Warship landing-oriented inertia / radar combination integrity monitoring method

Abstract

The invention discloses a carrier landing-oriented inertia / radar combination integrity monitoring method. Based on whether there is a fault, dividing the total integrity risk into a fault-free integrity risk and an integrity risk based on a fault detection mode; according to the integrity index requirement, the probability of danger misleading information under each risk type is set; establishing a filtering state equation model, and further setting a measurement equation model; calculating an integrity level protection threshold HPLFF under a fault-free condition according to the measurement output by the measurement equation model in combination with the probability of danger misleading information under each risk type; calculating an integrity level protection threshold HPLFD in a fault detection mode according to the measurement output by the measurement equation model in combination with the probability of danger misleading information under each risk type; and according to the HPLFF and the HPLFD, calculating the integrity protection level of the inertia / satellite combination output parameters. In a precise carrier landing link with the highest accident rate, a confidence index required based on a scene can be realized, and a horizontal protection threshold under the scheme can be output.

Description

Technical Field

[0001] The invention belongs to the technical field of combined navigation integrity, and in particular relates to a landing-oriented inertial / radar combined integrity monitoring method. Background Art

[0002] Precision landing is the link with the highest accident rate. Statistics show that about 80% of accidents involving carrier-based aircraft occur during the landing process, so how to ensure the reliability of the landing process is one of the key technologies to achieve precision landing. Given that the satellite navigation system radar is an important part of the ship-borne air navigation system, it plays an irreplaceable role in guiding carrier-based aircraft to land safely. The landing guidance radar measures the relative position information of the aircraft, such as azimuth, elevation and distance, in real time, and transmits it to the take-off and landing guidance command facilities through a wireless channel. Combined navigation based on inertial / radar can make up for the dangerous scenario where RTK relative positioning fails in a denied environment and cannot provide accurate relative positioning for landing.

[0003] The accuracy of radar ranging and direction finding is easily affected by atmospheric conditions, synchronization errors, light spot overlap accuracy, antenna height, etc. It is also easily interfered by other surrounding signal sources, such as power lines, buildings, radio equipment, etc. Therefore, the integrity of the inertial / radar combination should be monitored in real time during the landing process, and an alarm should be issued in time when a fault occurs. Summary of the invention

[0004] The purpose of the present invention is to provide a method for monitoring the integrity of an inertial / radar integrated navigation system for landing, to achieve a confidence index required based on a scenario, to output a horizontal protection threshold under the scheme, and to give an alarm in time when the position error exceeds the threshold, so as to ensure the integrity of the system.

[0005] Technical solution of the present invention: In order to achieve the above-mentioned invention object, a landing-oriented inertial / radar combination integrity monitoring method is proposed, comprising the following steps:

[0006] Step 1: Based on whether there is a fault, the total integrity risk is divided into the integrity risk without fault and the integrity risk under the fault detection mode;

[0007] Step 2: According to the integrity index requirements of the inertial / radar integrated navigation system, set the probability of dangerous misleading information under each risk type in step 1, and record them as the probability of no fault dangerous misleading information PHMI FF and PHMI based on the probability of dangerous misleading information for fault detection FD ;

[0008] Step 3: Establish a filter state equation model for the landing-oriented inertial / radar integrated navigation system, and further set the measurement equation model;

[0009] Step 4: Based on the measurement output by the measurement equation model in step 3 and the probability of dangerous misleading information under each risk type in step 2, calculate the integrity level protection threshold HPL under no fault FF ;

[0010] Step 5: Based on the measurement output by the measurement equation model in step 3 and the probability of dangerous misleading information under each risk type in step 2, calculate the integrity level protection threshold HPL in the fault detection mode FD ;

[0011] Step 6: Obtain the integrity level protection threshold HPL under no fault conditions obtained in step 4 FF , the integrity level protection threshold HPL under the fault detection mode obtained in step 5 FD , the integrity protection level of the inertial / satellite combination output parameters is calculated using the following formula:

[0012] HPL F =max(HPL GNSS ,HPL H0 ,HPL FD )

[0013] Among them, HPL GNSS It is the horizontal protection threshold of the satellite space signal, which is transmitted from the outside.

[0014] In a possible embodiment, in step 3, the specific process of establishing the filtering state equation model of the landing-oriented inertial / radar integrated navigation system includes:

[0015] The state quantity of the inertial navigation in the earth coordinate system e is defined as follows:

[0016]

[0017] Among them, δP e is the 3D position error in the e system, δV e is the 3D velocity error in the e system, is the 3D attitude misalignment angle error, δb a is the 3D accelerometer zero bias, δb g It is the 3D gyro bias;

[0018] The filter state equation model of the inertial / radar integrated navigation system is as follows:

[0019]

[0020] Next, we update the time and calculate the state one-step prediction X according to the following formula: k / k-1 , one-step prediction mean square error matrix P k / k-1 :

[0021] X K / K-1 =φ K / K-1 ·X K-1

[0022]

[0023] Among them, the noise parameter matrix Q K-1 According to the system noise matrix, the specific value can be set according to the actual situation.

[0024] In a possible embodiment, in step 3, the specific process of setting the measurement equation model includes:

[0025] Based on the status update results, perform measurement updates:

[0026] Let L D , D 、h D is the latitude, longitude and altitude of the radar ranging station, L ψ , ψ 、h ψ is the latitude, longitude and altitude of the radar station, L θ , θ 、h θ is the latitude, longitude and altitude of the radar station; ρ D ,ψ D ,θ D They are radar measurement parameters, slant range, azimuth, and pitch;

[0027] The latitude, longitude and altitude of the radar ranging station are converted to the earth's fixed rectangular coordinate system according to the following formula:

[0028] Obtain the coordinates of the radar ranging station in the earth's fixed rectangular coordinate system:

[0029]

[0030] Obtain the coordinates of the radar station in the earth's fixed rectangular coordinate system:

[0031]

[0032] Find the coordinates of the radar height station in the earth's fixed rectangular coordinate system:

[0033]

[0034] Among them, R N is the radius of the Earth, e is the eccentricity of the Earth;

[0035] Calculate the corrected inertial navigation position:

[0036] Lat=Lat_ins-X INS_k / k-1[1]

[0037] Lon=Lon_ins-X INS_k / k-1 [2]

[0038] height=hX INS_k / k-1 [3]

[0039] Among them, X INS_k / k-1 is the state transfer variable of one step, Lat_ins is the local latitude of inertial solution, Lon_ins is the local longitude of inertial solution, and h is the inertial atmospheric damping height; the corrected inertial navigation position is converted from the geographic system to the earth fixed rectangular coordinate system:

[0040] x=(R N +height)cos(Lat)cos(Lon)

[0041] y=(R N +height)cos(Lat)sin(Lon)

[0042] z=[R N (1-e 2 )+height]sin(Lat)

[0043] Based on the corrected inertial navigation position (x, y, z) and station position (x D ,y D ,z D ), calculate the intermediate measurement according to the following formula:

[0044]

[0045] in,

[0046] Calculate the bearing intermediate measurement according to the following formula:

[0047]

[0048] in,

[0049] Calculate the intermediate height measurement using the following formula:

[0050]

[0051] in, Based on the intermediate measurements in the above steps, the equivalent measurements are calculated as follows:

[0052] Based on the intermediate measurement of the distance measurement, calculate the equivalent slope distance:

[0053]

[0054] Based on the intermediate measurement of the azimuth, calculate the equivalent azimuth angle ψ C :

[0055]

[0056] In the above formula, Δλ=Lon-λ ψ ;

[0057] Based on the intermediate measurement of the azimuth, calculate the equivalent altitude angle:

[0058]

[0059] In summary, the calculation equation is as follows:

[0060]

[0061] The measurement transfer matrix is ​​calculated as follows:

[0062]

[0063] Among them, the B matrix is ​​calculated using the corrected inertial navigation position Lat, Lon, height:

[0064]

[0065] Substituting the above calculation results into the first 3×3 dimensions of the Hk matrix, which is a (3*number of states)-dimensional matrix, H k The rest of the elements are 0;

[0066] R k The matrix determines the parameters according to the actual situation;

[0067] Based on this, the covariance P of the measurement update result is calculated k and the state estimate X K :

[0068]

[0069] X K =X K / K-1 +K K ·Z K

[0070]

[0071] Among them, K K is the filter gain coefficient, R K is the measurement noise matrix.

[0072] In a possible embodiment, in step 4, the following steps are specifically included:

[0073] Step 4.1 Calculate the attitude projection matrix from the Earth-centered Earth-fixed coordinate system ECEF to the navigation system Then obtain the projection of the filter covariance in step 3 under the N system:

[0074]

[0075] in, The calculation method is as follows:

[0076]

[0077] Among them, R N is the radius of the earth, e is the eccentricity of the earth, Lat is the local latitude, and Lon is the local longitude;

[0078] The converted inertial / radar combined filter Pk N Matrix calculates the largest eigenvalue of the covariance matrix:

[0079]

[0080] Step 4.2 calculates the horizontal protection level of the fault-free integrity risk FF assumption based on the maximum eigenvalue of the covariance matrix:

[0081]

[0082] Among them, K FF Based on PHMI FF According to the Rayleigh distribution confidence interval calculation, the details are as follows:

[0083] K FF =raylinv(1-PHMI FF ).

[0084] In a possible embodiment, in step 5, the following steps are specifically included:

[0085] Step 5.1 Calculate the measurement residual: r K =Z K -H k X k / k-1

[0086] Step 5.2: Based on step 5.1, perform fault detection analysis:

[0087] In the absence of faults, r K =Z K -H k X k / k-1 Subordinate to the normal distribution with zero mean N(0,A K ), where A K is the variance of the normal distribution, and the specific calculation method is as follows:

[0088] A K =H K P K (H K ) T +R K

[0089] Step 5.3: For each dimension of the measurement residual r i Construct the detection function λ i =r i T r i , its detection threshold D i for:

[0090] D i =K FA *sqrt(A kii )

[0091] Where K FA It is determined based on the false alarm rate index PML of the system for radar failure. The specific calculation method is as follows:

[0092] K FAi =Q -1 (1-P ml i)

[0093] Project the above Di onto the geographic coordinate system, and we get D H .

[0094] Step 5.4 then obtains the horizontal detection threshold:

[0095]

[0096] Step 5.5 is based on step 5.4 to obtain the integrity protection level:

[0097] HPL FD Calculated by the inertial / radar combined detection capability, the detection threshold of the residual detection is projected into the geographic coordinate system

[0098] Calculate the integrity protection level:

[0099] HPL FD =D H +K Hmd σ H

[0100] Among them, K Hmd =raylinv(1-PHMI FD i)

[0101] σ H is the 1σ value of the inertial / radar combination horizontal error.

[0102] According to a second aspect of the present invention, a landing-oriented inertial / radar combination integrity monitoring device is proposed, which is used to implement the above-mentioned landing-oriented inertial / radar combination integrity monitoring method, and is characterized in that it includes: a filter state equation model construction unit of an inertial / radar combined navigation system, which is used to establish a filter state equation model of an inertial / radar combined navigation system for landing; a measurement equation model construction unit, which is used to further set a measurement equation model according to the filter state equation model of the inertial / radar combined navigation system for landing; and a fault-free integrity level protection threshold calculation unit, which is used to calculate a fault-free integrity level protection threshold HPL according to the measurement output by the measurement equation model and the probability of dangerous misleading information under each risk type. FF ; The integrity level protection threshold calculation unit in the fault detection mode is used to calculate the integrity level protection threshold HPL in the fault detection mode according to the measurement output by the measurement equation model and the probability of dangerous misleading information under each risk type. FD ; Inertial / satellite combined inertial navigation system integrity protection level calculation unit, used to calculate the inertial / satellite combined inertial navigation system integrity protection level HPL GNSS .

[0103] According to a third aspect of the present invention, a computer storage medium is provided, storing a computer program, and when the computer program is executed by a processor, the above-mentioned inertial / radar combined integrity monitoring method for landing is implemented.

[0104] According to a fourth aspect of the present invention, a computer program product comprising instructions is provided, and when the computer program product is run on a computer, the computer is caused to execute the above-mentioned inertial / radar combined integrity monitoring method for landing.

[0105] The beneficial technical effects of the present invention are as follows: the present invention provides a method for monitoring the integrity of an inertial / radar combined navigation for landing. In the precision landing link where the accident rate is the highest, the confidence index required based on the scenario can be realized, the horizontal protection threshold under the scheme can be output, and an alarm can be issued in time when the position error exceeds the threshold, so as to avoid the introduction of erroneous information into the landing process when the radar ranging and direction finding accuracy is adversely affected by the outside world and interfered by other surrounding signal sources, thereby causing safety risks. BRIEF DESCRIPTION OF THE DRAWINGS

[0106] Figure 1 A flow chart of a method according to a preferred embodiment of the present invention;

[0107] Figure 2 It is a detailed step diagram of a preferred embodiment of the present invention. DETAILED DESCRIPTION

[0108] 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.

[0109] The features and illustrative embodiments of various aspects of the present invention will be described in detail below. In the detailed description below, many specific details are proposed to provide a comprehensive understanding of the present invention. However, it is obvious to those skilled in the art that the present invention can be implemented without the need for some of these specific details. The following description of the embodiments is only to provide a better understanding of the present invention by illustrating examples of the present invention. The present invention is by no means limited to any specific settings and methods proposed below, but covers any improvements, replacements and modifications of structures, methods, devices without departing from the spirit of the present invention. In the accompanying drawings and the following description, known structures and technologies are not shown to avoid unnecessary ambiguity in the present invention.

[0110] 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.

[0111] A method for integrity monitoring of inertial / radar integrated navigation for landing, such as Figure 1 As shown, the following steps are included:

[0112] Step 1. Classify integrity risks.

[0113] The total integrity risk is divided into three categories: fault-free integrity risk (abbreviated as FF: FaultFree in this scenario below) and integrity risk based on fault detection mode (abbreviated as FD: Fault Detected in this scenario below).

[0114] Step 2: Classification of integrity indicators.

[0115] According to the system integrity index requirements, the total probability of causing system dangerous misleading information (Misleading Hazardous Information) is defined as PHMI (probability of Misleading Hazardous Information)

[0116] According to the scenario in step 1, the probability of dangerous misleading information is divided into: probability of no fault dangerous misleading information PHMIFF and PHMI based on the probability of dangerous misleading information for fault detection FD , the PHMI based on the probability of dangerous misleading information for fault detection FD According to the distance measurement, azimuth angle and altitude angle, they are: distance measurement PHMI FD1 , Azimuth PHMI FD2 , height angle PHMI FD3 .

[0117] Step 3. Establish the filtering state equation model of the inertial / radar integrated navigation system for landing, such as Figure 2 As shown, the measurement equation model is further set;

[0118] Step 3.1 The filter state variables of the integrated navigation system are updated as follows:

[0119] The state quantity of the inertial navigation in the earth coordinate system e is defined as follows:

[0120]

[0121] Among the state variables, δP e is the 3D position error in the e system, δV e is the 3D velocity error in the e system, is the 3D attitude misalignment angle error, δb a is the 3D accelerometer zero bias, δb g It is the 3D gyro bias.

[0122] The system state equation is as follows:

[0123]

[0124] Next, the time is updated and the state one-step prediction Xk / k-1 and one-step prediction mean square error matrix Pk / k-1 are calculated according to the following formula:

[0125] X K / K-1 =φ K / K-1 ·X K-1

[0126]

[0127] Among them, Q K-1 According to the system noise matrix, the specific value can be set according to the actual situation.

[0128] Step 3.2: Update the measurement based on the status update result:

[0129] Let L D , D 、h D is the latitude, longitude and altitude of the radar ranging station, L ψ, ψ 、h ψ is the latitude, longitude and altitude of the radar station, L θ , θ 、h θ is the latitude, longitude and altitude of the radar station;

[0130] ρ D ,ψ D ,θ D Radar measurement parameters: "slant range", "azimuth" and "pitch";

[0131] Convert the station latitude, longitude and height to the earth fixed rectangular coordinate system:

[0132]

[0133] Among them, R N is the radius of the Earth, and e is the eccentricity of the Earth.

[0134] Calculate the corrected inertial navigation position:

[0135] Lat=Lat_ins-X INS_k / k-1 [1]

[0136] Lon=Lon_ins-X INS_k / k-1 [2]

[0137] height=hX NS_k / k-1 [3]

[0138] Among them, Lat_ins is the local latitude of the inertial solution, and Lon_ins is the local longitude of the inertial solution. Convert the corrected position from the geographic system to the earth-fixed rectangular coordinate system:

[0139] x=(R N +height)cos(Lat)cos(Lon)

[0140] y=(R N +height)cos(Lat)sin(Lon)

[0141] z=[R N (1-e 2 )+height]sin(Lat)

[0142] Based on the inertial correction position (x, y, z) in the above earth-fixed rectangular coordinate system and the station position (x D ,y D ,z D ), further calculate the ranging intermediate measurement:

[0143]

[0144] in,

[0145] Similarly, calculate the intermediate measurement of the orientation:

[0146]

[0147] in,

[0148] Similarly, calculate the intermediate measurement of height:

[0149]

[0150] in,

[0151] Based on the intermediate measurements in the above steps, the equivalent measurements are calculated as follows:

[0152] Based on the intermediate measurement of the distance measurement, calculate the equivalent slope distance:

[0153]

[0154] Based on the intermediate measurement of the azimuth, calculate the equivalent azimuth angle ψ C :

[0155]

[0156] In the above formula, Δλ=Lon-λ ψ ;

[0157] Based on the intermediate measurement of the azimuth, calculate the equivalent altitude angle:

[0158]

[0159] In summary, the calculation equation is as follows:

[0160]

[0161] The measurement transfer matrix is ​​calculated as follows:

[0162]

[0163] Among them, the B matrix is ​​calculated using the corrected inertial navigation position Lat, Lon, height:

[0164]

[0165] Substituting the above calculation results into the first 3×3 dimensions of the Hk matrix, which is a (3*number of states)-dimensional matrix, the remaining elements of Hk are 0;

[0166] R kThe matrix determines the parameters according to the actual situation;

[0167] Based on this, the covariance P of the measurement update result is calculated k and the state estimate X K :

[0168]

[0169] X K =X K / K-1 +K K ·Z K

[0170] P k =(IK k *H k )*P k / k-1 *(IK k *H k ) T +K k *R k *K k T .

[0171] Among them, K K is the filter gain coefficient, R K is the measurement noise matrix.

[0172] Step 4. Figure 2 As shown, the integrity index is further decomposed, and the integrity level protection threshold HPL under no fault is calculated according to the measurement output of the measurement equation model in step 3. FF ;

[0173] Step 4.1 Calculate the attitude projection matrix from the Earth-centered Earth-fixed coordinate system ECEF to the navigation system Then obtain the projection of the filter covariance in step 3 under the N system:

[0174]

[0175] in, The calculation method is as follows:

[0176]

[0177] Among them, R N is the radius of the earth, e is the eccentricity of the earth, Lat is the local latitude, and Lon is the local longitude;

[0178] The converted inertial / satellite combined filter Pk N Matrix calculates the largest eigenvalue of the covariance matrix:

[0179]

[0180] Step 4.2 calculates the horizontal protection level of the fault-free integrity risk FF assumption based on the maximum eigenvalue of the covariance matrix:

[0181]

[0182] Among them, K H0 Based on PHMI FF According to the Rayleigh distribution confidence interval calculation, the details are as follows:

[0183] K FF =raylinv(1-PHMI FF ).

[0184] Step 5. Calculate the integrity level protection threshold HPL in the fault detection mode according to the measurement output of the measurement equation model in step 3 FD ;

[0185] Step 5.1 Calculate the measurement residual: r K =Z K -H k X k / k-1

[0186] Step 5.2: Based on step 5.1, perform fault detection analysis:

[0187] In the absence of faults, r K =Z K -H k X k / k-1 Subordinate to the normal distribution with zero mean N(0,A K ), where A K is the variance of the normal distribution, and the specific calculation method is as follows:

[0188] A K =H K P K (H K ) T +R K

[0189] Step 5.3: For each dimension of the measurement residual r i Construct the detection function λ i =r i T r i , its detection threshold D i for:

[0190] D i =K FA *sqrt(A kii )

[0191] Where KFA According to the false alarm rate index P of the system for radar failure ML The specific calculation method is as follows:

[0192] K FAi =Q -1 (1-P ml i)

[0193] The above D i Projected onto the geographic coordinate system, we get D H .

[0194] Step 5.4 then obtains the horizontal detection threshold:

[0195]

[0196] Step 5.5 is based on step 5.4 to obtain the integrity protection level:

[0197] HPL FD The detection threshold of residual detection is calculated based on the inertial / radar combined detection capability, and the integrity protection level is calculated by projecting the detection threshold of residual detection into the geographic coordinate system:

[0198] HPL FD =D H +K Hmd σ H

[0199] Among them, K Hmd =raylinv(1-PHMI FD i / P Fault i), P Fault i is the probability that the measurement failure in this dimension contains bias, i.e. PHMI FD i / P Fault i is the probability of missed fault alarm.

[0200] σ H is the 1σ value of the inertial / radar combination horizontal error.

[0201] Step 6: The integrity level protection threshold HPL under no fault condition obtained in step 4 FF , the integrity level protection threshold HPL under the fault detection mode obtained in step 5 FD Calculate the integrity protection level of the inertial / satellite combination output parameters:

[0202] HPL F =max(HPL GNSS ,HPL H0 ,HPL FD )

[0203] Among them, HPL GNSSIt is the horizontal protection threshold of the satellite space signal, which is transmitted from the outside.

Claims

1. A landing-oriented inertial / radar combination integrity monitoring method, characterized in that: The steps include: Step 1: Based on whether there is a fault, the total integrity risk is divided into the integrity risk without fault and the integrity risk under the fault detection mode; Step 2: According to the integrity index requirements of the inertial / radar integrated navigation system, set the probability of dangerous misleading information under each risk type in step 1, and record them as the probability of no fault dangerous misleading information PHMI FF and PHMI based on the probability of dangerous misleading information for fault detection FD ; Step 3: Establish a filter state equation model for the landing-oriented inertial / radar integrated navigation system, and further set the measurement equation model; Step 4: Based on the measurement output by the measurement equation model in step 3 and the probability of dangerous misleading information under each risk type in step 2, calculate the integrity level protection threshold HPL under no fault FF ; Step 5: Based on the measurement output by the measurement equation model in step 3 and the probability of dangerous misleading information under each risk type in step 2, calculate the integrity level protection threshold HPL in the fault detection mode FD ; Step 6: Obtain the integrity level protection threshold HPL under no fault conditions obtained in step 4 FF , the integrity level protection threshold HPL under the fault detection mode obtained in step 5 FD , the integrity protection level of the inertial / satellite combination output parameters is calculated using the following formula: HPL F =max(HPL GNSS ,HPL H0 ,HPL FD ) Among them, HPL GNSS It is the horizontal protection threshold of the satellite space signal, which is transmitted from the outside.

2. The landing-oriented inertial / radar combination integrity monitoring method according to claim 1, characterized in that: In step 3, the specific process of establishing the filtering state equation model of the landing-oriented inertial / radar integrated navigation system includes: defining the state quantity of the inertial navigation in the earth coordinate system e as follows: Among them, δP e is the 3D position error in the e system, δV e is the 3D velocity error in the e system, is the 3D attitude misalignment angle error, δb a is the 3D accelerometer zero bias, δb g It is the 3D gyro bias; The filter state equation model of the inertial / radar integrated navigation system is as follows: Next, we update the time and calculate the state one-step prediction X according to the following formula: k / k-1 , one-step prediction mean square error matrix P k / k-1 : X K / K-1 =φ K / K-1 ·X K-1 Among them, the noise parameter matrix Q K-1 According to the system noise matrix, the specific value can be set according to the actual situation.

3. The landing-oriented inertial / radar combination integrity monitoring method according to claim 2, characterized in that: In step 3, the specific process of setting the measurement equation model includes: performing measurement update based on the state update result: Let L D , D 、h D is the latitude, longitude and altitude of the radar ranging station, L ψ , ψ 、h ψ is the latitude, longitude and altitude of the radar station, L θ , θ 、h θ is the latitude, longitude and altitude of the radar station; ρ D ,ψ D ,θ D They are radar measurement parameters, slant range, azimuth, and pitch; The latitude, longitude and altitude of the radar ranging station are converted to the earth's fixed rectangular coordinate system according to the following formula: Obtain the coordinates of the radar ranging station in the earth's fixed rectangular coordinate system: Obtain the coordinates of the radar station in the earth's fixed rectangular coordinate system: Find the coordinates of the radar height station in the earth's fixed rectangular coordinate system: Among them, R N is the radius of the Earth, e is the eccentricity of the Earth; Calculate the corrected inertial navigation position: Lat=Lat_ins-X INS_k / k-1 [1] Lon=Lon_ins-X INS_k / k-1 [2] height=h-X INS_k / k-1 [3] Among them, X INS_k / k-1 is the state transfer variable of one step, Lat_ins is the local latitude of inertial solution, Lon_ins is the local longitude of inertial solution, and h is the inertial atmospheric damping height; the corrected inertial navigation position is converted from the geographic system to the earth fixed rectangular coordinate system: x=(R N +height)cos(Lat)cos(Lon) y=(R N +height)cos(Lat)sin(Lon) z=[R N (1-e 2 )+height]sin(Lat) Based on the corrected inertial navigation position (x, y, z) and station position (x D ,y D ,z D ), calculate the intermediate measurement according to the following formula: in, Calculate the bearing intermediate measurement according to the following formula: in, Calculate the intermediate height measurement using the following formula: in, Based on the intermediate measurements in the above steps, the equivalent measurements are calculated as follows: Based on the intermediate measurement of the distance measurement, calculate the equivalent slope distance: Based on the intermediate measurement of the azimuth, calculate the equivalent azimuth angle ψ C : In the above formula, Δλ=Lon-λ ψ ; Based on the intermediate measurement of the azimuth, calculate the equivalent altitude angle: In summary, the calculation equation is as follows: The measurement transfer matrix is ​​calculated as follows: Among them, the B matrix is ​​calculated using the corrected inertial navigation position Lat, Lon, height: Substituting the above calculation results into the first 3×3 dimensions of the Hk matrix, which is a (3*number of states)-dimensional matrix, H k The rest of the elements are 0; R k The matrix determines the parameters according to the actual situation; Based on this, the covariance P of the measurement update result is calculated k and the state estimate X K : X K =X K / K-1 +K K ·Z K P k =(I-K k *H k )*P k / k-1 *(I-K k *H k ) T +K k *R k *K k T 。 Among them, K K is the filter gain coefficient, R K is the measurement noise matrix.

4. The landing-oriented inertial / radar combination integrity monitoring method according to claim 3, characterized in that: In step 4, the following steps are specifically included: Step 4.1 Calculate the attitude projection matrix from the Earth-centered Earth-fixed coordinate system ECEF to the navigation system Then obtain the projection of the filter covariance in step 3 under the N system: in, The calculation method is as follows: Among them, R N is the radius of the earth, e is the eccentricity of the earth, Lat is the local latitude, and Lon is the local longitude; The converted inertial / radar combined filter Pk N Matrix calculates the largest eigenvalue of the covariance matrix: Step 4.2 calculates the horizontal protection level of the fault-free integrity risk FF assumption based on the maximum eigenvalue of the covariance matrix: Among them, K FF Based on PHMI FF According to the Rayleigh distribution confidence interval calculation, the details are as follows: K FF =raylinv(1-PHMI FF )。 5. The landing-oriented inertial / radar combination integrity monitoring method according to claim 3, characterized in that: In the step 5, the following steps are specifically included: Step 5.1 Calculate the measurement residual: r K =Z K -H k X k / k-1 Step 5.2: Based on step 5.1, perform fault detection analysis: In the absence of faults, r K =Z K -H k X k / k-1 Subordinate to the normal distribution with zero mean N(0,A K ), where A K is the variance of the normal distribution, and the specific calculation method is as follows: A K =H K P K (H K ) T +R K Step 5.3: For each dimension of the measurement residual r i Construct the detection function λ i =r i T r i , its detection threshold D i for: <h2 style=";text-align:left;direction:ltr">D<h2 style=";text-align:left;direction:ltr"> i <h2 style=";text-align:left;direction:ltr"> =K<h2 style=";text-align:left;direction:ltr"> FA <h2 style=";text-align:left;direction:ltr"> *sqrt(A<h2 style=";text-align:left;direction:ltr"> kii <h2 style=";text-align:left;direction:ltr"> ) Where K FA According to the false alarm rate index P of the system for radar failure ML The specific calculation method is as follows: K FAi =Q -1 (1-P ml i) The above D i Projected onto the geographic coordinate system, we get D H . Step 5.4 then obtains the horizontal detection threshold: Step 5.5 is based on step 5.4 to obtain the integrity protection level: HPL FD Calculated by the inertial / radar combined detection capability, the detection threshold of the residual detection is projected into the geographic coordinate system Calculate the integrity protection level: HPL FD =D H +K Hmd σ H Among them, K Hmd =raylinv(1-PHMI FD i) σ H is the 1σ value of the inertial / radar combination horizontal error.

6. A landing-oriented inertial / radar combination integrity monitoring device, used to implement a landing-oriented inertial / radar combination integrity monitoring method according to any one of claims 1 to 5, characterized in that: include: The filter state equation model construction unit of the inertial / radar integrated navigation system is used to establish the filter state equation model of the inertial / radar integrated navigation system for landing; the measurement equation model construction unit is used to further set the measurement equation model according to the filter state equation model of the inertial / radar integrated navigation system for landing; the integrity level protection threshold calculation unit under no fault is used to calculate the integrity level protection threshold HPL under no fault according to the measurement output by the measurement equation model and the probability of dangerous misleading information under each risk type. FF ; The integrity level protection threshold calculation unit in the fault detection mode is used to calculate the integrity level protection threshold HPL in the fault detection mode according to the measurement output by the measurement equation model and the probability of dangerous misleading information under each risk type. FD ; Inertial / satellite combined inertial navigation system integrity protection level calculation unit, used to calculate the inertial / satellite combined inertial navigation system integrity protection level HPL GNSS .

7. A computer storage medium, characterized in that: A computer program is stored, and when the computer program is executed by a processor, a landing-oriented inertial / radar combination integrity monitoring method as described in any one of claims 1 to 5 is implemented.

8. A computer program product comprising instructions, characterized in that When the computer program product runs on a computer, the computer is enabled to execute the landing-oriented inertial / radar combination integrity monitoring method as described in any one of claims 1 to 5.

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