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MODELING AND FLIGHT TESTING OF DIFFERENTIAL THRUST AND THRUST VECTORING ON A SMALL UAV by Zachary J. Merceruio The primary objectives of this research are to mathematically model the propulsion forces applied to the aircraft during nominal, differential thrust, and thrust vectored flight configurations, and verify this modeling through simulation and flight testing experiments. This thesis outlines the modeling process, simulator development, design, and implementation of a propulsion assisted control system for the WVU Flight Control Systems Lab (FCSL) research aircraft. Differential thrust and thrust vectoring introduce additional propulsive terms in the aircraft force equations that are not present when the thrust line passes through the center of gravity. These additional forces were modeled and incorporated into a simulator of the research aircraft. The effects from differential thrust were small and difficult to quantify. The thrust vectoring effects were also found to be small with the elevator having significantly more pitch control over the vectored motors at the simulated flight conditions. Differential thrust was implemented using the on-board computer to command a different thrust level to each motor. The desired thrust differential was programed into a flight scheme based on simulation data, and activated during flight via a control switch on the transmitter. The thrust vectoring mechanism was designed using SolidWorks ® , built and tested outside of the aircraft, and finally incorporated into the aircraft. A high torque servo was used to rotate the motor mounting bar and vector the motors to a desired deflection. Utilizing this mechanism, the thrust vectoring was flight tested, mimicking scenarios tested in simulation. The signal to noise ratio was very low, making it difficult to identify the small changes in the aircraft parameters caused by the vectored thrust. iii Acknowledgments I would firstly like to thank my family for the love and support they have given me throughout this journey. Without their support this would have been a much more difficult undertaking. I would like to thank my committee chairman, Dr. Yu Gu, for offering me a direction, and then guiding me along that path. Also I would like to thank my research advisor and committee member, Dr. Marcello Napolitano, for giving me the opportunity to work in the Flight Controls Research Lab. I would like to thank my remaining committee members, Dr. Gary Morris and Dr. Srikanth Gururajan, for giving me advice and guidance, better preparing me for life after graduate school. I would like to extend a thank you to the flight testing crew who helped me to achieve my research goals. Thank you to the pilots who flew the research platforms beautifully: Mike Eden, Mike Spencer, and Dave Ellis. Thank you to the students who facilitated this research through

This paper presents the results of a parameter identification study for determining a mathematical model of the WVU YF-22 unmanned research aircraft under both nominal and actuator failure conditions to simulate malfunctions on primary control surfaces. Specifically, both linear and nonlinear mathematical models were derived from measured flight data acquired from pilot and automated computer-injected maneuvers under nominal and failure conditions. From analysis, the stability and control derivatives were extracted to determine the aerodynamic forces and moments. The aerodynamic derivatives were then introduced into a simulation model implemented within a Simulink-based environment, and studies were conducted to validate the accuracy of the identified models. Additionally, "Empirical" and DATCOM analyses were conducted for the WVU YF-22 to further validate the nonlinear model obtained through the parameter identification study. Nomenclature a = linear acceleration (m/s 2 ) A = decoupled (failure) state matrix b = wing span (m) B = decoupled (failure) input matrix C = aerodynamic coefficient c = mean aerodynamic chord (m) e = error H = altitude (m) i = surface deflection (deg) I = moment of inertia (kg m 2 ) J = product of inertia (kg m 2 ) m = aircraft mass (kg) TP 1 2 p = roll rate (deg/s) q = pitch rate (deg/s) q = dynamic pressure (psi) r = yaw rate (deg/s) S = wing surface area (m 2 ) T = thrust (N) V = velocity (m/s) Greek Letters α = angle of attack (deg) β = angle of sideslip (deg) δ = control surface deflection (deg) θ = pitch angle (deg)  = roll angle (deg) ψ = yaw angle (deg) ρ = air density (kg/m 3 ) Subscripts A = aileron D = drag H = stabilator l = rolling moment L = lift, left m = pitching moment n = yawing moment R = rudder, right xx = about the x-axis (body) xz = about the x and z axes (body) Y = side force yy = about the y-axis (body) zz = about the z-axis (body) Downloaded by PURDUE UNIVERSITY on July 30, 2015 | http://arc.aiaa.org |