ABSTRACT. Domestic production of general aviation (GA) aircraft has declined markedly in recent years and an effort is underway to revitalize this industry by ...
IMPLEMENTATION OF AN ATTITUDE-HEADING REFERENCE SYSTEM STATE ESTIMATE FILTER USING DISSIMILAR SENSOR TYPES AND SAMPLING RATES PETER J. GORDER, Ph.D., J. GARTH THOMPSON, Ph.D., and HOSAM FATHY gorder, jgt, hosam @ksu.edu Department of Mechanical and Nuclear Engineering Kansas State University 302 Rathbone Hall, Manhattan, KS 66506
ABSTRACT Domestic production of general aviation (GA) aircraft has declined markedly in recent years and an effort is underway to revitalize this industry by opening new markets for aviation, shifting the perception of general aviation as recreation to general aviation as a travel option. This can be achieved only by providing a product that is safer, easier to fly in all weather conditions, and no more expensive than today’s low end general aviation aircraft. The critical path to this goal is the development of an inexpensive yet reliable advanced flight system, incorporating various levels of automation and control augmentation to simplify the operation of the aircraft. One key element in the control of an aircraft is the ability to accurately deduce the aircraft’s attitude and angular velocity. The work reported in this paper describes the development of a state estimate filter to accept measurements from dissimilar sensor types sampled at different rates and produce a quasi-continuous best estimate of the vehicle attitude and angular velocity. A novel configuration incorporating Global Positioning System based sensors, inexpensive tuning fork gyros, and distributed accelerometers has been developed. This paper extends the theoretical work presented previously at the IASTED International Conference on Control and Applications, 1998. It describes the hardware implementation of the system and discusses the validation testing procedure.
KEYWORDS Attitude-Heading Reference System, Advanced Flight Control Systems, GPS-based sensors, State Estimation
INTRODUCTION The recovery of the General Aviation (GA) industry, which has declined dramatically in the past 25 years, will rely on the opening of new markets for aviation, shifting
the perception of general aviation as recreation to general aviation as a travel option. This can be achieved only by providing a product that is safer, easier to fly in all weather conditions, and no more expensive than today’s low end general aviation aircraft. Indeed, the NASA Aeronautics Advisory Committee Task Force on General Aviation concluded that “without innovation enabled by technological advancement, GA within the United States will fail to respond to opportunities for expanded use and is destined to continue its decline” [1]. The critical path to this goal is the development of an inexpensive yet reliable advanced flight system, incorporating various levels of automation and control augmentation to simplify the operation of the aircraft. One key element in the control of an aircraft is the ability to accurately deduce the aircraft’s attitude and angular velocity. The work documented in this paper extends the theoretical work towards the development of a novel system that incorporates a set of dissimilar sensors sampled at different rates to provide a reliable and reliably accurate quasi-continuous estimate of the aircraft’s attitude and angular velocity [2 ]. It describes the hardware implementation of the system and discusses the validation testing procedure.
BACKGROUND To acheive the desired ease of operation in a GA aircraft, it has been proposed by the Advanced General Aviation Technology Experiments (AGATE) program, a joint NASA/FAA/Industry consortium, that a decoupled flight control system (DCFCS) be implemented along with advanced display formats. It has been demonstrated that such a system greatly simplifies the operation of an aircraft for novice pilots [3]. Required for the functioning of a DCFCS is accurate attitude and attitude rate information. Developing means of acquiring this information is an active research area including work involved in the discovery of new types of sensors that can be used to
ascertain attitude information, such as Global Positioning System (GPS) based sensors [4]. GPS based attitude determination, however, is limited in that the sample rates are currently limited to ten (10) samples per second in the more expensive units and one to two samples per second in more moderately priced units, too slow to use this information in a feedback control loop. Inexpensive rate gyroscopes contain an unsteady bias in the angular velocity measurement that, if integrated to estimate attitude, would compound into unacceptable inaccuracies in the attitude measurement. A novel configuration incorporating these sensors augmented with accelerometers distributed about the airframe and used to measure the angular and translational accelerations of the aircraft has been developed [5], and integral to the functioning of the system is a filtering algorithm that mitigates the limitations of the different sensor types and provides a reliable and reliably accurate quasi-continuous estimate of the aircraft’s attitude and angular velocity[2] (Fig. 1). The remainder of this paper describes the hardware implementation of this system concept, along with the validation testing of the system. Calculated Angular Acceleration (Accelerometers)
Calculated Quaternion Representation of Aircraft Attitude (GPS-based Sensors)
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scale jet propelled remotely piloted aircraft named Aladdin. This aircraft was constructed as part of a student design competition by a team of students from Kansas State University, the University of Kansas, Pittsburgh State University, and Wichita State University. It has been demonstrated to be flight worthy in radio-controlled test flights. It will be equipped with the AHRS described herein to demonstrate not only the AHRS, but ultimately a DCFCS as well.
IMPLEMENTATION
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Figure 2. Laboratory Test Article
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The underlying motivation behind the attitude reference system described herein was to acheive high reliability at a reasonable cost. Considering the cost constraint, it was proposed that a system consisting of a distributed set of dissimilar sensors be employed such that the full compliment of sensors provided redundant information. To this end, a sensor package including five (5) GPS-based sensors distributed about the aircraft, a set of inexpensive and, correspondingly, less accurate rate gyroscopes mounted close to the aircraft center of gravity (CG), and a set of four (4) 3-axis accelerometers, also distributed about the aircraft, was designed. The layout of
Figure 1. State Estimate Filter
SYSTEM TEST PLATFORMS To validate the proposed attitude and attitude rate estimation system, hereafter refered to as the attitudeheading reference system (AHRS), two test platforms were devised and constructed. First, a three-axis gimballed platform was constructed to test the AHRS in a laboratory setting (Fig. 2). The platform’s attitude is driven by three servo motors controlled with a joy stick and an encoder is mounted on each axis of rotation to accurately measure the actual platform attitude so that comparisons can be made with the estimates from the AHRS. In addition, a flight test platform has been constructed. Shown in Fig. 3 is a wireframe diagram of the 1/4
Figure 3. 1/4 Scale “Aladdin” Model Aircraft
RESULTS Validation testing for the AHRS in general, and the state estimate filter specifically, has been laid out in three phases. The first phase of the validation testing consisted of off-line numerical simulation tests, where the sensor data was simulated, including offset bias and noise. Figure 5 presents a sample of the attitude estimation results. The offset between the measured and actual attitudes indicates an artificially added offset bias to the attitude determined using the GPS sensor data. The jagged nature of this signal is indicative of the 10 Hz data rate for the GPS receivers.. There were significant biases and noise added to the “measured” angular velocities and accelerations as well. Yet, note how well the filter smooths the attitude estimate and tracks the attitude measurement. These simulation tests include worst case bias and noise conditions, but no sensor failures. additional testing will be conducted to test the performance degradation in the face of individual sensor failures. The second phase of testing involves the implementation of the system on the laboratory test article to demonstrate the performance of the AHRS using actual sensor measurements. Again, sensor failures will be induced to test the performance degradation. Although intended to be complete at this point in time, hardware difficulties have delayed progress. It is anticipated that the laboratory testing will be completed by early spring. The final phase of testing will be on board the Aladdin model aircraft. It is intended that these test
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the sensors on the Aladdin test aircraft is coincident with the sensor layout on the laboratory test article (Fig. 2) where the elevated section on the test article platform corresponds to the elevated horizontal tail on Aladdin. The aircraft orientation on the platform is diagonal to provide the greatest wingspan possible on the test article. The GPS antennae are placed on the wing tips, the tops of the twin vertical tails, and in the fuselage, providing significant redundancy in the information necessary for the computation of attitude. The CMC ALLSTAR GPS receivers provide carrier phase information, used to deduce attitude, at 10 Hz.. The three (3) orthogonally mounted tuning fork rate gyros, part number QRS 14-64109 from Systron Donner Inertial Division, are placed in the fuselage. These analog devices, while lacking the accuracy that would be required for use as a primary source of attitude information, are sampled at 100 Hz to “fill in” between GPS attitude calculations. Finally, the 3axis accelerometer packages, part number ADXL05EM3 from Analog Devices, Inc., are located in the nose, tail, and wing tips. These analog devices are also sampled at 100 Hz, and can be used to deduce translational as well as rotational accelerations of the aircraft [5]. All of these sensors provide the inputs for the state estimate filtering algorithm shown in Fig. 1.
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Time Figure 5. Validation Test Results - Roll Attitude Estimation flights be purely demonstration of the AHRS, and there are no plans for inducing sensor failures artificially. Test flights are anticipated in the late spring of this year.
CONCLUSIONS An innovation critical to the development of safer, easier to fly general avaition aircraft, is a reliable, accurate, yet inexpensive attitude-heading reference system. Such a system is proposed herein consisting of an array of dissimilar sensors providing inputs to a state estimation filter that mitigates the limitations of each. Simulation testing has verified the viability of this system, and hardware testing is anticipated in the coming months.
REFERENCES [1] Anon., “Industry, NASA, the FAA, and Universities Join Forces to Revitalize the General Aviation Industry,” The AGATE Flier c/o Office of Communications, Research Triangle Institute, August 1994. [2] Gorder, Peter J., J. Garth Thompson, and Hosam Fathy, “State Estimate Filtering using Dissimilar Sensor Types and Sampling Rates Applied to an AttitudeHeading Reference System,” IASTED International Conference on Control and Applications, Honolulu, HI, Aug. 11-14, 1998. [3] Stewart, Eric C., “A Piloted Simulation Study of Advanced Controls and Displays for General Aviation Airplanes,” NASA-TM-111545, Jan. 1994. [4] C.E. Cohen, “Attitude Determination Using GPS,” Ph.D. Dissertation, Dept. of Aeronautics and Astronautics, Stanford University (Dec. 1992). [5] Gorder, P.J., J.G. Thompson, W.H. Thompson, H. Fathy, and E. Schultz, “Design and Implementation of a State-Estimation System Utilizing a Low-Cost Distributed Sensor Array,” SAE World Aviation Congress, Anaheim, CA, Sept. 28-30, 1998.
