Civil aircraft inertial satellite integrity performance test system and method based on semi-physical test platform

Abstract

The invention provides a civil aircraft inertial satellite integrity performance test system and method based on a semi-physical test platform, and belongs to the technical field of aviation navigation. The method can meet the requirement of civil aircraft standards for testing the integrity performance of the inertial satellite, and the integrity performance of the inertial satellite is tested by simulating a flight path, IMU data and error input and satellite data and error input. The test content mainly comprises false alarm rate, fault-free precision performance, rare normal verification, detection / elimination verification and integrated sliding. According to the method, inertial satellite integrity performance test verification meeting civil aircraft standards can be realized, so that an inertial satellite combination system obtains airworthiness authentication.

Description

Technical Field

[0001] The present invention relates to the field of civil aviation navigation technology, and in particular to a civil aircraft inertial satellite integrity performance test system and method based on a semi-physical test platform. Background Art

[0002] With the rapid progress in the field of civil aviation navigation, civil aviation organizations of various countries are promoting the development of civil aircraft navigation systems, and the inertial satellite integrated navigation system is an important navigation system for civil aircraft. Inertial navigation systems and satellite navigation systems have natural complementary advantages, so the inertial satellite integrated navigation system can provide more complete performance for civil aircraft.

[0003] The integrity of inertial satellite integrated navigation has always attracted the attention of the international community and is also an important indicator in the field of life safety of civil aircraft. Integrity refers to the ability to promptly warn users of deviations during the flight phase of the aircraft, including the accuracy, continuity and availability indicators of each route phase. The inertial satellite navigation system used in civil aircraft must meet the airworthiness requirements, so it must be tested and verified. Summary of the invention

[0004] The purpose of the present invention is to propose a civil aircraft inertial satellite integrity performance test system and method based on a semi-physical test platform, which is used to test the compliance of the integrity solution capability of the inertial satellite navigation system with civil aircraft standards.

[0005] The technical solution of the present invention: According to the first aspect of the present invention, a civil aircraft inertial satellite integrity performance test system based on a semi-physical test platform is proposed, which is applied to the test of an inertial satellite navigation system, and includes: a flight trajectory simulation module, an IMU data and error input module, a satellite data and error input module, a false alarm rate test module, a fault-free accuracy performance test module, a rare normal verification module, a detection / elimination verification module, and an integrated taxiing module; the flight trajectory simulation module is used to simulate and generate flight parameters required according to flight requirements; the IMU data and error input module is used to reversely solve and generate angular increments and velocity increments output by sensors according to the flight parameters required, and simulate the injection of errors to obtain angular increments and velocity increments containing errors, and convert the angular increments and velocity increments containing errors into digital signals and transmit them to the inertial satellite navigation system; the satellite data and error input module is used to simulate and generate satellite data according to the flight parameters required using broadcast ephemeris, and at the same time, obtain error-containing data by introducing a satellite space segment propagation error model. The invention relates to a method for detecting a satellite data containing errors, converting the satellite data containing errors into radio frequency signals and sending them to an inertial satellite navigation system; the inertial satellite navigation system calculates pseudoranges, pseudorange rates, GNSS / IRS combination information and integrity parameters according to the received angular increments and velocity increments containing errors and the satellite data containing errors; the false alarm rate test module is used to statistically calculate the false alarm rate according to multiple groups of integrity parameters; the fault-free accuracy performance test module is used to calculate the snapshot 95% horizontal accuracy and the inertial satellite filtering accuracy according to multiple groups of GNSS / IRS combination information; the rare normal verification module is used to verify whether the fault-free rare normal HPL in the integrity parameters can correctly limit the horizontal position error according to the GNSS / IRS combination information and the parameters required for the flight; the detection / elimination verification module is used to verify the integrity detection / elimination algorithm of the inertial satellite navigation system after injecting faults into the satellite data and error input module; the integrated taxiing module is used to simulate the calculation of the false alarm rate, the missed alarm rate and the failure elimination probability under the satellite failure state.

[0006] In a possible embodiment, the flight requirements include initial information, flight path and flight time. The initial information includes the initial position, specifically including latitude, longitude and altitude; the initial attitude, specifically including heading angle, roll angle and pitch angle; the flight required parameters include instant latitude and longitude, instant attitude and instant time.

[0007] In a possible embodiment, the injected error includes zero bias error, proportional factor error, random walk error, non-orthogonal installation error, Gaussian white noise error, and first-order Markov process error.

[0008] In a possible embodiment, the satellite space segment propagation error model includes an ionosphere error model, a troposphere error model, a satellite clock / ephemeris error model and a receiver noise error model.

[0009] According to a second aspect of the present invention, a method for testing the integrity performance of an inertial satellite of a civil aircraft based on a semi-physical test platform is proposed. The method adopts the above-mentioned inertial satellite integrity performance testing system of a civil aircraft based on a semi-physical test platform, and specifically comprises the following steps:

[0010] S1: Generate flight parameters according to flight requirements;

[0011] S2: Generate output information of the sensor by reverse calculation according to the parameters required for flight, and simulate injection error to obtain angular increment and velocity increment containing error, convert the angular increment and velocity increment containing error into digital signals and transmit them to the inertial satellite navigation system;

[0012] S3: Generate satellite data using broadcast ephemeris based on the flight parameters. At the same time, obtain satellite data with errors by introducing the satellite space segment propagation error model. Convert the satellite data with errors into radio frequency signals and send them to the inertial satellite navigation system. Satellite data include pseudorange, pseudorange rate, satellite position, satellite speed, and number of visible satellites.

[0013] S4: The inertial satellite navigation system obtains GNSS / IRS combination information and integrity parameters based on the received angular increment and velocity increment containing errors and the satellite data containing errors;

[0014] S5: Repeat steps S1-S4 to obtain multiple sets of GNSS / IRS combination information and integrity parameters;

[0015] S6: Calculate the false alarm rate based on multiple sets of integrity parameters;

[0016] S7: Calculate snapshot 95% horizontal accuracy and inertial satellite filtering accuracy based on multiple sets of GNSS / IRS combined information;

[0017] S8: Verify whether the fault-free rare normal HPL in the integrity parameters can correctly limit the horizontal position error based on the GNSS / IRS combined information and the flight required parameters;

[0018] S9: injecting integrity faults into the satellite data and error input module to verify whether the integrity detection / elimination algorithm of the inertial satellite navigation system is effective;

[0019] S10: Simulate the satellite outage state and calculate the false alarm rate, missed alarm rate and failure elimination probability.

[0020] In a possible embodiment, step S6 specifically includes the following steps:

[0021] S61: generating at least 40 sets of satellite data with different geometric structures in step 3;

[0022] S62: For each set of satellite data with different geometric structures, set the satellite speed to 0, fly for at least 82,500 hours in the preset flight trajectory, and count the number of alarms N in the integrity parameters. alarm ; Through 40×82500 hours of testing, the adequacy of the false alarm rate test can be guaranteed; Setting a fixed satellite geometry structure can avoid the impact of satellite observation jumps caused by changes in satellite positions, ensuring the integrity of the false alarm rate test;

[0023] S63: Calculate the false alarm rate according to the following formula (1): Formula (I), N all Indicates the total number of combined navigation.

[0024] In a possible embodiment, in step S7, the snapshot 95% horizontal accuracy is calculated according to the following formula:

[0025]

[0026] Among them, d i is the instantaneous two-dimensional horizontal position error (meters), which is calculated based on the instantaneous longitude and latitude and the combined information in the flight parameters required. Among them, (x1, y1) is the instantaneous latitude and longitude in the flight parameters required, (x2, y2) is the real-time position information in the combined information; N is the total number of combined navigation, and HDOP is the instantaneous horizontal precision factor in the combined information.

[0027] In a possible embodiment, in step S7, the inertial satellite filtering accuracy is calculated using the following formula according to the inertial satellite navigation system combined filter:

[0028] Among them, p 11 、p 22 is the covariance matrix P in the Kalman filter.

[0029] In a possible embodiment, step S8 specifically includes the following steps:

[0030] S81: Calculate the horizontal position error based on the instantaneous latitude and longitude and combined information in the flight parameters required. Among them, (x1, y1) is the instant latitude and longitude in the flight required parameters, and (x2, y2) is the real-time position information in the combined information;

[0031] S82: According to the HPL in the integrity parameter, compare with HEL;

[0032] S83: If HPL > HEL, the fault - free rare normal HPL (H0) can correctly limit the horizontal position error; otherwise, the fault - free rare normal HPL (H0) cannot correctly limit the horizontal position error.

[0033] In a possible embodiment, in the step S9, it specifically includes the following steps:

[0034] S91: Inject integrity faults into the satellite data and error input module in the pseudorange domain. The integrity faults are divided into integrity step faults and integrity ramp faults.

[0035] S92: The integrity parameters are calculated by the inertial satellite navigation system. If HPL > HAL in the integrity parameters, it indicates that a fault in the satellite space segment signal is currently detected. Among them, HAL is the horizontal alert limit for the current flight route phase, which is uniformly specified by the International Civil Aviation Organization.

[0036] S93: Test and verify the performance of the standard receiver autonomous integrity monitoring (RAIM). The standard RAIM algorithm built in the detection / exclusion verification module excludes the faults and calculates the HPL after the faults are excluded. RAIM ;

[0037] S94: The inertial satellite navigation system calculates the HPL after the faults are excluded. FD ;

[0038] S95: After the faults are excluded, calculate the current HEL. FD , Among them, (x1, y1) is the instant longitude and latitude in the flight required parameters, and (x2, y2) is the real - time position information in the combined information after the faults are excluded.

[0039] S96: Compare HEL FD , HPL RAIM , HPL FD and HAL. If HEL FD < HPL FD < HPL RAIM < HAL, it indicates that the integrity fault detection and exclusion algorithm in the device under test is effective.

[0040] In a possible embodiment, in the step S10, it specifically includes the following steps:

[0041] S101: Generate satellite data with at least two different geometric structures in the step 3. In each scenario, remove enough visible satellites so that the number of visible satellites is less than 4, ensuring that the inertial satellite navigation system is in a coasting state.

[0042] S102: For each taxiing geometry, at least 1,650 tests shall be conducted to count the number of false alarms, missed alarms, and failed troubleshooting.

[0043] S103: Calculate the false alarm rate, missed alarm rate and failure elimination probability; false alarm rate, Missed alarm rate, Failure exclusion probability, Among them, N fa is the number of alarms when there is no fault, N md is the number of times no alarm is given when there is a fault, N fd N is the number of times when there is a fault and the failure to eliminate it occurs. all is the total number of trials.

[0044] Beneficial technical effects of the present invention: The present invention proposes a civil aircraft inertial satellite integrity performance test system based on a semi-physical test platform, which can provide a physical platform and system architecture for the test of the integrity algorithm. At the same time, the present invention proposes a complete set of civil aircraft inertial satellite integrity performance test methods, which fills the gap in the current complete test method for civil aircraft inertial satellite integrity performance and meets the civil aircraft inertial satellite integrity test requirements and airworthiness requirements. BRIEF DESCRIPTION OF THE DRAWINGS

[0045] 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.

[0046] Figure 1 It is a schematic diagram of the process of the present invention for testing the integrity performance of a civil aircraft inertial satellite based on a semi-physical testing platform; DETAILED DESCRIPTION

[0047] 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.

[0048] 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.

[0049] 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.

[0050] Example 1

[0051] The method may include the following steps:

[0052] Step 1: Set the initial information, initial position (33.48°, 108.5°, 80m); initial attitude (90°, 0°, 0°); flight path, Xi'an-Shanghai; flight time 2h; obtain the required flight parameters, including real-time latitude and longitude, real-time attitude and real-time time;

[0053] Step 2: Calculate the output information of the sensor, and reversely calculate and generate the output information of the sensor according to the simulated flight trajectory, as well as simulate the sensor error input;

[0054] Step 3, generating output information of the sensor, including the angular increment output by the three-axis gyroscope and the linear velocity increment output by the three-axis accelerometer;

[0055] Step 4, simulate the sensor error input, the deterministic error includes the constant zero bias of the gyroscope and accelerometer, the non-orthogonal installation error and the proportional factor error; the random error includes the random wandering error of the gyroscope and accelerometer, and a Markov process error;

[0056] Step 5, using the broadcast ephemeris to generate satellite data at the simulation time; at the same time, simulating satellite errors through the model;

[0057] Step 6, satellite data includes pseudorange, pseudorange rate, satellite position, satellite speed, and number of visible satellites;

[0058] Step 7, satellite errors include ionosphere error, troposphere error, satellite ephemeris and clock error, and receiver error;

[0059] Step 8, the ionospheric error model is the International Reference 2001 (IRI-2001) model;

[0060] Step 9, the tropospheric error model is a first-order Gaussian Markov process with a correlation time of 30 minutes;

[0061] Step 10, the satellite ephemeris and clock error model is a first-order Gaussian Markov process with a 2-hour correlation time and σ=2m;

[0062] Step 11, the receiver error model is σ multipath [i] = 0.13 + 0.53e (-θ[i] / 10deg) , θ[i] is the elevation angle of the ith satellite; this model is applicable to satellite elevation angles above 5°;

[0063] Step 12, the inertial satellite navigation system calculates the GNSS / IRS combination information and integrity parameters based on the received angular increment and velocity increment containing errors and the satellite data containing errors;

[0064] Step 13, repeating steps 1 to 12 to obtain multiple sets of GNSS / IRS combination information and integrity parameters;

[0065] Step 14, false alarm rate test, testing the false alarm rate in the inertial satellite integrity performance;

[0066] Step 15, generating 40 different geometries in step 13, each of which is simulated for a total of N=82500 hours of operation time;

[0067] Step 16, for each set of satellite data with different geometric structures, set the satellite speed to 0 and fly for at least 82,500 hours in the preset flight trajectory;

[0068] Step 17, counting the total number of false alarms, the total number of alarms on all allowed geometric shapes is 32; for each allowed geometric shape, the alarm shall not exceed 1;

[0069] Step 18, calculate the false alarm rate to be 2.6e-9;

[0070] Step 19, trouble-free accuracy performance test, including snapshot 95% horizontal accuracy test and inertial satellite filter test;

[0071] Step 20, snapshot 95% horizontal accuracy test, The precision statistic is 2drms1=8m;

[0072] Step 21, compare the accuracy statistic 2drms1 with the accuracy of the corresponding route stage, 2drms1<16m, meeting the accuracy requirement;

[0073] Step 22, inertial satellite filter test, should be tested according to the 2drms2 accuracy limit provided by the combined filter. Among them, p 11 、p 22 is the covariance matrix P in Kalman filtering; 2drms2=6m;

[0074] Step 23, 2drms2<16m, meeting the accuracy requirement;

[0075] Step 24, rare normal verification, can be performed simultaneously with the false alarm rate test to verify that the fault-free rare normal HPL (H0) correctly limits the horizontal position error;

[0076] Step 25, calculate the horizontal position error based on the instantaneous latitude and longitude and the combined information in the flight required parameters, Among them, (x1, y1) is the instantaneous latitude and longitude in the flight parameters required, and (x2, y2) is the real-time position information in the combined information; HEL = 7.63m;

[0077] Step 26, compare the HPL in the integrity parameter with the HEL; the integrity parameter HPL = 10.5m;

[0078] Step 27, HPL>HEL, no fault rare normal HPL (H0) can correctly limit the horizontal position error;

[0079] Step 28, detection / elimination verification, testing and verifying the integrity fault detection algorithm and elimination algorithm;

[0080] Step 29, injecting an integrity step fault into the satellite data and error input module in the pseudorange domain;

[0081] Step 30, the inertial satellite navigation system calculates the integrity parameter and calculates HPL = 20m. ICAO uniformly stipulates that the HAL of the LPV-200 approach phase is 16m; HPL>HAL in the integrity parameter indicates that there is a fault in the satellite space segment signal, and fault detection and elimination are performed;

[0082] Step 31, test and verify the standard RAIM performance, use the standard RAIM algorithm to eliminate the fault and calculate the HPL after the fault is eliminated RAIM =15m;

[0083] Step 32, the inertial satellite navigation system calculates the HPL after the fault is eliminated FD =9.8m;

[0084] Step 33, after troubleshooting, calculate the current HEL FD , Among them, (x1, y1) is the instant longitude, latitude and altitude in the flight required parameters, and (x2, y2) is the real-time position information in the combined information after troubleshooting; HEL FD = 5.6m;

[0085] Step 34, compare HEL FD , HPL RAIM , HPL FD and HAL. If HEL FD < HPL FD < HPL RAIM < HAL, it indicates that the integrity fault detection and troubleshooting algorithm in the device under test is effective;

[0086] Step 35, integrated taxi test, test and verify the false alarm rate, missed alarm rate and failure troubleshooting probability during GNSS interruption;

[0087] Step 36, generate satellite data with at least two different geometric structures in the above Step 12. In each scenario, remove enough visible satellites so that the number of visible satellites is less than 4, and ensure that the inertial satellite navigation system is in the taxi state;

[0088] Step 37, for each taxi geometry, at least 1650 tests should be carried out, and count the number of false alarms, missed alarms and failure troubleshooting faults;

[0089] Step 38, calculate the false alarm rate, missed alarm rate and failure troubleshooting probability; false alarm rate P fa = 1.9e-9; missed alarm rate P md = 2e-5; failure troubleshooting probability P fd = 0;

[0090] Step 39: In Appendix 10 of Volume 1 of the civil aviation standards of the International Civil Aviation Organization, the false alarm rate in the LPV-200 operation stage is specified as 3.33e-6, the missed alarm rate is 0.001, and the failure troubleshooting probability is 1e-7; after testing, the false alarm rate, missed alarm rate and failure troubleshooting probability meet the standard requirements, and the integrity performance of this inertial satellite meets the requirements.

[0091] 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 think of various equivalent modifications or replacements within the technical scope disclosed by the present invention, and these modifications or replacements should all be covered within the protection scope of the present invention.

Claims

1. A civil aircraft inertial satellite integrity performance test system based on a semi-physical test platform, applied to the test of inertial satellite navigation system, characterized in that: include: Flight trajectory simulation module, IMU data and error input module, satellite data and error input module, false alarm rate test module, fault-free accuracy performance test module, rare normal verification module, detection / elimination verification module, integrated taxiing module; the flight trajectory simulation module is used to simulate and generate flight parameters required according to flight requirements; the IMU data and error input module is used to reversely solve and generate angular increments and velocity increments output by sensors according to flight parameters required, and simulate injection of errors to obtain angular increments and velocity increments containing errors, convert the angular increments and velocity increments containing errors into digital signals and transmit them to an inertial satellite navigation system; the satellite data and error input module is used to simulate and generate satellite data using broadcast ephemeris according to flight parameters required, and at the same time, obtain satellite data containing errors by introducing a satellite space segment propagation error model, convert the satellite data containing errors into radio frequency signals and transmit them to an inertial satellite navigation system The inertial satellite navigation system calculates the pseudorange, pseudorange rate, GNSS / IRS combination information and integrity parameters based on the received angular increment and velocity increment containing errors and the satellite data containing errors; the false alarm rate test module is used to statistically calculate the false alarm rate based on multiple groups of integrity parameters; the fault-free accuracy performance test module is used to calculate the snapshot 95% horizontal accuracy and the inertial satellite filtering accuracy based on multiple groups of GNSS / IRS combination information; the rare normal verification module is used to verify whether the fault-free rare normal HPL in the integrity parameters can correctly limit the horizontal position error based on the GNSS / IRS combination information and the parameters required for the flight; the detection / elimination verification module is used to verify the integrity detection / elimination algorithm of the inertial satellite navigation system after injecting a fault into the satellite data and error input module; the integrated taxiing module is used to simulate the calculation of the false alarm rate, missed alarm rate and failure elimination probability under the satellite failure state.

2. According to claim 1, a civil aircraft inertial satellite integrity performance test system based on a semi-physical test platform is characterized in that: The flight requirements include initial information, flight path and flight time. The initial information includes initial position, specifically latitude, longitude and altitude; initial attitude, specifically heading angle, roll angle and pitch angle; the flight required parameters include instant latitude and longitude, instant attitude and instant time.

3. The civil aircraft inertial satellite integrity performance test system based on a semi-physical test platform according to claim 1 is characterized in that: The injection error includes zero bias error, proportional factor error, random walk error, non-orthogonal installation error, Gaussian white noise error, and first-order Markov process error.

4. The civil aircraft inertial satellite integrity performance test system based on a semi-physical test platform according to claim 1 is characterized in that: The satellite space segment propagation error model includes an ionosphere error model, a troposphere error model, a satellite clock / ephemeris error model and a receiver noise error model.

5. A civil aircraft inertial satellite integrity performance test method based on a semi-physical test platform, using a civil aircraft inertial satellite integrity performance test system based on a semi-physical test platform as claimed in any one of claims 1 to 4, characterized in that: The specific steps include: S1: Generate flight parameters according to flight requirements; S2: Generate output information of the sensor by reverse calculation according to the parameters required for flight, and simulate injection error to obtain the angular increment and velocity increment containing error, and convert the angular increment and velocity increment containing error into digital signals and transmit them to the inertial satellite navigation system; S3: Generate satellite data using broadcast ephemeris based on the flight parameters. At the same time, obtain satellite data with errors by introducing the satellite space segment propagation error model. Convert the satellite data with errors into radio frequency signals and send them to the inertial satellite navigation system. Satellite data include pseudorange, pseudorange rate, satellite position, satellite speed, and number of visible satellites. S4: The inertial satellite navigation system obtains GNSS / IRS combination information and integrity parameters based on the received angular increment and velocity increment containing errors and the satellite data containing errors; S5: Repeat steps S1-S4 to obtain multiple sets of GNSS / IRS combination information and integrity parameters; S6: Calculate the false alarm rate based on multiple sets of integrity parameters; S7: Calculate snapshot 95% horizontal accuracy and inertial satellite filtering accuracy based on multiple sets of GNSS / IRS combined information; S8: Verify whether the fault-free rare normal HPL in the integrity parameters can correctly limit the horizontal position error based on the GNSS / IRS combined information and the flight required parameters; S9: injecting integrity faults into the satellite data and error input module to verify whether the integrity detection / elimination algorithm of the inertial satellite navigation system is effective; S10: Simulate the satellite outage state and calculate the false alarm rate, missed alarm rate and failure elimination probability.

6. The method for testing the integrity performance of a civil aircraft inertial satellite based on a semi-physical test platform according to claim 5, characterized in that: The step S6 specifically includes the following steps: S61: generating at least 40 sets of satellite data with different geometric structures in step 3; S62: For each set of satellite data with different geometric structures, set the satellite speed to 0, fly for at least 82,500 hours in the preset flight trajectory, and count the number of alarms N in the integrity parameters. alarm ; S63: Calculate the false alarm rate according to the following formula (1): N all Indicates the total number of combined navigation times.

7. The method for testing the integrity performance of a civil aircraft inertial satellite based on a semi-physical test platform according to claim 5, characterized in that: In step S7, the 95% horizontal accuracy of the snapshot is calculated according to the following formula: Among them, d i is the instantaneous two-dimensional horizontal position error (meters), N is the number of sampling points, and HDOP is the instantaneous horizontal precision factor.

8. The method for testing the integrity performance of a civil aircraft inertial satellite based on a semi-physical test platform according to claim 5, characterized in that: In step S7, the inertial satellite filtering accuracy is calculated using the following formula according to the inertial satellite navigation system combined filter: Among them, p 11 、p 22 is the covariance matrix P in the Kalman filter.

9. The method for testing the integrity performance of a civil aircraft inertial satellite based on a semi-physical test platform according to claim 5, characterized in that: In the step S8, the following steps are specifically included: S81: Calculate the horizontal position error based on the instantaneous latitude and longitude and combined information in the flight parameters required. Among them, (x1, y1) is the instant latitude and longitude in the flight parameters required, and (x2, y2) is the real-time position information in the combined information; S82: According to the HPL in the integrity parameter, compare with HEL; S83: If HPL>HEL, the fault-free, rare, and normal HPL (H0) can correctly limit the horizontal position error; otherwise, the fault-free, rare, and normal HPL (H0) cannot correctly limit the horizontal position error.

10. The method for testing the integrity performance of a civil aircraft inertial satellite based on a semi-physical test platform according to claim 5, characterized in that: In the step S9, the following steps are specifically included: S91: injecting an integrity fault into the satellite data and error input module in the pseudorange domain, where the integrity fault is divided into an integrity step fault and an integrity ramp fault; S92: The integrity parameters are calculated by the inertial satellite navigation system. If HPL>HAL in the integrity parameters, it indicates that a fault has been detected in the satellite space segment signal. HAL is the horizontal warning limit of the current flight path phase, which is uniformly specified by ICAO. S93: Testing and verifying the performance of the standard receiver autonomous integrity monitoring (RAIM), wherein the standard RAIM algorithm built into the detection / elimination verification module eliminates the fault and calculates the HPL after the fault is eliminated. RAIM ; S94: HPL after troubleshooting calculated by the inertial satellite navigation system FD ; S95: After troubleshooting, calculate the current HEL FD , Among them, (x1, y1) is the instant latitude and longitude in the flight parameters required, and (x2, y2) is the real-time position information in the combined information after the fault is eliminated; S96: Compare HEL FD , HPL RAIM , HPL FD and HAL. If HEL FD < HPL FD < HPL RAIM < HAL, it indicates that the integrity fault detection and elimination algorithm in the device under test is effective.

11. The method for testing the integrity performance of a civil aircraft inertial satellite based on a semi-physical test platform according to claim 5, characterized in that: In the step S10, the following steps are specifically included: S101: generating satellite data of at least two different geometric structures in step 3, and in each scenario, removing enough visible satellites so that the number of visible satellites is less than 4, ensuring that the inertial satellite navigation system is in a gliding state; S102: For each taxiing geometry, at least 1,650 tests shall be conducted to count the number of false alarms, missed alarms, and failed troubleshooting. S103: Calculate the false alarm rate, missed alarm rate and failure elimination probability; false alarm rate, Missed alarm rate, Failure exclusion probability, Among them, N fa is the number of alarms when there is no fault, N md is the number of times no alarm is given when there is a fault, N fd N is the number of times when there is a fault and the failure to eliminate it occurs. all is the total number of trials.

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