Attitude Determination Using a MEMS-Based Flight Information Measurement Unit

Obtaining precise attitude information is essential for aircraft navigation and control. This paper presents the results of the attitude determination using an in-house designed low-cost MEMS-based flight information measurement unit. This study proposes a quaternion-based extended Kalman filter to integrate the traditional quaternion and gravitational force decomposition methods for attitude determination algorithm. The proposed extended Kalman filter utilizes the evolution of the four elements in the quaternion method for attitude determination as the dynamic model, with the four elements as the states of the filter. The attitude angles obtained from the gravity computations and from the electronic magnetic sensors are regarded as the measurement of the filter. The immeasurable gravity accelerations are deduced from the outputs of the three axes accelerometers, the relative accelerations, and the accelerations due to body rotation. The constraint of the four elements of the quaternion method is treated as a perfect measurement and is integrated into the filter computation. Approximations of the time-varying noise variances of the measured signals are discussed and presented with details through Taylor series expansions. The algorithm is intuitive, easy to implement, and reliable for long-term high dynamic maneuvers. Moreover, a set of flight test data is utilized to demonstrate the success and practicality of the proposed algorithm and the filter design

Human motion tracking based on complementary Kalman filter

Miniaturized Inertial Measurement Unit (IMU) has been widely used in many motion capturing applications. In order to overcome stability and noise problems of IMU, a lot of efforts have been made to develop appropriate data fusion method to obtain reliable orientation estimation from IMU data. This article presents a method which models the errors of orientation, gyroscope bias and magnetic disturbance, and compensate the errors of state variables with complementary Kalman filter in a body motion capture system. Experimental results have shown that the proposed method significantly reduces the accumulative orientation estimation errors

An Application of UAV Attitude Estimation Using a Low-Cost Inertial Navigation System

Unmanned Aerial Vehicles (UAV) are playing an increasing role in aviation. Various methods exist for the computation of UAV attitude based on low cost microelectromechanical systems (MEMS) and Global Positioning System (GPS) receivers. There has been a recent increase in UAV autonomy as sensors are becoming more compact and onboard processing power has increased significantly. Correct UAV attitude estimation will play a critical role in navigation and separation assurance as UAVs share airspace with civil air traffic. This paper describes attitude estimation derived by post-processing data from a small low cost Inertial Navigation System (INS) recorded during the flight of a subscale commercial off the shelf (COTS) UAV. Two discrete time attitude estimation schemes are presented here in detail. The first is an adaptation of the Kalman Filter to accommodate nonlinear systems, the Extended Kalman Filter (EKF). The EKF returns quaternion estimates of the UAV attitude based on MEMS gyro, magnetometer, accelerometer, and pitot tube inputs. The second scheme is the complementary filter which is a simpler algorithm that splits the sensor frequency spectrum based on noise characteristics. The necessity to correct both filters for gravity measurement errors during turning maneuvers is demonstrated. It is shown that the proposed algorithms may be used to estimate UAV attitude. The effects of vibration on sensor measurements are discussed. Heuristic tuning comments pertaining to sensor filtering and gain selection to achieve acceptable performance during flight are given. Comparisons of attitude estimation performance are made between the EKF and the complementary filter

A Kalman Filter Based Attitude Heading Reference System Using a Low Cost Inertial Measurement Unit

This paper describes, the development of a sensor fusion algorithm-based Kalman lter ar- chitecture, in combination with a low cost Inertial Measurement Unit (IMU) for an Attitude Heading Reference System (AHRS). A low cost IMU takes advantage of the use of MEMS technology enabling cheap, compact, low grade sensors. The use of low cost IMUs is primar- ily targeted towards Unmanned Aerial Vehicle (UAV) applications due to the requirements for small package size, light weight, and low energy consumption. The high dynamics nature of smaller airframes, coupled with the typical vibration induced noise of UAVs require an e cient, reliable, and robust AHRS for vehicle control. To eliminate the singularities at 90 on the pitch and roll axes, and to keep the computational e ciency high, quaternions are used for state attitude representation

Fault Tolerant Flight Control: An Application of the Fully Connected Cascade Neural Network

The endurance of an aircraft can be increased in the presence of failures by utilising flight control systems that are tolerant to failures. Such systems are known as fault tolerant flight control systems (FTFCS). FTFCS can be implemented by developing failure detection, identification and accommodation (FDIA) schemes. Two of the major types of failures in an aircraft system are the sensor and actuator failures. In this research, a sensor failure detection, identification and accommodation (SFDIA); and an actuator failure detection, identification and accommodation (AFDIA) schemes are developed. These schemes are developed using the artificial neural network (ANN). A number of techniques can be found in the literature that address FDIA in aircraft systems. These techniques are, for example, Kalman filters, fuzzy logic and ANN. This research uses the fully connected cascade (FCC) neural network (NN) for the development of the SFDIA and AFDIA schemes. Based on the study presented in the literature, this NN architecture is compact and efficient in comparison to the multi-layer perceptron (MLP) NN, which is a popular choice for NN applications. This is the first reported instance of the use of the FCC NN for fault tolerance applications, especially in the aerospace domain. For this research, the X-Plane 9 flight simulator is used for data collection and as a test bed. This simulator is well known for its realistic simulations and is certified by the Federal Aviation Administration (FAA) for pilot training. The developed SFDIA scheme adds endurance to an aircraft in the presence of failures in the aircraft pitch, roll and yaw rate gyro sensors. The SFDIA scheme is able to replace a faulty gyro sensor with a FCC NN based estimate, with as few as 2 neurons. In total, 105 failure experiments were conducted, out of which only 1 went undetected. In the developed AFDIA scheme, a FCC NN based roll controller is employed, which uses just 5 neurons. This controller can adapt on-line to the post failure dynamics of the aircraft following a 66\% loss of wing surface. With 66\% of the wing surface missing, the NN based roll controller is able to maintain flight. This is a remarkable display of endurance by the AFDIA scheme, following such a severe failure. The results presented in this research validate the use of FCC NNs for SFDIA and AFDIA applications

Design of a MEMS-Based Flight Information Measurement Unit for UAV Application

[[abstract]]Obtaining precise attitude information is essential for aircraft navigation and control. This paper presents the results of the attitude determination using an in-house designed low-cost MEMS-based flight information measurement unit. This study proposes a quaternion-based extended Kalman filter to integrate the traditional quaternion and gravitational force decomposition methods for attitude determination algorithm. The proposed extended Kalman filter utilizes the evolution of the four elements in the quaternion method for attitude determination as the dynamic model, with the four elements as the states of the filter. The attitude angles obtained from the gravity computations and from the electronic magnetic sensors are regarded as the measurement of the filter. The immeasurable gravity accelerations are deduced from the outputs of the three axes accelerometers, the relative accelerations, and the accelerations due to body rotation. The constraint of the four elements of the quaternion method is treated as a perfect measurement and is integrated into the filter computation. Approximations of the time-varying noise variances of the measured signals are discussed and presented with details through Taylor series expansions. The algorithm is intuitive, easy to implement, and reliable for long-term high dynamic maneuvers. Moreover, a set of flight test data is utilized to demonstrate the success and practicality of the proposed algorithm and the filter design.[[notice]]補正完畢[[journaltype]]國外[[booktype]]紙本[[countrycodes]]GB