INTEGRATION OF INSTRUMENTATION PAYLOAD ON A REMOTELY PILOTED AERIAL VEHICLE

INTEGRATION OF INSTRUMENTATION PAYLOAD ON A REMOTELY PILOTED AERIAL VEHICLE R. Shivkumar*, H. Arya**, K. Sudhakar*** Department of Aerospace Engineering, IIT Bombay, Powai, Mumbai 4 76 National Systems Conference - 2, Systems Society of India, December 7-9, 2, Bangalore ABSTRACT Several academic institutions are presently engaged in research activities in connection with remotely piloted aerial vehicles. However, there is very little work of this nature undertaken by academic institutions within the country. The Department of Aerospace Engineering, IIT Bombay has decided to make inroads in this field and has undertaken a project to implement an onboard miniature instrumentation system for remotely piloted aerial vehicles. It is envisaged that such an activity will be a precursor to pursuit of design, construction and operation of remotely piloted vehicles with autonomous capabilities in the future. Also availability of such a platform at an academic institution will expose students to the real world problems of dealing with systems and related problems. The project aims to construct an aerial vehicle, in this case a radio controlled aircraft, and integrate a miniature generic instrumentation system onboard. Eight parameters have been identified for measurement. The package can be enhanced to include more parameters if required at a later date. Experimental work that was undertaken as part of this work includes, design, fabrication and wind tunnel calibration of pitot and static probes, a wind vane system for measurement of angle of attack and an interface unit to facilitate recording of aircraft pitch rate. A total of nine flights were possible and flight data could be recorded. The data quality was satisfactory except for accelerometer data, which was too noisy to be usable. ------------------------------------------------- *MTech Student ** Research Scientist, *** Professor Unfortunately towards the end of the ninth sortie the aircraft was destroyed in a crash. A second aircraft was built thereafter and is now ready for flight tests. The problem of noise in accelerometer data has been overcome to a great extent by implementation of filters. The project is scheduled to conclude by end of December 2. INTRODUCTION The Department of Aerospace Engineering, IIT Bombay, intends to enhance its flight mechanics and flight control activities by developing remotely piloted aerial vehicles as experimental platforms. It is envisaged that such an activity will help gain valuable insight into design, fabrication and operation of such vehicles. Further, these vehicles may serve as experimental flight mechanics platforms and can be used to impart valuable practical training to students of Aerospace Engineering. Specifically this activity will help build knowledge and expertise in the following areas. (i) Design and construction of unmanned aerial vehicles. (ii) Sensor selection, calibration and integration of miniature data acquisition systems on small aerial vehicles. (iii) Development of system identification methods. (iv) Design and implementation simple controllers, and pave the way for more ambitious controllers perhaps integrated with onboard GPS for autonomous flight and autonomous safe recovery and landing. 1

(v) (vi) (vii) Design and fabrication of airframe components using lightweight composite materials. Design and development of custom electronic packages like, data acquisition and flight control hardware (in collaboration with sister faculties). Carriage of special payloads (e.g. remote sensing systems, communications relay equipment, environmental monitoring systems etc). Currently there is large interest being shown by several Universities in similar activities. At the University of Sydney, Australia, research in Unmanned Aerial Vehicles (UAV s) has produced promising results towards the development of fully autonomous capabilities. They have implemented a GPS based autopilot system on their aerial vehicle christened Ariel, which is capable of autonomous flight along a pre-determined trajectory [1]. Stanford University s DragonFly project is aiming to equip a large radio controlled model airplane with avionics. The autopilot of the vehicle will then be able to autonomously direct the plane through a series of demanding aerobatic maneuvers such as barrel rolls and loops. To accomplish this objective the vehicle autopilot will use data from an onboard GPS and inertial measurement system [2]. AIMS OF THIS PROJECT (i) Identification of sensors and configuration of a generic data acquisition system for implementation on small aerial vehicles. (ii) Fabrication of a suitable model aircraft to carry the instrumentation payload. (iii) Flight test the vehicle and acquire flight data under steady state and maneuvering flight. (iv) Try to excite different modes of the aircraft (if possible) and deduce (v) information on aircraft characteristics. Implement a rudimentary flight control system for maintaining demanded airspeed and/or altitude. AERIAL VEHICLE A proven radio controlled aircraft design has been selected for this purpose. The model identified is called the KADET designed by Claude McCullough (SIG Manufacturing Company USA). The aircraft, which has a stable high wing configuration with a wing span of 1.46 m is particularly suited for the envisaged role due to its roomy fuselage that is capable of accommodating the instrumentation payload. The aircraft is controlled remotely by a ground-based pilot. A standard model aircraft radio control system has been used for the purpose. The vehicle has controls for operating throttle, ailerons, elevator and rudder (coupled with nose wheel steering to facilitate ground handling). The aircraft has a 7.5 cc single cylinder piston engine as its powerplant. However, the aircraft needed the following modifications to facilitate the integration of instrumentation. Fig. 1 shows the completed aerial vehicle. (i) The wing span had to be increased by.15 m to maintain design wing loading (of 5.2 kg/m 2 ) owing to an increase in all up weight due to the instrumentation payload. (ii) The fuselage width was increased by approximately 2 cm to facilitate installation and easy access to the instrumentation payload. (iii) The throttle, elevator and rudder actuation servos were relocated at the rear of the fuselage to create space for the payload. INSTRUMENTATION The following parameters have been identified for measurement on the aircraft for in-flight recording of flight data. (i) Angle of Attack (α). (ii) Indicated Airspeed (V). 3

(iii) (iv) (v) (vi) Pressure Altitude (h). Three axes Acceleration (Nx, Ny and Nz). Elevator Position (δe). Pitch Rate (q). fabricated and installed on the aircraft wing tip, well outside the influence of the propeller slipstream. Details of the arrangement are elaborated under experimental work. The installation of the vane assembly on the port wing of the aircraft can be seen in Fig. 3. AOA Vane Pitot Static Probes IAS and Altitude Sensors Figure 1. Instrumented aerial vehicle Selection of the various components of the instrumentation package was based on aspects like small size, low weight, low operating voltage and low power consumption and low cost (preferably). The block schematic of the instrumentation is shown in Fig. 2. To be implemented Figure 2. Instrumentation Block Schematic Angle of Attack (a): For measurement of angle of attack a wind vane assembly coupled to a precision potentiometer was Figure 3. Installation of AOA vane and pitot static probes on port wing. Indicated Airspeed (V): For the measurement of indicated airspeed of the aircraft, a differential pressure transducer model 163PC1D75 [4] of M/s Honeywell Sensing and Control, Canada has been used (refer Fig. 3). The operating range of the device is ± 2.5 inches of water and corresponds to an equivalent airspeed of approximately 32 m/s (which is considered adequate to cover the range of operating speeds of the aircraft). The sensor has been coupled to pitot and static probes. Details regarding fabrication and calibration of the pitot and static probes is described under experimental work. Pressure Altitude Sensor (h): A differential pressure transducer, model 142PC1D [4], with a range upto 28 inches ( to 71.12 cm) of water has been used. The range corresponds to an altitude band (at sea level) of approximately 87 m. One of the ports of the transducer has been connected to a sealed tank. An arrangement to equalize the pressure of this tank with respect to the ambient atmospheric pressure on ground before each flight has been incorporated. The other port is open to the atmosphere via the static probe. The differential pressure measured by this system is therefore 4

indicative of the height of the aircraft above ground. This arrangement will have adequate short-term accuracy for typical flight durations of 15 to 2 minutes. The schematic of the altitude measurement arrangement is shown in Fig. 4. Figure 4. Schematic arrangement for altitude measurement Tri-axial Accelerometer (Nx, Ny, Nz): For the measurement of accelerations along all three axes of the aircraft, a tri-axial accelerometer has been used. The transducer, model ADXL15-EM3, is manufactured by M/s Analog Devices, USA [5]. The operating range of the device is ± 4 g along all three axes. The installation of the accelerometer in the aircraft fuselage (close to CG location) is shown in Fig. 5. Elevator Position Sensor (de): For the measurement of elevator position, a potentiometric transducer has been employed. The elevator has been coupled to the transducer using a suitable linkage. The sensor installation is shown in Fig. 6. Pitch Rate Sensor (q): Efforts to identify a miniature rate gyro for measurement of pitch rate did not fructify. However, during market survey a piezoelectric gyro based flight stabilization unit was identified. This unit is specially designed for use on radio controlled aircraft as a single axis stabilization system and is not intended for use as an angular rate sensor. However, it was decided to adapt this system for measurement of pitch rate. An electromechanical interface was developed for this purpose and the scheme is elaborated under experimental work. Figures 5 and 7 show the gyro sensor installation and interface respectively. Potentiometer Accelerometer Figure 6. Elevator potentiometer installation Gyro Rate Gyro Figure 5. Accelerometer and rate gyro installation Interface Figure 7. Interface for Gyro Unit Data Acquisition Unit: An off-the-shelf miniature data acquisition unit was identified for recording flight data onboard the aircraft. The unit is the Model TFX 11 5

data logger from Onset Computer Corporation, USA [6]. The device can accept upto 19 channels of analog data and operate at a maximum sampling rate of 3.2 kilo samples/sec. Onboard memory consists of 128 kilo bytes of RAM and 472 kilobytes of flash EEPROM, which can be used for data as well as programs. In addition, as a part of this project, avionics sub-system requirements have been developed for work to be sub-contracted. Based on these requirements, design and fabrication of the following sub-systems was undertaken by the Department of Electrical Engineering, IIT Bombay, through student projects. (i) Miniature data logging system. (ii) Miniature flight control system. This was done in a bid to fulfil long term objectives of building in-house expertise. It was felt that the above activity would contribute to achieving self-sufficiency in building aerial vehicles and related systems within this Institute. These systems are currently undergoing laboratory tests and will be used on the vehicle after being proved flight worthy. EXPERIMENTAL WORK The following experimental work was undertaken. (i) Fabrication of wind vane assembly for measurement of angle of attack. (ii) Fabrication of pitot static probes for measurement of indicated airspeed and altitude. (iii) Fabrication of a stub wing to test the above in a wind tunnel. (iv) Fabrication of an interface for the gyro system to enable recording of pitch rate. Angle of Attack Vane (AOA): For measurement of angle of attack, it was decided to mount the vane assembly on the wing tip well outside the influence of propeller slipstream. However, to emancipate the vane from the effects of wing tip vortices the vane was mounted on an extended shaft to maintain a reasonable separation between the vane and the wing tip. It was decided to try out the arrangement on a stub wing and calibrate the α vane in the wind tunnel. The vane assembly mounted on the stub wing is shown in Fig. 8. The vane shaft is supported with a ball bearing unit to reduce errors due to friction. AOA Vane Probes Figure 8. Stub wing with AOA vane and pitot static probes Pitot and Static Probes: Standard pitot static probes consist of a coaxial arrangement of two tubes. The inner one measures total pressure and is open to the direction of airflow. The outer tube that measures static pressure has holes tangential to the airflow. This arrangement though compact, presented certain manufacturing difficulties due to the co-axial arrangement of tubes. It was decided to fabricate a simpler arrangement consisting of two separate tubes, one each for total and static pressure measurement. Four diametrically opposite holes were drilled on the surface of the static probe. Stainless steel tubes of 11.5 cm length (exposed length) and 3.2 mm diameter were used for the purpose. The schematic drawing of the static tube is shown in Fig. 9. The pitot pressure probe was simply a length of tube open to the airflow. To minimize errors in measurement of static pressure, it is recommended that the ratio of x to d (refer Fig. 9) should be in the region of 1 and above (reference [3] pp. 53). A ratio of ten was implemented during fabrication of the static probe. Ideally, the probe should be as long as possible, to minimize the influence of the probe supports (in this case the aircraft wing) on the measured value of static pressure. However, it was been decided to fix the length at 11.5 cm (4.5 ) more for convenience rather than 6

any other reason. The installation of the probes on the stub wing to enable wind tunnel testing is shown in Fig. 8. Figure 9. Static Tube Schematic Stub Wing: The wind tunnel test section of, two feet by two feet (61 cm), dictated the size of the stub wing. To ensure that an aspect ratio of at least four is maintained the actual aircraft wing was scaled down to 73%. Thus wing chord of 29.2 cm was reduced to 21.3 cm on the stub wing. The span was fixed at 42 cm, which together with its image in the tunnel wall yielded an aspect ratio of approximately four. The schematic arrangement of the stub wing in the wind tunnel is shown in Fig. 1. The stub wing was constructed with balsa wood and covered with commercially available polyurethane self-adhesive heat shrink film (Monokote). An aluminum tube was embedded spanwise into the stub wing. This enabled the wing angle of attack to be varied from outside the tunnel by rotating the tube. All pressure connections and electrical wires of the pitot static probes and the angle of attack vane were brought out of the wind tunnel through this tube. The arrangement of the stub wing in the wind tunnel is shown in Fig. 11. Figure 11. Stub Wing in Wind Tunnel Wind Tunnel Tests: A platform was constructed on the aluminum tube embedded in the stub wing and extending out of the wind tunnel wall. The platform was rigged to have zero inclination with respect to the wing chord. Thus, angle of attack of the stub wing was measurable at all times by placing an inclinometer on this platform. The wing incidence was varied between 6 to 16 during the tests. All pressures were measured in mm of water using a micromanometer. The airspeed measured by the pitot static test probes was plotted against the tunnel speed for different inclinations of the wing (wind tunnel speed was varied between 5 m/s to 3 m/s). The plot is shown in Fig. 12. Figure 1. Schematic Arrangement of Stub Wing in Wind Tunnel Figure 12. Measured Velocity versus Tunnel Velocity for Different AOA 7

It can be seen that the slope of these plots increases with increase in angle of attack. Ideally, they should have coalesced, however they are close and linear. The percentage error in measured airspeed was then plotted versus angle of attack and is shown in Fig. 13. It is evident from the plot that the errors are within +6% and 3%. Therefore, the probe arrangement was considered acceptable for measurement of airspeed on the actual aircraft. angular rates. The block schematic of the system is shown in Fig. 14. Figure 14. Schematic Arrangement for Rate Gyro Calibration FLIGHT TESTS Figure 13. Percentage error in measured velocity versus AOA Gyro Interface: To enable recording of angular rate using the available gyro system an electromechanical interface was developed. The gyro system consists of a piezoelectric sensor and an amplifier unit. The amplifier is capable of driving a servo, which in turn can move a control surface in the appropriate direction to annul the effect of disturbances. The system can be rigged to act as a roll, pitch or yaw damping system. In the present application however, a precision potentiometer was coupled to the output shaft of a servo, which in turn was connected to the gyro amplifier. The arrangement together with the sensor unit was then placed on an angular rate calibration turntable. The servo movement, which is proportional to the angular rate, was then picked off by the potentiometer and its output was calibrated for different Flight tests were conducted during the end of September 2. During the tests level fights at different speeds were flown in addition to simple maneuvers like loops and stalls. Also pulse inputs were applied to the elevator in a bid to excite short period pitch oscillations. A total of nine flights were possible. The first three flights were flown essentially to get hands-on experience in flying and trimming the aircraft. In these flights the data acquisition unit was removed as a precaution. Thereafter instrumented flights were attempted. On examination of the flight data it was immediately apparent that the accelerometer data was too noisy to be usable. Quality of the other recorded parameters was found to be satisfactory. Data for one level flight and one loop are presented in figures 15 and 16. From figure 16 the loop diameter can be determined (approx. 26 m). In the ninth flight, prior to landing, while executing of a turn to port, there was a severe and uncommanded pitch down by the aircraft at a height of approximately 1 to 12 meters. The aircraft thereafter crashed in a nose down attitude and the airframe was completely destroyed. The reason for the crash could not be established. Examination of the wreck did not reveal any clues. In fact the elevator actuation servo was serviceable even after the crash. It is felt that the uncommanded pitch down may have been due to a 8

elev deg aoa deg q deg/sec alt m elev deg aoa deg alt m momentary radio interference problem and insufficient height prevented safe recovery of the aircraft. However, the onboard radio control equipment, sensors package and data logger survived undamaged. It decided to build a second platform and continue the project. 2-2 21 22 23 24 25 26 27 5 21 22 23 24 25 26 27 1-1 IAS km/hr1 21 22 23 24 25 26 27 5-5 21 1 22 23 24 25 26 27 5 21 22 23 24 25 26 27 Time (sec) 2 Figure 15. Level flight data. -2 525 525.5 526 526.5 527 527.5 528 528.5 529 529.5 53 5 525 525.5 526 526.5 527 527.5 528 528.5 529 529.5 53 1-1 IAS km/hr1 5 525 525.5 526 526.5 527 527.5 528 528.5 529 529.5 53 525 1 525.5 526 526.5 527 527.5 528 528.5 529 529.5 53 q deg/sec1 5 525 525.5 526 526.5 527 527.5 528 528.5 529 529.5 53 Time (sec) Figure 16. Loop data SECOND AERIAL VEHICLE Construction of the second instrumented platform has been is completed. Certain modifications were incorporated in this vehicle to enhance ease of operation. The prominent changes are: (i) Modularizing the instrumentation package to enable quick installation and removal of the package (refer Fig. 17). (ii) Pitot static probes have been made removable. This feature has made ground handling of the aircraft easier and will allow field replacement of damaged probes if required. (iii) The elevator area has been reduced by 3%. This is because elevator travel for the entire speed range was range was seen to be only in the range of 1 to 1.5 degrees. Reduction in elevator power will make the (iv) changes discernable. The earlier airspeed sensor with a range of 2.5" of water used to saturate at high speeds especially in dives. It has now been replaced a higher range sensor (5" of water). A picture of the second instrumented platform is shown in Fig. 18. Figure 17. Modular Instrumentation Package 9

Subsequent data acquisition during ground run revealed significant data smoothening on all accelerometer channels. Fig. 19 shows time history of Nz data for unfiltered and filtered channels. Fig. 2 shows the spectral plot of the data. It is readily apparent from the figures the extent to which filtering has helped in smoothening the data. Figure 18. Second Instrumented Platform DATA FILTERING Accelerometer data acquired during flight tests pointed to the need for filtering. It was felt that the principal source of noise in the data was engine vibration. To confirm the same, ground runs were conducted in which the accelerometer data was sampled at close to 4 samples per second. The data was acquired at different throttle settings and engine rpm was measured using a tachometer. The spectral content of the data was then determined by FFT techniques. The plots revealed that the principal source of noise was the engine indeed. Thereafter first order filters with a cutoff frequency of 5 Hz were implemented on the three accelerometer channels (the cutoff frequency of 5 Hz was chosen as it is unlikely that any of vehicle dynamics will exceed this frequency). Unfiltered Nz G Filtered Nz G.5.4.3.2.1 2 4 6 8 1 12 14 16 18 2 frequency Hz.5.4.3.2.1 ~99 RPM ie 165 Hz 2 4 6 8 1 12 14 16 18 2 frequency Hz Figure 2. Spectral plots of unfiltered and filtered Nz data. CURRENT STATUS AND FUTURE PLANS The second platform is ready and flight trials are scheduled for second week of November 2. This activity is part of a MTech project, which is time bound to conclude by end of December 2. The building and ground testing of the second platform took almost five weeks. This loss of time has caused shelving of plans to implement an airspeed/altitude controller onboard the vehicle. This activity will however be taken up as subsequent projects. ACKNOWLEDGEMENTS Figure 19. Time history of unfiltered and filtered Nz data The authors are grateful to ARDB for the support they have received for this project through the Center for Aerospace Systems Engineering, IIT Bombay, which has been setup with ARDB funding. Acknowledgements are also due to 1

Mr. Sudam Hanumant, FTC HAL (Bangalore) and FTC, ASTE, IAF for their support. REFERENCES 1. Wong, K.C., Aerospace Industry Opportunities in Australia - Unmanned Aerial Vehicles (UAV s) - Are They Ready This Time? Are We?, URL, http://www.aero.usyd.edu.au/wwwdocs/uav. html. 2. DragonFly UAV Project, Stanford University, Stanford, California, USA, URL, http://gpsuav.stanford.edu/project.overview /researchproject.overview.html. 3. Doeblin, E.O., Measurement Systems Application and Design, Fourth Edition, McGraw Hill International Edition 199. 4. M/s Honeywell Sensing and Control International, Canada, URL, http://www.honeywell.com/products/index.ht ml. 5. M/s Analog Devices, USA,URL, http://www.analog.com. 6. M/s Onset Computer Corporation, USA, URL, http://www.onsetcomp.com. 11