is merely intended to provide a better understanding of the invention by illustrating examples. The invention is not limited to any specific setups and methods provided below, but covers all improvements, substitutions, etc., to product structures and methods without departing from the spirit of the invention. In the various drawings and the following description, well-known structures and techniques are not shown to avoid unnecessarily obscuring the invention.

[0055] It should be noted that, unless otherwise specified, the embodiments of the present invention and the features thereof can be combined with each other, and the various embodiments can be referenced and cited in each other. The present invention will now be described in detail with reference to the accompanying drawings and 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 described. Figure 2As shown, this is applied to a triple-redundant inertial reference system. Each redundancy in the triple-redundant inertial reference system has the same components, including a chassis, an inertial measurement unit (IMU), and electronic circuit components including a CPU and an FPGA. The IMU senses the angular velocity and acceleration information of the aircraft through gyroscopes and accelerometers, forming a sensing link; the CPU receives the angular velocity and acceleration information and performs attitude calculations, forming a calculation link; the I / O partition outputs navigation information to the aircraft platform and other devices through external interfaces, forming a transmission link.

[0058] Specifically, the steps include the following:

[0059] Step 1: As Figure 3 As shown, for monitoring the "sensor link", the gyroscope first senses the carrier's motion and outputs the number of gyroscope pulses. After calculation, the angular velocity signal ω of the carrier relative to the navigation system is obtained. gyro The IRS sends the angular velocity signal to the Flight Control Computer (FCM) for voting. Simultaneously, the FCM incorporates the angular velocity information ω output from the DMRS. DMRS After the vote, ω DMRS It will be used to monitor ω DMRS Information is used to determine whether a common-mode fault has occurred in the gyroscope within the IRS.

[0060] Step 2: Based on the asynchronous time of the IRS and DMRS angular velocity information received in the FCM, the asynchronous time caused by the low-pass filtering delay of the IRS and DMRS angular velocities, and the asynchronous time caused by the sampling, calculation, and output of the IRS and DMRS angular velocities, the maximum time of asynchrony of the two angular velocity information is obtained as T. Then, the angular velocity error caused by the asynchronous time is obtained. Adding the maximum error of the angular velocity output of the two, a monitoring threshold can be set accordingly. If the difference between the two exceeds the monitoring threshold, it is considered that the gyroscope has a common-mode fault, and the system alarms.

[0061] Step 3: The IRS will obtain the carrier's location information from the satellite receiver, including longitude λ. GNSS Latitude L GNSS Height H GNSS By combining the initial attitude information of the carrier during the alignment process, including pitch angle θ, roll angle γ, and yaw angle ψ, the initial attitude transfer matrix can be obtained. Transformation matrix between navigation system n and Earth system e

[0062]

[0063] According to the differential equation:

[0064]

[0065] Matrix can be implemented The update. Among them, Let be the transformation matrix of the machine system b relative to the inertial system i. In the formula... The angular velocity output by the gyroscope is ω. gyro And then according to

[0066]

[0067] The attitude φ is obtained directly. gyro , where ω ie This is the Earth's rotational angular velocity. The calculations rely on location information provided by satellites.

[0068] Step 4: Considering scenarios where satellite navigation position information may be discontinuous, resulting in the loss of satellite position information, information such as vacuum velocity is output from the atmospheric data system. Vacuum velocity is approximated as the "Ground velocity," and position is recursively calculated using vacuum velocity. This replaces satellite position information in the event of satellite failure, compensating for Earth curvature errors and maintaining the integral calculation of angular velocity information output by the gyroscope. Let the atmospheric vacuum velocity be... With angle of attack α and sideslip angle β, the airspeed output by the atmospheric data system is a scalar, which needs to be converted to the b-series using the angle of attack and sideslip angle information. The formula is:

[0069]

[0070] After converting the velocity to the navigation coordinate system, the velocity 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: The above formula The attitude matrix is ​​the one from the previous frame; it will be updated after the position is updated. The current position is obtained by integrating the velocity information from the navigation coordinate system, and the attitude matrix is ​​then updated to obtain the attitude φ. Air .

[0073] Step 5: Inertial attitude calculation φ using angular velocity information output from the gyroscope and acceleration information output from the accelerometer. Its basic process is divided into pose update

[0074]

[0075] Speed ​​Update

[0076]

[0077] Location update

[0078]

[0079] in, and The position information used is calculated jointly by the gyroscope and the accelerometer. The rotation of the navigation frame n relative to the inertial frame i, including the Earth's rotational angular velocity. And the IRS moves near the Earth's surface due to the rotation of the navigation system caused by the curvature of the Earth's surface.

[0080]

[0081] In the formula, v N v E These represent the northward and eastward velocities under the geographical system.

[0082] Step 6: When satellite data is available, prioritize using φ. gyro The accelerometer φ is monitored. A monitoring threshold is set based on the angular velocity integral attitude error and the attitude accuracy of the IRS navigation mode. If the difference between the two exceeds the monitoring threshold, a common-mode fault in the accelerometer is considered possible, and the system sends an alarm message to the FCM for further assessment. Similarly, when satellite information is intermittently unavailable, φ is used... Air Monitor attitude φ.

[0083] Step 7: After completing the monitoring of the "sensor link", such as... Figure 4 As shown, common-mode faults in the "resolution link" are monitored. Since the attitude calculation algorithm is implemented in the navigation computer, a separate FPGA is added to implement a different attitude calculation algorithm than that in S5. This algorithm calculates the attitude and monitors the attitude results in the "resolution link." A complementary filtering algorithm is considered. Based on the direct integration of the gyro angular velocity in S3, the attitude φ can be obtained. gyro During uniform motion, by converting the triaxial acceleration to a geocentric coordinate system, the attitude angle φ measured by the accelerometer can be obtained by considering the resultant acceleration of the carrier as gravitational acceleration. acc ,according to

[0084] φ OF =k*φ gyro +(1-k)φ acc

[0085] Obtain the complementary filter attitude result φ OF In the formula, k is the complementary filter weight coefficient.

[0086] Step 8: 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, determine the attitude monitoring threshold. If the difference between the two exceeds the monitoring threshold, it is considered that a common-mode fault has occurred in the attitude calculation link, and the system provides alarm information to the FCM.

[0087] Step 9: After monitoring the "resolved link", the IRS will process the solution results of the navigation computer through the FPGA chip and output them to the FCM via the ARINC429 bus. In order to avoid common-mode failure of the "transmission link" and thus output errors, a monitoring FPGA chip or protocol chip is added to realize the loop 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 two data are inconsistent, the system alarms and disables 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 person skilled in the art can easily conceive of various equivalent modifications or substitutions within the technical scope disclosed in the present invention, and these modifications or substitutions should 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: Includes the following steps: S1, firstly, the gyroscope senses the motion of the carrier and outputs the number of gyroscope pulses. After calculation, the angular velocity signal ω of the carrier relative to the navigation system is obtained. gyro The IRS sends the angular velocity signal to the FCM for voting, and at the same time, the FCM incorporates the angular velocity information ω output by the DMRS. DMRS Using the angular velocity information ω after voting DMRS Monitor angular velocity information to determine if a common-mode fault has occurred in the gyroscope in the IRS. If a common-mode fault is detected, an alarm is issued. Otherwise, proceed to step S2; S2, based on the angular velocity information ω after the vote... DMRS And the position information obtained from the satellite receiver, or the monitoring attitude information of the IRS obtained from vacuum velocity calculation; In step S2, based on the angular velocity information ω after voting... DMRS The specific process of calculating the IRS monitoring attitude information from the position information obtained by the satellite receiver includes: IRS obtains the carrier's location information, including longitude λ, from the satellite receiver. GNSS Latitude L GNSS Height H GNSS By combining the initial attitude information of the carrier during the alignment process, including pitch angle θ, roll angle γ, and yaw angle ψ, the initial attitude transfer matrix is ​​obtained. Transformation matrix between navigation system n and Earth system e According to the differential equation: Implement matrix The update; among them, Let be the transformation matrix of the machine system b relative to the inertial system i; where The angular velocity output by the gyroscope is ω. gyro And then according to The satellite monitoring attitude information φ is obtained directly through calculation. gyro , where ω ie The angular velocity of Earth's rotation. The calculation relies on the position information provided by the satellite; in step S2, based on the angular velocity information ω after voting... DMRS The specific process of calculating the vacuum velocity to obtain the monitoring attitude information of the IRS includes: Let the atmospheric vacuum velocity be With angle of attack α and sideslip angle β, the airspeed output by the atmospheric data system is a scalar, which needs to be converted to the b-series using the angle of attack and sideslip angle information. The formula is: After converting the velocity to the navigation coordinate system, the velocity can be integrated to obtain the current position longitude λ. Air Latitude L Air Height H Air The conversion formula is as follows: Specifically: by integrating the velocity information from the navigation coordinate system, the current position information is obtained, and the attitude matrix is ​​updated to obtain the vacuum velocity monitoring attitude information φ. Air S3, using the angular velocity information output by the gyroscope and the acceleration information output by the accelerometer. The inertial attitude φ is obtained; S4, a monitoring threshold is set. If the difference between the monitored attitude information and the inertial attitude exceeds the monitoring threshold, it is considered that the accelerometer may have a common-mode fault, and the system will send an alarm message to the FCM.

2. The common-mode fault monitoring method for an inertial reference system according to claim 1, characterized in that: In step S1, the ω after voting is adopted. DMRS The specific process of monitoring angular velocity information and determining whether the gyroscope in the IRS has a common-mode fault includes: obtaining the maximum time of angular velocity information asynchrony between the IRS and DMRS in the FCM, the time of asynchrony caused by the low-pass filtering delay of the angular velocities of the IRS and DMRS, and the time of asynchrony caused by the sampling, calculation, and output of the angular velocities of the IRS and DMRS, and then obtaining the angular velocity error caused by the time of asynchrony between the two. Adding the maximum error of the angular velocity output of the two, a monitoring threshold is set accordingly. If the difference between the two exceeds the monitoring threshold, it is determined that the gyroscope has a common-mode fault.

3. The common-mode fault monitoring method for an inertial reference system according to claim 1, characterized in that: In step S3, the basic process is divided into pose update. Speed ​​Update Location update in, and The position information used is calculated based on the gyroscope and accelerometer. The rotation of the navigation frame n relative to the inertial frame i, including the Earth's rotational angular velocity. And the IRS moves near the Earth's surface due to the rotation of the navigation system caused by the curvature of the Earth's surface. In the formula, v N v E These represent the northward and eastward velocities under the geographical system.

4. The common-mode fault monitoring method for an inertial reference system according to claim 1, characterized in that: In step S3, when satellite data is available, satellite monitoring attitude information φ is used preferentially. gyro The inertial attitude φ is monitored; when satellite information is unavailable due to intermittent delays, the vacuum velocity is used to monitor the attitude information φ. Air Monitor inertial calculation of attitude φ.

5. The common-mode fault monitoring method for an inertial reference system according to claim 1, characterized in that: It also includes resolution link monitoring. By adding a separate FPGA, it implements an attitude resolution algorithm that is dissimilar to that in S4, calculates the attitude, and monitors the attitude results in the resolution link. It uses a complementary filtering algorithm and obtains the attitude φ by directly integrating the gyro angular velocity from S3. gyro During uniform motion, the triaxial acceleration is converted to a geocentric coordinate system. Since the resultant acceleration of the carrier is gravitational acceleration, the attitude angle φ measured by the accelerometer is obtained. acc ,according to f OF =k*φ gyro +(1-k)φ acc Obtain the complementary filter attitude result φ OF In the formula, 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 fault has occurred in the attitude calculation link, and the system provides alarm information to the FCM.

6. The common-mode fault monitoring method for an inertial reference system according to claim 5, characterized in that: It also includes transmission link monitoring. After the resolution link monitoring, the IRS will process the resolution result of the navigation computer through the FPGA chip and output it to the FCM through the ARINC429 bus. Add a monitoring FPGA chip or protocol chip to realize the loop detection of the output attitude information. After the resolution result of the navigation computer is input into two FPGA chips respectively, it is latched and then the data consistency is compared. If the two data are inconsistent, the system alarms and disables the ARINC429 bus output through hardware means.

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