Rotation modulation inertial navigation full-state simulation method

Abstract

The invention belongs to the technical field of inertial navigation, and particularly relates to a rotation modulation inertial navigation full-state simulation method. The method is applied to a rotation modulation inertial navigation system and comprises the following steps: step 1, generating flight path data in a typical flight environment; 2, determining a transposition strategy of the rotation modulation inertial navigation system, and reversely calculating an angle increment and a speed increment of a sensitive system of the rotation modulation inertial navigation system by combining the flight path data in the step 1; step 3, injecting an inertial device error of the rotation modulation inertial navigation system into the angle increment and the speed increment obtained in the step 2 to obtain the actually output angle increment and speed increment; step 4, the rotation modulation inertial navigation system performs inertial calculation by using the actually output angle increment and speed increment obtained in the step 3 to generate navigation trajectory data; and 5, performing consistency comparison verification on the navigation track data generated in the step 4 and the flight track data in the step 1, and if the consistency comparison verification is not passed, returning to the step 2 to update the transposition strategy until the consistency comparison verification is passed. And the shafting arrangement, the rotation strategy and the inertia calculation process of the rotation modulation inertial navigation under any aircraft maneuver are comprehensively verified.

Description

Technical Field

[0001] The invention belongs to the technical field of inertial navigation, and in particular relates to a rotation modulation inertial navigation full-state simulation method. Background Art

[0002] With the development of military equipment and the advancement of national defense science and technology, the aviation field has put forward new requirements for autonomous navigation capabilities such as long flight time and high precision. The rotation modulation inertial navigation system drives the inertial measurement unit (IMU) to rotate through the motor drive frame, modulates the IMU constant value and slowly changing device errors, and can provide high-precision position and speed information to meet the aircraft's autonomous navigation capability requirements.

[0003] The rotating inertial navigation system has multiple axis arrangement schemes such as single axis, dual axis, and tri-axis according to design indicators such as accuracy and weight. The rotation strategies and angular motion isolation schemes used in different axis inertial navigation systems are also different, but their core principles are all based on the platform inertial navigation angular motion isolation and strapdown inertial navigation error model.

[0004] At present, the design and verification of the rotating inertial navigation axis arrangement and rotation strategy mainly rely on research and development experience. There is a lack of a comprehensive, unified and mature simulation platform, which makes it difficult to carry out model-based forward design of the rotating inertial navigation and to guide young engineers to become familiar with and master the rotating inertial navigation design process. At the same time, the existing simulation platforms are mostly based on trajectory error propagation solutions, lack of an inertial navigation original output forward solution platform, and cannot fully verify the inertial navigation solution process. Summary of the invention

[0005] The purpose of the present invention is to provide a full-state simulation method for a rotational modulation inertial navigation system, which can verify the axis arrangement and rotation strategy of the rotational modulation inertial navigation system through injecting flight trajectory data and original errors of inertial devices in any flight environment. The present invention can be used to guide the forward design of the rotational modulation inertial navigation system.

[0006] Technical solution of the present invention: In order to achieve the above-mentioned invention object, according to the first aspect of the present invention, a rotation modulation inertial navigation full-state simulation method is proposed, which is applied to the rotation modulation inertial navigation system, and comprises the following steps:

[0007] Step 1. Generate flight trajectory data under typical flight environment;

[0008] Step 2. Determine the rotational modulation inertial navigation system's indexing strategy, and reversely calculate the angular increment and velocity increment of the rotational modulation inertial navigation system's sensitive system in combination with the flight trajectory data in step 1;

[0009] Step 3. Inject the inertial device error of the rotation modulation inertial navigation system into the angular increment and velocity increment obtained in step 2 to obtain the actual output angular increment and velocity increment;

[0010] Step 4: The rotation modulation inertial navigation system uses the actual output angular increment and velocity increment obtained in step 3 to perform inertial solution and generate navigation trajectory data;

[0011] Step 5: Perform consistency comparison and verification on the navigation trajectory data generated in step 4 and the flight trajectory data in step 1. If the consistency comparison and verification fails, return to step 2 to update the transfer strategy until the consistency comparison and verification passes.

[0012] In a possible embodiment, in step 1, the flight trajectory data includes attitude, speed, and position.

[0013] In a possible embodiment, in step 1, the flight trajectory in a typical flight environment is composed of a segmented combination of acceleration motion, rolling motion, pitching motion, and azimuth motion.

[0014] In a possible embodiment, in step 1, the generation process of the flight trajectory data includes the following steps:

[0015] Let posture E = [θγψ] T 、Attitude angular velocity vector ω=[ω θ ω γ ω ψ ] T , Machine system speed Machine system acceleration Aircraft position p = [Lλh].

[0016] Accelerated Motion:

[0017] Rolling motion:

[0018] Pitch movement:

[0019] Azimuth motion: According to aerodynamics, the azimuth motion of an aircraft is accompanied by rolling motion, so its centripetal acceleration a c , gravitational acceleration g, heading angular rate ω ψ , the rolling angle γ has the relationship of formula 1

[0020]

[0021] From this we can get the azimuth motion:

[0022] Then the differential equations for the aircraft position are:

[0023]

[0024] The initial value is set as:

[0025] E(t0)=[θ0γ0ψ0] T , p(t0)=[L0λ0h0] T

[0026] By segmenting the input ω θ ,ω γ ,ω ψ and a y , using the differential equation system to solve the time-varying state equation, the flight trajectory data can be generated to obtain the aircraft attitude [θγψ] T , speed [V e V n V u ] T , position [Lλh] T .

[0027] In a possible embodiment, in step 2, the specific process of inversely calculating the angular increment of the sensitive system includes:

[0028] Calculate the transformation matrix from the inertial sensitive system to the navigation system;

[0029]

[0030] Where:

[0031] is the conversion matrix from aircraft system b to navigation system n;

[0032] is the conversion matrix from the inertial sensitive system s to the navigation system n;

[0033] is the transformation matrix from the internal frame moving coordinate system 4 to the inertial sensitive system s;

[0034] is the transformation matrix from the inner frame static coordinate system 3 to the inner frame dynamic coordinate system 4;

[0035] is the transformation matrix from the middle frame dynamic coordinate system 2 to the inner frame static coordinate system 3;

[0036] is the transformation matrix from the middle frame static coordinate system 1 to the middle frame dynamic coordinate system 2;

[0037] is the transformation matrix from the machine system b system to the middle frame static coordinate system 1 system;

[0038] in, It is caused by the installation of the shaft system during assembly, and is set as a unit matrix in the present invention; then the above formula can be simplified as:

[0039]

[0040] The rotation matrix is ​​a unit orthogonal matrix, so Formula 4 can be transformed into:

[0041]

[0042] in: Represents the transformation matrix from the s system to the n system at time m; Represents the transformation matrix from system b to system n at time m;

[0043] right Perform the following transformations:

[0044]

[0045] Combining Formula 5 and Formula 6, we can get:

[0046]

[0047] in:

[0048] represents the transformation matrix between time m-1 and time m in the s system, In the left formula, is the equivalent rotation vector at time m, Δθ m is the angular increment at time m;

[0049] is the transformation matrix from the n system to the s system at time m-1, which can be obtained by formula 5;

[0050] Represents the transformation matrix of the n-system at time m-1 and time m, is the equivalent rotation vector, let but

[0051] is the Earth's rotation angular velocity, Where [V e V n V u ] T is the projection of the aircraft velocity vector in the geographic system, R M is the meridian radius of the earth, R N is the radius of the Earth's circumpolar circle, and L is the current latitude;

[0052] By the following formula:

[0053]

[0054]

[0055] Available The angle increment can be obtained by the following formula. Let Δθ0 = 0 to find the gyro angle increment at any time;

[0056]

[0057] In a possible embodiment, in step 2, the specific process of inversely calculating the velocity increment of the sensitive system s includes:

[0058] The velocity update algorithm can be obtained by solving the specific force equation:

[0059]

[0060] in: represents the projection of the aircraft velocity vector v at time m in the n system, represents the projection of the aircraft velocity vector v at time m-1 in the n system, It represents the projection of the aircraft speed increment at time m in the n system. It represents the projection of the velocity increment calculation error at time m in the n system;

[0061]

[0062] Δv m is the velocity increment at time m; Δθ m is the angular increment at time m;

[0063]

[0064] The expression can be written as:

[0065]

[0066] From the above formula, we can get:

[0067]

[0068] Through the relationship between the shell system b system and the sensitive system s system Solve for the s-frame velocity increment:

[0069]

[0070] In a possible embodiment, in step 3, the inertial device error includes: addendum, gyro scale coefficient error, installation deflection error, and zero position error.

[0071] In a possible embodiment, in step 3, the specific process of injecting the inertial device error includes: gyro residual scaling coefficient error δk gii (i=x,y,z), installation angle error And zero drift error Then the gyro angle increment output after the error is injected can be expressed as:

[0072]

[0073] in:

[0074] is from the orthogonal coordinate system s to the non-orthogonal coordinate system s g The coordinate transformation matrix of the system, where μ gx , μ gy , μ gz They are the non-orthogonal angles of the x, y, and z axes respectively;

[0075] is the angular increment in the non-orthogonal coordinate system, Δθ m is the angle increment in the orthogonal system, obtained by formula 9;

[0076] δk g is the gyro scale coefficient error;

[0077] is the zero drift error in non-orthogonal system;

[0078] Similarly, the incremental speed output after error injection can be obtained:

[0079]

[0080] in:

[0081] It is the output of the calculated specific force in the non-orthogonal coordinate system. is the specific force output in the orthogonal system, obtained by formula 14;

[0082] δk g is the gyro scale coefficient error;

[0083] is the added zero bias error.

[0084] In a possible embodiment, in step 4, the specific process of performing inertial solution on the rotation modulation inertial navigation system to generate navigation trajectory data includes:

[0085] Posture Update:

[0086]

[0087]

[0088] in: and The calculation method is detailed in Formula 6; at the initial moment, let is the unit matrix; then It can be solved iteratively through angle increment and velocity increment;

[0089] Speed ​​Update:

[0090]

[0091] in, and The calculation method is detailed in Formula 10; at the initial moment, let If is 0, It can be solved iteratively through angle increment and velocity increment;

[0092] Location Updates:

[0093]

[0094]

[0095]

[0096] In a possible embodiment, in step 5, the step of performing consistency comparison includes: determining the accuracy indicators of position, speed, and attitude; calculating the position accuracy, speed accuracy, and attitude accuracy of the navigation results, the position accuracy is calculated using method 2 in GJB729-CEP, and the speed and attitude accuracy are calculated using root mean square; performing indicator conformity judgment based on the accuracy indicators of position, speed, and attitude, if the indicator requirements are not met, analyzing the source of the error, adjusting the injected error value, or adjusting the transposition strategy to repeat the simulation.

[0097] Beneficial technical effects of the present invention:

[0098] The present invention proposes a full-state simulation method for rotational modulated inertial navigation. By designing complex motion profiles, axis arrangement schemes and transfer strategies, the transformation relationship between the inertial measurement coordinate system and the shell system is established. Through end-to-end full-process simulation, the axis arrangement, rotation strategy and inertial solution process of the rotational modulated inertial navigation under any aircraft maneuvers are fully verified.

[0099] Prior to this, due to the coupling between the rotational modulation inertial navigation aircraft maneuver and the inertial navigation's own rotation strategy, it was difficult to directly solve the original gyro data under the sensitive system through the flight trajectory, making full-state simulation without a feasible solution. The present invention pioneered a reverse solution method, making end-to-end simulation verification of the rotational modulation inertial navigation possible, filling the gap in full-state simulation verification of the rotational modulation inertial navigation, and guiding the forward design of the rotational modulation inertial navigation in a more scientific and efficient manner. BRIEF DESCRIPTION OF THE DRAWINGS

[0100] 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.

[0101] Figure 1 is a method flow chart of an embodiment of the present invention;

[0102] Figure 2 is a flight trajectory combination diagram of embodiment 1 of the present invention;

[0103] Figure 3 Schematic diagram of the angular increment of the rotational modulation inertial navigation system according to Embodiment 1 of the present invention;

[0104] Figure 4 Schematic diagram of the speed increment of the rotary modulated inertial navigation system according to Embodiment 1 of the present invention;

[0105] Figure 5 1 is a schematic diagram of a navigation error combination for consistency comparison verification according to Embodiment 1 of the present invention;

[0106] Figure 6 is a flight trajectory combination diagram of embodiment 2 of the present invention;

[0107] Figure 7 Schematic diagram of the angular increment of the rotational modulation inertial navigation system according to Embodiment 2 of the present invention;

[0108] Figure 8 Schematic diagram of the speed increment of the rotary modulated inertial navigation system according to Embodiment 2 of the present invention;

[0109] Fig. 9 It is a schematic diagram of the navigation error combination for consistency comparison verification in Example 2 of the present invention. DETAILED DESCRIPTION

[0110] 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.

[0111] 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.

[0112] 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.

[0113] Example 1

[0114] A dual-axis rotation modulation inertial navigation full-state simulation method specifically includes:

[0115] Step 1: Generate a typical flight trajectory

[0116] like Figure 2 As shown, assume that the motion trajectory of the aircraft consists of the following four flight actions, and there is only one flight action at the same time.

[0117] (1) Aircraft acceleration:

[0118] (2) Aircraft rolling motion:

[0119] (3) Aircraft pitch motion:

[0120] (4) Aircraft azimuth movement:

[0121] The differential equation for the aircraft position is:

[0122] Set the initial value of the state variable to:

[0123] E(t0)=

[000] T ,v b (t0) =

[000] T ,p(t0)=[34.1989108.8291400] T

[0124] By setting the aircraft to perform the above four flight actions in sections and solving the time-varying state equation by integrating the differential equation group, the generation of flight trajectory data can be completed, and the aircraft Euler angle [θγψ] can be obtained. T , speed [V e V n V u ] T , position [Lλh] T As shown below:

[0125] Step 2: Generate angular increment and velocity increment in inertial sensitive system

[0126] The strategy for determining the position of the dual-axis rotation modulation inertial navigation 16 is as follows:

[0127]

[0128]

[0129] Combine the flight trajectory data in step 1 and reversely calculate the angular increment and velocity increment of the sensitive system s;

[0130] Calculate the transformation matrix from the sensitive system to the navigation system;

[0131]

[0132] Where:

[0133] is the conversion matrix from aircraft system b to navigation system n;

[0134] is the conversion matrix from the inertial sensitive system s to the navigation system n;

[0135] is the transformation matrix from the internal frame moving coordinate system 4 to the inertial sensitive system s;

[0136] is the transformation matrix from the inner frame static coordinate system 3 to the inner frame dynamic coordinate system 4;

[0137] is the transformation matrix from the middle frame dynamic coordinate system 2 to the inner frame static coordinate system 3;

[0138] is the transformation matrix from the middle frame static coordinate system 1 to the middle frame dynamic coordinate system 2;

[0139] is the transformation matrix from the machine system b system to the middle frame static coordinate system 1 system;

[0140] in, It is caused by the installation of the shaft system during assembly, which is not considered here and is set as a unit matrix; then the above formula can be simplified to:

[0141]

[0142] Where: By designing a specific rotation strategy, it can be obtained;

[0143] The rotation matrix is ​​a unit orthogonal matrix, so Equation 4 can be transformed into:

[0144]

[0145] in: Represents the transformation matrix from the s system to the n system at time m; It represents the transformation matrix from the b system to the n system at time m, which can be obtained in real time from the flight trajectory.

[0146] right Perform the following transformations:

[0147]

[0148] Combining Formula 5 and Formula 6, we can get:

[0149]

[0150] in:

[0151] represents the transformation matrix between time m-1 and time m in the s system, In the left formula, is the equivalent rotation vector at time m, Δθ m is the angular increment at time m (the quantity to be solved);

[0152] is the transformation matrix from the n system to the s system at time m-1, which can be obtained by formula 5;

[0153] Represents the transformation matrix of the n-system at time m-1 and time m, is the equivalent rotation vector, let but

[0154] is the Earth's rotation angular velocity, Where [V eV n V u ] T is the projection of the aircraft velocity vector in the geographic system, R M is the meridian radius of the earth, R N is the radius of the earth's ecliptic circle, and L is the current latitude.

[0155] By the following formula:

[0156]

[0157]

[0158] Available The angle increment can be obtained by the following formula. Let Δθ0 = 0 to calculate the gyro angle increment at any time, such as Figure 3 As shown;

[0159]

[0160] The specific process of inversely calculating the velocity increment of the sensitive system s includes:

[0161] The velocity update algorithm can be obtained by solving the specific force equation:

[0162]

[0163] in: represents the projection of the aircraft velocity vector v at time m in the n system, represents the projection of the aircraft velocity vector v at time m-1 in the n system, It represents the projection of the actual speed increment of the aircraft at time m in the n system. It represents the projection of the velocity increment calculation error at time m in the n system.

[0164]

[0165] in:

[0166] Δv m is the velocity increment at time m; Δθ m is the angular increment at time m;

[0167]

[0168] The expression can be written as:

[0169]

[0170] From the above formula, we can get:

[0171]

[0172] Through the relationship between the shell system b system and the sensitive system s system Solve the s-system velocity increment, such as Figure 4 As shown:

[0173]

[0174] Step 3: Angular increment and velocity increment inject inertial device error

[0175] Gyro residual scale factor error δk gii (i=x,y,z)

[0176] Gyro installation deflection error

[0177] Gyro zero drift error Then the gyro angle increment output after the error is injected can be expressed as:

[0178]

[0179] in:

[0180] is the angular increment in the non-orthogonal coordinate system, Δθ m is the angle increment in the orthogonal system, obtained by formula 9; is from the orthogonal coordinate system s to the non-orthogonal coordinate system s g The coordinate transformation matrix of the system, in

[0181] δk g is the gyro scale coefficient error, where

[0182] is zero drift in non-orthogonal system,

[0183] Similarly, the incremental speed output after error injection can be obtained:

[0184]

[0185] in:

[0186] It is the output of the calculated specific force in the non-orthogonal coordinate system. is the specific force output in the orthogonal system, obtained by formula 13;

[0187] is from the orthogonal coordinate system s to the non-orthogonal coordinate system s g The coordinate transformation matrix of the system, in

[0188] δk g is the gyro scale coefficient error, where

[0189] To add zero bias,

[0190] Step 4: Forward solution using the inertial data with injected errors

[0191] The specific process of performing inertial solution and generating navigation trajectory data by the rotation modulation inertial navigation system includes:

[0192] Posture Update:

[0193]

[0194]

[0195] in: and The calculation method is detailed in Formula 6; at the initial moment, let is the unit matrix; then The solution can be iterated through angle increment and velocity increment.

[0196] Speed ​​Update:

[0197]

[0198] in, and The calculation method is detailed in Formula 10; at the initial moment, let If is 0, The solution can be iterated through angle increment and velocity increment.

[0199] Location Updates:

[0200]

[0201] Step 5: Compare the navigation data with the flight trajectory data for consistency

[0202] Index requirements:

[0203] Attitude: 0.01°(RMS)

[0204] Speed: 0.8m / s (RMS)

[0205] Location: 500m (MAX)

[0206] Calculate the error between the navigation data and the simulation trajectory, and adjust the transposition strategy or device error according to the error to make the navigation result closest to the simulation trajectory. Figure 6 The navigation result error is shown:

[0207] Speed ​​RMS indicator calculation:

[0208] where v i Indicates the speed value (eastward or northward) at each sampling point;

[0209] Attitude RMS indicator calculation:

[0210] where v i Indicates the angle value (pitch, roll, heading) of each sampling point;

[0211] Position maximum indicator calculation:

[0212]

[0213] Δlat i =(lat i -lat1)×30×1852

[0214] Δlon i =(lon i -lon1)×cos(lat i )×30×1852

[0215] From the above formula, we can get:

[0216] RMS θ : 0.002RMS γ : 0.005RMS ψ : 0.004

[0217] RMS ve : 0.3RMS vn : 0.33

[0218] POS max :775

[0219] The position error is greater than the index requirement, the transposition plan does not meet the requirements, and the transposition strategy needs to be readjusted.

[0220] Example 2

[0221] A dual-axis rotation modulation inertial navigation full-state simulation method specifically includes:

[0222] Step 1: Generate a typical flight trajectory

[0223] like Figure 6 As shown, assume that the motion trajectory of the aircraft consists of the following four flight actions, and there is only one flight action at the same time.

[0224] (1) Aircraft acceleration:

[0225] (2) Aircraft rolling motion:

[0226] (3) Aircraft pitch motion:

[0227] (4) Aircraft azimuth movement:

[0228] The differential equation for the aircraft position is:

[0229] Set the initial value of the state variable to:

[0230] E(t0)=

[000] T ,v b (t0) =

[000] T ,p(t0)=[34.1989108.8291400] T

[0231] By setting the aircraft to perform the above four flight actions in sections and solving the time-varying state equation by integrating the differential equation group, the generation of flight trajectory data can be completed, and the aircraft Euler angle [θγψ] can be obtained. T , speed [V e V n V u ] T , position [Lλh] T As shown below:

[0232] Step 2: Generate angular increment and velocity increment in inertial sensitive system

[0233] The strategy for determining the position of the dual-axis rotation modulation inertial navigation 32 is as follows:

[0234]

[0235]

[0236] Combine the flight trajectory data in step 1 and reversely calculate the angular increment and velocity increment of the sensitive system s;

[0237] Calculate the transformation matrix from the sensitive system to the navigation system;

[0238]

[0239] Where:

[0240] is the conversion matrix from aircraft system b to navigation system n;

[0241] is the conversion matrix from the inertial sensitive system s to the navigation system n;

[0242] is the transformation matrix from the internal frame moving coordinate system 4 to the inertial sensitive system s;

[0243] is the transformation matrix from the inner frame static coordinate system 3 to the inner frame dynamic coordinate system 4;

[0244] is the transformation matrix from the middle frame dynamic coordinate system 2 to the inner frame static coordinate system 3;

[0245] is the transformation matrix from the middle frame static coordinate system 1 to the middle frame dynamic coordinate system 2;

[0246] is the transformation matrix from the machine system b system to the middle frame static coordinate system 1 system;

[0247] in, It is caused by the installation of the shaft system during assembly, which is not considered here and is set as a unit matrix; then the above formula can be simplified to:

[0248]

[0249] Where: By designing a specific rotation strategy, it can be obtained;

[0250] The rotation matrix is ​​a unit orthogonal matrix, so Equation 4 can be transformed into:

[0251]

[0252] in: Represents the transformation matrix from the s system to the n system at time m; It represents the transformation matrix from the b system to the n system at time m, which can be obtained in real time from the flight trajectory.

[0253] right Perform the following transformations:

[0254]

[0255] Combining Formula 5 and Formula 6, we can get:

[0256]

[0257] in:

[0258] represents the transformation matrix between time m-1 and time m in the s system, In the left formula, is the equivalent rotation vector at time m, Δθ m is the angular increment at time m (the quantity to be solved);

[0259] is the transformation matrix from the n system to the s system at time m-1, which can be obtained by formula 5;

[0260] Represents the transformation matrix of the n-system at time m-1 and time m, is the equivalent rotation vector, let but

[0261] is the Earth's rotation angular velocity, Where [V e V n V u ] T is the projection of the aircraft velocity vector in the geographic system, R M is the meridian radius of the earth, R N is the radius of the earth's ecliptic circle, and L is the current latitude.

[0262] By the following formula:

[0263]

[0264]

[0265] Available The angle increment can be obtained by the following formula. Let Δθ0 = 0 to calculate the gyro angle increment at any time, such as Figure 7 As shown;

[0266]

[0267] The specific process of inversely calculating the velocity increment of the sensitive system s includes:

[0268] The velocity update algorithm can be obtained by solving the specific force equation:

[0269]

[0270] in: represents the projection of the aircraft velocity vector v at time m in the n system, represents the projection of the aircraft velocity vector v at time m-1 in the n system, It represents the projection of the actual speed increment of the aircraft at time m in the n system. It represents the projection of the velocity increment calculation error at time m in the n system.

[0271]

[0272] in:

[0273] Δv m is the velocity increment at time m;

[0274] Δθ m is the angular increment at time m;

[0275]

[0276] The expression can be written as:

[0277]

[0278] From the above formula, we can get:

[0279]

[0280] Through the relationship between the shell system b system and the sensitive system s system Solve the s-system velocity increment, such as Figure 8 As shown:

[0281]

[0282] Step 3: Angular increment and velocity increment inject inertial device error

[0283] Gyro residual scale factor error δk gii (i=x,y,z)

[0284] Gyro installation deflection error

[0285] Gyro zero drift error Then the gyro angle increment output after the error is injected can be expressed as:

[0286]

[0287] in:

[0288] is the angular increment in the non-orthogonal coordinate system, Δθ m is the angle increment in the orthogonal system, obtained by formula 9;

[0289] is from the orthogonal coordinate system s to the non-orthogonal coordinate system s g The coordinate transformation matrix of the system, in

[0290] δk gis the gyro scale coefficient error, where

[0291] is zero drift in non-orthogonal system,

[0292] Similarly, the incremental speed output after error injection can be obtained:

[0293]

[0294] in:

[0295] It is the output of the calculated specific force in the non-orthogonal coordinate system. is the specific force output in the orthogonal system, obtained by formula 13; is from the orthogonal coordinate system s to the non-orthogonal coordinate system s g The coordinate transformation matrix of the system, in

[0296] δk g is the gyro scale coefficient error, where

[0297] To add zero bias,

[0298] Step 4: Forward solution using the inertial data with injected errors

[0299] The specific process of performing inertial solution and generating navigation trajectory data by the rotation modulation inertial navigation system includes:

[0300] Posture Update:

[0301]

[0302]

[0303] in: and The calculation method is detailed in Formula 6; at the initial moment, let is the unit matrix; then The solution can be iterated through angle increment and velocity increment.

[0304] Speed ​​Update:

[0305]

[0306] in, and The calculation method is detailed in Formula 10; at the initial moment, let If is 0, The solution can be iterated through angle increment and velocity increment.

[0307] Location Updates:

[0308]

[0309]

[0310]

[0311] Step 5: Compare the navigation data with the flight trajectory data for consistency

[0312] Index requirements:

[0313] Attitude: 0.01°(RMS)

[0314] Speed: 0.8m / s (RMS)

[0315] Location: 500m (MAX)

[0316] Calculate the error between the navigation data and the simulation trajectory, and adjust the transposition strategy or device error according to the error to make the navigation result closest to the simulation trajectory. Fig. 9 The navigation result error is shown:

[0317] Speed ​​RMS indicator calculation:

[0318] Where vi represents the velocity value (eastward or northward) of each sampling point;

[0319] Attitude RMS indicator calculation:

[0320] where v i Indicates the angle value (pitch, roll, heading) of each sampling point;

[0321] Position maximum indicator calculation:

[0322]

[0323] Δlat i =(lat i -lat1)×30×1852

[0324] Δlon i =(lon i -lon1)×cos(lat i )×30×1852

[0325] From the above formula, we can get:

[0326] RMS θ : 0.0014RMSγ : 0.0012RMS ψ : 0.003

[0327] RMS ve : 0.15RMS vn : 0.16

[0328] POS max :247

[0329] The transposition scheme meets the index requirements.

[0330] Although the embodiments disclosed in the present invention are as above, the contents described are only embodiments adopted to facilitate understanding of the present invention and are not intended to limit the present invention. Any technician in the field to which the present invention belongs can make any modifications and changes in the form and details of implementation without departing from the spirit and scope disclosed in the present invention, but the patent protection scope of the present invention shall still be subject to the scope defined in the attached claims.

Claims

1. A rotation modulation inertial navigation full-state simulation method, characterized in that: Applied to the rotation modulation inertial navigation system, including the following steps: Step 1. Generate flight trajectory data under typical flight environment; Step 2. Determine the rotational modulation inertial navigation system's indexing strategy, and reversely calculate the angular increment and velocity increment of the rotational modulation inertial navigation system's sensitive system in combination with the flight trajectory data in step 1; Step 3. Inject the inertial device error of the rotation modulation inertial navigation system into the angular increment and velocity increment obtained in step 2 to obtain the actual output angular increment and velocity increment; Step 4: The rotation modulation inertial navigation system uses the actual output angular increment and velocity increment obtained in step 3 to perform inertial solution and generate navigation trajectory data; Step 5: Perform consistency comparison and verification on the navigation trajectory data generated in step 4 and the flight trajectory data in step 1. If the consistency comparison and verification fails, return to step 2 to update the transfer strategy until the consistency comparison and verification passes.

2. A rotation modulation inertial navigation full-state simulation method according to claim 1, characterized in that: In step 1, the flight trajectory data includes attitude, speed, and position.

3. The method for full-state simulation of a rotational modulation inertial navigation system according to claim 1, characterized in that: In step 1, the flight trajectory in a typical flight environment is composed of a segmented combination of acceleration motion, rolling motion, pitching motion, and azimuth motion.

4. A rotation modulation inertial navigation full-state simulation method according to claim 3, characterized in that: In step 1, the generation process of flight trajectory data includes the following steps: Let posture E = [θ γ ψ] T 、Attitude angular velocity vector ω=[ω θ ω γ ω ψ ] T , Machine system speed Machine system acceleration Aircraft position p = [L λ h]. Accelerated Motion: Rolling motion: Pitch movement: Azimuth motion: According to aerodynamics, the azimuth motion of an aircraft is accompanied by rolling motion, so its centripetal acceleration a c , gravitational acceleration g, heading angular velocity ω ψ , the rolling angle γ has the relationship of formula 1 From this we can get the azimuth motion: Then the differential equations for the aircraft position are: The initial value is set as: E(t0)=[θ0γ0ψ0] T , p(t0)=[L0λ0h0] T By segmenting the input ω θ ,ω γ ,ω ψ and a y , using the differential equation system to solve the time-varying state equation, the flight trajectory data can be generated to obtain the aircraft attitude [θγψ] T , speed [V e V n V u ] T , position [Lλh] T .

5. A rotation modulation inertial navigation full-state simulation method according to claim 4, characterized in that: In step 2, the specific process of inversely calculating the angular increment of the sensitive system includes: Calculate the transformation matrix from the inertial sensitive system s to the navigation system n; Where: is the conversion matrix from aircraft system b to navigation system n; is the conversion matrix from the inertial sensitive system s to the navigation system n; is the transformation matrix from the internal frame moving coordinate system 4 to the inertial sensitive system s; is the transformation matrix from the inner frame static coordinate system 3 to the inner frame dynamic coordinate system 4; is the transformation matrix from the middle frame dynamic coordinate system 2 to the inner frame static coordinate system 3; is the transformation matrix from the middle frame static coordinate system 1 to the middle frame dynamic coordinate system 2; is the transformation matrix from the machine system b system to the middle frame static coordinate system 1 system; in, It is caused by the installation of the shaft system during assembly, and is set as a unit matrix in the present invention; then the above formula can be simplified as: The rotation matrix is ​​a unit orthogonal matrix, so Formula 4 can be transformed into: in: Represents the transformation matrix from the s system to the n system at time m; represents the transformation matrix from system b to system n at time m; Perform the following transformations: Combining Formula 5 and Formula 6, we can get: in: It represents the transformation matrix between the inertial sensitive system s at time m-1 and time m, In the left formula, is the equivalent rotation vector at time m, Δθ m is the angular increment at time m; is the transformation matrix from the navigation system n to the inertial sensitive system s at time m-1, which can be obtained by formula 5; Represents the conversion matrix between the navigation system n at time m-1 and time m, is the equivalent rotation vector, let but is the Earth's rotation angular velocity, Where [V e V n V u ] T is the projection of the aircraft velocity vector in the geographic system, R M is the meridian radius of the earth, R N is the radius of the Earth's circumpolar circle, and L is the current latitude; By the following formula: Available The angle increment can be obtained by the following formula. Let Δθ0 = 0 to find the gyro angle increment at any time; 6. A rotation modulation inertial navigation full-state simulation method according to claim 5, characterized in that: In step 2, the specific process of inversely calculating the velocity increment of the sensitive system s includes: The velocity update algorithm can be obtained by solving the specific force equation: in: represents the projection of the aircraft velocity vector v at time m in the n system, represents the projection of the aircraft velocity vector v at time m-1 in the n system, It represents the projection of the aircraft speed increment at time m in the n system. It represents the projection of the velocity increment calculation error at time m in the n system; Δv m is the velocity increment at time m; Δθ m is the angular increment at time m; The expression can be written as: From the above formula, we can get: Through the relationship between the shell system b system and the sensitive system s system Solve the sensitive system s system velocity increment:

7. A rotation modulation inertial navigation full-state simulation method according to claim 6, characterized in that: In step 3, the specific process of injecting the inertial device error includes: the gyro residual scaling coefficient error δk gii (i=x,y,z), installation angle error And zero drift error Then the gyro angle increment output after the error is injected can be expressed as: in: is from the orthogonal coordinate system s to the non-orthogonal coordinate system s g The coordinate transformation matrix of the system, where μ gx , μ gy , μ gz They are the non-orthogonal angles of the x, y, and z axes respectively; is the angular increment in the non-orthogonal coordinate system, Δθ m is the angle increment in the orthogonal system, obtained by formula 9; δk g is the gyro scale coefficient error; is the zero drift error in non-orthogonal system; Similarly, the incremental speed output after error injection can be obtained: in: It is the output of the calculated specific force in the non-orthogonal coordinate system. is the specific force output in the orthogonal system, obtained by formula 14; δk g is the gyro scale coefficient error; is the added zero bias error.

8. A rotation modulation inertial navigation full-state simulation method according to claim 7, characterized in that: In step 4, the specific process of performing inertial solution on the rotation modulation inertial navigation system and generating navigation trajectory data includes: Posture Update: in: and The calculation method is detailed in Formula 6; at the initial moment, let is the unit matrix; then It can be solved iteratively through angle increment and velocity increment; Speed ​​Update: in, and The calculation method is detailed in Formula 10; at the initial moment, let If is 0, It can be solved iteratively through angle increment and velocity increment; Location Updates:

9. The method for full-state simulation of a rotational modulation inertial navigation system according to claim 1, characterized in that: The steps for consistency comparison include: determining the accuracy indicators of position, velocity and attitude; calculating the position accuracy, velocity accuracy and attitude accuracy of the navigation results. The position accuracy is calculated using method 2 in GJB729-CEP, and the velocity and attitude accuracy are calculated using root mean square; according to the accuracy indicators of position, velocity and attitude, the indicator conformity criterion is judged. If the indicator requirements are not met, the source of the error is analyzed, and the injected error value is adjusted or the transfer strategy is adjusted to repeat the simulation.

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