Gyro fault detection and isolation method for strapdown inertial navigation system based on navigation attitude angle

Abstract

The application discloses a kind of based on navigation attitude angle's ten table redundant inertial measurement unit gyro fault detection and isolation method, it is related to aerospace technology field, including: at least six groups of independent navigation attitude solution loop are established under the established inertial measurement unit measurement coordinate system, the absolute value of comparison result is obtained by navigation attitude solution loop, it is judged whether the size of each comparison result absolute value is less than the preset threshold value, if all do not exceed the preset threshold value, the solution result of a group of navigation attitude solution loop is used as the attitude angle used by control system;If there is one or more comparison results absolute value exceeds or equal to the preset threshold value, the gyro of fault is isolated, selects a group from the navigation attitude solution loop corresponding to the gyro not isolated as new reference, and the solution result of the navigation attitude solution loop corresponding to new reference is used as the attitude angle used by control system.The application can improve the accuracy of gyro fault monitoring and isolation.

Description

Technical Field

[0001] This application relates to the field of aerospace technology, and in particular to a method for fault detection and isolation of a ten-table redundant inertial navigation system gyroscope based on navigation attitude angle. Background Technology

[0002] With the rapid increase in the number of space launch missions in China, the ultra-high density of launches and the diverse launch mission requirements have placed higher demands on the research and development efficiency, reliability, fault tolerance, and economic cost of launch vehicle control systems. Redundancy configuration is a common technical approach to improve the reliability and fault tolerance of navigation systems, generally divided into system-level redundancy and device-level redundancy. System-level redundancy includes dual-INS master-slave redundancy and triple-INS redundancy, etc.; device-level redundancy includes single-INS multi-table redundancy, etc. Among them, the single-set ten-table configuration of INS has advantages over system-level redundancy in terms of size, weight, cost, and simplicity, and the single-set ten-table configuration of INS has gradually become the main inertial device of subsequent launch vehicle models.

[0003] Launch vehicles, due to their high dynamic requirements, demand fault diagnosis systems capable of rapidly and accurately processing fault data and promptly locating and isolating fault information, thereby ensuring the successful completion of flight missions through navigation information reconstruction. Currently, various fault diagnosis methods for inertial devices have been developed, including threshold comparison, generalized likelihood ratio, mean test, singular value decomposition, and neural network methods.

[0004] As a key component of launch vehicle navigation, inertial navigation systems often rely heavily on gyroscope technology to improve system reliability. By employing different gyroscope combinations from a ten-dial inertial navigation system to calculate navigation attitude in an inertial frame and comparing the consistency of attitude angle calculation results from different combinations in real time, gyroscope fault detection and isolation can be achieved.

[0005] In fault detection, the threshold value directly affects whether a fault can be detected quickly enough and the degree of false alarms, thus directly impacting the system's fault diagnosis performance and task reliability. A fault detection threshold that is too large increases the false alarm rate, while a threshold that is too small increases the false alarm rate. These error rates have a significant impact on the system's fault diagnosis performance. Therefore, determining a reasonable fault detection threshold is crucial to the effectiveness of fault detection.

[0006] Traditional methods for detecting and isolating faults in gyroscopes in a 10-meter redundant inertial measurement unit (TIM) mainly involve diagnosing the faults at the instrument pulse conversion stage. These methods are highly susceptible to external environmental factors, which can lead to false alarms or missed alarms due to gyroscope pulse faults. Summary of the Invention

[0007] The purpose of this application is to provide a method for fault detection and isolation of a ten-table redundant inertial navigation system gyroscope based on navigation attitude angle, which can improve the accuracy of gyroscope fault monitoring and isolation.

[0008] To achieve the above objectives, this application provides a method for fault detection and isolation of a ten-instrument redundant inertial navigation gyroscope based on navigation attitude angles, comprising the following steps: S1: After rocket ignition, an inertial navigation measurement coordinate system is established, and at least six independent navigation attitude calculation loops are established under the inertial navigation measurement coordinate system. Navigation attitude calculation begins, and S2 is executed; S2: Output data is obtained through the axial instrument combination participating in the calculation in each navigation attitude calculation loop. The output data is converted to the inertial navigation measurement coordinate system to obtain the output value of the inertial navigation measurement coordinate system. The navigation attitude is updated according to the output value of the inertial navigation measurement coordinate system to obtain the updated output data, and the updated output data is used as the calculation result; wherein, the output data includes: the output navigation attitude angle and the output acceleration; the output value of the inertial navigation measurement coordinate system 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 calculation result of one navigation attitude calculation loop from the at least six independent navigation attitude calculation loops is compared with the calculation result of the at least six independent navigation attitude calculation loops. The remaining navigation attitude calculation loops in the calculation loop are compared, the comparison results are calculated, and the absolute value of the comparison results is obtained. S4 is executed. S4: It is determined whether the absolute value of each comparison result is less than the preset threshold value. If the absolute value of all comparison results does not exceed the preset threshold value, the calculation results of a set of navigation attitude calculation loops are used as the attitude angles used by the control system. S6 is executed. If the absolute value of one or more comparison results exceeds or equals the preset threshold value, S5 is executed. S5: For the gyroscopes in the axial instrument group participating in the calculation of the navigation attitude calculation loop whose absolute value of the comparison result exceeds or equals the preset threshold value, fault analysis is performed. After identifying the faulty gyroscope, the faulty gyroscope is isolated. A set of navigation attitude calculation loops corresponding to the unisolated gyroscopes is selected as the new reference. The calculation results of the navigation attitude calculation loop corresponding to the new reference are used as the attitude angles used by the control system. S6 is executed. S6: When the navigation calculation cycle meets the preset conditions of the navigation calculation cycle, S2~S5 are executed until the flight mission ends.

[0009] As shown above, there are six navigation attitude calculation loops: the first navigation attitude calculation loop, the second navigation attitude calculation loop, the third navigation attitude calculation loop, the fourth navigation attitude calculation loop, the fifth navigation attitude calculation loop, and the sixth navigation attitude calculation loop.

[0010] As described above, the first group of navigation attitude calculation loops includes the following axial instrument combinations: instruments on the X-axis, Y-axis, and Z-axis of the inertial measurement system (INS); the second group of navigation attitude calculation loops includes the following axial instrument combinations: instruments on the X-axis, Y-axis, Z-axis, and S-axis of the INS; the third group of navigation attitude calculation loops includes the following axial instrument combinations: instruments on the X-axis, Y-axis, Z-axis, and T-axis of the INS; the fourth group of navigation attitude calculation loops... The axial instrument combinations involved in the attitude calculation loop include: instruments on the X-axis, Y-axis, S-axis, and T-axis of the inertial measurement system (INS); the axial instrument combinations involved in the attitude calculation loop include: instruments on the X-axis, Z-axis, S-axis, and T-axis of the INS; the axial instrument combinations involved in the attitude calculation loop include: instruments on the X-axis, Z-axis, S-axis, and T-axis of the INS; the axial instrument combinations involved in the attitude calculation loop include: instruments on the Y-axis, Z-axis, S-axis, and T-axis of the INS.

[0011] As mentioned above, the instruments include at least a gyroscope and an accelerometer.

[0012] As shown above, the navigation attitude update is performed using the traditional quaternion update calculation method, and the updated attitude angles are: ,in, The pitch angle for the i-th navigation attitude calculation loop; The yaw angle for the i-th navigation attitude calculation loop; The roll angle of the i-th navigation attitude calculation loop. .

[0013] As shown above, the calculation results of the first group of navigation attitude calculation loops are compared with the calculation results of the second group of navigation attitude calculation loops, the third group of navigation attitude calculation loops, the fourth group of navigation attitude calculation loops, the fifth group of navigation attitude calculation loops, and the sixth group of navigation attitude calculation loops, respectively. The comparison results are calculated and the absolute value of the comparison results is obtained.

[0014] As shown above, when there are six navigation attitude calculation loops, the calculation result of the first navigation attitude calculation loop is: The solution results for the remaining navigation attitude calculation loops are: , The expression for the absolute value of the comparison result is:

[0015] ;

[0016] in, This is a comparison result between the solution result of the j-th navigation attitude solution loop and the settlement result of the first navigation attitude solution loop; The pitch angle for the first navigation attitude calculation loop; The yaw angle for the first set of navigation attitude calculation loops; The roll angle for the first set of navigation attitude calculation loops; The pitch angle for the j-th navigation attitude calculation loop; The yaw angle for the j-th navigation attitude calculation loop; The roll angle of the j-th navigation attitude calculation loop.

[0017] As shown above, the preset threshold value ,in, To fix the angular deviation to zero bias, t represents the slope of the attitude angle deviation modulus diverging with navigation time; t represents the time length for establishing the inertial navigation system.

[0018] As shown above, the navigation calculation cycle is 5ms.

[0019] As shown above, the sampling period of the 10-meter inertial navigation system is 5ms; the error compensation period of the 10-meter inertial navigation system is 5ms.

[0020] The beneficial effects achieved by this application are as follows:

[0021] (1) Based on the ten-table inertial navigation system, this application proposes a fault diagnosis process and algorithm, information reconstruction scheme and fault diagnosis threshold design method through threshold comparison. This method identifies the gyroscope fault modes that affect the success or failure of the control system through the navigation attitude calculation results in the inertial navigation system measurement coordinate system. While ensuring no omissions and no misjudgments, it reduces the complexity of the discrimination scheme and has good engineering application prospects.

[0022] (2) The method for fault detection and isolation of gyroscope based on navigation attitude angle of this application adopts multiple independent navigation attitude calculation loops to obtain corresponding calculation results. One set of calculation results is compared with the other calculation results to obtain the absolute value of the comparison result. The absolute value of each comparison result is analyzed by a reasonable preset threshold value, which can improve the accuracy of gyroscope fault monitoring and isolation.

[0023] (3) The method for fault detection and isolation of the gyroscope based on navigation attitude angle of the present application adopts multi-channel attitude calculation (wherein, a group of navigation attitude calculation loops is a channel attitude calculation), and has navigation attitude memory function. When a fault occurs, the gyroscope fault isolation and navigation attitude reset can be performed immediately. Attached Figure Description

[0024] To more clearly illustrate the technical solutions in the embodiments of this application or the prior art, the drawings used in the description of the embodiments or the prior art will be briefly introduced below. Obviously, the drawings described below are only some embodiments recorded in this application. For those skilled in the art, other drawings can be obtained based on these drawings.

[0025] Figure 1 This is a schematic diagram of one embodiment of a ten-inertial-meter inertial navigation system;

[0026] Figure 2 The coordinate relationship diagram of the gyroscope of the 10-inertial-meter inertial navigation system;

[0027] Figure 3 The coordinate relationship diagram of the accelerometers in a ten-gauge inertial measurement unit;

[0028] Figure 4 This is a flowchart of one embodiment of a method for fault detection and isolation of a ten-table redundant inertial navigation system gyroscope based on navigation attitude angles. Detailed Implementation

[0029] The technical solutions of the embodiments of the present invention will be clearly and completely described below with reference to the accompanying drawings. Obviously, the described embodiments are only some, not all, of the embodiments of the present invention. Based on the embodiments of the present invention, all other embodiments obtained by those skilled in the art without creative effort are within the scope of protection of the present invention.

[0030] like Figure 1-4 As shown, this application provides a method for fault detection and isolation of a ten-table redundant inertial navigation system gyroscope based on navigation attitude angles, including the following steps:

[0031] S1: After rocket ignition, establish an inertial measurement coordinate system, establish at least six independent navigation attitude calculation loops under the inertial measurement coordinate system, start navigation attitude calculation, and execute S2.

[0032] The ten-meter inertial navigation system (INS) is set on the installation reference plane, and a point O is selected on the azimuth reference plane of the INS as the origin of the INS measurement coordinate system. The X-axis, Y-axis, Z-axis, oblique S-axis, and oblique T-axis are set with the origin as the starting point. The X-axis is perpendicular to the azimuth reference plane, and the Y-axis and Z-axis are both located on the azimuth reference plane, with the Z-axis perpendicular to the Y-axis. The angles between the S-axis and the X-axis, Y-axis, and Z-axis are 127°, 47°, and 66°, respectively; the angles between the T-axis and the X-axis, Y-axis, and Z-axis are 46.3°, 71°, and 49.8°, respectively.

[0033] At least one instrument is installed on each of the X-axis, Y-axis, Z-axis, S-axis, and T-axis. The number of instruments is set according to the actual situation, and this application preferably has one instrument installed on each of the X-axis, Y-axis, Z-axis, S-axis, and T-axis.

[0034] The instruments include at least a gyroscope and an accelerometer, but are not limited to gyroscopes and accelerometers. In this application, gyroscopes and accelerometers are preferred.

[0035] Specifically, the number of independent navigation attitude calculation loops established in the inertial measurement coordinate system is set according to the actual situation. In this application, it is preferred to have six groups, namely: the first group of navigation attitude calculation loops, the second group of navigation attitude calculation loops, the third group of navigation attitude calculation loops, the fourth group of navigation attitude calculation loops, the fifth group of navigation attitude calculation loops and the sixth group of navigation attitude calculation loops.

[0036] Figure 1 This is a schematic diagram of one embodiment of a ten-inertial-meter inertial navigation system (INS). By designing combinations of instruments along different axes within the ten-inertial-meter INS, six independent navigation attitude calculation loops are established. Through analysis and comparison of the six sets of navigation attitude angles, gyroscope fault isolation and navigation attitude reset are performed. As shown in Table 1, the specific combinations of instruments along different axes of the INS measurement coordinate system are as follows:

[0037] The first navigation attitude calculation loop (i.e., when the inertial group axis is the reference in Table 1) includes the following axial instrument combinations: instruments on the X-axis, instruments on the Y-axis, and instruments on the Z-axis.

[0038] The second navigation attitude calculation loop (i.e., when the inertial group axis is redundant 1 in Table 1) includes the following axial instrument combinations: instruments on the X-axis, instruments on the Y-axis, instruments on the Z-axis, and instruments on the S-axis.

[0039] The third navigation attitude calculation loop (i.e., when the inertial group axis is redundant 2 in Table 1) includes the following axial instrument combinations: instruments on the X-axis, instruments on the Y-axis, instruments on the Z-axis, and instruments on the T-axis.

[0040] The fourth navigation attitude calculation loop (i.e., when the inertial group axis is redundant 3 in Table 1) includes the following axial instrument combinations: instruments on the X-axis, instruments on the Y-axis, instruments on the S-axis, and instruments on the T-axis.

[0041] The fifth navigation attitude calculation loop (i.e., when the inertial group axis is redundant 4 in Table 1) includes the following axial instrument combinations: instruments on the X-axis, instruments on the Z-axis, instruments on the S-axis, and instruments on the T-axis.

[0042] The axial instrument combinations involved in the calculation of the sixth navigation attitude calculation loop (i.e., when the inertial group axis is redundant 5 in Table 1) include: instruments on the Y-axis, instruments on the Z-axis, instruments on the S-axis, and instruments on the T-axis.

[0043]

[0044] Table 1: Selection of Axial Combinations for Inertial Measurement Systems

[0045] S2: Obtain output data from the axial instrument clusters participating in the solution in each navigation attitude calculation loop, convert the output data to the inertial measurement system (INS) coordinate system, obtain the INS coordinate system output value, update the navigation attitude based on the INS coordinate system output value, 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 INS 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 then execute S3.

[0046] Furthermore, the traditional quaternion update calculation method is adopted to transform the output navigation attitude angle into the inertial measurement coordinate system to obtain the output value of the inertial measurement coordinate system. The navigation attitude is updated according to the output value of the inertial measurement coordinate system to obtain the updated attitude angle. However, it is not limited to the traditional quaternion update calculation method. This application prefers the traditional quaternion update calculation method.

[0047] Specifically, the updated attitude angles obtained by performing navigation attitude updates using the traditional quaternion update calculation method are: ,in, The pitch angle for the i-th navigation attitude calculation loop; The yaw angle for the i-th navigation attitude calculation loop; The roll angle of the i-th navigation attitude calculation loop. .

[0048] Furthermore, such as Figure 2 As shown, the relationship between the navigation attitude angle output by the gyroscope and the inertial measurement coordinate system is as follows:

[0049] ;

[0050] in, This refers to the navigation attitude angle actually output by the gyroscope in the instrument along the measurement axis when the X-axis is the measurement axis; This refers to the actual navigation attitude angle output by the gyroscope along the measurement axis when the Y-axis is the measurement axis; This refers to the navigation attitude angle actually output by the gyroscope along the measurement axis when the Z-axis is the measurement axis; This refers to the actual navigation attitude angle output by the gyroscope along the measurement axis when the S-axis is the measurement axis. This refers to the navigation attitude angle actually output by the gyroscope along the measurement axis when the T-axis is the measurement axis; This represents the angular rate component of the gyroscope along the X-axis in the inertial measurement coordinate system. This represents the angular rate component of the gyroscope along the Y-axis in the inertial measurement coordinate system. This represents the angular rate component of the gyroscope along the Z-axis in the inertial measurement coordinate system. The installation error angle of the gyroscope on the Y-axis relative to the X-axis of the inertial measurement coordinate system; The installation error angle of the gyroscope on the X-axis relative to the Y-axis of the inertial measurement coordinate system; The installation error angle of the gyroscope relative to the inertial measurement coordinate system on the X-axis; The installation error angle of the Z-axis gyroscope relative to the X-axis of the inertial measurement coordinate system; The installation error angle of the Z-axis gyroscope relative to the Y-axis of the inertial measurement coordinate system; The installation error angle of the gyroscope on the Y-axis relative to the Z-axis of the inertial measurement coordinate system; The angle between the gyroscope on the inclined S-axis and the X-axis of the inertial measurement coordinate system; The angle between the gyroscope on the inclined S-axis and the Y-axis of the inertial measurement coordinate system; The angle between the gyroscope on the inclined S-axis and the Z-axis of the inertial measurement coordinate system; The angle between the gyroscope on the T-axis (which is tilted) and the X-axis of the inertial measurement coordinate system. The angle between the gyroscope on the T-axis (which is tilted) and the Y-axis of the inertial measurement coordinate system. The angle between the gyroscope on the T-axis (which is tilted) and the Z-axis of the inertial measurement coordinate system. Calculate the cosine;

[0051] For matrix ;

[0052] For matrix .

[0053] Specifically, , , , , and The values ​​are the factory calibration parameters of the inertial navigation system.

[0054] Furthermore, based on the relationship between the navigation attitude angles output by the gyroscope and the inertial measurement coordinate system, the transformed navigation attitude angles in the output values ​​of the inertial measurement coordinate system of the i-th navigation attitude solution loop are calculated.

[0055] When i=1, it indicates that the navigation attitude calculation loop is the first group of navigation attitude calculation loops. The expression for the transformed navigation attitude angle in the inertial measurement coordinate system output value of the first group of navigation attitude calculation loops is:

[0056] ;

[0057] ;

[0058] in, When i=1, and the X-axis is the measurement axis, the actual navigation attitude angle output by the gyroscope in the instrument along the measurement axis; When i=1, the actual navigation attitude angle output by the gyroscope along the measurement axis when the Y-axis is the measurement axis; When i=1, and the Z-axis is the measurement axis, the actual navigation attitude angle output by the gyroscope along the measurement axis;

[0059] ;

[0060] For matrix ;

[0061] When i=1, the converted navigation attitude angle; for The transpose of .

[0062] When i=2, it indicates that the navigation attitude calculation loop is the second group of navigation attitude calculation loops. The expression for the transformed navigation attitude angle in the inertial measurement coordinate system output value of the second group of navigation attitude calculation loops is:

[0063] ;

[0064] ;

[0065] in, When i=2, and the X-axis is the measurement axis, the actual navigation attitude angle output by the gyroscope in the instrument along the measurement axis; When i=2, and the Y-axis is the measurement axis, the actual navigation attitude angle output by the gyroscope along the measurement axis; When i=2, and the Z-axis is the measurement axis, the actual navigation attitude angle output by the gyroscope along the measurement axis; When i=2, and the S-axis is the measurement axis, the actual navigation attitude angle output by the gyroscope along the measurement axis;

[0066] ;

[0067] For matrix ;

[0068] When i=2, the converted navigation attitude angle; for The transpose of .

[0069] When i=3, it indicates that the navigation attitude calculation loop is the third group of navigation attitude calculation loops. The expression for the transformed navigation attitude angle in the inertial measurement coordinate system output value of the third group of navigation attitude calculation loops is:

[0070] ;

[0071] ;

[0072] in, When i=3, and the X-axis is the measurement axis, the actual navigation attitude angle output by the gyroscope in the instrument along the measurement axis; When i=3, the actual navigation attitude angle output by the gyroscope along the measurement axis when the Y-axis is the measurement axis; When i=3, the actual navigation attitude angle output by the gyroscope along the measurement axis when the Z-axis is the measurement axis; When i=3, the actual navigation attitude angle output by the gyroscope along the measurement axis when the T-axis is the measurement axis;

[0073] ;

[0074] For matrix ;

[0075] When i=3, the converted navigation attitude angle; for The transpose of .

[0076] When i=4, it indicates that the navigation attitude calculation loop is the fourth group of navigation attitude calculation loops. The expression for the transformed navigation attitude angle in the inertial measurement coordinate system output value of the fourth group of navigation attitude calculation loops is:

[0077] ;

[0078] ;

[0079] in, When i=4, and the X-axis is the measurement axis, the actual navigation attitude angle output by the gyroscope in the instrument along the measurement axis; When i=4, the actual navigation attitude angle output by the gyroscope along the measurement axis when the Y-axis is the measurement axis; When i=4, and the S-axis is the measurement axis, the actual navigation attitude angle output by the gyroscope along the measurement axis; When i=4, the actual navigation attitude angle output by the gyroscope along the measurement axis when the T-axis is the measurement axis;

[0080] ;

[0081] For matrix ;

[0082] When i=4, the converted navigation attitude angle; for The transpose of .

[0083] When i=5, it indicates that the navigation attitude calculation loop is the fifth group of navigation attitude calculation loops. The expression for the transformed navigation attitude angle in the inertial measurement coordinate system output value of the fifth group of navigation attitude calculation loops is:

[0084] ;

[0085] ;

[0086] in, When i=5, and the X-axis is the measurement axis, the actual navigation attitude angle output by the gyroscope in the instrument along the measurement axis; When i=5, the actual navigation attitude angle output by the gyroscope along the measurement axis when the Z-axis is the measurement axis; When i=5, the actual navigation attitude angle output by the gyroscope along the measurement axis when the S-axis is the measurement axis; When i=5, the actual navigation attitude angle output by the gyroscope along the measurement axis when the T-axis is the measurement axis;

[0087] ;

[0088] For matrix ;

[0089] When i=5, the converted navigation attitude angle; for The transpose of .

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

[0091] ;

[0092] ;

[0093] in, When i=6, and the Y-axis is the measurement axis, the actual navigation attitude angle output by the gyroscope in the instrument along the measurement axis; When i=6, the actual navigation attitude angle output by the gyroscope along the measurement axis when the Z-axis is the measurement axis; When i=6, the actual navigation attitude angle output by the gyroscope along the measurement axis when the S-axis is the measurement axis; When i=6, the actual navigation attitude angle output by the gyroscope along the measurement axis when the T-axis is the measurement axis;

[0094] ;

[0095] For matrix ;

[0096] When i=6, the converted navigation attitude angle; for The transpose of .

[0097] Furthermore, such as Figure 3 As shown, the relationship between the acceleration output by the accelerometer and the coordinate system of the inertial measurement system is as follows:

[0098] ;

[0099] in, This refers to the actual acceleration output by the accelerometer along the measurement axis when the X-axis is the measurement axis. This represents the actual acceleration output by the accelerometer along the measurement axis when the Y-axis is the measurement axis. This represents the actual acceleration output by the accelerometer along the measurement axis when the Z-axis is the measurement axis. This represents the actual acceleration output by the accelerometer along the measurement axis when the S-axis is the measurement axis. This represents the actual acceleration output by the accelerometer along the measurement axis when the axis is the measurement axis. The apparent acceleration component of the X-axis in the inertial measurement coordinate system; The apparent acceleration component along the Y-axis of the inertial measurement coordinate system; The apparent acceleration component of the Z-axis in the inertial measurement coordinate system; The installation error angle of the accelerometer on the X-axis relative to the Y-axis of the inertial measurement coordinate system; The installation error angle of the accelerometer on the X-axis relative to the Z-axis of the inertial measurement coordinate system; The installation error angle of the accelerometer on the Y-axis relative to the Z-axis of the inertial measurement coordinate system; The angle between the accelerometer on the inclined S-axis and the X-axis of the inertial measurement coordinate system; The angle between the accelerometer on the inclined S-axis and the Y-axis of the inertial measurement coordinate system; The angle between the inclined S-axis accelerometer and the Z-axis of the inertial measurement coordinate system; The angle between the accelerometer on the T-axis (which is tilted) and the X-axis of the inertial measurement coordinate system; The angle between the accelerometer on the T-axis (which is tilted) and the Y-axis of the inertial measurement coordinate system; The angle between the accelerometer on the T-axis (which is tilted) and the Z-axis of the inertial measurement coordinate system;

[0100] For matrix ;

[0101] for .

[0102] Specifically, , and The parameters are the factory calibration parameters for the inertial navigation system. , , , , and The parameters are the factory calibration parameters for the inertial navigation system.

[0103] Furthermore, based on the relationship between the acceleration output by the accelerometer and the inertial measurement coordinate system, the transformed acceleration in the output value of the inertial measurement coordinate system of the i-th navigation attitude calculation loop is calculated.

[0104] When i=1, it indicates that the navigation attitude calculation loop is the first group of navigation attitude calculation loops. The expression for the transformed acceleration in the inertial measurement coordinate system output value of the first group of navigation attitude calculation loops is:

[0105] ;

[0106] ;

[0107] in, When i=1, the actual acceleration output by the accelerometer in the instrument along the measuring axis when the X-axis is the measuring axis; When i=1, the actual acceleration output by the accelerometer along the measurement axis when the Y-axis is the measurement axis; When i=1, the actual acceleration output by the accelerometer along the measurement axis when the Z-axis is the measurement axis;

[0108] ;

[0109] For matrix ;

[0110] The converted acceleration when i=1; for The transpose of .

[0111] When i=2, it indicates that the navigation attitude calculation loop is the second group of navigation attitude calculation loops. The expression for the transformed acceleration in the inertial measurement coordinate system output value of the second group of navigation attitude calculation loops is:

[0112] ;

[0113] ;

[0114] in, When i=2, the actual acceleration output by the accelerometer in the instrument along the measuring axis when the X-axis is the measuring axis; When i=2, the actual acceleration output by the accelerometer along the measurement axis when the Y-axis is the measurement axis; When i=2, the actual acceleration output by the accelerometer along the measurement axis when the Z-axis is the measurement axis; When i=2, the actual acceleration output by the accelerometer along the measurement axis when the S-axis is the measurement axis;

[0115] ;

[0116] For matrix ;

[0117] The converted acceleration when i=2; for The transpose of .

[0118] When i=3, it indicates that the navigation attitude calculation loop is the third group of navigation attitude calculation loops. The expression for the transformed acceleration in the inertial measurement coordinate system output value of the third group of navigation attitude calculation loops is:

[0119] ;

[0120] ;

[0121] in, When i=3, the actual acceleration output by the accelerometer in the instrument along the measuring axis when the X-axis is the measuring axis; When i=3, the actual acceleration output by the accelerometer along the measurement axis when the Y-axis is the measurement axis; When i=3, the actual acceleration output by the accelerometer along the measurement axis when the Z-axis is the measurement axis; When i=3, the actual acceleration output by the accelerometer along the measurement axis when the T-axis is the measurement axis;

[0122] ;

[0123] For matrix ;

[0124] The converted acceleration when i=3; for The transpose of .

[0125] When i=4, it indicates that the navigation attitude calculation loop is the fourth group of navigation attitude calculation loops. The expression for the transformed acceleration in the inertial measurement coordinate system output value of the fourth group of navigation attitude calculation loops is:

[0126] ;

[0127] ;

[0128] in, When i=4, the actual acceleration output by the accelerometer in the instrument along the measuring axis when the X-axis is the measuring axis; When i=4, the actual acceleration output by the accelerometer along the measurement axis when the Y-axis is the measurement axis; When i=4, the actual acceleration output by the accelerometer along the measurement axis when the S-axis is the measurement axis; When i=4, the actual acceleration output by the accelerometer along the measurement axis when the T-axis is the measurement axis;

[0129] ;

[0130] For matrix ;

[0131] The converted acceleration when i=4; for The transpose of .

[0132] When i=5, it indicates that the navigation attitude calculation loop is the fifth group of navigation attitude calculation loops. The expression for the transformed acceleration in the inertial measurement coordinate system output value of the fifth group of navigation attitude calculation loops is:

[0133] ;

[0134] ;

[0135] in, When i=5, the actual acceleration output by the accelerometer in the instrument along the measuring axis when the X-axis is the measuring axis; When i=5, the actual acceleration output by the accelerometer along the measurement axis when the Z-axis is the measurement axis; When i=5, the actual acceleration output by the accelerometer along the measurement axis when the S-axis is the measurement axis; When i=5, the actual acceleration output by the accelerometer along the measurement axis when the T-axis is the measurement axis;

[0136] ;

[0137] For matrix ;

[0138] The converted acceleration when i=5; for The transpose of .

[0139] When i=6, it indicates that the navigation attitude calculation loop is the sixth group of navigation attitude calculation loops. The expression for the transformed acceleration in the inertial measurement coordinate system output value of the sixth group of navigation attitude calculation loops is:

[0140] ;

[0141] ;

[0142] in, When i=6, the actual acceleration output by the accelerometer in the instrument along the measuring axis when the Y-axis is the measuring axis; When i=6, the actual acceleration output by the accelerometer along the measurement axis when the Z-axis is the measurement axis; When i=6, the actual acceleration output by the accelerometer along the measurement axis when the S-axis is the measurement axis; When i=6, the actual acceleration output by the accelerometer along the measurement axis when the T-axis is the measurement axis;

[0143] ;

[0144] For matrix ;

[0145] The converted acceleration when i=6; for The transpose of .

[0146] S3: Compare the solution result of one of the at least six independent navigation attitude solution loops with the remaining navigation attitude solution loops in the at least six independent navigation attitude solution loops, calculate the comparison result, obtain the absolute value of the comparison result, and execute S4.

[0147] Furthermore, when there are six navigation attitude calculation loops, the calculation results of the first navigation attitude calculation loop are compared with the calculation results of the second, third, fourth, fifth and sixth navigation attitude calculation loops, respectively. The comparison results are calculated, and the absolute value of the comparison results is obtained. Then, S4 is executed.

[0148] Furthermore, when there are six navigation attitude calculation loops, the calculation result of the first navigation attitude calculation loop is: The solution results for the remaining navigation attitude calculation loops are: , .

[0149] The formula for calculating the comparison results is:

[0150] ;

[0151] in, This is a comparison result between the solution result of the j-th navigation attitude solution loop and the settlement result of the first navigation attitude solution loop; The pitch angle for the first navigation attitude calculation loop; The yaw angle for the first set of navigation attitude calculation loops; The roll angle for the first set of navigation attitude calculation loops; The pitch angle for the j-th navigation attitude calculation loop; The yaw angle for the j-th navigation attitude calculation loop; The roll angle of the j-th navigation attitude calculation loop.

[0152] Furthermore, the expression for the absolute value of the comparison result is:

[0153] ,in .

[0154] S4: Determine whether the absolute value of each comparison result is less than the preset threshold. If the absolute value of all comparison results does not exceed the preset threshold, then use the calculation result of a set of navigation attitude calculation loops as the attitude angle used by the control system and execute S6. If the absolute value of one or more comparison results exceeds or equals the preset threshold, then execute S5.

[0155] Furthermore, the specific value of the preset threshold is set according to the actual situation. Preferably, this application specifies: a preset threshold value... ,in, To fix the angular deviation to zero bias, t represents the slope of the attitude angle deviation modulus diverging with navigation time; t represents the time length for establishing the inertial navigation system.

[0156] Specifically, to prevent misjudgment or omission, The selection of parameters needs to be designed in close conjunction with the angular velocity characteristics of the entire flight trajectory, and the parameters are obtained by simulation calculation and other methods.

[0157] like None of them exceeded (i.e., were less than or equal to) the preset threshold value. If the outputs of all combined instruments in the ten-meter inertial measurement unit are normal, then the following approach is adopted. Attitude angles are used as a control system.

[0158] like There are one or more Exceeding (i.e., greater than) the preset threshold value If so, the existing methods are used to identify gyroscope faults. At the same time, the gyroscope's angular velocity information, maximum value and consistency judgment are combined to screen out the axis system where the faulty gyroscope is located and isolate the gyroscope.

[0159] S5: For gyroscopes in the axial instrument clusters participating in the navigation attitude calculation loop where the absolute value of the comparison result exceeds or equals the preset threshold, perform fault analysis. After identifying the faulty gyroscope, isolate the faulty gyroscope. Select a set of navigation attitude calculation loops corresponding to the unisolated gyroscopes as a new reference. Use the calculation result of the navigation attitude calculation loop corresponding to the new reference as the attitude angle used by the control system. Execute S6.

[0160] Specifically, based on the direction of the faulty gyroscope, redundant groups are eliminated. If the fault is caused by three orthogonal gyroscopes (i.e., gyroscopes on the X-axis, Y-axis, and Z-axis), then one of the redundancy methods (i.e., the third group of navigation attitude calculation loop, the fourth group of navigation attitude calculation loop, and the fifth group of navigation attitude calculation loop) that does not contain the faulty gyroscope is selected as the reference for subsequent navigation attitude calculation. At the same time, the attitude angle calculated by this group is used as the attitude angle for the control system.

[0161] S6: When the navigation calculation cycle meets the preset conditions of the navigation calculation cycle, execute S2~S5 until the flight mission ends.

[0162] Furthermore, the preset conditions for the navigation calculation cycle are set according to the actual situation. In this application, the preferred condition is that the navigation calculation cycle is Pms, where P is a natural number. In this application, the preferred condition is P=5, that is, the navigation calculation cycle is 5ms.

[0163] Furthermore, the specific value of the sampling period of the inertial navigation system (INS) is set according to the actual situation. In this application, the preferred value is 5ms.

[0164] Furthermore, the specific value of the error compensation period of the ten-meter inertial navigation system is set according to the actual situation. In this application, the preferred error compensation period is 5ms.

[0165] The beneficial effects achieved by this application are as follows:

[0166] (1) Based on the ten-table inertial navigation system, this application proposes a fault diagnosis process and algorithm, information reconstruction scheme and fault diagnosis threshold design method through threshold comparison. This method identifies the gyroscope fault modes that affect the success or failure of the control system through the navigation attitude calculation results in the inertial navigation system measurement coordinate system. While ensuring no omissions and no misjudgments, it reduces the complexity of the discrimination scheme and has good engineering application prospects.

[0167] (2) The method for fault detection and isolation of gyroscope based on navigation attitude angle of this application adopts multiple independent navigation attitude calculation loops to obtain corresponding calculation results. One set of calculation results is compared with the other calculation results to obtain the absolute value of the comparison result. The absolute value of each comparison result is analyzed by a reasonable preset threshold value, which can improve the accuracy of gyroscope fault monitoring and isolation.

[0168] (3) The method for fault detection and isolation of the gyroscope based on navigation attitude angle of the present application adopts multi-channel attitude calculation (wherein, a group of navigation attitude calculation loops is a channel attitude calculation), and has navigation attitude memory function. When a fault occurs, the gyroscope fault isolation and navigation attitude reset can be performed immediately.

[0169] Although preferred embodiments of this application have been described, those skilled in the art, upon learning the basic inventive concept, can make other changes and modifications to these embodiments. Therefore, the scope of protection of this application is intended to be interpreted as including the preferred embodiments as well as all changes and modifications falling within the scope of this application. Obviously, those skilled in the art can make various alterations and variations to this application without departing from the spirit and scope of this application. Thus, if these modifications and variations of this application fall within the scope of protection of this application and its equivalents, this application also intends to include these modifications and variations.

Claims

1. A navigation attitude angle based gyro fault detection and isolation method for a strapdown inertial package, characterized in that, Includes the following steps: S1: After rocket ignition, establish the inertial measurement coordinate system, establish at least six independent navigation attitude calculation loops under the inertial measurement coordinate system, start navigation attitude calculation, and execute S2; S2: Obtain output data from the axial instrument clusters participating in the solution in each navigation attitude calculation loop, convert the output data to the inertial measurement system (INS) coordinate system, obtain the INS coordinate system output value, update the navigation attitude based on the INS coordinate system output value, 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 INS 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, then execute S3; S3: Compare the solution result of one of the at least six independent navigation attitude solution loops with the remaining navigation attitude solution loops in the at least six independent navigation attitude solution loops, calculate the comparison result, obtain the absolute value of the comparison result, and execute S4. S4: Determine whether the absolute value of each comparison result is less than the preset threshold. If the absolute value of all comparison results does not exceed the preset threshold, then use the calculation result of a set of navigation attitude calculation loops as the attitude angle used by the control system and execute S6. If the absolute value of one or more comparison results exceeds or equals the preset threshold, then execute S5. S5: For gyroscopes in the axial instrument clusters participating in the navigation attitude calculation loop where the absolute value of the comparison result exceeds or equals the preset threshold value, perform fault analysis. After identifying the faulty gyroscope, isolate the faulty gyroscope. Select a set of navigation attitude calculation loops corresponding to the unisolated gyroscopes as a new reference. Use the calculation result of the navigation attitude calculation loop corresponding to the new reference as the attitude angle used by the control system. Execute S6. S6: When the navigation calculation cycle meets the preset conditions of the navigation calculation cycle, execute S2~S5 until the flight mission ends; Wherein, the preset threshold value Wherein, is the angle deviation fixed zero deviation, is the slope of the attitude angle deviation module divergence with navigation time; t is the time length of establishing inertial navigation. If are all less than or equal to a preset threshold value It is considered that each combination instrument output of the inertial measurement unit is normal, and the solving result of the first navigation attitude solving loop is used as the control system using attitude angle; like There are one or more Greater than the preset threshold value If so, the existing methods are used to identify gyroscope faults. At the same time, the gyroscope's angular velocity information, maximum value and consistency judgment are combined to screen out the axis system where the faulty gyroscope is located and isolate the gyroscope. The absolute value of the comparison result. .

2. The method for fault detection and isolation of a ten-table redundant inertial navigation system gyroscope based on navigation attitude angles according to claim 1, characterized in that, The navigation attitude calculation loop consists of six groups: the first group of navigation attitude calculation loops, the second group of navigation attitude calculation loops, the third group of navigation attitude calculation loops, the fourth group of navigation attitude calculation loops, the fifth group of navigation attitude calculation loops, and the sixth group of navigation attitude calculation loops.

3. The navigation attitude angle based ten-table redundant inertial package gyro fault detection and isolation method according to claim 2, characterized in that, The first set of navigation attitude calculation loops includes the following axial instrument combinations: instruments on the X-axis of the inertial measurement coordinate system, instruments on the Y-axis of the inertial measurement coordinate system, and instruments on the Z-axis of the inertial measurement coordinate system. The axial instrument combination involved in the second navigation attitude calculation loop includes: instruments on the X-axis of the inertial measurement coordinate system, instruments on the Y-axis of the inertial measurement coordinate system, instruments on the Z-axis of the inertial measurement coordinate system, and instruments on the S-axis of the inertial measurement coordinate system. The axial instrument combination involved in the third navigation attitude calculation loop includes: instruments on the X-axis of the inertial measurement coordinate system, instruments on the Y-axis of the inertial measurement coordinate system, instruments on the Z-axis of the inertial measurement coordinate system, and instruments on the T-axis of the inertial measurement coordinate system. The axial instrument combination involved in the fourth navigation attitude calculation loop includes: instruments on the X-axis of the inertial measurement coordinate system, instruments on the Y-axis of the inertial measurement coordinate system, instruments on the S-axis of the inertial measurement coordinate system, and instruments on the T-axis of the inertial measurement coordinate system. The fifth navigation attitude calculation loop includes the following axial instrument combinations: instruments on the X-axis, Z-axis, S-axis, and T-axis of the inertial measurement system. The axial instrument group involved in the calculation in the sixth navigation attitude calculation loop includes: instruments on the Y-axis of the inertial measurement coordinate system, instruments on the Z-axis of the inertial measurement coordinate system, instruments on the S-axis of the inertial measurement coordinate system, and instruments on the T-axis of the inertial measurement coordinate system.

4. The navigation attitude angle based ten-table redundant inertial package gyro fault detection and isolation method according to claim 3, characterized in that, The instruments should include at least a gyroscope and an accelerometer.

5. The navigation attitude angle based ten-table redundant inertial package gyro fault detection and isolation method according to claim 4, characterized in that, The navigation attitude is updated using the traditional quaternion update calculation method, and the updated attitude angles are: ,in, The pitch angle for the i-th navigation attitude calculation loop; The yaw angle for the i-th navigation attitude calculation loop; The roll angle of the i-th navigation attitude calculation loop. .

6. The method for fault detection and isolation of a ten-table redundant inertial navigation system gyroscope based on navigation attitude angles according to claim 5, characterized in that, The calculation results of the first group of navigation attitude calculation loops are compared with the calculation results of the second group of navigation attitude calculation loops, the third group of navigation attitude calculation loops, the fourth group of navigation attitude calculation loops, the fifth group of navigation attitude calculation loops, and the sixth group of navigation attitude calculation loops. The comparison results are calculated and the absolute values ​​of the comparison results are obtained.

7. The navigation attitude angle based gyro fault detection and isolation method for a strapdown inertial package of claim 6, wherein, When there are six navigation attitude calculation loops, the calculation result of the first navigation attitude calculation loop is: The solution results for the remaining navigation attitude calculation loops are: , ; The expression for the absolute value of the comparison result is: ； in, This is a comparison result between the solution result of the j-th navigation attitude solution loop and the settlement result of the first navigation attitude solution loop; The pitch angle for the first navigation attitude calculation loop; The yaw angle for the first set of navigation attitude calculation loops; The roll angle for the first set of navigation attitude calculation loops; The pitch angle for the j-th navigation attitude calculation loop; The yaw angle for the j-th navigation attitude calculation loop; The roll angle of the j-th navigation attitude calculation loop.

8. The method for fault detection and isolation of a ten-table redundant inertial navigation system gyroscope based on navigation attitude angles according to claim 1, characterized in that, The navigation calculation cycle is 5ms.

9. The navigation attitude angle based decoupled ring laser gyro fault detection and isolation method of claim 1, wherein, The sampling period of the 10-meter inertial navigation system is 5ms; the error compensation period of the 10-meter inertial navigation system is 5ms.

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