Ten-table redundant inertial measurement unit gyroscope fault detection and isolation method based on navigation attitude angle

Abstract

The invention discloses a ten-table redundant inertial measurement unit gyroscope fault detection and isolation method based on a navigation attitude angle, and relates to the technical field of spaceflight, and the method comprises the following steps: establishing at least six groups of independent navigation attitude calculation loops under an established inertial measurement unit measurement coordinate system, and obtaining absolute values of comparison results through the navigation attitude calculation loops, whether the absolute value of each comparison result is smaller than a preset threshold value or not is judged, and if the absolute value of each comparison result does not exceed the preset threshold value, the resolving results of a group of navigation attitude resolving loops are adopted as attitude angles used by a control system; if the absolute value of one or more comparison results exceeds or is equal to a preset threshold value, isolating the faulted gyroscopes, and selecting one group of navigation attitude resolving loops corresponding to the non-isolated gyroscopes as a new reference; and taking a resolving result of the navigation attitude resolving loop corresponding to the new reference as an attitude angle used by a control system. According to the invention, the accuracy of gyroscope fault monitoring and isolation can be improved.

Description

Technical Field

[0001] The present application relates to the field of aerospace technology, and in particular to a method for detecting and isolating gyro faults of a ten-meter redundant inertial group based on navigation attitude angle. Background Art

[0002] With the rapid increase in the number of China's space launch missions, the ultra-high density launch frequency and diverse launch mission requirements have put forward higher requirements on the research and development efficiency, reliability, fault tolerance and economic cost of the launch vehicle control system. Redundant configuration is a common technical approach to improve the reliability and fault tolerance of the navigation system, which is usually divided into two methods: system-level redundancy and device-level redundancy. System-level redundancy includes dual inertial group master-slave redundancy and triple inertial group redundancy; device-level redundancy includes single inertial group multi-meter redundancy. Among them, the inertial group with a single set of ten meters has more advantages than the system-level redundancy in terms of volume, mass, cost and simplicity. The inertial group with a single set of ten meters has gradually become the main inertial device for subsequent launch models.

[0003] The launch vehicle is highly dynamic and requires the fault diagnosis system to process fault data quickly and accurately and locate and isolate fault information in a timely manner, and ensure the smooth completion of the flight mission through navigation information reconstruction. At present, a variety of fault diagnosis methods for inertial devices have been developed, including threshold comparison method, generalized likelihood ratio method, mean test method, singular value decomposition method and neural network method.

[0004] Inertial navigation system is a key device for launch vehicle navigation, and rest technology is usually an important method to improve system reliability. Using different gyro combinations of ten-meter inertial group to solve navigation attitude in the inertial system and performing consistency comparison of attitude angle solution results of different combinations in real time can realize gyro fault detection and isolation.

[0005] In the process of fault detection, the size of the threshold is directly related to whether the fault can be detected in a very short time and the degree of false alarm of the system, that is, it directly affects the fault diagnosis performance and task reliability of the system. When the fault detection threshold value is too large, the missed detection rate of the system fault detection will increase; conversely, when the fault detection threshold value is too small, the false alarm rate of the system fault detection will increase. These error rates have a great impact on the fault diagnosis performance of the system. It can be seen that how to reasonably determine the fault detection threshold is the key to the effectiveness of fault detection.

[0006] The traditional method of detecting and isolating gyro faults in the ten-meter redundant inertia group mainly performs diagnosis in the instrument pulse conversion link, which is easily affected by external environmental factors and may lead to false alarms or missed alarms of gyro pulse faults. Summary of the invention

[0007] The purpose of the present application is to provide a method for detecting and isolating gyro faults in a ten-meter redundant inertial group based on navigation attitude angle, which can improve the accuracy of gyro fault monitoring and isolation.

[0008] To achieve the above-mentioned purpose, the present application provides a method for detecting and isolating a gyro fault of a ten-meter redundant inertial group based on a navigation attitude angle, comprising the following steps: S1: after the rocket is ignited, an inertial group measurement coordinate system is established, at least six groups of independent navigation attitude solution loops are established in the inertial group measurement coordinate system, navigation attitude solution is started, and S2 is executed; S2: output data is obtained through the axial instrument combination participating in the solution in each group of navigation attitude solution loops, the output data is converted to the inertial group measurement coordinate system, the output value of the inertial group measurement coordinate system is obtained, the navigation attitude is updated according to the output value of the inertial group measurement coordinate system, the updated output data is obtained, and the updated output data is used as the solution result; wherein the output data includes: the output navigation attitude angle and the output acceleration; the inertial group measurement coordinate system output value includes: the converted navigation attitude angle and the converted acceleration; the updated output data includes: the updated attitude angle and the updated acceleration, and S3 is executed; S3: the solution results of one group of navigation attitude solution loops in the at least six groups of independent navigation attitude solution loops are respectively compared with the solution results of the at least six groups of independent navigation attitude solution loops. The remaining navigation attitude solution loops in the solution loop are compared, the comparison results are calculated, and the absolute values ​​of the comparison results are obtained, and S4 is executed; S4: whether the absolute values ​​of each comparison result are determined to be less than the preset threshold value, if the absolute values ​​of all the comparison results do not exceed the preset threshold value, a set of solution results of the navigation attitude solution loop is used as the attitude angle used by the control system, and S6 is executed; if there is one or more comparison results whose absolute values ​​exceed or are equal to the preset threshold value, S5 is executed; S5: the gyroscope in the axial instrument combination participating in the solution in the navigation attitude solution loop whose absolute value of the comparison result exceeds or is equal to the preset threshold value is analyzed for faults, after the faulty gyroscope is determined, the faulty gyroscope is isolated, and a group is selected from the navigation attitude solution loop corresponding to the gyroscope that has not been isolated as a new reference, and the solution results of the navigation attitude solution loop corresponding to the new reference are used as the attitude angle used by the control system, and S6 is executed; S6: when the navigation calculation cycle meets the preset conditions of the navigation calculation cycle, S2 to S5 are executed until the flight mission ends.

[0009] As above, there are six groups of navigation attitude solution loops, namely: the first group of navigation attitude solution loops, the second group of navigation attitude solution loops, the third group of navigation attitude solution loops, the fourth group of navigation attitude solution loops, the fifth group of navigation attitude solution loops and the sixth group of navigation attitude solution loops.

[0010] As above, the axial instrument combination involved in the solution of the first group of navigation attitude solution loops includes: the instrument on the X-axis of the inertial group measurement coordinate system, the instrument on the Y-axis of the inertial group measurement coordinate system, and the instrument on the Z-axis of the inertial group measurement coordinate system; the axial instrument combination involved in the solution of the second group of navigation attitude solution loops includes: the instrument on the X-axis of the inertial group measurement coordinate system, the instrument on the Y-axis of the inertial group measurement coordinate system, the instrument on the Z-axis of the inertial group measurement coordinate system, and the instrument on the S-axis of the inertial group measurement coordinate system; the axial instrument combination involved in the solution of the third group of navigation attitude solution loops includes: the instrument on the X-axis of the inertial group measurement coordinate system, the instrument on the Y-axis of the inertial group measurement coordinate system, the instrument on the Z-axis of the inertial group measurement coordinate system, and the instrument on the T-axis of the inertial group measurement coordinate system; the fourth group of navigation attitude solution loops includes: the instrument on the X-axis of the inertial group measurement coordinate system, the instrument on the Y-axis of the inertial group measurement coordinate system, the instrument on the Z-axis of the inertial group measurement coordinate system, and the instrument on the T-axis of the inertial group measurement coordinate system. The axial instrument combination involved in the solution of the heading attitude solution loop includes: the instrument on the X-axis of the inertial group measurement coordinate system, the instrument on the Y-axis of the inertial group measurement coordinate system, the instrument on the S-axis of the inertial group measurement coordinate system, and the instrument on the T-axis of the inertial group measurement coordinate system; the axial instrument combination involved in the solution of the fifth group of navigation attitude solution loop includes: the instrument on the X-axis of the inertial group measurement coordinate system, the instrument on the Z-axis of the inertial group measurement coordinate system, the instrument on the S-axis of the inertial group measurement coordinate system, and the instrument on the T-axis of the inertial group measurement coordinate system; the axial instrument combination involved in the solution of the sixth group of navigation attitude solution loop includes: the instrument on the Y-axis of the inertial group measurement coordinate system, the instrument on the Z-axis of the inertial group measurement coordinate system, the instrument on the S-axis of the inertial group measurement coordinate system, and the instrument on the T-axis of the inertial group measurement coordinate system.

[0011] As above, wherein the instrument comprises at least: a gyroscope and an accelerometer.

[0012] As above, the navigation attitude update is completed according to the traditional quaternion update calculation method, and the updated attitude angle obtained is: i =[φ i ,ψ i ,γ i ], where φ i is the pitch angle of the i-th group of navigation attitude solution loop; ψ i is the yaw angle of the i-th group of navigation attitude solution loop; i The roll angle of the loop solution for the i-th set of navigation attitudes, i∈[1,6].

[0013] As above, the solution results of the first group of navigation attitude solution loops are compared with the solution results of the second group of navigation attitude solution loops, the third group of navigation attitude solution loops, the fourth group of navigation attitude solution loops, the fifth group of navigation attitude solution loops and the sixth group of navigation attitude solution loops, the comparison results are calculated, and the absolute values ​​of the comparison results are obtained.

[0014] As above, when there are six groups of navigation attitude solution loops, the solution result of the first group of navigation attitude solution loops is φ1 = [φ1, ψ1, γ1], and the solution results of the remaining navigation attitude solution loops are j∈[2,6]; the expression of the absolute value of the comparison result is: Among them, Δφ j is the comparison result between the solution result of the jth group of navigation attitude solution loop and the settlement result of the first group of navigation attitude solution loop; φ1 is the pitch angle of the first group of navigation attitude solution loop; ψ1 is the yaw angle of the first group of navigation attitude solution loop; γ1 is the roll angle of the first group of navigation attitude solution loop; φ j is the pitch angle of the jth group of navigation attitude solution loop; ψ j is the yaw angle of the jth group of navigation attitude solution loop; j Calculate the roll angle of the loop for the jth set of navigation attitude solutions.

[0015] As above, where the preset threshold value δ Lim =δ0+k s ·t, where δ0 is the fixed zero bias of the angle deviation, k s is the slope of the attitude angle deviation modulus diverging with navigation time; t is the time length for establishing inertial navigation.

[0016] As above, the navigation calculation cycle is 5 ms.

[0017] As above, the sampling period of the inertia group of the ten-meter inertia group is 5ms; the error compensation period of the ten-meter inertia group is 5ms.

[0018] The beneficial effects achieved by this application are as follows:

[0019] (1) Based on the ten-meter inertial group, this application proposes a fault diagnosis process and algorithm, an information reconstruction scheme and a method for fault diagnosis threshold design through the threshold comparison method. This method identifies the gyroscope fault mode that affects the success or failure of the control system through the navigation attitude solution results in the inertial group measurement coordinate system. While ensuring no missed judgment or misjudgment, it reduces the complexity of the judgment scheme and has good engineering application prospects.

[0020] (2) The present invention discloses a method for detecting and isolating gyro faults based on a ten-meter redundant inertial group of navigation attitude angles. It adopts multiple independent navigation attitude solution loops to obtain corresponding solution results, compares one of the multiple solution results with the other solution results, obtains the absolute value of the comparison result, and analyzes the absolute value of each comparison result through a reasonable preset threshold value, which can improve the accuracy of gyro fault monitoring and isolation.

[0021] (3) The present application's method for detecting and isolating gyro faults in a ten-meter redundant inertial group based on navigation attitude angles adopts multi-channel attitude solution (wherein a group of navigation attitude solution loops is one channel attitude solution) and has a navigation attitude memory function. When a fault occurs, the gyro fault isolation and navigation attitude reset can be performed immediately. BRIEF DESCRIPTION OF THE DRAWINGS

[0022] In order to more clearly illustrate the embodiments of the present application or the technical solutions in the prior art, the drawings required for use in the embodiments or the description of the prior art will be briefly introduced below. Obviously, the drawings described below are only some embodiments recorded in the present application. For ordinary technicians in this field, other drawings can also be obtained based on these drawings.

[0023] Figure 1 It is a schematic diagram of an embodiment of a ten-meter inertial group;

[0024] Figure 2 The coordinate relationship diagram of the gyroscope of the ten-meter inertia group;

[0025] Figure 3 This is the coordinate relationship diagram of the accelerometer of the ten-meter inertia group;

[0026] Figure 4 The present invention is a flow chart of an embodiment of a method for detecting and isolating gyro faults of a ten-meter redundant inertial group based on navigation attitude angle. DETAILED DESCRIPTION

[0027] The following is a clear and complete description of the technical solutions in the embodiments of the present invention 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 those skilled in the art without creative work are within the scope of protection of the present invention.

[0028] like Figure 1-4 As shown, the present application provides a method for detecting and isolating gyro faults of a ten-meter redundant inertial group based on navigation attitude angle, comprising the following steps:

[0029] S1: After the rocket is ignited, the inertial group measurement coordinate system is established. At least six independent navigation attitude solution loops are established in the inertial group measurement coordinate system, the navigation attitude solution is started, and S2 is executed.

[0030] The ten-meter inertial group is set on the installation reference plane, and a point O is selected on the azimuth reference plane of the ten-meter inertial group as the origin of the inertial group measurement coordinate system. The X-axis, Y-axis, Z-axis, inclined S-axis and inclined T-axis are set with the origin as the starting point. The X-axis is perpendicular to the azimuth reference plane, the Y-axis and the Z-axis are both located on the azimuth reference plane, and the Z-axis is perpendicular to the Y-axis. The angles between the S-axis and the X-axis, the Y-axis and the Z-axis are 127°, 47° and 66° respectively; the angles between the T-axis and the X-axis, the Y-axis and the Z-axis are 46.3°, 71° and 49.8° respectively.

[0031] At least one instrument is provided on each of the X-axis, Y-axis, Z-axis, S-axis and T-axis. The number of instruments is set according to actual conditions, and the present application preferably provides one instrument on each of the X-axis, Y-axis, Z-axis, S-axis and T-axis.

[0032] The instrument at least includes: a gyroscope and an accelerometer, but is not limited to a gyroscope and an accelerometer. The present application preferably uses a gyroscope and an accelerometer.

[0033] Specifically, the specific number of independent navigation attitude solution loops established in the inertial group measurement coordinate system is set according to actual conditions. The present application preferably has six groups, namely: the first group of navigation attitude solution loops, the second group of navigation attitude solution loops, the third group of navigation attitude solution loops, the fourth group of navigation attitude solution loops, the fifth group of navigation attitude solution loops and the sixth group of navigation attitude solution loops.

[0034] Figure 1 This is a schematic diagram of an embodiment of a ten-meter inertial group. By designing a combination of instruments in different axes of the ten-meter inertial group, six independent navigation attitude solution loops are established. By analyzing and comparing the six groups of navigation attitude angles, gyro fault isolation and navigation attitude reset are performed. As shown in Table 1, the combination of instruments on different axes of the inertial group measurement coordinate system is specifically:

[0035] The axial instrument combination involved in the calculation of the first group of navigation attitude calculation loops (i.e., when the inertial group axis in Table 1 is taken as the reference) includes: instruments on the X-axis, instruments on the Y-axis, and instruments on the Z-axis.

[0036] The axial instrument combination involved in the solution of the second group of navigation attitude solution loop (i.e., when the inertial group axis in Table 1 is redundancy 1) includes: instruments on the X-axis, instruments on the Y-axis, instruments on the Z-axis, and instruments on the S-axis.

[0037] The axial instrument combination involved in the solution of the third group of navigation attitude solution loop (i.e., when the inertial group axis in Table 1 is redundancy 2) includes: the instrument on the X-axis, the instrument on the Y-axis, the instrument on the Z-axis and the instrument on the T-axis.

[0038] The axial instrument combination involved in the calculation of the fourth group of navigation attitude calculation loop (ie, when the inertial group axis is redundancy 3 in Table 1) includes: the instrument on the X-axis, the instrument on the Y-axis, the instrument on the S-axis and the instrument on the T-axis.

[0039] The axial instrument combination involved in the solution of the fifth group of navigation attitude solution loop (ie, when the inertial group axis is redundant 4 in Table 1) includes: the instrument on the X-axis, the instrument on the Z-axis, the instrument on the S-axis and the instrument on the T-axis.

[0040] The axial instrument combination involved in the sixth group of navigation attitude solution loop (ie, when the inertial group axis is redundancy 5 in Table 1) includes: the instrument on the Y axis, the instrument on the Z axis, the instrument on the S axis and the instrument on the T axis.

[0041]

[0042] Table 1: Inertial group axial combination selection

[0043] S2: Obtain output data through the axial instrument combination participating in the solution in each group of navigation attitude solution loop, convert the output data into the inertial group measurement coordinate system, obtain the output value of the inertial group measurement coordinate system, update the navigation attitude according to the output value of the inertial group measurement coordinate system, obtain the updated output data, and use the updated output data as the solution result; wherein the output data includes: the output navigation attitude angle and the output acceleration; the inertial group measurement coordinate system output value includes: the converted navigation attitude angle and the converted acceleration; the updated output data includes: the updated attitude angle and the updated acceleration, and execute S3.

[0044] Furthermore, the traditional quaternion update calculation method is used to convert the output navigation attitude angle into the inertial group measurement coordinate system to obtain the output value of the inertial group measurement coordinate system, and the navigation attitude is updated according to the output value of the inertial group measurement coordinate system to obtain the updated attitude angle, but it is not limited to the traditional quaternion update calculation method. The present application preferably uses the traditional quaternion update calculation method.

[0045] Specifically, the updated attitude angle obtained by completing the navigation attitude update according to the traditional quaternion update calculation method is: i =[φ i ,ψ i ,γ i ], where φ i is the pitch angle of the i-th group of navigation attitude solution loop; ψ i is the yaw angle of the i-th group of navigation attitude solution loop; i The roll angle of the loop solution for the i-th set of navigation attitudes, i∈[1,6].

[0046] Further, such as Figure 2As shown in the figure, the relationship between the navigation attitude angle output by the gyroscope and the inertial group measurement coordinate system is:

[0047]

[0048] Among them, ω x When the X axis is the measuring axis, the navigation attitude angle actually output by the gyroscope in the instrument along the measuring axis; ω y is the navigation attitude angle actually output by the gyroscope along the measurement axis when the Y axis is the measurement axis; ω z When the Z axis is the measurement axis, the navigation attitude angle actually output by the gyroscope along the measurement axis; ω S is the navigation attitude angle actually output by the gyroscope along the measurement axis when the S axis is the measurement axis; ω T is the navigation attitude angle actually output by the gyroscope along the measurement axis when the T axis is the measurement axis; ω xM is the angular velocity component of the X-axis of the gyroscope in the inertial measurement coordinate system; ω yM is the angular velocity component of the gyroscope on the Y axis of the inertial measurement coordinate system; ω zM is the angular velocity component of the Z axis of the gyroscope in the inertial measurement coordinate system; E yx E is the installation error angle of the Y-axis gyroscope relative to the X-axis of the inertial measurement coordinate system; xy E is the installation error angle of the X-axis gyroscope relative to the Y-axis of the inertial measurement coordinate system; xz E is the installation error angle of the X-axis gyroscope relative to the inertial measurement coordinate system; zx E is the installation error angle of the Z-axis gyroscope relative to the X-axis of the inertial measurement coordinate system; zy E is the installation error angle of the Z-axis gyroscope relative to the Y-axis of the inertial measurement coordinate system; yz θ is the installation error angle of the Y-axis gyroscope relative to the Z-axis of the inertial measurement coordinate system; gxS θ is the angle between the inclined S-axis gyroscope and the X-axis of the inertial measurement coordinate system; gyS θ is the angle between the inclined S-axis gyroscope and the Y-axis of the inertial measurement coordinate system; gzS θ is the angle between the inclined S-axis gyroscope and the Z-axis of the inertial measurement coordinate system; gxT θ is the angle between the inclined T-axis gyroscope and the X-axis of the inertial measurement coordinate system; gyT θ is the angle between the inclined T-axis gyroscope and the Y-axis of the inertial measurement coordinate system; gzT is the angle between the inclined T-axis gyroscope and the Z-axis of the inertial measurement coordinate system; cos(·) is the cosine calculation; H g For the matrix ω M For the matrix

[0049] Specifically, θgxS ,θ gyS ,θ gzS ,θ gxT ,θ gyT and θ gzT The values ​​are the factory calibration parameters of the ten-meter inertia group.

[0050] Furthermore, the converted navigation attitude angle in the inertial group measurement coordinate system output value of the i-group navigation attitude solution loop is calculated according to the relationship between the navigation attitude angle output by the gyroscope and the inertial group measurement coordinate system.

[0051] When i=1, it indicates that the navigation attitude solution loop is the first group of navigation attitude solution loops. The expression of the converted navigation attitude angle in the output value of the inertial measurement coordinate system of the first group of navigation attitude solution loops is:

[0052]

[0053] Among them, ω x1 When i=1, when the X axis is the measuring axis, the navigation attitude angle actually output by the gyroscope in the instrument along the measuring axis; ω y1 When i=1, when the Y axis is the measurement axis, the navigation attitude angle actually output by the gyroscope along the measurement axis; ω z1 When i=1, when the Z axis is the measurement axis, the navigation attitude angle actually output by the gyroscope along the measurement axis; H g1 For the matrix ω M1 When i=1, the converted navigation attitude angle; H g1 The transpose of .

[0054] When i=2, it indicates that the navigation attitude solution loop is the second group of navigation attitude solution loops. The expression of the converted navigation attitude angle in the inertial measurement coordinate system output value of the second group of navigation attitude solution loops is:

[0055]

[0056]

[0057] Among them, ω x2 When i=2, when the X axis is the measuring axis, the navigation attitude angle actually output by the gyroscope in the instrument along the measuring axis; ω y2 When i=2, when the Y axis is the measurement axis, the navigation attitude angle actually output by the gyroscope along the measurement axis; ω z2 When i=2, when the Z axis is the measurement axis, the navigation attitude angle actually output by the gyroscope along the measurement axis; ω S2 When i=2, when the S axis is the measurement axis, the navigation attitude angle actually output by the gyroscope along the measurement axis; H g2 For the matrix ω M2 When i=2, the converted navigation attitude angle; H g2 The transpose of .

[0058] When i=3, it indicates that the navigation attitude solution loop is the third group of navigation attitude solution loops. The expression of the converted navigation attitude angle in the output value of the inertial measurement coordinate system of the third group of navigation attitude solution loops is:

[0059]

[0060] Among them, ω x3 When i=3, when the X axis is the measuring axis, the navigation attitude angle actually output by the gyroscope in the instrument along the measuring axis; ω y3 When i=3, when the Y axis is the measurement axis, the navigation attitude angle actually output by the gyroscope along the measurement axis; ω z3 When i=3, when the Z axis is the measurement axis, the navigation attitude angle actually output by the gyroscope along the measurement axis; ω T3 When i=3, when the T axis is the measurement axis, the navigation attitude angle actually output by the gyroscope along the measurement axis; H g3 For the matrix ω M3 When i=3, the converted navigation attitude angle; H g3 The transpose of .

[0061] When i=4, it indicates that the navigation attitude solution loop is the fourth group of navigation attitude solution loops. The expression of the converted navigation attitude angle in the output value of the inertial measurement coordinate system of the fourth group of navigation attitude solution loops is:

[0062]

[0063] Among them, ω x4 When i=4, when the X axis is the measuring axis, the navigation attitude angle actually output by the gyroscope in the instrument along the measuring axis; ω y4 When i=4, when the Y axis is the measurement axis, the navigation attitude angle actually output by the gyroscope along the measurement axis; ω S4 When i=4, when the S axis is the measurement axis, the navigation attitude angle actually output by the gyroscope along the measurement axis; ω T4 When i=4, when the T axis is the measurement axis, the navigation attitude angle actually output by the gyroscope along the measurement axis; H g4 For the matrix ω M4 When i=4, the converted navigation attitude angle; H g4 The transpose of .

[0064] When i=5, it indicates that the navigation attitude solution loop is the fifth group of navigation attitude solution loops. The expression of the converted navigation attitude angle in the inertial measurement coordinate system output value of the fifth group of navigation attitude solution loops is:

[0065]

[0066] Among them, ω x5 When i=5, when the X axis is the measuring axis, the navigation attitude angle actually output by the gyroscope in the instrument along the measuring axis; ω z5 When i=5, when the Z axis is the measurement axis, the navigation attitude angle actually output by the gyroscope along the measurement axis; ω S5 When i=5, when the S axis is the measurement axis, the navigation attitude angle actually output by the gyroscope along the measurement axis; ω T5 When i=5, when the T axis is the measurement axis, the navigation attitude angle actually output by the gyroscope along the measurement axis; H g5 For the matrix ω M5 When i=5, the converted navigation attitude angle; H g5 The transpose of .

[0067] When i=6, it indicates that the navigation attitude solution loop is the sixth group of navigation attitude solution loops. The expression of the converted navigation attitude angle in the output value of the inertial measurement coordinate system of the sixth group of navigation attitude solution loops is:

[0068]

[0069] Among them, ω y6 When i=6, when the Y axis is the measuring axis, the navigation attitude angle actually output by the gyroscope in the instrument along the measuring axis; ω z6 When i=6, when the Z axis is the measurement axis, the navigation attitude angle actually output by the gyroscope along the measurement axis; ω S6 When i=6, when the S axis is the measurement axis, the navigation attitude angle actually output by the gyroscope along the measurement axis; ω T6 When i=6, when the T axis is the measurement axis, the navigation attitude angle actually output by the gyroscope along the measurement axis; H g6 For the matrix ω M6 When i=6, the converted navigation attitude angle; H g6 The transpose of .

[0070] Further, such as Figure 3As shown in the figure, the relationship between the acceleration output by the accelerometer and the inertial measurement coordinate system is:

[0071]

[0072] Among them, A x When the X axis is the measuring axis, the acceleration actually output by the accelerometer in the instrument along the measuring axis; A y A is the acceleration actually output by the accelerometer along the measuring axis when the Y axis is the measuring axis; z When the Z axis is the measuring axis, the acceleration actually output by the accelerometer along the measuring axis; A S When the S axis is the measuring axis, the acceleration actually output by the accelerometer along the measuring axis; A T A is the acceleration actually output by the accelerometer along the measuring axis when the axis is the measuring axis; xM A is the apparent acceleration component of the X-axis of the inertial measurement coordinate system; yM A is the apparent acceleration component of the Y axis of the inertial measurement coordinate system; zM K is the apparent acceleration component of the Z axis of the inertial measurement coordinate system; xy K is the installation error angle of the X-axis accelerometer relative to the Y-axis of the inertial measurement coordinate system; xz K is the installation error angle of the X-axis accelerometer relative to the Z-axis of the inertial measurement coordinate system; yz θ is the installation error angle of the Y-axis accelerometer relative to the Z-axis of the inertial measurement coordinate system; AxS θ is the angle between the inclined S-axis accelerometer and the X-axis of the inertial measurement coordinate system; AyS θ is the angle between the inclined S-axis accelerometer and the Y-axis of the inertial measurement coordinate system; AzS θ is the angle between the inclined S-axis accelerometer and the Z-axis of the inertial measurement coordinate system; AxT θ is the angle between the inclined T-axis accelerometer and the X-axis of the inertial measurement coordinate system; AyT θ is the angle between the inclined T-axis accelerometer and the Y-axis of the inertial measurement coordinate system; AzT H is the angle between the inclined T-axis accelerometer and the Z-axis of the inertial measurement coordinate system; a For the matrix A M for

[0073] Specifically, K xy , K xz and K yz The factory calibration parameters of the IMU. AxS ,θ AyS ,θ AzS ,θ AxT ,θ AyT and θ AzTCalibrate the parameters of the inertial group before leaving the factory.

[0074] Furthermore, the converted acceleration in the inertial group measurement coordinate system output value of the i-group navigation attitude solution loop is calculated according to the relationship between the acceleration output by the accelerometer and the inertial group measurement coordinate system.

[0075] When i=1, it indicates that the navigation attitude solution loop is the first group of navigation attitude solution loops. The expression of the converted acceleration in the output value of the inertial measurement coordinate system of the first group of navigation attitude solution loops is:

[0076]

[0077] Among them, A x1 When i=1, when the X axis is the measuring axis, the acceleration actually output by the accelerometer in the instrument along the measuring axis; A y1 When i=1, when the Y axis is the measuring axis, the acceleration actually output by the accelerometer along the measuring axis; A z1 When i=1, when the Z axis is the measuring axis, the acceleration actually output by the accelerometer along the measuring axis; H a1 For the matrix A M1 When i=1, the converted acceleration; H a1 The transpose of .

[0078] When i=2, it indicates that the navigation attitude solution loop is the second group of navigation attitude solution loops. The expression of the converted acceleration in the output value of the inertial measurement coordinate system of the second group of navigation attitude solution loops is:

[0079]

[0080] Among them, A x2 When i=2, when the X axis is the measuring axis, the acceleration actually output by the accelerometer in the instrument along the measuring axis; A y2 When i=2, when the Y axis is the measuring axis, the acceleration actually output by the accelerometer along the measuring axis; A z2 When i=2, when the Z axis is the measuring axis, the acceleration actually output by the accelerometer along the measuring axis; A S2 When i=2, when the S axis is the measuring axis, the acceleration actually output by the accelerometer along the measuring axis; H a2 For the matrix A M2 When i=2, the converted acceleration; H a2 The transpose of .

[0081] When i=3, it indicates that the navigation attitude solution loop is the third group of navigation attitude solution loops. The expression of the converted acceleration in the output value of the inertial measurement coordinate system of the third group of navigation attitude solution loops is:

[0082]

[0083] Among them, A x3 When i=3, when the X axis is the measuring axis, the acceleration actually output by the accelerometer in the instrument along the measuring axis; A y3 When i=3, when the Y axis is the measuring axis, the acceleration actually output by the accelerometer along the measuring axis; A z3 When i=3, when the Z axis is the measuring axis, the acceleration actually output by the accelerometer along the measuring axis; A T3 When i=3, when the T axis is the measuring axis, the acceleration actually output by the accelerometer along the measuring axis; H a3 For the matrix A M3 When i=3, the converted acceleration; H a3 The transpose of .

[0084] When i=4, it indicates that the navigation attitude solution loop is the fourth group of navigation attitude solution loops. The expression of the converted acceleration in the output value of the inertial measurement coordinate system of the fourth group of navigation attitude solution loops is:

[0085]

[0086] Among them, A x4 When i=4, when the X axis is the measuring axis, the acceleration actually output by the accelerometer in the instrument along the measuring axis; A y4 When i=4, when the Y axis is the measuring axis, the acceleration actually output by the accelerometer along the measuring axis; A S4 When i=4, when the S axis is the measuring axis, the acceleration actually output by the accelerometer along the measuring axis; A T4 When i=4, when the T axis is the measuring axis, the acceleration actually output by the accelerometer along the measuring axis; H a4 For the matrix A M4 When i=4, the converted acceleration; H a4 The transpose of .

[0087] When i=5, it indicates that the navigation attitude solution loop is the fifth group of navigation attitude solution loops. The expression of the converted acceleration in the output value of the inertial measurement coordinate system of the fifth group of navigation attitude solution loops is:

[0088]

[0089] Among them, A x5 When i=5, when the X axis is the measuring axis, the acceleration actually output by the accelerometer in the instrument along the measuring axis; A z5 When i=5, when the Z axis is the measuring axis, the acceleration actually output by the accelerometer along the measuring axis; A S5 When i=5, when the S axis is the measuring axis, the acceleration actually output by the accelerometer along the measuring axis; A T5 When i=5, when the T axis is the measuring axis, the acceleration actually output by the accelerometer along the measuring axis; H a5 For the matrix A M5 When i=5, the converted acceleration; H a5 The transpose of .

[0090] When i=6, it indicates that the navigation attitude solution loop is the sixth group of navigation attitude solution loops. The expression of the converted acceleration in the output value of the inertial measurement coordinate system of the sixth group of navigation attitude solution loops is:

[0091]

[0092] Among them, A y6 When i=6, when the Y axis is the measuring axis, the acceleration actually output by the accelerometer in the instrument along the measuring axis; A z6 When i=6, when the Z axis is the measuring axis, the acceleration actually output by the accelerometer along the measuring axis; A S6 When i=6, when the S axis is the measuring axis, the acceleration actually output by the accelerometer along the measuring axis; A T6 When i=6, when the T axis is the measuring axis, the acceleration actually output by the accelerometer along the measuring axis; H a6 For the matrix A M6 When i=6, the converted acceleration; H a6 The transpose of .

[0093] S3: using the solution results of one set of navigation attitude solution loops in at least six sets of independent navigation attitude solution loops to compare with the remaining navigation attitude solution loops in at least six sets of independent navigation attitude solution loops, respectively, calculating the comparison results, and obtaining the absolute value of the comparison results, and executing S4.

[0094] Further, when there are six groups of navigation attitude solution loops, the solution results of the first group of navigation attitude solution loops are compared with the solution results of the second group of navigation attitude solution loops, the third group of navigation attitude solution loops, the fourth group of navigation attitude solution loops, the fifth group of navigation attitude solution loops and the sixth group of navigation attitude solution loops, respectively, the comparison results are calculated, and the absolute value of the comparison results is obtained, and S4 is executed.

[0095] Furthermore, when there are six groups of navigation attitude solution loops, the solution result of the first group of navigation attitude solution loops is φ1 = [φ1, ψ1, γ1], and the solution results of the remaining navigation attitude solution loops are j∈[2,6].

[0096] The formula for calculating the comparison result is:

[0097] Δφ j =[φ j -φ1ψ j -ψ1γ j -γ1];

[0098] Among them, Δφ j is the comparison result between the solution result of the jth group of navigation attitude solution loop and the settlement result of the first group of navigation attitude solution loop; φ1 is the pitch angle of the first group of navigation attitude solution loop; ψ1 is the yaw angle of the first group of navigation attitude solution loop; γ1 is the roll angle of the first group of navigation attitude solution loop; φ j is the pitch angle of the jth group of navigation attitude solution loop; ψ j is the yaw angle of the jth group of navigation attitude solution loop; j Calculate the roll angle of the loop for the jth set of navigation attitude solutions.

[0099] Furthermore, the expression of the absolute value of the comparison result is:

[0100] where j∈[2,6].

[0101] S4: Determine whether the absolute value of each comparison result is less than the preset threshold value. If the absolute values ​​of all comparison results do not exceed the preset threshold value, a set of solution results of the navigation attitude solution loop is used as the attitude angle used by the control system, and execute S6; if the absolute value of one or more comparison results exceeds or is equal to the preset threshold value, execute S5.

[0102] Furthermore, the specific value of the preset threshold value is set according to the actual situation. The present application is preferably: the preset threshold value δ Lim =δ0+k s ·t, where δ0 is the fixed zero bias of the angle deviation, k sis the slope of the attitude angle deviation modulus diverging with navigation time; t is the time length for establishing inertial navigation.

[0103] Specifically, to prevent misjudgment or missed judgment, δ0, k s The selection needs to be closely combined with the angular velocity characteristics of the entire flight trajectory for design, which are the parameters obtained by simulation calculation and other methods.

[0104] If |Δφ2|~|Δφ6| do not exceed (i.e., less than or equal to) the preset threshold value δ Lim , it is considered that the outputs of each instrument combination of the ten-meter inertial group are normal, and φ1=[φ1, ψ1, γ1] is used as the attitude angle used by the control system.

[0105] If |Δφ2|~|Δφ6| has one or more |Δφ i | Exceeds (i.e., is greater than) the preset threshold value δ Lim , the existing method is used to identify the gyroscope fault, and the angular velocity information, maximum value and consistency judgment of the gyroscope are combined to screen out the axis system where the faulty gyroscope is located, and isolate the gyroscope.

[0106] S5: Perform fault analysis on the gyroscopes in the axial instrument combination participating in the navigation attitude solution loop whose absolute value of the comparison result exceeds or is equal to the preset threshold value. After the faulty gyroscope is determined, the faulty gyroscope is isolated, and a group is selected from the navigation attitude solution loop corresponding to the gyroscopes that have not been isolated as a new reference. The solution result of the navigation attitude solution loop corresponding to the new reference is used as the attitude angle used by the control system, and execute S6.

[0107] Specifically, according to the direction of the faulty gyroscope, the redundant group is eliminated. If the three orthogonal gyroscopes (i.e., the gyroscopes on the X-axis, Y-axis, and Z-axis) fail, one of the 3rd to 5th redundant modes (i.e., the third navigation attitude solution loop, the fourth navigation attitude solution loop, and the fifth navigation attitude solution loop) that do not contain the gyroscope is selected as the basis for subsequent navigation attitude solution, and the attitude angle solved by this group is used as the attitude angle used by the control system.

[0108] S6: When the navigation calculation cycle meets the preset conditions of the navigation calculation cycle, execute S2 to S5 until the flight mission ends.

[0109] Furthermore, the preset conditions of the navigation calculation cycle are set according to actual conditions. The preferred embodiment of the present application is: the navigation calculation cycle is Pms, P is a natural number, and the preferred embodiment of the present application is: P=5, that is, the navigation calculation cycle is 5ms.

[0110] Furthermore, the specific value of the sampling period of the inertia group of the ten-meter inertia group is set according to the actual situation. The preferred value of the present application is: the sampling period of the inertia group is 5ms.

[0111] Furthermore, the specific value of the error compensation period of the ten-meter inertia group is set according to the actual situation. In the present application, the error compensation period is preferably 5 ms.

[0112] The beneficial effects achieved by this application are as follows:

[0113] (1) Based on the ten-meter inertial group, this application proposes a fault diagnosis process and algorithm, an information reconstruction scheme and a method for fault diagnosis threshold design through the threshold comparison method. This method identifies the gyroscope fault mode that affects the success or failure of the control system through the navigation attitude solution results in the inertial group measurement coordinate system. While ensuring no missed judgment or misjudgment, it reduces the complexity of the judgment scheme and has good engineering application prospects.

[0114] (2) The present invention discloses a method for detecting and isolating gyro faults based on a ten-meter redundant inertial group of navigation attitude angles. It adopts multiple independent navigation attitude solution loops to obtain corresponding solution results, compares one of the multiple solution results with the other solution results, obtains the absolute value of the comparison result, and analyzes the absolute value of each comparison result through a reasonable preset threshold value, which can improve the accuracy of gyro fault monitoring and isolation.

[0115] (3) The present application's method for detecting and isolating gyro faults in a ten-meter redundant inertial group based on navigation attitude angles adopts multi-channel attitude solution (wherein a group of navigation attitude solution loops is one channel attitude solution) and has a navigation attitude memory function. When a fault occurs, the gyro fault isolation and navigation attitude reset can be performed immediately.

[0116] Although the preferred embodiments of the present application have been described, those skilled in the art may make additional changes and modifications to these embodiments once they are aware of the basic creative concepts. Therefore, the scope of protection of the present application is intended to be interpreted as including the preferred embodiments and all changes and modifications that fall within the scope of the present application. Obviously, those skilled in the art may make various changes and modifications to the present application without departing from the spirit and scope of the present application. Thus, if these modifications and variations of the present application fall within the scope of the protection scope of the present application and its equivalent technology, the present application is also intended to include these modifications and variations.

Claims

1. A method for detecting and isolating gyro faults of a ten-meter redundant inertial group based on navigation attitude angle, characterized in that: The steps include: S1: After the rocket is ignited, the inertial measurement coordinate system is established, and at least six independent navigation attitude solution loops are established in the inertial measurement coordinate system, and the navigation attitude solution is started, and S2 is executed; S2: Obtain output data through the axial instrument combination participating in the solution in each group of navigation attitude solution loop, convert the output data into the inertial group measurement coordinate system, obtain the output value of the inertial group measurement coordinate system, update the navigation attitude according to the output value of the inertial group measurement coordinate system, obtain the updated output data, and use the updated output data as the solution result; wherein the output data includes: the output navigation attitude angle and the output acceleration; the inertial group measurement coordinate system output value includes: the converted navigation attitude angle and the converted acceleration; the updated output data includes: the updated attitude angle and the updated acceleration, and execute S3; S3: using the solution results of one set of navigation attitude solution loops in the at least six sets of independent navigation attitude solution loops to compare with the remaining navigation attitude solution loops in the at least six sets of independent navigation attitude solution loops, respectively, calculating the comparison results, and obtaining the absolute value of the comparison results, and executing S4; S4: Determine whether the absolute value of each comparison result is less than a preset threshold value. If the absolute values ​​of all comparison results do not exceed the preset threshold value, a set of navigation attitude solution loop solution results are used as the attitude angle used by the control system, and S6 is executed; if the absolute value of one or more comparison results exceeds or is equal to the preset threshold value, S5 is executed; S5: Perform fault analysis on the gyroscopes in the axial instrument combination involved in the navigation attitude solution loop whose absolute value of the comparison result exceeds or is equal to the preset threshold value, isolate the faulty gyroscope after determining the faulty gyroscope, select a group from the navigation attitude solution loop corresponding to the unisolated gyroscope as a new reference, use the solution result of the navigation attitude solution loop corresponding to the new reference as the attitude angle used by the control system, and execute S6; S6: When the navigation calculation cycle meets the preset conditions of the navigation calculation cycle, execute S2 to S5 until the flight mission ends.

2. The method for detecting and isolating gyro faults of a ten-meter redundant inertial group based on navigation attitude angle according to claim 1 is characterized in that: There are six groups of navigation attitude solution loops, namely: the first group of navigation attitude solution loops, the second group of navigation attitude solution loops, the third group of navigation attitude solution loops, the fourth group of navigation attitude solution loops, the fifth group of navigation attitude solution loops and the sixth group of navigation attitude solution loops.

3. The method for detecting and isolating gyro faults of a ten-meter redundant inertial group based on navigation attitude angle according to claim 2 is characterized in that: The axial instrument combination involved in the solution of the first group of navigation attitude solution loop includes: the instrument on the X-axis of the inertial group measurement coordinate system, the instrument on the Y-axis of the inertial group measurement coordinate system, and the instrument on the Z-axis of the inertial group measurement coordinate system; The axial instrument combination involved in the solution in the second group of navigation attitude solution loop includes: the instrument on the X-axis of the inertial group measurement coordinate system, the instrument on the Y-axis of the inertial group measurement coordinate system, the instrument on the Z-axis of the inertial group measurement coordinate system, and the instrument on the S-axis of the inertial group measurement coordinate system; The axial instrument combination involved in the solution of the third group of navigation attitude solution loop includes: the instrument on the X-axis of the inertial group measurement coordinate system, the instrument on the Y-axis of the inertial group measurement coordinate system, the instrument on the Z-axis of the inertial group measurement coordinate system, and the instrument on the T-axis of the inertial group measurement coordinate system; The axial instrument combination involved in the solution in the fourth group of navigation attitude solution loop includes: an instrument on the X-axis of the inertial group measurement coordinate system, an instrument on the Y-axis of the inertial group measurement coordinate system, an instrument on the S-axis of the inertial group measurement coordinate system, and an instrument on the T-axis of the inertial group measurement coordinate system; The axial instrument combination involved in the solution in the fifth group of navigation attitude solution loop includes: an instrument on the X-axis of the inertial group measurement coordinate system, an instrument on the Z-axis of the inertial group measurement coordinate system, an instrument on the S-axis of the inertial group measurement coordinate system, and an instrument on the T-axis of the inertial group measurement coordinate system; The axial instrument combination participating in the solution of the sixth group of navigation attitude solution loop includes: an instrument on the Y axis of the inertial group measurement coordinate system, an instrument on the Z axis of the inertial group measurement coordinate system, an instrument on the S axis of the inertial group measurement coordinate system, and an instrument on the T axis of the inertial group measurement coordinate system.

4. The method for detecting and isolating gyro faults of a ten-meter redundant inertial group based on navigation attitude angle according to claim 3 is characterized in that: The instrument includes at least: a gyroscope and an accelerometer.

5. The method for detecting and isolating gyro faults of a ten-meter redundant inertial group based on navigation attitude angle according to claim 4 is characterized in that: According to the traditional quaternion update calculation method, the navigation attitude update is completed, and the updated attitude angle is obtained: i =[φ i ,ψ i ,γ i ], where φ i is the pitch angle of the i-th group of navigation attitude solution loop; ψ i is the yaw angle of the i-th group of navigation attitude solution loop; i The roll angle of the loop solution for the i-th set of navigation attitudes, i∈[1,6].

6. The method for detecting and isolating gyro faults of a ten-meter redundant inertial group based on navigation attitude angle according to claim 5 is characterized in that: The solution results of the first group of navigation attitude solution loops are compared with the solution results of the second group of navigation attitude solution loops, the third group of navigation attitude solution loops, the fourth group of navigation attitude solution loops, the fifth group of navigation attitude solution loops and the sixth group of navigation attitude solution loops, the comparison results are calculated, and the absolute values ​​of the comparison results are obtained.

7. The method for detecting and isolating gyro faults of a ten-meter redundant inertial group based on navigation attitude angle according to claim 6 is characterized in that: When there are six groups of navigation attitude solution loops, the solution result of the first group of navigation attitude solution loops is φ1=[φ1,ψ1,γ1], and the solution results of the remaining navigation attitude solution loops are φ j =[φ j ,ψ j ,γ j ], j∈[2,6]; The expression for the absolute value of the comparison result is: Among them, Δφ j is the comparison result between the solution result of the jth group of navigation attitude solution loop and the settlement result of the first group of navigation attitude solution loop; φ1 is the pitch angle of the first group of navigation attitude solution loop; ψ1 is the yaw angle of the first group of navigation attitude solution loop; γ1 is the roll angle of the first group of navigation attitude solution loop; φ j is the pitch angle of the jth group of navigation attitude solution loop; ψ j is the yaw angle of the jth group of navigation attitude solution loop; j Calculate the roll angle of the loop for the jth set of navigation attitude solutions.

8. The method for detecting and isolating gyro faults of a ten-meter redundant inertial group based on navigation attitude angle according to claim 1, characterized in that: The preset threshold value δ Lim =δ0+k s ·t, where δ0 is the fixed zero bias of the angle deviation, k s is the slope of the attitude angle deviation modulus diverging with navigation time; t is the time length for establishing inertial navigation.

9. The method for detecting and isolating gyro faults of a ten-meter redundant inertial group based on navigation attitude angle according to claim 1, characterized in that: The navigation calculation cycle is 5ms.

10. The method for detecting and isolating gyro faults of a ten-meter redundant inertial group based on navigation attitude angle according to claim 1, characterized in that: The sampling period of the ten-meter inertia group is 5ms; the error compensation period of the ten-meter inertia group is 5ms.

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