CN115683092A - DME/DME/SINS tightly-combined navigation system relocation method - Google Patents

Abstract

The DME/DME/SINS tightly-combined navigation system repositioning method disclosed by the invention has the anti-interference capability and can meet the requirements of navigation instantaneity and reliability. The invention is realized by the following technical scheme: according to the measured output value of the inertial device, SINS navigation calculation is completed, the attitude, the speed and the position of the airplane are output, the pseudo-slant distance from the airplane to the DME station is calculated according to the calculated pseudo-slant distance, the calculated pseudo-slant distance and the slant distance obtained by DME airborne equipment are differentiated to form slant distance measurement information, and the altitude output by the atmospheric data machine are differentiated to form altitude measurement information; establishing a Kalman filtering state equation and a mathematical relation between a slope distance, a height error and a state variable on the basis of an SINS attitude, speed and position error equation so as to establish a Kalman filtering measurement equation; and (3) the SINS attitude, speed and position are corrected by using the estimated attitude, speed and position errors in a feedback manner, so that the SINS navigation error is suppressed, and navigation information with higher precision is output.

Description

Relocation Method of DME/DME/SINS Tightly Integrated Navigation System

technical field

The invention relates to a rapid relocation technology of a DME/DME/SINS tight integrated navigation system during the flight of a carrier under high dynamic, weak signal and other working environments.

Background technique

With the continuous development of navigation technology, the requirements for the accuracy and stability of the navigation system are getting higher and higher. Modern aviation industry involves more and more technologies, and they are becoming more and more complex. Aircraft is no longer a simple means of transportation, but a collection of multiple technologies. There are certain limitations in the performance of a single navigation system, which often cannot Meet the positioning requirements in special environments. The integrated navigation system can give full play to the advantages of each system and make the advantages of each navigation system complement each other, so as to ensure the accuracy of the whole system and improve the stability of the system. In view of this advantage, the integrated navigation system has been developed rapidly. Among them, strapdown inertial navigation system (SINS) and GNSS integrated navigation system are the most widely used. Due to the shortcomings of GNSS, such as poor dynamic response capability, susceptibility to electronic interference, signal occlusion and poor integrity, the quality of GNSS signals is poor in some special environments. In this case, GNSS/SINS integrated navigation cannot Normal operation, the system enters the pure SINS calculation mode. Due to the existence of error sources such as inertial device measurement errors, its navigation and positioning errors accumulate over time, and relying on SINS to work independently cannot meet the needs of long-term navigation. Therefore, it is urgent to design a backup navigation system of the GNSS/SINS integrated navigation system to meet the needs of aircraft route navigation under GNSS denial conditions.

The ground-based radio navigation system is an international civil aviation navigation system, and it is the basic system for civil aviation organization and implementation of flight, safety production and operation management. The positioning accuracy of the land-based radio navigation system is relatively poor, but its signal transmission power is large, and it has the advantages that the satellite navigation system does not have, such as being less susceptible to interference and having a higher data update rate. The distance measuring machine (DME) is a high-precision non-autonomous time-ranging short-range navigation system. It is an important ground-based ranging navigation system for civil aviation. It can provide aircraft with a slant distance to the DME station. The DME rangefinder adopts the query-response distance measurement method, which is a general ranging guidance equipment for civil aviation all over the world. It can be used as both route navigation equipment and airport navigation equipment. DME obtains distance information by measuring the propagation time of radio waves in space. The ranging process is divided into search, tracking and memory stages. The DME system consists of two parts: the airborne interrogator and the ground transponder. By accurately measuring the total time of sending and receiving signals and the fixed delay of the ground response part, the distance information from the aircraft to the ground DME station is calculated. DME/DME navigation is widely used in the en-route navigation stage in performance-based navigation applications. DME/DME navigation is a positioning method. Based on the boundedness of the slant distance measured by DME, the positioning error of DME/DME is also bounded. The inertial navigation system is a navigation parameter calculation system based on the assembled three-axis gyroscope and accelerometer as sensitive devices. The system establishes a navigation coordinate system based on the output of the gyroscope, and calculates the velocity and position of the carrier based on the output of the accelerometer. According to the different methods of constructing the navigation coordinate system, the inertial navigation system can be divided into two types: platform type and strapdown type. Among them, the system that uses mathematical algorithms to determine the navigation coordinate system is called strapdown inertial navigation system, or SINS. When the GNSS signal is interfered in a special environment and the quality decreases, the navigation accuracy of the GNSS/SINS integrated navigation system decreases, which cannot meet the needs of route navigation.

At present, most domestic civil aviation aircraft are equipped with foreign-made flight management systems, and some models of flight management systems already have DME/DME/SINS integrated navigation functions, but the specific technical details are unknown, and it is difficult to find relevant technical information . There are few researches on DME/DME/SINS integrated navigation system in China, and there is only one literature on it. Under the condition of height reference, the integrated navigation problem is reduced to two-dimensional planar integrated navigation. Apparently, the partial simplification of the earth into a plane brings a principle error at the modeling level of the state space model, which will directly affect the navigation accuracy, so the practical application value of this research is not great. To sum up, it is necessary to carry out more in-depth research on this integrated navigation system with broad application prospects. The simplest and direct way to study DME/DME/SINS integrated navigation system is to make the difference between the position result of DME/DME positioning and the position calculated by SINS to form measurement information, and establish the Kalman filter state space based on the SINS error propagation equation The model is used to estimate and compensate the navigation error of the system through the Kalman filter to realize integrated navigation. This combination method is usually called the loose combination navigation mode, which is completely consistent with the GNSS/SINS loose combination mode. Its advantage is that the implementation method is simple, but the existing problems are that the algorithm navigation accuracy is not high enough and there is a mutation measurement noise interference condition However, the navigation accuracy is seriously lost, and even the navigation requirements cannot be met within a certain period of time. Therefore, the loose integrated navigation mode is not an ideal working mode for DME/DME/SINS integrated navigation. In order to solve the problems of loose combination, a large number of researchers have carried out a lot of research on the filtering algorithm of integrated navigation information fusion. There are other problems: first, in order to ensure the effect of the filtering algorithm, the setting of the initial filtering parameters is extremely strict; second, the improved filtering algorithm greatly increases the amount of computation, which poses new challenges for the real-time performance of the algorithm. challenge. Therefore, the feasibility of the strategy of optimizing the performance of the loosely integrated navigation system by improving the filtering algorithm in practical applications still needs a lot of verification. In addition to the research on filtering algorithms, some researchers have applied related methods of fault diagnosis and isolation to loosely integrated navigation systems, such as wavelet transform, M estimation, etc. These methods also have problems in the real-time performance of the algorithm.

The tight integrated navigation mode is the most widely used working mode of the GNSS/SINS integrated navigation system, and the tight integrated mode can also be used to design the DME/DME/SINS integrated navigation system. Compared with the loose combination mode that uses external position or velocity assistance, the tight combination uses deeper observations to form measurement information, such as pseudorange and pseudorange rate. Due to the combination of deeper observations, the integrated navigation system in the tight combination mode has stronger anti-interference ability and higher navigation accuracy in practical applications. However, it is impossible to design a DME/DME/SINS tight integrated navigation system by directly referring to the GNSS/SINS tight integrated navigation mode. There is no measurement information such as pseudorange rate during the combination process. For the DME/DME/SINS tight integrated navigation system, integrated navigation cannot be realized only by relying on two pseudo-range observation information. Combined with the actual airborne navigation equipment, the airborne altitude information provided by the air data machine can be used as additional observation information, which can be combined with The dual pseudo-range observation information together can assist SINS to realize tight integrated navigation. The DME/DME/SINS integrated navigation system suppresses the navigation error accumulated by SINS over time by establishing Kalman filter state equation and measurement equation, Kalman filter state estimation and navigation parameter error correction mode.

To sum up, in view of the lack of research on DME/DME/SINS integrated navigation system in China, and considering that DME/DME/SINS has important application prospects as the backup of GNSS/SINS integrated navigation system, it is necessary to carry out research on it. Systematic research. In the process of realization, considering the difficulty of engineering realization of the system and the anti-interference ability and navigation accuracy of the integrated navigation system, in-depth research should be carried out on the tight integrated navigation mode of the DME/DME/SINS integrated navigation system.

Contents of the invention

In order to ensure the navigation accuracy of the airborne navigation system under GNSS denial conditions, the present invention provides a DME/DME/SINS tight integrated navigation solution with anti-jamming capability and capable of meeting navigation real-time and reliability requirements, aiming at ensuring GNSS denial Provide a reliable backup navigation system for the GNSS/SINS integrated navigation system to ensure the flight safety of the aircraft during flight.

The above-mentioned purpose of the present invention can be obtained by the following measures, a kind of DME/DME/SINS tightly integrated navigation system relocation method is characterized in that comprising the steps: to point to the north position SINS attitude error equation, speed error equation and position error equation Based on the three-axis attitude error, the speed error in the northeast sky direction, the position error, the gyro constant drift and the accelerometer constant zero bias as state variables, the state equation of the Kalman filter is established; according to the SINS solution of the strapdown inertial navigation system calculate the pseudo-slant distance from the aircraft to the DME dual stations of the rangefinder, make a difference between the slant distance measured by the DME airborne measuring equipment and the pseudo-slant distance to form the slant distance measurement information, and compare the height measured by the air data machine with the pseudo-slant distance The altitude output by SINS is made to form altitude measurement information; the conversion relationship between measurement information and state variables is deduced, and the Kalman filter measurement equation is established; SINS navigation calculation is performed according to the measurement value output by the inertial device, and the navigation parameters are output; With the assistance of two types of navigation information, the slant distance output by the DME and the altitude of the air data machine, the real-time estimation of the system state variables is realized through the Kalman filter. The Kalman filter uses a feedback correction strategy to feedback and correct the attitude error, velocity error and Position error, real-time correction of SINS navigation error, determination of the final position, speed and attitude of the aircraft, establishment of Kalman filter state equation and measurement equation, calculation of measurement information, and Kalman filter state estimation and feedback correction. Synthesis, get the results of DME/DME/SINS tight integrated navigation system navigation sensor data acquisition, data processing, information fusion and relocation, use the estimated attitude, velocity and position error feedback to correct the SINS attitude, velocity and position, so as to suppress SINS navigation error, output higher precision navigation information, and display.

The present invention has following beneficial effects:

The present invention is aimed at the DME/DME/SINS tight combination navigation system, and there is no complete solution that can be implemented in practical applications. DME/DME is used instead of GNSS, and the better error smoothing characteristics of SINS are used to effectively smooth the DME/DME positioning error.

The present invention uses the DME+DME+SINS combined navigation scheme with a complete theoretical framework and simple and reliable algorithm content, which improves the performance of the airborne navigation system under GNSS denial conditions and lays the foundation for further research on the subsequent DME/DME/SINS combined navigation scheme. Base.

The system has strong anti-interference ability and high navigation accuracy. The DME/DME/SINS integrated navigation scheme designed by the present invention adopts a tight integrated navigation mode. Compared with using DME/DME to obtain the position of the aircraft through positioning, and then making a difference between the position and the position calculated by SINS to form the loose combination navigation mode of measurement information, the present invention uses deeper slant distance and pseudo-slope distance as The difference forms a tight combination mode of measurement information, which has stronger anti-interference ability and higher navigation accuracy in practical applications.

Provides an important backup for the GNSS/SINS integrated navigation system. The execution of the DME/DME/SINS integrated navigation scheme proposed by the present invention is completely independent of the GNSS/SINS integrated navigation system. Under the condition that the GNSS signal is good, the navigation result provided by the GNSS/SINS integrated navigation system can be referred to. Under the condition of denial, the DME/DME/SINS integrated navigation system can provide an important backup for the GNSS/SINS integrated navigation system, and meet the flight safety of the aircraft route flight stage under the condition of GNSS denial.

The invention can realize information fusion of navigation data output by SINS and DME, improve the anti-jamming ability and navigation precision of the integrated navigation system, and provide important backup for the GNSS/SINS integrated navigation system under the condition of GNSS signal rejection.

The present invention is suitable for the navigation information fusion of the original data output by DME dual sets, atmospheric data machine and inertial device, which can effectively realize DME/DME/SINS combined navigation, and the designed scheme has low external dependence and Strong anti-interference ability. With strong anti-interference ability and high navigation accuracy, it can be used as an important backup of GNSS/SINS integrated navigation system under GNSS denial conditions, providing reliable navigation information for the en route flight phase.

Description of drawings

The present invention will be further described below in conjunction with drawings and embodiments.

Fig. 1 is the schematic diagram of the principles of the DME/DME/SINS tight combined navigation scheme of the present invention;

Fig. 2 is a schematic diagram of a scenario in which the DME/DME/SINS tight integrated navigation solution is implemented in the present invention.

Detailed ways

Refer to Figure 1 and Figure 2. According to the present invention, based on the north position SINS attitude error equation, speed error equation and position error equation, based on the three-axis attitude error, the speed error in the northeast sky direction, the position error, the gyro constant value drift and the accelerometer constant value zero offset As the state variable, the state equation of the Kalman filter is established; according to the position solved by the strapdown inertial navigation system SINS, the pseudo-slant distance from the aircraft to the DME dual station of the rangefinder is calculated, and the slope distance measured by the DME airborne measuring equipment is compared with the The pseudo slant distance difference forms the slant distance measurement information, and the height measured by the air data machine and the height output by the SINS are compared to form the height measurement information; the conversion relationship between the measurement information and the state variable is derived, and the Kalman filter measurement is established Equation; SINS navigation calculation is performed according to the measurement value output by the inertial device, and the navigation parameters are output; with the assistance of the two types of navigation information output by the DME, the slant distance and the altitude of the air data machine, the real-time control of the system state variables is realized through the Kalman filter. Estimation, the Kalman filter adopts the strategy of feedback correction, feedback corrects the attitude error, velocity error and position error, corrects the SINS navigation error in real time, determines the final position, velocity and attitude of the aircraft, and combines the Kalman filter state equation and measurement The establishment of equations, the calculation of measurement information, and the Kalman filter state estimation and feedback correction are combined to obtain the results of data acquisition, data processing, information fusion and relocation of navigation sensors in the DME/DME/SINS compact integrated navigation system. The obtained attitude, velocity and position error feedback corrects the SINS attitude, velocity and position, thereby suppressing the SINS navigation error, outputting higher-precision navigation information, and displaying it.

Under normal working conditions, the onboard GNSS receiver of the aircraft receives at least 4 satellite navigation signals, and when the positioning is completed, it outputs the position, speed and attitude of the aircraft through loosely integrated navigation with the strapdown inertial navigation system SINS. When the situation shown in the figure occurs, that is, when the navigation signals of satellite 2 and satellite 4 are rejected (satellites are attacked or signals are artificially interfered), it is impossible to continue outputting aircraft navigation parameters through GNSS/SINS integrated navigation at this time, only Rely on SINS itself to calculate and output system navigation parameters. With the extension of time, due to the existence of error sources such as inertial device measurement errors, SINS navigation errors accumulate seriously over time, until it cannot meet the needs of aircraft flight navigation. If the DME airborne equipment can receive the response signals from DME station 1 and DME station 2 (at least two DME stations) after sending the inquiry signal when the satellite navigation signal is rejected, you can use the combination of DME/DME/SINS The navigation method outputs the position, speed and attitude of the aircraft, so as to provide a backup for the aircraft route navigation under the condition of GNSS signal rejection.

加速度计输出比力角速度f^b。捷联惯性导航系统SINS根据惯性器件的输出测量值进行导航解算，包括姿态解算、速度解算和位置解算，输出飞机位置、速度和姿态。根据当前时刻机载SINS输出的位置，同时结合两个地面DME台站的位置进行伪斜距计算，计算飞机至地面DME台站的伪斜距。Inertial devices include gyroscopes and accelerometers, where the gyroscope outputs angular velocity

The accelerometer outputs the specific force angular velocity f ^b . The strapdown inertial navigation system SINS performs navigation calculations based on the output measurement values of the inertial devices, including attitude calculations, velocity calculations and position calculations, and outputs aircraft position, speed and attitude. According to the position output by the airborne SINS at the current moment, combined with the positions of the two ground DME stations to calculate the pseudo-slope distance, calculate the pseudo-slope distance from the aircraft to the ground DME station.

λ_s和h_s，地面DME台站地心地固坐标系下的位置为x_D、y_D和z_D，根据捷联惯性导航系统SINS输出的位置为x_s、y_s和z_s，按如下方式的伪斜距计算公式计算得到伪斜距l_s：

DME airborne measurement equipment includes: DME ground transponder and DMEDME airborne interrogator for two-way interrogation and response interaction. The latitude, longitude and altitude output by SINS are respectively

λ _s and h _s , the positions of the ground DME station in the earth-centered ground-fixed coordinate system are x _D , y _D and z _D , and the positions output by the strapdown inertial navigation system SINS are x _s , y _s and z _s , as follows The pseudo-slope distance calculation formula of the method is calculated to obtain the pseudo-slope distance l _s :

If the DME airborne interrogator sends an interrogation signal when the GNSS signal is denied, it can receive response signals from at least two DME station DME ground transponders. Assuming that the DME airborne interrogator can receive the response signals from the DME ground transponders of the two ground DME stations of DME station 1 and DME station 2, then according to the position information output by SINS, it can be calculated according to the above pseudo-slant distance calculation formula Get the pseudo-slant distances l _s1 and l _s2 from the aircraft to the two ground DME stations, and enter the Kalman filter.

建立状态矩阵F，系统噪声驱动阵G，量测矩阵H，依据系统激励噪声序列w、量测噪声序列v和状态变量X，得到卡尔曼滤波器状态空间模型的状态方程

和量测方程

其中：φ_E、φ_N和φ_U表示姿态误差，δv_E、δv_N和δv_U表示速度误差，δx、δy和δz表示位置误差，

和

表示陀螺漂移，

和

表示加速度计零偏。The Kalman filter constructs the Kalman filter state space model according to the height h _a and two slant distance values output by the air data machine, and calculates the measured value Z entering the Kalman filter as:

Establish the state matrix F, the system noise driving matrix G, and the measurement matrix H, and obtain the state equation of the Kalman filter state space model based on the system excitation noise sequence w, the measurement noise sequence v and the state variable X

and measurement equation

Among them: φ _E , φ _N and φ _U represent attitude errors, δv _E , δv _N and δv _U represent velocity errors, δx, δy and δz represent position errors,

and

represents the gyro drift,

and

Indicates the zero bias of the accelerometer.

According to the SINS attitude error equation, velocity error equation and position error equation, the state matrix of the Kalman filter is established as shown in the following formula:

In the formula:

M _ap =M ₁ +M ₂

M _va =f ⁿ ×

M _vp ＝(v ⁿ ×)(2M ₁ +M ₂ )

为捷联矩阵；ω_ie为地球自转角速度；R_M和R_N分别为子午圈和卯酉圈曲率半径；vⁿ为速度，v_E和v_N为东向、北向速度；

为地球自转角速度在导航坐标系下的投影；

为导航坐标系相对于地球坐标系的角速度在导航坐标系下的投影，

为导航坐标系相对于惯性系的旋转角速度在导航坐标系下的投影；fⁿ表示比力在导航坐标系下的投影；×表示对向量取反对称矩阵。In the formula:

is the strapdown matrix; ω _ie is the earth's rotation angular velocity; R _M and R _N are the curvature radii of the meridian circle and the Maoyou circle respectively; v ⁿ is the velocity, and v _E and v _N are the eastward and northward velocities;

is the projection of the earth's rotation angular velocity in the navigation coordinate system;

is the projection of the angular velocity of the navigation coordinate system relative to the earth coordinate system under the navigation coordinate system,

is the projection of the rotational angular velocity of the navigation coordinate system relative to the inertial system in the navigation coordinate system; f ⁿ represents the projection of the specific force in the navigation coordinate system; × represents the antisymmetric matrix of the vector.

In an optional embodiment, the geographic coordinate system is selected as the navigation coordinate system, and the system noise driving matrix G shown in the following formula is obtained:

After a series of mathematical derivations, the measurement matrix H with the following form can be obtained:

In the formula, matrix N=PM, wherein: matrix P and matrix M are calculated as follows:

Where: x _D1 , y _D1 and z _D1 are the positions of ground DME station 1; x _D2 , y _D2 and z _D2 are the positions of ground DME station 2; e is the eccentricity of the ellipse.

The estimated value of the 15-dimensional state variable can be obtained at each filtering moment. According to the results of the observability analysis, only the attitude error, velocity error and position error can be accurately estimated, so only the estimated attitude is used at the filtering estimation moment Error, velocity error and position error feedback correct the attitude, velocity and position of SINS to suppress the navigation error accumulated by SINS over time.

The embodiments of the present invention have been described in detail above, and the present invention has been described using specific implementation methods herein. The descriptions of the above embodiments are only used to help understand the method and equipment of the present invention; meanwhile, for those of ordinary skill in the art, According to the idea of the present invention, there will be changes in the specific implementation and scope of application. To sum up, the contents of this specification should not be construed as limiting the present invention.

Claims (10)

1. A DME/DME/SINS tight integrated navigation system relocation method is characterized in that comprising the steps: based on the north position SINS attitude error equation, velocity error equation and position error equation, based on the three-axis attitude error, northeast The speed error, position error, gyro constant value drift and accelerometer constant value zero bias in the direction of the sky are state variables, and the state equation of the Kalman filter is established; according to the position calculated by the strapdown inertial navigation system SINS, the distance between the aircraft and the rangefinder is calculated. The pseudo-slant distance of the DME dual stations, the difference between the slant distance measured by the DME airborne measuring equipment and the pseudo-slant distance forms the slant distance measurement information, and the difference between the altitude measured by the air data machine and the altitude output by SINS forms the height measurement Information; deduce the conversion relationship between measurement information and state variables, and establish the Kalman filter measurement equation; perform SINS navigation calculations based on the output measurement values of inertial devices, and output navigation parameters; calculate pseudo-slope distances at DME stations and atmospheric data With the assistance of two types of navigation information, the Kalman filter is used to realize the real-time estimation of the system state variables, and the feedback correction strategy is adopted to feedback and correct the attitude error, velocity error and position error, and the SINS navigation error is corrected in real time to determine The final position, velocity and attitude of the aircraft, the establishment of the Kalman filter state equation and the measurement equation, the calculation of the measurement information, and the Kalman filter state estimation and feedback correction are integrated to obtain the DME/DME/SINS tight integrated navigation system The results of navigation sensor data collection, data processing, information fusion and relocation, use the estimated attitude, velocity and position error feedback to correct the SINS attitude, velocity and position, so as to suppress the SINS navigation error and output higher precision navigation information. 2. The DME/DME/SINS tight integrated navigation system relocation method as claimed in claim 1, is characterized in that: the aircraft airborne GNSS receiver receives the minimum 4 satellite navigation signals, and under the situation of completing positioning, the strapdown inertia The navigation system SINS outputs the position, speed and attitude of the aircraft through loosely integrated navigation. 3. DME/DME/SINS tightly combined navigation system relocation method as claimed in claim 2 is characterized in that: when the navigation signal of going out satellite 2 and satellite 4 is rejected, promptly satellite is attacked or signal is interfered by man, Rely on SINS itself to calculate and output system navigation parameters. 4. DME/DME/SINS tightly integrated navigation system relocation method as claimed in claim 3, is characterized in that: if after DME airborne equipment sends inquiry signal when satellite navigation signal refuses, receive DME station 1 and The response signal of DME station 2 outputs the position, speed and attitude of the aircraft through DME/DME/SINS integrated navigation, so as to provide a backup for the aircraft route navigation under the condition of GNSS signal rejection. 5. DME/DME/SINS tightly integrated navigation system relocation method as claimed in claim 1, is characterized in that: inertial device comprises gyroscope and accelerometer, wherein, gyroscope output angular velocity

The accelerometer outputs the specific force angular velocity f ^b . 6. DME/DME/SINS tightly integrated navigation system relocation method as claimed in claim 1 is characterized in that: strapdown inertial navigation system SINS carries out navigation solution according to the output measurement value of inertial device, comprises attitude solution, velocity Calculation and position calculation, output aircraft position, speed and attitude, according to the position output by SINS, combined with the position of two ground DME stations to calculate the pseudo-slope distance, calculate the pseudo-slope distance from the aircraft to the ground DME station. 7. DME/DME/SINS tightly combined navigation system relocation method as claimed in claim 1, is characterized in that: DME airborne measurement equipment comprises: carry out the DME ground transponder and the DME airborne interrogator of two-way query response interaction, The latitude, longitude and altitude output by SINS are respectively

λ _s and h _s , the positions of ground DME stations in the earth-centered earth-fixed coordinate system are x _D , y _D and z _D , and the position information of the DME station is transformed into the earth-centered earth-fixed coordinate system. According to the strapdown inertial navigation system SINS The output positions are x _s , y _s and z _s , and the pseudo-slope distance l _s is calculated according to the pseudo-slope distance calculation formula as follows:

8. The DME/DME/SINS tight integrated navigation system relocation method as claimed in claim 7, is characterized in that: DME airborne interrogator can receive two ground DME station grounds of DME station 1 and DME station 2 For the response signal of the transponder, according to the position information output by SINS, the pseudo-slant distance l _s1 and l _s2 from the aircraft to the two ground DME stations are calculated according to the pseudo-slant distance calculation formula, and the DME station 1 and DME station The two slant distances of station 2 are l _D1 and l _D2 . 9. DME/DME/SINS tightly combined navigation system relocation method as claimed in claim 1, is characterized in that: Kalman filter constructs Kalman filter according to the height _ha and two slant range values that air data machine outputs The filter state space model calculates the measured value Z entering the Kalman filter as:

Establish the state matrix F, the system noise driving matrix G, and the measurement matrix H, and obtain the state equation of the Kalman filter state space model based on the system excitation noise sequence w, the measurement noise sequence v and the state variable X

and measurement equation

In the formula, φ _E , φ _N and φ _U represent attitude errors, δv _E , δv _N and δv _U represent velocity errors, δx, δy and δz represent position errors,

and

represents the gyro drift,

and

Indicates the zero bias of the accelerometer. 10. DME/DME/SINS tightly combined navigation system relocation method as claimed in claim 9 is characterized in that: according to SINS attitude error equation, velocity error equation and position error equation, the state matrix of the Kalman filtering shown below is set up :

and

M _ap =M ₁ +M ₂ , M _va =f ⁿ ×

M _vp = (v ⁿ ×)(2M ₁ +M ₂ ),

in:

Indicates the latitude; n indicates the navigation coordinate system;

is the strapdown matrix; ω _ie is the earth's rotation angular velocity; R _M and R _N are the curvature radii of the meridian circle and the Maoyou circle respectively; v ⁿ is the velocity, and v _E and v _N are the eastward and northward velocities;

is the projection of the earth's rotation angular velocity in the navigation coordinate system;

is the projection of the angular velocity of the navigation coordinate system relative to the earth coordinate system under the navigation coordinate system,

is the projection of the rotational angular velocity of the navigation coordinate system relative to the inertial system in the navigation coordinate system; f ⁿ represents the projection of the specific force in the navigation coordinate system; × represents the antisymmetric matrix of the vector.

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