An aircraft attitude measurement method using dual-redundancy technology

Abstract

This invention provides a dual-redundancy method for measuring aircraft attitude. It establishes a first measurement channel for the aircraft's heading angle by jointly measuring and calculating with a triaxial accelerometer and a Hall effect magnetometer. A second measurement channel for the heading angle is then established using a MESE rate gyroscope and a MESE angle gyroscope. The errors from these two channels are then combined and nonlinearly transformed, and the error is fed back to the rate gyroscope for secondary measurement correction. Finally, the combined aircraft attitude measurement signal is obtained through integration. Furthermore, when the triaxial accelerometer, magnetometer, rate gyroscope, or angle gyroscope malfunctions, information from the normally functioning sensors is used, and the gain of the combined measurement feedback channel is switched on and off to form a dual-redundancy measurement, thus improving the reliability and accuracy of the overall attitude measurement.

Description

Technical Field

[0001] This invention relates to the field of aircraft attitude angle measurement, and more specifically, to an aircraft attitude measurement method employing dual-redundancy technology. Background Technology

[0002] In the field of aircraft attitude measurement, gyroscopes are commonly used to measure the attitude angles of aircraft. However, the measurement error of gyroscopes accumulates and drifts over time. Using high-precision gyroscopes is economically prohibitive for many low-cost aircraft. While a combination of a three-axis accelerometer and a magnetometer can calculate the aircraft's attitude under level or constant speed conditions, the overall accuracy is not high, although the error does not accumulate over time. Traditional methods use Kalman filtering to fuse these two sets of data, but the fusion accuracy still depends on prior knowledge of the covariance of random errors and noise during flight. Furthermore, when one sensor malfunctions, the Kalman filter fusion result will be insufficient to provide accurate attitude angle information for the aircraft control system. Based on these background reasons, we propose a method that uses nonlinear error transformation and secondary correction and combined integration of the rate gyroscope signal to achieve dual-redundancy fusion of information from two measurement channels. Moreover, even with partial sensor failure, the method can still measure the aircraft's heading angle by switching the feedback channel gain on and off, thus making the method highly reliable and practically applicable in engineering.

[0003] It should be noted that the information in the background section above is only used to enhance the understanding of the background of the present invention, and therefore may include information that does not constitute prior art known to those skilled in the art. Summary of the Invention

[0004] The purpose of this invention is to provide an aircraft attitude measurement method using dual-redundancy technology, thereby overcoming the problem of low reliability in aircraft heading and attitude angle measurement caused by limitations and defects in related technologies.

[0005] According to one aspect of the present invention, a method for measuring the attitude of an aircraft using dual-redundancy technology is provided, comprising the following steps:

[0006] Step S10: First, a triaxial accelerometer is installed on the longitudinal axis of the aircraft to measure the components of gravitational acceleration in the x, y, and z axes during uniform motion. The three acceleration components are then normalized to obtain the normalized components of gravitational acceleration in the x, y, and z axes. Next, the angle of gravitational acceleration relative to the y-axis, as well as the tilt and azimuth angles of the aircraft, are solved using inverse trigonometric functions. A Hall effect magnetometer is installed on the longitudinal axis of the aircraft to measure the components of magnetic induction intensity in the x, y, and z axes. Based on the tilt and azimuth angles of the aircraft, the x-axis component and y-axis component of the magnetic field intensity in the direction parallel to the ground are solved. Finally, the magnetometer measurement value of the aircraft's yaw angle is obtained through arctangent transformation.

[0007] Step S20: Install a MEMS rate gyroscope on the aircraft to measure the yaw rate of the aircraft and obtain the yaw rate measurement signal. Install a MEMS angle gyroscope on the aircraft to measure the yaw angle and obtain the yaw angle gyroscope measurement value. Then compare the yaw angle gyroscope magnetic force measurement error signal with the magnetometer measurement value signal of the aircraft yaw angle to obtain the yaw angle gyroscope magnetic force measurement error signal. Then perform a nonlinear transformation on the yaw angle gyroscope magnetic force measurement error signal to obtain the angle gyroscope measurement error nonlinear transformation signal. Then superimpose the proportionally amplified signal to obtain the angle gyroscope compensation signal. Then superimpose the aircraft yaw rate measurement signal obtained by the rate gyroscope to obtain the aircraft yaw rate compensation correction signal.

[0008] Step S30: Set the initial value of the yaw angle magnetic gyroscope combined measurement signal to 0, and then compare it with the magnetometer measurement signal of the aircraft's yaw angle to obtain the yaw angle combined measurement error signal; then perform a nonlinear transformation on the yaw angle combined measurement error signal to obtain the combined measurement error nonlinear transformation signal; then superimpose the proportionally amplified signal to obtain the angular rate gyroscope compensation signal; then superimpose the aforementioned aircraft yaw angle velocity compensation correction signal to obtain the aircraft yaw angle velocity secondary compensation correction signal; finally, perform integration to obtain the yaw angle magnetic gyroscope combined measurement signal.

[0009] Step S40: Based on the status of the three-axis accelerometer, Hall effect magnetometer, MESE rate gyroscope and MESE angle gyroscope, the aircraft heading angle dual redundancy design is carried out in six cases, and the fault-free signal is selected to form the final aircraft heading angle dual redundancy measurement signal.

[0010] In one exemplary embodiment of the present invention, a triaxial accelerometer is installed on the longitudinal axis of the aircraft body to measure the components of gravitational acceleration in the x, y, and z axes during uniform motion. The three acceleration components are then normalized to obtain the normalized components of gravitational acceleration in the x, y, and z axes. The angle of gravitational acceleration relative to the y-axis, as well as the aircraft's tilt and azimuth angles, are then calculated using inverse trigonometric functions. A Hall effect magnetometer is installed along the longitudinal axis of the aircraft body to measure the components of magnetic induction intensity in the x, y, and z axes. Based on the aircraft's tilt and azimuth angles, the x-axis component and y-axis component of the magnetic field strength in the direction parallel to the ground are calculated. Finally, the magnetometer measurement value for the aircraft's yaw angle is obtained through arctangent transformation, including:

[0011]

[0012]

[0013]

[0014] β=arccos(a y );

[0015] φ=arctan(a y / a z );

[0016] ρ=-arcsin(a x );

[0017] H a =H x cosρ+H y sinρsinφ+H z sinρcosφ;

[0018] H b =H y cosφ-H z sinφ;

[0019]

[0020] Where R x R y With R z To represent the components of gravitational acceleration along the x, y, and z axes as measured by a triaxial accelerometer, a x a y With a z ρ represents the unitized components of gravitational acceleration in the x, y, and z axes; β is the angle of gravitational acceleration relative to the y-axis; φ is the tilt angle of the spacecraft; ρ is the azimuth angle of the spacecraft; H xH y H z The components of magnetic flux density along the x, y, and z axes are measured using a Hall effect magnetometer; H a H represents the x-axis component of the magnetic field strength in the direction parallel to the ground. b ψ represents the y-axis component parallel to the ground, and ψ is the magnetometer measurement of the aircraft's yaw angle.

[0021] In one exemplary embodiment of the present invention, a MEMS rate gyroscope is installed on the aircraft to measure the yaw rate of the aircraft, obtaining a yaw rate measurement signal. A MEMS angle gyroscope is installed on the aircraft to measure the yaw angle, obtaining a yaw angle gyroscope measurement value. This value is then compared with the magnetometer measurement value of the yaw angle to obtain a yaw angle gyroscope magnetic force measurement error signal. The yaw angle gyroscope magnetic force measurement error signal is then subjected to a nonlinear transformation to obtain an angle gyroscope measurement error nonlinear transformation signal. This signal is then superimposed with a proportionally amplified signal to obtain an angle gyroscope compensation signal. Finally, the yaw rate measurement signal obtained from the rate gyroscope is superimposed to obtain a yaw rate compensation correction signal, including:

[0022] e1 = ψ - φ1;

[0023]

[0024] Δω1=f1+k4e1;

[0025] ω2=ω1+Δω1;

[0026] Where φ1 is the measured value of the yaw angle gyroscope, e1 is the yaw angle gyroscope magnetic force measurement error signal; f1 is the nonlinear transformation signal of the angle gyroscope measurement error; k1, k2, k3, and a1 are constant parameters of the nonlinear transformation; k4 is a constant proportional amplification parameter; Δω1 is the angle gyroscope compensation signal; ω2 is the yaw angle velocity compensation correction signal of the aircraft; and ω1 is the yaw angle rate measurement signal of the aircraft.

[0027] In one exemplary embodiment of the present invention, the initial value of the yaw angle magnetic gyroscope combined measurement signal is set to 0, and then compared with the magnetometer measurement signal of the aircraft yaw angle to obtain the yaw angle combined measurement error signal; then, the yaw angle combined measurement error signal is subjected to a nonlinear transformation to obtain a combined measurement error nonlinear transformation signal; then, a proportionally amplified signal is superimposed to obtain an angular rate gyroscope compensation signal; then, the aforementioned aircraft yaw angle velocity compensation correction signal is superimposed to obtain a secondary compensation correction signal for the aircraft yaw angle velocity; finally, integration is performed to obtain the yaw angle magnetic gyroscope combined measurement signal as follows:

[0028] e2 = ψ - φ2;

[0029]

[0030] Δω2=f2+k8e2;

[0031] ω3=ω2+Δω2;

[0032] φ2=∫ω3dt;

[0033] Where φ2 is initially set to 0, e2 is the yaw angle combined measurement error signal; f2 is the combined measurement error nonlinear transformation signal; k5, k6, and k7 are constant parameters of the nonlinear transformation; k8 is the constant proportional amplification parameter; Δω2 is the angular rate gyroscope compensation signal; ω3 is the aircraft yaw angle velocity secondary compensation correction signal; and φ2 is the yaw angle magnetic gyroscope combined measurement signal.

[0034] In one exemplary embodiment of the present invention, based on the states of the triaxial accelerometer, Hall effect magnetometer, MESE rate gyroscope, and MESE angle gyroscope, a dual-redundancy design for the aircraft's heading angle is implemented under six different conditions. Fault-free signals are selected to form the final dual-redundancy measurement signal for the aircraft's heading angle, which includes:

[0035]

[0036] Where i a =0 indicates a malfunction in the triaxial accelerometer; i a =1 indicates that the triaxial accelerometer is functioning normally; where i b =0 indicates a malfunction in the Hall effect magnetometer; i b =1 indicates that the Hall effect magnetometer is functioning normally; where i c =0 indicates a malfunction in the MESE rate gyroscope status; i c =1 indicates that the MESE rate gyroscope is functioning normally; where i d =0 indicates a malfunction in the MESE angle gyroscope status; i d =1 indicates that the MESE angle gyroscope is functioning normally; where φ a1 The φ2 signal is obtained when k1, k2, k3, k4, k5, k6, k7, and k8 are all 0, which is the combined measurement output signal when all feedback is disconnected; φ a2 The φ2 signal is obtained by keeping the original values ​​of k5, k6, k7, and k8 unchanged, while selecting k1, k2, k3, and k4 as 0. In other words, it is the combined measurement output signal with the feedback part of the faulty angle gyroscope disconnected and the others unchanged.

[0037] Beneficial effects

[0038] This invention discloses an aircraft attitude measurement method employing dual-redundancy technology. Its main innovations are as follows: First, it establishes two attitude angle measurement channels using a three-degree-of-freedom accelerometer, magnetometer, rate gyroscope, and angle gyroscope. When some sensors fail, the feedback gain of the combined measurement channels can be switched on and off to fuse information from sensors in normal operation, improving attitude angle information and achieving dual-redundancy measurement. This significantly improves system reliability and enables normal yaw angle measurement even with partial sensor failures. Second, the dual-channel dual-redundancy measurement scheme allows for high-precision measurement using a low-cost MESE gyroscope, greatly reducing the economic cost of using high-precision gyroscopes. Third, the dual-redundancy technology can utilize magnetometer information to correct gyroscope drift, solving the problem of gyroscope cumulative error. Fourth, it employs a method of matching error nonlinear feedback with the rate gyroscope to achieve the fusion of dual-redundancy information from the two channels. Traditional fusion methods, such as Kalman filtering, render the fused information unusable if one channel fails. However, by using a two-way feedback method, we were able to ensure that the angle gyroscope could still function normally even when the channel was disconnected and the gain of some feedback channels was set to 0, with only a slight decrease in accuracy.

[0039] It should be understood that the above general description and the following detailed description are exemplary and explanatory only, and are not intended to limit the invention. Attached Figure Description

[0040] The accompanying drawings, which are incorporated in and constitute a part of this specification, illustrate embodiments consistent with the invention and, together with the description, serve to explain the principles of the invention. It is obvious that the drawings described below are merely some embodiments of the invention, and those skilled in the art can obtain other drawings based on these drawings without any inventive effort.

[0041] Figure 1 This is a flowchart of an aircraft attitude measurement method using dual-redundancy technology provided by the present invention;

[0042] Figure 2 This is a magnetometer measurement curve of the aircraft yaw angle (unit: degrees) provided by the method in the embodiments of the present invention;

[0043] Figure 3 This is the yaw angle gyroscope measurement curve of the aircraft (unit: degrees) provided by the method in the embodiments of the present invention;

[0044] Figure 4 This is the yaw angle gyroscope magnetic force measurement error signal curve (unit: degrees) of the method provided in the embodiments of the present invention;

[0045] Figure 5This is the aircraft yaw rate compensation correction signal curve (unitless) provided by the method in the embodiments of the present invention;

[0046] Figure 6 This is the yaw angle combination measurement error signal (unit: degrees) of the method provided in the embodiments of the present invention;

[0047] Figure 7 This is the secondary compensation correction signal curve of the aircraft yaw rate (unitless) provided by the method in the embodiments of the present invention;

[0048] Figure 8 This is the yaw angle magnetic gyroscope combination measurement signal curve (unit: degrees) of the method provided in the embodiments of the present invention;

[0049] Figure 9 This is the dual-redundant measurement signal curve of the aircraft heading angle (unit: degrees) provided by the method in the embodiments of the present invention. Detailed Implementation

[0050] Exemplary embodiments will now be described more fully with reference to the accompanying drawings. However, these exemplary embodiments can be implemented in many forms and should not be construed as limited to the examples set forth herein; rather, these embodiments are provided to make the invention more comprehensive and complete, and to fully convey the concept of the exemplary embodiments to those skilled in the art. The described features, structures, or characteristics can be combined in any suitable manner in one or more embodiments. In the following description, numerous specific details are provided to give a full understanding of embodiments of the invention. However, those skilled in the art will recognize that the technical solutions of the invention may be practiced with one or more of these specific details omitted, or other methods, components, apparatus, steps, etc., may be employed. In other instances, well-known technical solutions are not shown or described in detail to avoid obscuring various aspects of the invention.

[0051] This invention provides a dual-redundancy method for measuring aircraft attitude. It establishes a first measurement channel for the aircraft's heading angle by jointly measuring and calculating with a triaxial accelerometer and a Hall effect magnetometer. A second measurement channel for the heading angle is then established using a MESE rate gyroscope and a MESE angle gyroscope. The errors from these two channels are then combined and nonlinearly transformed, and the error is fed back to the rate gyroscope for secondary measurement correction. Finally, the combined aircraft attitude measurement signal is obtained through integration. Furthermore, when the triaxial accelerometer, magnetometer, rate gyroscope, or angle gyroscope malfunctions, information from the normally functioning sensors is used, and the gain of the combined measurement feedback channel is switched on and off to form a dual-redundancy measurement, thus improving the reliability and accuracy of the overall attitude measurement.

[0052] The following will, with reference to the accompanying drawings, further explain and illustrate an aircraft attitude measurement method employing dual-redundancy technology according to the present invention. (Reference) Figure 1 As shown, this aircraft attitude measurement method using dual-redundancy technology may include the following steps:

[0053] Step S10: First, a triaxial accelerometer is installed on the longitudinal axis of the aircraft to measure the components of gravitational acceleration in the x, y, and z axes during uniform motion. The three acceleration components are then normalized to obtain the normalized components of gravitational acceleration in the x, y, and z axes. Next, the angle of gravitational acceleration relative to the y-axis, as well as the tilt and azimuth angles of the aircraft, are solved using inverse trigonometric functions. A Hall effect magnetometer is installed on the longitudinal axis of the aircraft to measure the components of magnetic induction intensity in the x, y, and z axes. Based on the tilt and azimuth angles of the aircraft, the x-axis component and y-axis component of the magnetic field intensity in the direction parallel to the ground are solved. Finally, the magnetometer measurement value of the aircraft's yaw angle is obtained through arctangent transformation.

[0054] Specifically, this can be broken down into the following four steps. The first step is to install a triaxial accelerometer on the longitudinal axis of the spacecraft to measure the components of gravitational acceleration along the x, y, and z axes during uniform motion. These three acceleration components are then normalized to obtain the normalized components of gravitational acceleration along the x, y, and z axes as follows:

[0055]

[0056]

[0057]

[0058] Where R x R y With R z To represent the components of gravitational acceleration along the x, y, and z axes as measured by a triaxial accelerometer, a x a y With a z This represents the unitized components of gravitational acceleration in the x, y, and z axes.

[0059] The second step involves using inverse trigonometric functions to solve for the angle between gravitational acceleration and the y-axis, as well as the aircraft's tilt and azimuth angles, as follows:

[0060] β=arccos(a y );

[0061] φ=arctan(a y / a z );

[0062] ρ=-arcsin(a x );

[0063] Where β is the angle between gravitational acceleration and the y-axis; φ is the tilt angle of the aircraft; and ρ is the azimuth angle of the aircraft.

[0064] The third step involves installing a Hall effect magnetometer along the longitudinal axis of the aircraft to measure the components of the magnetic field strength along the x, y, and z axes. Then, based on the aircraft's tilt and azimuth angles, the x-axis and y-axis components of the magnetic field strength parallel to the ground are calculated as follows:

[0065] H a =H x cosρ+H y sinρsinφ+H z sinρcosφ;

[0066] H b =H y cosφ-H z sinφ;

[0067] Where H x H y H z The components of magnetic flux density along the x, y, and z axes are measured using a Hall effect magnetometer; H a H represents the x-axis component of the magnetic field strength in the direction parallel to the ground. b The y-axis component is parallel to the ground.

[0068] Fourth, based on the angle between gravitational acceleration and the y-axis, and the tilt and azimuth angles of the aircraft, the magnetometer measurement value of the aircraft's yaw angle is obtained through arctangent transformation, as follows:

[0069]

[0070] Where ψ is the magnetometer measurement of the aircraft's yaw angle.

[0071] Step S20: Install a MEMS rate gyroscope on the aircraft to measure the yaw rate of the aircraft and obtain the yaw rate measurement signal. Install a MEMS angle gyroscope on the aircraft to measure the yaw angle and obtain the yaw angle gyroscope measurement value. Then compare the yaw angle gyroscope magnetic force measurement error signal with the magnetometer measurement value signal of the aircraft yaw angle to obtain the yaw angle gyroscope magnetic force measurement error signal. Then perform a nonlinear transformation on the yaw angle gyroscope magnetic force measurement error signal to obtain the angle gyroscope measurement error nonlinear transformation signal. Then superimpose the proportionally amplified signal to obtain the angle gyroscope compensation signal. Then superimpose the aircraft yaw rate measurement signal obtained by the rate gyroscope to obtain the aircraft yaw rate compensation correction signal.

[0072] Specifically, it can be broken down into the following four steps. First, install a MEMS rate gyroscope on the aircraft to measure the yaw rate of the aircraft and obtain the yaw rate measurement signal. Then, install a MEMS angle gyroscope on the aircraft to measure the yaw angle and obtain the yaw angle gyroscope measurement value.

[0073] The second step involves comparing the yaw angle gyroscope measurement value with the magnetometer measurement value of the yaw angle to obtain the yaw angle gyroscope magnetic force measurement error signal, as follows:

[0074] e1 = ψ - φ1;

[0075] Where φ1 is the measured value of the yaw angle gyroscope, and e1 is the yaw angle gyroscope magnetic force measurement error signal.

[0076] The third step is to perform a nonlinear transformation on the yaw angle gyroscope magnetic force measurement error signal, resulting in the following nonlinear transformation signal for the angle gyroscope measurement error:

[0077]

[0078] Where f1 is the nonlinear transformation signal of the angle gyroscope measurement error; k1, k2, k3, and a1 are constant parameters of the nonlinear transformation.

[0079] The fourth step involves superimposing the nonlinear transformation signal of the angle gyroscope measurement error onto a proportionally amplified signal to obtain the angle gyroscope compensation signal; then, superimposing this signal onto the aircraft yaw rate measurement signal obtained from the rate gyroscope, the resulting aircraft yaw rate compensation correction signal is as follows:

[0080] Δω1=f1+k4e1;

[0081] ω2=ω1+Δω1;

[0082] Where k4 is a constant proportional amplification parameter, Δω1 is the angle gyroscope compensation signal; ω2 is the aircraft yaw rate compensation correction signal; and ω1 is the aircraft yaw rate measurement signal.

[0083] Step S30: Set the initial value of the yaw angle magnetic gyroscope combined measurement signal to 0, and then compare it with the magnetometer measurement signal of the aircraft's yaw angle to obtain the yaw angle combined measurement error signal; then perform a nonlinear transformation on the yaw angle combined measurement error signal to obtain the combined measurement error nonlinear transformation signal; then superimpose the proportionally amplified signal to obtain the angular rate gyroscope compensation signal; then superimpose the aforementioned aircraft yaw angle velocity compensation correction signal to obtain the aircraft yaw angle velocity secondary compensation correction signal; finally, perform integration to obtain the yaw angle magnetic gyroscope combined measurement signal.

[0084] Specifically, this can be broken down into the following four steps. Step 1: Set the initial value of the yaw angle magnetic gyroscope combination measurement signal to 0, and then compare it with the magnetometer measurement signal of the aircraft's yaw angle to obtain the yaw angle combination measurement error signal as follows:

[0085] e2 = ψ - φ2;

[0086] The initial value of φ2 is set to 0, and e2 is the yaw angle combined measurement error signal.

[0087] The second step is to perform a nonlinear transformation on the combined yaw angle measurement error signal, resulting in the following nonlinear transformed signal of the combined measurement error:

[0088]

[0089] Where f2 is the nonlinear transformation signal of the combined measurement error; k5, k6, and k7 are constant parameters of the nonlinear transformation.

[0090] The third step involves superimposing the nonlinear transformation signal of the combined measurement error onto the proportionally amplified signal to obtain the angular rate gyroscope compensation signal as follows:

[0091] Δω2=f2+k8e2;

[0092] Where k8 is a constant proportional amplification parameter; Δω2 is the angular rate gyroscope compensation signal.

[0093] The fourth step involves superimposing the angular rate gyroscope compensation signal onto the aforementioned yaw rate compensation correction signal to obtain the secondary yaw rate compensation correction signal. This is then integrated to obtain the yaw angle magnetic gyroscope combined measurement signal, as follows:

[0094] ω3=ω2+Δω2;

[0095] φ2=∫ω3dt;

[0096] Where ω3 is the secondary compensation correction signal for the yaw rate of the aircraft; φ2 is the measurement signal of the yaw angle magnetic gyroscope combination;

[0097] Step S40: Based on the status of the three-axis accelerometer, Hall effect magnetometer, MESE rate gyroscope and MESE angle gyroscope, the aircraft heading angle dual redundancy design is carried out in six cases, and the fault-free signal is selected to form the final aircraft heading angle dual redundancy measurement signal.

[0098] Specifically, the dual-redundant measurement signal for the aircraft's heading angle is calculated under the following six conditions. The first condition is when the three-axis accelerometer or the Hall effect magnetometer malfunctions, and the MESE rate gyroscope malfunctions, while the MESE angle gyroscope is functioning normally. In this case, the output signal of the MESE angle gyroscope is selected as the final dual-redundant measurement signal for the aircraft's heading angle.

[0099] The second scenario is when the three-axis accelerometer or the Hall effect magnetometer malfunctions, and the MESE rate gyroscope is normal while the MESE angle gyroscope malfunctions. In this case, the φ2 signal obtained when the feedback coefficients k1, k2, k3, k4, k5, k6, k7, and k8 in the combined measurement are all 0, which is the combined measurement output signal when all feedback is disconnected, is selected as the final dual-redundant measurement signal for the aircraft heading angle.

[0100] The third scenario is when the three-axis accelerometer or the Hall effect magnetometer malfunctions, but the MESE rate gyroscope and the MESE angle gyroscope are functioning normally. In this case, the average value of the yaw angle magnetic gyroscope combined measurement signal when all feedback is disconnected and the output signal of the MESE angle gyroscope are selected as the final dual-redundant measurement signal for the aircraft's heading angle.

[0101] The fourth scenario is that the three-axis accelerometer and the Hall effect magnetometer are both functioning normally, but the MESE rate gyroscope is malfunctioning while the MESE angle gyroscope is functioning normally. In this case, the output signal of the MESE angle gyroscope is selected as the final dual-redundant measurement signal for the aircraft's heading angle.

[0102] The fifth scenario is that the three-axis accelerometer and the Hall effect magnetometer are both functioning normally, while the MESE rate gyroscope is functioning normally, but the MESE angle gyroscope is malfunctioning. In this case, the combined measurement output signal with k5, k6, k7, and k8 kept unchanged and k1, k2, k3, and k4 set to 0 is used as the final dual-redundant measurement signal for the aircraft heading angle.

[0103] The sixth scenario is that the three-axis accelerometer and the Hall effect magnetometer are both functioning normally, as are the MESE rate gyroscope and the MESE angle gyroscope. In this case, the combined measurement signal of the yaw angle magnetogyroscope is selected as the final dual-redundant measurement signal for the aircraft's heading angle.

[0104] The above six scenarios are summarized in the table below:

[0105]

[0106] Where i a =0 indicates a malfunction in the triaxial accelerometer; i a=1 indicates that the triaxial accelerometer is functioning normally; where i b =0 indicates a malfunction in the Hall effect magnetometer; i b =1 indicates that the Hall effect magnetometer is functioning normally; where i c =0 indicates a malfunction in the MESE rate gyroscope status; i c =1 indicates that the MESE rate gyroscope is functioning normally; where i d =0 indicates a malfunction in the MESE angle gyroscope status; i d =1 indicates that the MESE angle gyroscope is functioning normally; where φ a1 The φ2 signal is obtained when k1, k2, k3, k4, k5, k6, k7, and k8 are all 0, which is the combined measurement output signal when all feedback is disconnected; φ a2 The φ2 signal is obtained by keeping the original values ​​of k5, k6, k7, and k8 unchanged, while selecting k1, k2, k3, and k4 as 0. In other words, it is the combined measurement output signal with the feedback part of the faulty angle gyroscope disconnected and the others unchanged.

[0107] Case Implementation and Computer Simulation Results Analysis

[0108] In step S10, the magnetometer measurement value of the aircraft yaw angle obtained by the magnetometer is as follows: Figure 2 As shown.

[0109] In step S20, k1 = 0.15, k2 = 0.02, k3 = 0.18, a1 = 0.05, and k4 = 0.4 are selected to obtain the aircraft yaw angle gyroscope measurement values ​​as follows: Figure 3 As shown; the yaw angle gyroscope magnetic force measurement error signal is as follows: Figure 4 As shown; the aircraft yaw rate compensation correction signal is as follows: Figure 5 As shown.

[0110] In step S30, k5 = 0.24, k6 = 0.15, k7 = 0.12, and k8 = 0.3 are selected to obtain the yaw angle combined measurement error signal as follows: Figure 6 As shown; the secondary compensation correction signal for the aircraft's yaw rate is as follows: Figure 7 As shown, the yaw angle magnetic gyroscope combination measurement signal is as follows: Figure 8 As shown.

[0111] In step S40, select i a =1, i b =1, i c =1, i d In the case where = 0, the dual-redundant measurement signal of the aircraft's heading angle is obtained as follows: Figure 9 As shown.

[0112] Depend on Figure 2 It can be seen that the measurement error of a single magnetometer is relatively large, but the error does not show a trend of cumulative drift. Figure 3 It can be seen that a standalone angle gyroscope provides high dynamic accuracy, but it tends to exhibit increasing drift. Figure 8 It can be seen that the final combined measurement results have good dynamic accuracy, and there is no cumulative drift in the error. Figure 9 It can be seen that when the angle gyroscope malfunctions, the yaw angle measurement function can still be improved by using the rate gyroscope and magnetometer. Although the accuracy decreases, the measurement effect is still relatively good. Therefore, the above experimental results show that the method provided by the present invention is effective and has high engineering application value.

Claims

1. A method for measuring the attitude of an aircraft using dual-redundancy technology, characterized by the following steps: Step S10: First, a triaxial accelerometer is installed on the longitudinal axis of the aircraft to measure the components of gravitational acceleration in the x, y, and z axes during uniform motion. These three acceleration components are then normalized to obtain the normalized components of gravitational acceleration in the x, y, and z axes. Next, the angle of gravitational acceleration relative to the y-axis, as well as the aircraft's tilt and azimuth angles, are calculated using inverse trigonometric functions. A Hall effect magnetometer is installed on the longitudinal axis of the aircraft to measure the components of magnetic induction intensity in the x, y, and z axes. Based on the aircraft's tilt and azimuth angles, the x-axis component and y-axis component of the magnetic field strength in the direction parallel to the ground are calculated. Finally, the magnetometer measurement values ​​for the aircraft's yaw angle are obtained through arctangent transformation, as follows: β=arccos(a y ); φ=arctan(a y / a z ); ρ=-arcsin(a x ); H a =H x cosρ+H y sinρsinφ+H z sinρcosφ; H b =H y cosφ-H z sinφ; Where R x R y With R z To represent the components of gravitational acceleration along the x, y, and z axes as measured by a triaxial accelerometer, a x a y With a z ρ represents the unitized components of gravitational acceleration in the x, y, and z axes; β is the angle of gravitational acceleration relative to the y-axis; φ is the tilt angle of the spacecraft; ρ is the azimuth angle of the spacecraft; H x H y H z The components of magnetic flux density along the x, y, and z axes are measured using a Hall effect magnetometer; H a H represents the x-axis component of the magnetic field strength in the direction parallel to the ground. b ψ represents the y-axis component parallel to the ground, and ψ is the magnetometer measurement of the aircraft's yaw angle. Step S20: Install a MEMS rate gyroscope on the aircraft to measure the yaw rate, obtaining the yaw rate measurement signal. Install a MEMS angle gyroscope on the aircraft to measure the yaw angle, obtaining the yaw angle gyroscope measurement value. Then, compare this value with the magnetometer measurement value of the yaw angle to obtain the yaw angle gyroscope magnetic force measurement error signal. Perform a nonlinear transformation on the yaw angle gyroscope magnetic force measurement error signal to obtain the angle gyroscope measurement error nonlinear transformation signal. Then, superimpose the proportionally amplified signal to obtain the angle gyroscope compensation signal. Finally, superimpose the yaw rate measurement signal obtained from the rate gyroscope to obtain the yaw rate compensation correction signal as follows: e1 = ψ - φ1; Δω1=f1+k4e1; ω2=ω1+Δω1; Where φ1 is the measured value of the yaw angle gyroscope, e1 is the yaw angle gyroscope magnetic force measurement error signal; f1 is the nonlinear transformation signal of the angle gyroscope measurement error; k1, k2, k3, and a1 are constant parameters of the nonlinear transformation; k4 is a constant proportional amplification parameter; Δω1 is the angle gyroscope compensation signal; ω2 is the yaw rate compensation correction signal of the aircraft; and ω3 is the yaw rate measurement signal of the aircraft. Step S30: Set the initial value of the yaw angle magnetic gyroscope combined measurement signal to 0, then compare it with the magnetometer measurement signal of the aircraft's yaw angle to obtain the yaw angle combined measurement error signal; then perform a nonlinear transformation on the yaw angle combined measurement error signal to obtain the combined measurement error nonlinear transformation signal; then superimpose the proportionally amplified signal to obtain the angular rate gyroscope compensation signal; then superimpose the aforementioned aircraft yaw angle velocity compensation correction signal to obtain the aircraft yaw angle velocity secondary compensation correction signal; finally, perform integration to obtain the yaw angle magnetic gyroscope combined measurement signal as follows: e2 = ψ - φ2; Δω2=f2+k8e2; ω3=ω2+Δω2; φ2=∫ω3dt; Where φ2 is initially set to 0, e2 is the yaw angle combined measurement error signal; f2 is the combined measurement error nonlinear transformation signal; k5, k6, and k7 are constant parameters of the nonlinear transformation; k8 is the constant proportional amplification parameter; Δω2 is the angular rate gyroscope compensation signal; ω3 is the aircraft yaw angle velocity secondary compensation correction signal; and φ2 is the yaw angle magnetic gyroscope combined measurement signal. Step S40: Based on the states of the three-axis accelerometer, Hall effect magnetometer, MEMS rate gyroscope, and MEMS angle gyroscope, a dual-redundancy design for the aircraft's heading angle is implemented under six different scenarios. Fault-free signals are selected to form the final dual-redundancy measurement signal for the aircraft's heading angle, as follows: Where i a =0 indicates a malfunction in the triaxial accelerometer; i a =1 indicates that the triaxial accelerometer is functioning normally; where i b =0 indicates a malfunction in the Hall effect magnetometer; i b =1 indicates that the Hall effect magnetometer is functioning normally; Where i c =0 indicates a malfunction in the MEMS rate gyroscope; i c =1 indicates that the MEMS rate gyroscope is in normal condition; where i d =0 indicates a malfunction in the MEMS angle gyroscope; i d =1 indicates that the MEMS angle gyroscope is in normal condition; where φ a1 The φ2 signal is obtained when k1, k2, k3, k4, k5, k6, k7, and k8 are all 0, which is the combined measurement output signal when all feedback is disconnected; φ a2 The φ2 signal is obtained by keeping the original values ​​of k5, k6, k7, and k8 unchanged, while selecting k1, k2, k3, and k4 as 0. In other words, it is the combined measurement output signal with the feedback part of the faulty angle gyroscope disconnected and the others unchanged.

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