A method and system for determining redundant strategy of ten-meter inertial group

Abstract

The present application relates to the field of aerospace technology, and in particular to a method and system for distinguishing the redundant strategy of a ten-meter inertial group, including: establishing seven groups of inertial element combinations with different axes through redundant combination, and converting the output of each group to the inertial group measurement coordinate system; performing the first combined navigation attitude solution according to the output of the first group after conversion and the output of the GNSS receiver; performing the pure inertial navigation attitude solution according to the output of the second to seventh groups after conversion; performing the second combined navigation attitude solution according to the output of the first group after conversion and the output of the GNSS receiver; distinguishing the consistency of the solution results of the second to seventh groups with the second solution results of the first group; judging whether the judgment results exceed the design threshold value; if none of them exceed the design threshold value, using the first solution result of the first group as the navigation result used for guidance and attitude control. The present application can improve the accuracy of fault detection for the ten-meter redundant inertial group.

Description

Technical Field

[0001] The present application relates to the field of aerospace technology, and in particular to a method and system for determining a redundancy strategy of a ten-meter inertial group. Background Art

[0002] With the rapid increase in the number of 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 ways: system-level redundancy and device-level redundancy. System-level redundancy includes dual inertial group master-slave redundancy, triple inertial group redundancy, etc.; device-level redundancy includes single inertial group multi-meter redundancy, etc. Among them, the single set of ten-meter configuration inertial group has more advantages than system-level redundancy in terms of volume, mass, cost and simplicity, and gradually becomes the main inertial device of subsequent launch vehicles.

[0003] Due to the high dynamic requirements of launch vehicles, the fault diagnosis system can quickly and accurately process fault data and complete fault information location and isolation in a timely manner, and ensure the smooth completion of the flight mission through navigation information reconstruction. Among them, inertial devices are key equipment for launch vehicle navigation. At present, a variety of methods have been developed for fault diagnosis of inertial devices, such as threshold comparison method, generalized likelihood ratio method, mean test method, singular value decomposition method and neural network method.

[0004] The traditional method of detecting and isolating gyro faults in the ten-meter redundant inertial group is mainly to diagnose in the instrument pulse conversion link, and there is no good analysis method for the failure modes such as slow change and slight change of the inertial instrument zero position error that affect the long-term navigation accuracy. It is easily affected by external environmental factors, resulting in false alarm or missed alarm of gyro pulse fault.

[0005] Therefore, how to improve the accuracy of fault detection for a redundant set of ten meters is a technical problem that those skilled in the art need to solve urgently. Summary of the invention

[0006] The present application provides a method and system for determining a redundant strategy of a ten-meter inertial group, so as to improve the accuracy of fault detection of the ten-meter redundant inertial group.

[0007] In order to solve the above technical problems, this application provides the following technical solutions:

[0008] A redundancy strategy identification method for a ten-meter strapdown inertial group comprises the following steps: establishing seven groups of inertial element combinations in different axes by redundantly combining the ten-meter strapdown inertial group, and converting the output of each group of inertial element combinations to an inertial group measurement coordinate system; at the beginning of an integrated navigation cycle, performing the first integrated navigation attitude solution in the integrated navigation cycle according to the output of the first group of inertial element combinations converted to the inertial group measurement coordinate system and the output of a GNSS receiver; performing pure inertial navigation attitude solution according to the output of the second to seventh groups of inertial element combinations converted to the inertial group measurement coordinate system; at the end of the integrated navigation cycle, performing the first integrated navigation attitude solution according to the output of the first group of inertial element combinations converted to the inertial group measurement coordinate system The output in the measurement coordinate system and the output of the GNSS receiver are used to perform a second integrated navigation attitude solution within the integrated navigation cycle; the pure inertial navigation attitude solution results of the second to seventh inertial element combinations are subjected to consistency judgment with the second integrated navigation attitude solution result of the first inertial element combination; it is judged whether the six groups of judgment results exceed the corresponding design threshold values; if all the six groups of judgment results do not exceed the corresponding design threshold values, the first integrated navigation attitude solution result within the integrated navigation cycle is used as the navigation result used for guidance and attitude control; at the beginning of the next integrated navigation cycle, the navigation attitude solution and consistency judgment are repeated until the flight mission is completed.

[0009] The method for determining the redundant strategy of the ten-meter strapdown inertial group as described above, preferably, also includes the following steps: if any of the six groups of determination results exceeds the designed threshold value, the axis of the ten-meter strapdown inertial group failure is screened out and isolated; the inertial element combination containing the failed axial direction is eliminated; at the beginning of the next combined navigation cycle, the navigation attitude solution and consistency determination are repeated until the flight mission is completed.

[0010] In the above-mentioned method for determining the redundant strategy of the ten-meter inertial group, preferably, when the output of each inertial element combination is converted to the inertial group measurement coordinate system: the H of the gyroscope of each inertial element combination is calculated. g When H g Remove the rows corresponding to the output of the gyroscope in the inertial element combination that does not participate in the inertial navigation solution from the expression of , and you will get the H of the gyroscope in the inertial element combination. g ; Calculate the H of the accelerometer for each set of inertial elements a When H a Remove the rows corresponding to the output of the accelerometers in the inertial element combination that are not involved in the inertial navigation solution from the expression of , and you will get the H of the accelerometers in the inertial element combination. a ; Among them, H g is the calibration and error comprehensive parameter of the gyroscope combined with inertial components, H a Calibration and comprehensive error parameters of the accelerometer combined with inertial elements.

[0011] In the above-mentioned method for determining the redundant strategy of the ten-meter inertial element group, preferably, after calculating the H of each group of inertial element combinations, g and H a After that, through the formula Calculate the angular rate ω of each set of inertial components in the inertial measurement coordinate system M and apparent acceleration A M ;ω b A is the actual total output of the gyroscope for each set of inertial elements; b is the actual total output of the accelerometer for each set of inertial elements; the angular rate ω M and apparent acceleration A M As output in the inertial measurement coordinate system.

[0012] In the above-mentioned ten-meter inertial group redundant strategy determination method, it is preferred that the angular velocity and apparent acceleration are converted into position, velocity and attitude angle through inertial navigation solution.

[0013] In the above-mentioned method for determining the redundancy strategy of the ten-meter strapdown inertial group, preferably, PPS pulse per second is used to perform time synchronization between the data of the ten-meter strapdown inertial group and the data of the GNSS receiver.

[0014] As described above, in the method for determining the redundancy strategy of the ten-meter inertial group, preferably, the design threshold value includes: a speed design threshold value and a position design threshold value; wherein the speed design threshold value is δ v =δ v0 +k v ·t, the position design threshold is δ P =δ P 0+k P ·t;δ v0 is the initial threshold of speed deviation, δ P0 is the initial threshold of position deviation, k v is the slope of the velocity deviation diverging with navigation time, k P is the slope of the position deviation diverging with navigation time, δ v0 , δ P0 , k v , k P These are all the values ​​obtained by simulation calculation method, and t is the time length for establishing inertial navigation.

[0015] A ten-meter inertial group redundant strategy determination system comprises: a GNSS receiver, a GNSS antenna, a flight control computer and a ten-meter strapdown inertial group; wherein the ten-meter strapdown inertial group is arranged on a launch vehicle, and the ten-meter strapdown inertial group is connected to the flight control computer; the GNSS antenna is connected to the GNSS receiver, and the GNSS receiver is connected to the flight control computer; the ten-meter inertial group redundant strategy determination system executes any one of the ten-meter inertial group redundant strategy determination methods described above.

[0016] As described above, in the ten-meter strapdown inertial group redundant strategy identification system, preferably, the ten-meter strapdown inertial group and the flight control computer, the GNSS antenna and the GNSS receiver, and the GNSS receiver and the flight control computer are all connected via arrow cables.

[0017] Compared with the above background technology, this application adopts independent navigation attitude solution loops to judge the consistency of speed and position, and compares the speed and position information of the GNSS receiver. Through reasonable thresholds, it can realize the fault detection and isolation of the long-term navigation accuracy reduction caused by the large zero bias error of the inertial instrument, and improve the accuracy of inertial group fault monitoring and isolation. At the same time, multi-channel navigation solution is adopted, which has the function of navigation information memory. When a fault occurs, the inertial group fault isolation and navigation information reset can be carried out in the first time, improving the error correction capability of the navigation system. BRIEF DESCRIPTION OF THE DRAWINGS

[0018] In order to more clearly illustrate the embodiments of the present invention 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 invention. For ordinary technicians in this field, other drawings can also be obtained based on these drawings.

[0019] Figure 1 Schematic diagram of a redundancy strategy discrimination system for a ten-meter inertia group provided in an embodiment of the present application;

[0020] Figure 2 is a schematic diagram of a ten-meter strapdown inertial unit provided in an embodiment of the present application;

[0021] Figure 3 is a flow chart of a method for determining a redundancy strategy of a ten-table inertia group provided in an embodiment of the present application;

[0022] Figure 4 is a coordinate relationship diagram of a ten-meter inertial group gyroscope provided in an embodiment of the present application;

[0023] Figure 5 It is a coordinate relationship diagram of the ten-meter inertial group accelerometer provided in an embodiment of the present application. DETAILED DESCRIPTION

[0024] Embodiments of the present invention are described in detail below, examples of which are shown in the accompanying drawings, wherein the same or similar reference numerals throughout represent the same or similar elements or elements having the same or similar functions. The embodiments described below with reference to the accompanying drawings are exemplary and are only used to explain the present invention, and cannot be interpreted as limiting the present invention.

[0025] Embodiment 1

[0026] like Figure 1 The present application provides a ten-meter strapdown inertial group redundancy strategy determination system, comprising: a GNSS receiver 110, a GNSS antenna 120, a flight control computer 130 and a ten-meter strapdown inertial group 140; wherein, Figure 2 As shown, a ten-meter strapdown inertial group 140 is arranged on a launch vehicle, and is used to measure real-time angular velocity information and acceleration information in the directions of the X-axis, Y-axis, Z-axis, T-axis and S-axis, and the update frequency of such information is generally 200 Hz; and the ten-meter strapdown inertial group 140 is connected to a flight control computer 130 to send the measured angular velocity information and acceleration information to the flight control computer 130; a GNSS antenna 120 is connected to a GNSS receiver 110, and the GNSS receiver 110 measures real-time velocity information and position information in the directions of the X-axis, Y-axis and Z-axis of the position of the GNSS antenna 120, and the update frequency of such information is generally 10 Hz; and the GNSS receiver 110 is connected to the flight control computer 130 to send the measured velocity information and position information to the flight control computer 130; the flight control computer 130 processes the angular velocity information, acceleration information, velocity information and position information to obtain velocity information and position information; and sends the obtained velocity information and position information to a rocket control mechanism, so that the rocket control mechanism performs control according to the velocity information and position.

[0027] On the launch vehicle, the strapdown inertial unit 140 and the flight control computer 130, the GNSS antenna 120 and the GNSS receiver 110, the GNSS receiver 110 and the flight control computer 130, and the flight control computer 140 and the rocket control mechanism are all connected via onboard cables.

[0028] The strapdown inertial measurement unit is an inertial measurement device, mainly composed of inertial elements (such as gyroscopes and accelerometers). Gyroscopes are used to measure the angular velocity of objects, and accelerometers are used to measure the acceleration of objects. These inertial elements are directly fixed to the carrier, and the carrier's attitude, velocity, position and other parameters are determined by measuring the carrier's angular velocity and acceleration information, and then performing corresponding mathematical operations. The ten-meter strapdown inertial unit is configured with multiple identical inertial elements to increase the redundancy of the ten-meter strapdown inertial unit, for example: it may be multiple single-axis or multi-axis gyroscopes and accelerometers.

[0029] In the present application, ten strapdown inertial groups are combined redundantly to establish seven groups of inertial element combinations in different axes. Each group of inertial element combination forms an independent navigation attitude solution loop. The seven groups of inertial element combination form seven independent navigation attitude solution loops. By solving the navigation attitude in the seven groups of navigation attitude solution loops, gyro fault isolation and navigation attitude reset are performed.

[0030] The outputs of each group of inertial element combinations that participate in the navigation attitude solution are different, that is, some of the outputs of each group of inertial element combinations do not participate in the navigation attitude solution, as shown in Table 1.

[0031] Table 1: Axial combination selection of ten-meter strapdown inertial unit

[0032]

[0033] The navigation mode of the first group of inertial element combination is combined navigation, the outputs of the group of inertial elements in the directions of X-axis, Y-axis and Z-axis participate in the calculation, the outputs in the directions of T-axis and S-axis do not participate in the solution, and the group of inertial element combination is used as a reference combination; the navigation mode of the second group of inertial element combination is pure inertial navigation, the outputs of the group of inertial elements in the directions of X-axis, Y-axis and Z-axis participate in the calculation, the outputs in the directions of T-axis and S-axis do not participate in the solution, and the group of inertial element combination is used as a redundant combination; the navigation mode of the third group of inertial element combination is pure inertial navigation, the outputs of the group of inertial elements in the directions of X-axis, Y-axis, Z-axis and S-axis participate in the calculation, the output in the direction of T-axis does not participate in the solution, and the group of inertial element combination is used as a redundant combination; the navigation mode of the fourth group of inertial element combination is pure inertial navigation, the outputs of the group of inertial elements in the directions of X-axis, Y-axis, Z-axis and T-axis participate in the calculation, the outputs in the directions of The upward output participates in the calculation, the output in the S-axis direction does not participate in the solution, and this group of inertial component combinations is used as a redundant combination; the navigation mode of the fifth group of inertial component combinations is pure inertial navigation, the outputs of this group of inertial components in the X-axis, Y-axis, T-axis and S-axis directions participate in the calculation, the output in the Z-axis direction does not participate in the solution, and this group of inertial component combinations is used as a redundant combination; the navigation mode of the sixth group of inertial component combinations is pure inertial navigation, the outputs of this group of inertial components in the X-axis, Z-axis, T-axis and S-axis directions participate in the calculation, the output in the Y-axis direction does not participate in the solution, and this group of inertial component combinations is used as a redundant combination; the navigation mode of the seventh group of inertial component combinations is pure inertial navigation, the outputs of this group of inertial components in the Y-axis, Z-axis, T-axis and S-axis directions participate in the calculation, the output in the X-axis direction does not participate in the solution, and this group of inertial component combinations is used as a redundant combination.

[0034] Embodiment 2

[0035] like Figure 3 As shown, the present application provides a method for determining a redundancy strategy of a ten-table inertia group, comprising the following steps:

[0036] Step S305, the ten-meter strapdown inertial group is combined through redundancy to establish seven groups of inertial element combinations in different axes, and the output of each group of inertial element combinations is converted to the inertial group measurement coordinate system;

[0037] After the rocket is ignited, the launch inertial system navigation is established, and then the ten-meter strapdown inertial group of the carrier rocket is redundantly combined. According to Table 1, 7 independent inertial element combinations with different axes are established. Each group of inertial element combinations forms an independent inertial navigation solution loop, thereby independently performing the inertial navigation solution process. The cycle of sampling, error compensation and inertial navigation solution for each group of inertial element combinations is 5ms.

[0038] According to Table 1, the outputs of the seven groups of inertial elements are converted to the inertial group measurement coordinate system for inertial navigation solution. The inertial navigation solution adopts the traditional strapdown inertial navigation system update algorithm. The strapdown inertial navigation system (SINS) update algorithm is a method for processing the data of inertial elements (accelerometers and gyroscopes) to update the attitude, velocity and position of the carrier in real time.

[0039] Among them, the relationship between the output of the gyroscope and the output value of the inertial group measurement coordinate system is:

[0040]

[0041] Among them, Figure 4 As shown, ω x ,ω y ,ω z ,ω S ,ω T are the actual outputs of the gyroscope along the measurement axes (X, Y, Z, S and T); xM ,ω yM ,ω zM are the angular velocity components of the three axes of the inertial measurement coordinate system; E yx , E zx , E xy , E zy , E xz , E yz are the installation error angles of the gyroscope axes relative to the inertial measurement coordinate system; θ gxS ,θ gyS ,θ gzS ,θ gxT ,θ gyT ,θ gzT The angles between the two tilted gyroscopes and the three axes of the IMU measurement coordinate system are used to calibrate the parameters of the IMU before leaving the factory.

[0042] Will Simplified as H g , Hg is the calibration and error comprehensive parameters of a gyroscope composed of a group of inertial elements; Simplified as ω M ,ω M It is the angular velocity of a group of inertial elements in the inertial group measurement coordinate system.

[0043] The relationship between the output of the accelerometer and the output value of the inertial group measurement coordinate system is:

[0044]

[0045] Among them, Figure 5 As shown, A x , A y , A z , A S , A T are the actual outputs of the accelerometer along the measurement axes (X, Y, Z, S and T); xM , A yM , A zM are the apparent acceleration components of the three axes of the inertial measurement coordinate system; K xy , K xz , K yz are the installation error angles of the axes of the three orthogonal accelerometers relative to the measurement coordinate system of the inertial group, and are the factory calibration parameters of the inertial group; θ AxS ,θ AyS ,θ AzS ,θ AxT ,θ AyT ,θ AzT The angles between the two inclined accelerometers and the three axes of the IMU measurement coordinate system are the factory calibration parameters of the IMU.

[0046] Will Simplified as H a , H a The calibration and error comprehensive parameters of the accelerometer combined with a group of inertial elements; Simplified as A M , A M It is the apparent acceleration of a group of inertial elements in the inertial group measurement coordinate system.

[0047] Since part of the output of each inertial element combination in the present application does not participate in the inertial navigation solution, when converting the output of each inertial element combination to the inertial group measurement coordinate system (calculating the output values ​​of the gyroscope and accelerometer under different inertial element combinations in the inertial group measurement system according to Formula 1 and Formula 2), the output of the gyroscope that does not participate in the inertial navigation solution and the output of the accelerometer that does not participate in the inertial navigation solution need to be removed.

[0048] Specifically, according to the inertial element combination in Table 1, the H of the gyroscope of each inertial element combination is calculated. g When H g In the expression of , remove the rows corresponding to the output of the gyroscope in the inertial element combination that does not participate in the inertial navigation solution; calculate the H of the accelerometer of each inertial element combination a When H a In the expression of , the rows corresponding to the outputs of the accelerometers in the group of inertial elements that do not participate in the inertial navigation solution can be removed.

[0049] For example, the model of redundancy 1 in the second row of Table 1 is:

[0050]

[0051] Will Simplified expression: is the actual total output of the gyroscope of the second set of inertial elements; Simplified as H g2 , H g2 The calibration and error comprehensive parameters of the gyroscope of the second group of inertial components are calculated in H g2 When H g In the expression of T Remove the corresponding row and you get H g2 ;Will Simplified as ω M2 ,ω M2 It is the angular velocity of the second inertial element combination in the inertial group measurement coordinate system.

[0052]

[0053] Will Simplified expression: is the actual total output of the accelerometer of the second set of inertial elements; Simplified as H a2 , H a2 The calibration and error comprehensive parameters of the accelerometer of the second group of inertial components are calculated in H a2 When H a In the expression of T Remove the corresponding row and you get H a2 ;Will Simplified as A M2 , A M2 is the apparent acceleration of the second group of inertial elements in the inertial group measurement coordinate system.

[0054] After calculating the H of the second set of inertial components g2 and H a2Then, the angular velocity ω of the second set of inertial element combination in the inertial group measurement coordinate system is calculated by the following formula: M2 and apparent acceleration A M2 .

[0055]

[0056] in, H g2 The transpose of H a2 The transpose of .

[0057] According to the above method, the angular rates and apparent accelerations of other inertial element combinations in Table 1 in the inertial group measurement coordinate system are calculated to convert the outputs of each inertial element combination into the inertial group measurement coordinate system.

[0058] Step S310: at the beginning of an integrated navigation cycle, based on the output of the first group of inertial elements converted to the inertial group measurement coordinate system and the output of the GNSS receiver, perform the first integrated navigation attitude solution in the integrated navigation cycle;

[0059] The outputs of the gyroscopes and accelerometers in the three orthogonal X-axis, Y-axis, and Z-axis directions in the first group of Table 1 are converted to the inertial group measurement coordinate system to obtain the corresponding angular rate and apparent acceleration, and then the corresponding position, velocity, and attitude angle are obtained through inertial navigation solution, and the position, velocity, and attitude angle are sent to the flight control computer.

[0060] The present application adopts an inertial / satellite combined navigation method, the output period of the ten-meter strapdown inertial group is 5ms, the output period of the GNSS receiver is 100ms, the combined navigation period adopts the output period of the GNSS receiver of 100ms, and the GNSS receiver also sends a PPS (Pulse Per Second) second pulse to the flight control computer, and the flight control computer uses the PPS (Pulse Per Second) second pulse to synchronize the data of the ten-meter strapdown inertial group with the data of the GNSS receiver. After the time synchronization, at the beginning of a combined navigation cycle, the flight control computer uses the Kalman filtering method to perform combined navigation attitude solution through the position, velocity and attitude angle corresponding to the first group of inertial element combinations in the first row of Table 1 and the position, velocity and attitude angle sent to the flight control computer by the GNSS receiver. The obtained combined navigation attitude solution result is: the attitude angle is Φ0=[φ ψ γ]0, and the velocity is υ0=[v x v y v z ]0, the position is P0=[P x P y P z ]0, only closed-loop correction of speed and position is performed subsequently.

[0061] The first combined navigation attitude solution result in the combined navigation cycle is used as a reference, and the reference is used for resetting in subsequent navigation attitude solution processes in the combined navigation cycle.

[0062] Step S315, performing pure inertial navigation attitude solution according to the outputs of the second to seventh inertial element groups converted to the inertial group measurement coordinate system;

[0063] The inertial component combinations of groups 2 to 7 in Table 1 are converted to the inertial measurement coordinate system to obtain the corresponding angular rate and apparent acceleration. Each group uses pure inertial navigation attitude solution, and the pure inertial navigation attitude solution results are: the velocity of the i-th group is v i =[v x v y v z ] i , the position of the i-th group is P i =[P x P y P z ] i .

[0064] Step S320: at the end of an integrated navigation cycle, performing a second integrated navigation attitude solution within the integrated navigation cycle according to the output of the first group of inertial elements converted to the inertial group measurement coordinate system and the output of the GNSS receiver;

[0065] When 100ms of each integrated navigation cycle arrives, the output of the first group of inertial elements combined and converted to the inertial measurement coordinate system and the output of the GNSS receiver are used to perform the second integrated navigation attitude solution within the integrated navigation cycle. The integrated navigation attitude solution result is: the speed is Location is

[0066] Step S325, performing consistency judgment on the pure inertial navigation attitude solution results of the second to seventh inertial element combination groups and the second combined navigation attitude solution result of the first inertial element combination group;

[0067] The speed v of the second to seventh groups of inertial elements are respectively i =[v x v y v z ] i and position P i =[P x P y P z ] i The speed of the combination with the first group of inertial elements is v0 = [v x vy v z ] 01 and position P0 = [P x P y P z ] 01 Perform consistency judgment.

[0068] Specifically, speed determination Position determination Where i=2~7.

[0069] Step S330, judging whether the six groups of discrimination results exceed the corresponding design threshold values;

[0070] After consistency judgment, 6 groups of judgment results Δv are obtained i , ΔP i , where Δv i is the i-th speed judgment result, ΔP i is the result of the i-th position discrimination. Then, the discrimination result of each group Δv is judged. i , ΔP i Medium speed judgment result Δv i Whether the values ​​of are all less than the speed design threshold value v, and judge each group of judgment results Δυ i , ΔP i Middle position determination result ΔP i Whether the values ​​of are all smaller than the position design threshold value P.

[0071] Among them, the speed design threshold value is δ v =δ v0 +k v ·t, the position design threshold is δ P =δ P0 +k P ·t, where δ v0 is the initial threshold of speed deviation, δ P0 is the initial threshold of position deviation, k v is the slope of the velocity deviation diverging with navigation time, k P is the slope of the position deviation diverging with navigation time, δ v0 , δ P0 , k v , k P These are all values ​​of parameters obtained by simulation calculation and other methods, and t is the time length for establishing inertial navigation.

[0072] Step S335: If the six groups of discrimination results do not exceed the corresponding design threshold values, the first combined navigation attitude solution result in the combined navigation cycle is used as the navigation result used for guidance and attitude control;

[0073] If the six groups of discrimination results Δv2~Δv7, ΔP2~ΔP7 do not exceed the design threshold value, it is considered that the output of each group of inertial element combination of the ten-meter strapdown inertial unit redundant combination is normal, and the first combined navigation attitude solution result Φ0=[φ ψ γ]0, u0=[v x v y υ z ]0、P0=[P x P y P z ]0 as the navigation result used by the guidance system and attitude control system to control the rocket control mechanism.

[0074] Step S340: If any of the six groups of discrimination results exceeds the designed threshold value, the axis of the ten-meter strapdown inertial unit fault is selected and isolated;

[0075] According to step S310 to step S335, the speed deviation and position deviation consistency judgment and δ v ,δ P If Δv i >δ v or ΔP i >δ P , then perform gyroscope fault identification, screen out the axial direction where the faulty instrument is located, and isolate the instrument on that axial direction.

[0076] Step S345, eliminating the inertial element combination including the faulty axial direction;

[0077] If the deviations of the three orthogonal axial instruments are large, that is, the deviations of the results of the second group are large, the data of the fifth to seventh groups can be combined to determine the axis in which the specific instrument deviation is large, and then one of the inertial element combinations from the third to seventh groups that does not contain the instrument of this axis is selected as the basis for subsequent navigation attitude solution, and a combined navigation correction is performed every 100ms combined navigation cycle. At the same time, this group of navigation attitude solution results is used as the navigation results used by the guidance system and attitude control system.

[0078] When the next integrated navigation cycle begins, the above steps S310 to S345 are repeated until the flight mission is completed.

[0079] This application uses independent navigation attitude solution loops to judge the consistency of speed and position, and compares the speed and position information of the GNSS receiver. Through reasonable thresholds, it can detect and isolate the fault caused by the large zero bias error of the inertial instrument, which causes the navigation accuracy to drop for a long time, and improve the accuracy of inertial group fault monitoring and isolation. At the same time, it adopts multi-channel navigation solution with navigation information memory function. When a fault occurs, it can isolate the inertial group fault and reset the navigation information in the first time, improving the error correction capability of the navigation system.

[0080] It will be apparent to those skilled in the art that the invention is not limited to the details of the exemplary embodiments described above and that the invention can be implemented in other specific forms without departing from the spirit or essential features of the invention. Therefore, the embodiments should be considered exemplary and non-limiting in all respects, and the scope of the invention is defined by the appended claims rather than the foregoing description, and it is intended that all variations falling within the meaning and scope of the equivalent elements of the claims be included in the invention. Any reference numeral in a claim should not be considered as limiting the claim to which it relates.

[0081] In addition, it should be understood that although the present specification is described according to implementation modes, not every implementation mode contains only one independent technical solution. This description of the specification is only for the sake of clarity. Those skilled in the art should regard the specification as a whole. The technical solutions in each embodiment may also be appropriately combined to form other implementation modes that can be understood by those skilled in the art.

Claims

1. A method for determining a redundant strategy of a ten-meter inertia group, characterized in that: The steps include: The ten-meter strapdown inertial group is combined redundantly to establish seven groups of inertial element combinations in different axes, and the output of each group of inertial element combinations is converted to the inertial group measurement coordinate system; At the beginning of an integrated navigation cycle, performing the first integrated navigation attitude solution in the integrated navigation cycle according to the output of the first group of inertial elements combined and converted to the inertial group measurement coordinate system and the output of the GNSS receiver; Based on the outputs of the second to seventh inertial element combinations converted to the inertial group measurement coordinate system, pure inertial navigation attitude solution is performed; At the end of the integrated navigation cycle, performing a second integrated navigation attitude solution within the integrated navigation cycle according to the output of the first group of inertial elements combined and converted to the inertial group measurement coordinate system and the output of the GNSS receiver; Perform consistency judgment on the pure inertial navigation attitude solution results of the second to seventh inertial element combinations and the second combined navigation attitude solution result of the first inertial element combination; Determine whether the six groups of discrimination results exceed the corresponding design threshold values; If the six groups of discrimination results do not exceed the corresponding design threshold values, the first combined navigation attitude solution result in the combined navigation cycle is used as the navigation result used for guidance and attitude control; At the beginning of the next integrated navigation cycle, the navigation attitude solution and consistency judgment are repeated until the flight mission is completed.

2. The method for determining the redundancy strategy of a ten-meter inertia group according to claim 1 is characterized in that: The following steps are also included: If any of the six groups of discrimination results exceeds the designed threshold value, the axis where the ten-meter strapdown inertial unit fails will be selected and isolated; Eliminate the inertial component combination that contains the faulty axial direction; At the beginning of the next integrated navigation cycle, the navigation attitude solution and consistency judgment are repeated until the flight mission is completed.

3. The method for determining the redundancy strategy of a ten-meter inertia group according to claim 1 or 2, characterized in that: When the output of each set of inertial components is converted to the inertial group measurement coordinate system: Calculate the H of the gyroscope for each set of inertial components g When H g Remove the rows corresponding to the output of the gyroscope in the inertial element combination that does not participate in the inertial navigation solution from the expression of , and you will get the H of the gyroscope in the inertial element combination. g ; Calculate the H of the accelerometer for each set of inertial elements a When H a Remove the rows corresponding to the output of the accelerometers in the inertial element combination that are not involved in the inertial navigation solution from the expression of , and you will get the H of the accelerometers in the inertial element combination. a ; Among them, H g is the calibration and error comprehensive parameter of the gyroscope combined with inertial components, H a Calibration and comprehensive error parameters of the accelerometer combined with inertial elements.

4. The method for determining the redundancy strategy of a ten-meter inertia group according to claim 3 is characterized in that: After calculating the H of each group of inertial components g and H a back; By formula Calculate the angular rate ω of each set of inertial components in the inertial measurement coordinate system M and apparent acceleration A M ;ω b A is the actual total output of the gyroscope for each set of inertial elements; b is the actual total output of the accelerometer for each set of inertial elements; The angular rate ω M and apparent acceleration A M As output in the inertial measurement coordinate system.

5. The method for determining the redundancy strategy of a ten-meter inertia group according to claim 4 is characterized in that: The angular rate and apparent acceleration are converted into position, velocity and attitude angle through inertial navigation solution.

6. The method for determining the redundancy strategy of a ten-meter inertia group according to claim 1 or 2, characterized in that: PPS pulses per second are used to synchronize the data of the ten-meter strapdown inertial group and the data of the GNSS receiver.

7. The method for determining the redundancy strategy of a ten-meter inertia group according to claim 1 or 2, characterized in that: The design threshold values ​​include: speed design threshold value and position design threshold value; Among them, the speed design threshold value is δ υ =δ υ0 +k υ ·t, the position design threshold is δ P =δ P0 +k P ·t;δ υ0 is the initial threshold of speed deviation, δ P0 is the initial threshold of position deviation, k υ is the slope of the velocity deviation diverging with navigation time, k P is the slope of the position deviation diverging with navigation time, δ υ0 , δ P0 , k υ , k P These are all the values ​​obtained by simulation calculation method, and t is the time length for establishing inertial navigation.

8. A redundancy strategy identification system for a ten-meter inertial group, characterized in that: include: GNSS receiver, GNSS antenna, flight control computer and ten-meter strapdown inertial group; wherein the ten-meter strapdown inertial group is arranged on the launch vehicle, and the ten-meter strapdown inertial group is connected to the flight control computer; the GNSS antenna is connected to the GNSS receiver, and the GNSS receiver is connected to the flight control computer; The ten-meter inertia group redundancy strategy determination system executes the ten-meter inertia group redundancy strategy determination method according to any one of claims 1 to 7.

9. The ten-meter inertial group redundancy strategy identification system according to claim 8 is characterized in that: The ten-meter strapdown inertial group and the flight control computer, the GNSS antenna and the GNSS receiver, and the GNSS receiver and the flight control computer are all connected through arrow cables.

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