Research Article  Open Access
Shuangbiao Zhang, Xingcheng Li, Zhong Su, "InFlight SelfAlignment Method Aided by Geomagnetism for Moving Basement of Guided Munitions", Journal of Control Science and Engineering, vol. 2015, Article ID 425698, 8 pages, 2015. https://doi.org/10.1155/2015/425698
InFlight SelfAlignment Method Aided by Geomagnetism for Moving Basement of Guided Munitions
Abstract
Due to powerafterlaunch mode of guided munitions of high rolling speed, initial attitude of munitions cannot be determined accurately, and this makes it difficult for navigation and control system to work effectively and validly. An inflight selfalignment method aided by geomagnetism that includes a fast inflight coarse alignment method and an inflight alignment model based on Kalman theory is proposed in this paper. Firstly a fast inflight coarse alignment method is developed by using gyros, magnetic sensors, and trajectory angles. Then, an inflight alignment model is derived by investigation of the measurement errors and attitude errors, which regards attitude errors as state variables and geomagnetic components in navigation frame as observed variables. Finally, fight data of a spinning projectile is used to verify the performance of the inflight selfalignment method. The satisfying results show that (1) the precision of coarse alignment can attain below 5°; (2) the attitude errors by inflight alignment model converge to 24′ at early of the latter half of the flight; (3) the inflight alignment model based on Kalman theory has better adaptability, and show satisfying performance.
1. Introduction
Guided munitions improved from conventional artillery shells attract more global attention for their outstanding advantages such as low cost, high battle effectiveness, and high stability. In practical applications, the powerbeforelaunch mode is chosen for guided munitions. But this mode frequently induces the electrical circuits systems fixed in guided munitions to shock fiercely, and the electrical circuits systems are damaged to be invalid by momentary impact force. Thus, a more appropriate powerafterlaunch mode, as an ideal power mode, is raised to ensure the safety of the electrical circuits systems. The only drawback of this power mode is that initial attitude cannot be obtained accurately, and this makes it hard for navigation and control system to work effectively and validly. Therefore, to guarantee implement of guidance, it is imperative for the powerafterlaunch mode of guided munitions to exploit an effective approach of initial alignment.
Up to now, many achievements on inflight alignment and alignment of moving basement are undergoing by using Global Position System (GPS), due to the fact that GPS can provide accurate position for carriers in need. A strapdown inertial navigation system (SINS) inflight alignment algorithm proposed by Mei decomposes the attitude quaternion into three parts: inertial rate, separate earth motion, and a constant quaternion. The first two parts are solved based on Microelectromechanical Systems (MEMS) gyros’ measurement and GPS position information, respectively, and the constant quaternion is solved by using request method [1]. A coarse alignment method is presented by Qian et al. to acquire the initial position, velocity, and attitude information with the help of GPS. The method could be implemented in airplanes of high speed and high dynamic flight without any restrictive condition of dynamic process and flight stability [2]. A robust filter and various error models for SINS inflight alignment are presented by Yu and Lee to deal with the uncertainty by linearization of the system that degrades the performance of the filter. This method shows fine performance of estimation if enough time is allowed [3]. A filter designed by Pei et al. can analyze the uncertainty of measurement noise and realize inflight alignment of aircrafts [4]. A technique for constructing the initial alignment system for unmanned aerial vehicles is proposed by Veremeenko and Savel’Ev to improve the accuracy of alignment and reduce the alignment time [5]. An estimation algorithm is proposed by She et al. to calculate initial roll angle of guided munitions under small angle of attack with the help of GPS [6]. An adaptive filtering algorithm of an extended Kalman filter (EKF) combined with innovationbased adaptive estimation is proposed by Fang and Yang to realize initial alignment, and the algorithm introduces the calculated innovation covariance into the computation of the filter gain matrix directly [7]. Meanwhile, an estimation method is presented by Silson for fast initial coarse alignment of a ship’s strapdown inertial attitude reference system using only inertial measurement unit (IMU) measurements for quasistatic alignment and IMU measurements with GPS aiding for movingbase alignment [8]. An improved selfalignment scheme for SINS based on the EKF and augmented measurements is proposed by Acharya et al. Monte Carlo simulations show that the method provides fine azimuth alignment and improves rate of convergence of azimuth attitude error and reduction in sensitivity to gyro biases [9]. A novel method using the GPS measurements is investigated by Ma et al. to help to solve coarse alignment. Specific vectors are constructed with GPS measurements and used for attitude matrix computation [10]. In addition, a nonlinear approach for the inmotion alignment problem of inertial navigation system using Doppler speed measurements with unknown initial vehicle is proposed by Bimal Raj and Joshi, and an unscented Kalman filter for alignment shows better performance than that of extended Kalman filter [11]. Body frame velocity from Doppler Velocity Log (DVL) is used by Li et al. to construct the vector observations for initial alignment of underwater vehicles [12].
At our best knowledge, applications of satellite navigation technique are dependent on authorization, and it is easy to disturb signal of satellites by other strong signals on purpose. Moreover, GPS and Doppler technique are suitable for low dynamical carrier, especially for the carrier with low rolling speed or without rolling. In other words, there is no valid way to implement initial alignment of guided munitions with high spinning speed and powerafterlaunch mode.
In this paper, we represent an inflight selfalignment method of moving basement aided by geomagnetism, which includes a fast inflight coarse alignment method and an inflight alignment model based on Kalman theory. The rest of this paper is organized as follows. In the second section, we exploit a fast inflight coarse alignment method by using gyros, magnetic sensors, and trajectory angle. The third section describes an inflight alignment model by using Kalman theory to estimates the attitude angles. The case study is given in the fourth section. Finally, we conclude this paper in the fifth section.
2. A Fast InFlight Coarse Alignment Method
In general, the initial alignment consists of two stages, the coarse alignment and the fine alignment. The coarse alignment is the basic step to supply initial attitude for the fine alignment. This process is crucial but difficult to estimate attitude angle of munitions with high spinning speed and powerafterlaunch mode. Thus, in this section, we investigate a fast inflight coarse alignment method to calculate initial attitude by analyzing the angular kinematic model and attitude resolving algorithm of magnetic sensors.
For convenience of investigation, we choose eastnorthup frame as navigation frame, and the angular kinematic models are described as [13]where is the roll angle, is pitch angle, is yaw angle, , , and are the corresponding angle changing rates, and , , and are measurement data of MEMS gyros.
By transforming (1), the pitch angle can be formulated by
From (2), it is seen that the symbol of pitch angle is positive in the rising stage of guided munitions, and it is opposite in the falling stage. To implement coarse alignment and ensure enough time for fine alignment, the pitch angle must be determined in the rising stage. So its symbol of pitch angle is positive.
Assuming aerodynamic force and gravity that guided munitions suffer vary slightly, the influence of coning motion is ignored. The pitch angle, yaw angle, and their corresponding changing rates can be regarded as longperiodic variables. Meanwhile, the changing rates of pitch angle and yaw angle vary slowly and even keep invariant during very short time. We can formulate and by the ideal ballistic trajectory angles, inclination angle , and heading angle :where is time interval. Equation (3) can be used to calculate data on line. If some calculations need to be done offline, and can be calculated bywhere , , and are components of velocity in navigation frame, and they can be calculated aswhere , , and are components of location. In fact, the real ballistic trajectory angles are similar to the ideal ones.
Due to low cost of the MEMS sensors, they are familiar in navigation and control system of guided munitions. However, narrow measurement range and poor accuracy of existing MEMS gyros cannot meet measurement requirement of guided munitions with high rolling speed of axis. To solve the problem of calculation of roll angle, we make use of MEMS magnetic sensors that are fixed in guided munitions to help navigation computer calculate attitude.
The attitude algorithm of geomagnetic technology has been investigated for several years. It has high level of resolving precision and satisfying instantaneity, so we can resolve attitude in realtime. The formulations of yaw angle and roll angle are described as [14]where and are, respectively, inclination angle and declination angle of geomagnetic vector and , , and are measurement values of MEMS magnetic sensors in body frame. From (6), we can see that the yaw angle is related to measurement value of the longwise magnetic sensor, and the roll angle is related to lateral measurement values. By combining (2) and (6), the initial attitude can be determined, and the coarse alignment can be realized.
Generally speaking, the firing range of munitions is about several kilometers or scores of kilometers, and the geomagnetic vector changes slightly. Thus, the geomagnetic information can be stored in the navigation computer before launching. It is convenient to extract geomagnetic information when guided munitions are inflight.
3. An InFlight Alignment Model Based on Kalman Theory
Based on the coarse alignment, it is ready to make fine alignment for navigation and control system. In this section, we develop an inflight alignment model based on Kalman theory to estimate angle errors and achieve fine alignment.
The measurement errors of MEMS gyros sensing lateral control channels exist all the time because of their poor measurement precision. So the measurement values of gyros contain actual angular speed and error information, and the attitude resolved by (1) also contains actual attitude and error information. Thus, the actual formulations of (1) can be improved as where , , and are attitude errors, , , and are changing rates errors of attitude, and , , and are measurement errors of gyros.
As long as the attitude errors are enough small, which are assumed less than five degree, the trigonometric functions in (7) can be simplified by using and :
Thus, we can simplify (7) by substituting (8) into (7). Meanwhile, comparing (1) with simplified formulation of (7), the differential equations of attitude errors can be obtained:where , , and are drift errors of equivalent gyros. From (11), it is seen that yaw angle error may be singular if pitch angle is zero.
By dealing with (9)–(11), the state formulation based on Kalman theory can be described aswhere is state transition matrix and defined aswhere is state variable matrix, . is system noise drive matrix, . is system noise matrix, .
The main frame of mind to establish observing matrix is as follows. We transform the measurement values of MEMS magnetic sensors fixed in body frame to navigation frame and transform the geomagnetic vector to navigation frame. Then we regard the discrepancy of comparison between the two parts as the observed quantities to estimate the state values. Considering about attitude errors, the actual transform matrix contains the errors matrix. The observed equation can be formulated bywhere is error matrix, and it can be described as
Each element in is described as
The differences of magnetic components in navigation frame can be obtained by transforming (15):
Attitude errors that are contained in each element in the right side of (18) can be separated from the actual attitude, and the observed equation can be described directly aswhere is observed quantities, , is measurement matrix, , and is measurement noise, . In addition, each element of is formulated as
4. Case Study
For the sake of verification of the inflight selfalignment method, we utilize a real experimental data of a spinning projectile whose navigation computer monitored the whole fight data. The flight experiment is made in the northern hemisphere, and the angles of geomagnetic vector are rad and rad. The total flight time is 65 s, and the interval time is 0.01 s. The trajectory of the spinning projectile is shown in Figure 1, which is obtained by telemetry technology.
The main mind of this process is as follows. We apply the fast coarse alignment method firstly to calculate the initial attitude by using data during 0 s~5 s from Figures 1, 3, and 4 and then apply inflight alignment model based on Kalman theory to realize more accurate selfalignment by using data during 5 s~65 s.
Figure 2 shows measurement data of the MEMS triaxial gyro fixed in the projectile. A high rolling speed in axis of the projectile is 43.33 rad/s at first. Then the rolling speed suffers a fast reduction of speed in front half of the flight, and finally it maintains about 7.35 rad/s that is still regarded as a high spinning speed.
Figure 3 shows the measurement data of the MEMS triaxial magnetic sensor fixed in the projectile. The longwise component of the sensor is negative during rising stage and then becomes positive during falling stage. The transversal components vary alternately like a sinusoid as the projectile spins. In fact they are the components of projection of geomagnetic vector in body frame during the whole flight.
Figure 4 shows the changing speed of the trajectory angles that can be calculated by using the data of Figure 1. From Figure 4, we can see the varying range of is and the varying range of is . They change slowly in the whole time.
Figure 5 provides the attitude by using fast alignment method during 0 s~5 s. It is seen that seven sets are calculated by fast coarse alignment method. It is because not any value of by (2) is satisfied, and we choose the value under square root that is less than 1.
Table 1 shows the attitude error. It is clear that the attitude error is less than five degree and meets the assumption of (8). So the fast coarse alignment method is valid.

We choose the attitude of this projectile at 4.33 s as the initial attitude that is rad, rad, and rad, and the corresponding actual attitude is rad, rad, and rad.
Based on the coarse initial alignment, we continue to make fine alignment by using the inflight alignment mode based on Kalman theory. Figure 6 shows the process of the alignment from 5 s~65 s. It is seen that the attitude error reduces gradually, and the converging speed of rolling angle error is slower than others. When the projectile flies over the peak of the flight, the alignment error has fluctuation near 29 s due to occurrence of singular of . Through the alignment method working, the attitude converges soon. It is apparent that the model has a better adaptability.
Figure 7 shows the partial magnification of attitude error. At 34 s, the attitude error is less than 0.01 rad. At 39.26 s, the roll angle error is 0.006 rad, pitch angle error is zeros, and yaw angle error is −0.007 rad. It is clear that the maxium error is about 24′, which is high precision level for spining munintions. Although the pitch angle error is increasing after 39.26 s, the final error is 0.0020 rad at 65 s. At the same time, the roll angle error is 0.0067 rad, and yaw angle error is 0.0021 rad. It is apparent that fine alignment by the inflight alignment model based on Kalman theory shows satisfying performance, which is achieved finalliy.
5. Conclusions
To solve the problem of initial alignment of guided munitions, this paper introduces an inflight selfalignment method aided by geomagnetism composed of a fast coarse alignment method and an inflight alignment model by Kalman theory. The fast inflight coarse alignment method is developed by using gyros, magnetic sensors, and trajectory angle. The inflight alignment model by Kalman theory is developed through investigation of the measure errors and attitude errors, and this model regards attitude errors as state variables and geomagnetic components in navigation frame as observed variables. Finally, we utilize a data of a spinning projectile to verify the performance of alignment method. The results show that (1) the precision of coarse alignment can attain below 5°; (2) the attitude errors by inflight alignment model converge to 24′ at early of the latter half of the flight; (3) the inflight alignment model based on Kalman theory has better adaptability and show satisfying performance. Therefore, the inflight selfalignment method aided by geomagnetism can meet the requirement of guided munitions with high spinning speed and powerafterlaunch mode. Furthermore, this method is also suitable for carriers of low rotating speed. In the future, the guidance and control for guided munitions of the high rotating speed will be investigated.
Conflict of Interests
The authors declare that there is no conflict of interests regarding the publication of this paper.
Acknowledgment
The grant support from the National Science Foundation of China (nos. 61201417 and 11202023) is greatly acknowledged.
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Copyright
Copyright © 2015 Shuangbiao Zhang et al. This is an open access article distributed under the Creative Commons Attribution License, which permits unrestricted use, distribution, and reproduction in any medium, provided the original work is properly cited.