Future management of helicopter fleets will be more heavily based on individual component damage tracking and less on legacy usage monitoring (flight parameter-based) methods. This enhances health assessment capabilities by taking into account the actual loads on a component-by-component basis. However, accurate loads prediction in rotating frame components remains a challenge. Even with advanced computational fluid dynamics (CFD) techniques, prediction of the unsteady aerodynamic loads acting on the rotor blades is computationally intensive and problematic in terms of accurate loads prediction across the entire flight regime of the helicopter. High-speed flight can potentially introduce both shock and near-stall effects within a given rotor rotation. Low-speed flight can include blade-vortex interaction effects, wherein flow from a given blade (vorticity loading from tip vortices) impinges upon the preceding blade, causing unsteady aerodynamic loading that is difficult to quantity and predict numerically. Vehicle maneuvering can produce significantly higher blade pitching moments than steady flight. All of these regimes combine to represent the loading history of the rotor system. Therefore, accurate loads prediction methods, in terms of matching peak-to-peak, magnitude, phase, as well as vibratory/harmonic content, are required that capture all flight regimes for all critical structural components. This research focuses on the development of a loads prediction method, known as the Load Confluence Algorithm (LCA), and its application to the analysis of a large set of flight test data from the NASA/US Army UH-60A Airloads Program. The LCA combines measured response at a prescribed set of locations with a numerical model of the rotor system. For a given flight condition (steady flight, maneuvers, etc.) the numerical simulation's predicted loads distribution is iteratively incremented (by harmonic) until convergence with measured loads is reached at the prescribed locations (control points). Predicted loads response at non-instrumented locations is shown to be improved as well, thus enhancing fatigue lifing methods for these components. The procedure specifically investigates the harmonic content of the applied loads and the improved prediction of the harmonic components. The impact of the enhanced accuracy on loads predictions on component structural fatigue is illustrated by way of an example. Results show that, for a limited sensor set (two 3-axis sensors per blade), blade loads are accurately predicted across a full range of flight regimes. Hub loads are best modeled using the pushrod as the control point. Results also show that load magnitude has a tremendous influence on damage, with a 25% over-estimation of vibratory load resulting in a damage factor of nearly 3. This research highlights the importance of accurate loads prediction for a rotorcraft life tracking program. Small inaccuracies in loads lead to dramatic errors in damage assessment.

Active flow control has been extensively explored in wind tunnel studies but successful in-flight implementation of an active flow control technology still remains a challenge. This thesis presents implementation of active flow control technology onboard a 33% scale Extra 330S ARF aircraft, wind tunnel studies and flight testing of fluidic actuators. The design and construction of the pulsed blowing system for stall suppression (LE actuator) and continuous blowing system for roll control (TE actuator) and pitch control have been presented. Full scale wind tunnel testing in 7̕ X 10 Oran W. Nicks low speed wind tunnel shows that the TE actuators are about 50% effective as the conventional ailerons. The LE actuator is found to be capable of suppressing stall from 12° to about 22°. Comparison of characteristics of Active elevator and conventional elevator in 3' X 4' low speed wind tunnel show that, the active elevator is as effective as of conventional elevator deflected at 5°. Flight tests show that TE actuators are able to control the aircraft in flight in banked turns. The measured roll rates in-flight support the wind tunnel test findings.

Thesis (Ph.D)--Aerospace Engineering, Georgia Institute of Technology, 2009. / Committee Chair: Costello, Mark. Part of the SMARTech Electronic Thesis and Dissertation Collection.

The primary objective of this thesis was to study, implement, and test low-cost electronic flight control systems (FCS) in remotely piloted subscale research aircraft with relaxed static longitudinal stability. Even though this implementation was carried out in small, simplified test-bed aircraft, it was designed with the aim of being installed later in more complex demonstrator aircraft such as the Generic Future Fighter concept demonstrator project. The recent boom of the unmanned aircraft market has led to the appearance of numerous electronic FCS designed for small-scale vehicles and even hobbyist-type model aircraft. Therefore, the purpose was not to develop a new FCS from scratch, but rather to take advantage of the available technology and to examine the performance of different commercial off-the-shelf (COTS) low-cost systems in statically unstable aircraft models. Two different systems were integrated, calibrated and tested: a simple, gyroscope-based, single-axis controller, and an advanced flight controller with a complete suite of sensors, including a specifically manufactured angle-of-attack transducer. A flight testing methodology and appropriate flight-test data analysis tools were also developed. The satisfactory results are discussed for different flight control laws, and the controller tuning procedure is described. On the other hand, the different test-bed aircraft were analysed from a theoretical point of view by using common aircraft-design methods and conventional preliminary-design tools. The theoretical models were integrated into a flight dynamics simulator, which was compared with flight-test data obtaining a reasonable qualitative correlation. Possible FCS modifications are discussed and some future implementations are proposed, such as the integration of the angle-of-attack in the control laws.

Use of an inflatable/rigidizable wing is explored for Mars airplane designs. The BIG BLUE (Baseline Inflatable-wing Glider Balloon Launched Unmanned airplane Experiment) project was developed at the University of Kentucky, with an objective to demonstrate feasibility of this technology with a flight-test of an high-altitude glider with inflatable/rigidizable wings. The focus of this thesis research was to design and analyze the wing for this project. The wings are stowed in the fuselage, inflate during ascent, and rigidize with exposure to UV light. The design of wings was evaluated by using aerodynamic and finite element software and wind tunnel testing. The profile is chosen based upon aerodynamic results and consideration of manufacturability of the inflatable wing structures. Flow over prototypes of inflatable/rigidizable and ideal shaped wings were also examined in the wind tunnel. Flow visualization, lift and drag measurements, and wake survey testing methods were performed. Results from the wind tunnel testing are presented along with suggestions in improving the inflatable/rigidizable wings aerodynamic efficiency and use on a low Reynolds number platform. In addition, high altitude wing deployment tests and low altitude flight tests of the inflatable/rigidizable wing were conducted.

Flight test telemetry environments can be particularly challenging due to RF shadowing, interference, multipath propagation, antenna pattern variations, and large operating ranges. In cases where the link quality is unacceptable, applying multiple receiving assets to a single test article can significantly improve the overall link reliability. The process of combining multiple received streams into a single consolidated stream is called Best Source Selection (BSS). Recent developments in BSS technology include a description of the maximum likelihood detection approach for combining multiple bit sources, and an efficient protocol for providing the real-time data quality metrics necessary for optimal BSS performance. This approach is being standardized and will be included in Appendix 2G of IRIG-106-17. This paper describes the application of this technology and presents performance results obtained during flight testing.

The control of multiple unmanned aerial vehicles (UAVs) in a swarm or cooperative team scenario has been a topic of great interest for well over a decade, growing steadily with the advancements in UAV technologies. In the academic community, a majority of the studies conducted rely on simulation to test developed control strategies, with only a few institutions known to have nurtured the infrastructure required to propel multiple UAV control studies beyond simulation and into experimental testing. With the Cal Poly UAV FLOC Project, such an infrastructure was created, paving the way for future experimentation with multiple UAV control systems. The control system architecture presented was built on concepts developed in previous work by Cal Poly faculty and graduate students. An outer-loop formation flight controller based on a virtual waypoint implementation of potential function guidance was developed for use on an embedded microcontroller. A commercially-available autopilot system, designed for fully autonomous waypoint navigation utilizing low cost hardware and open source software, was modified to include the formation flight controller and an inter-UAV communication network. A hardware-in-the-loop (HIL) simulation was set up for multiple UAV testing and was utilized to verify the functionality of the modified autopilot system. HIL simulation results demonstrated leader-follower formation convergence to 15 meters as well as formation flight with three UAVs. Several sets of flight tests were conducted, demonstrating a successful leader-follower formation, but with relative distance convergence only reaching a steady state value of approximately 35 +/- 5 meters away from the leader.

Flight testing methods are developed to find the drag polar for small UAVs powered by electric motors with fixed-pitch propellers. Wind tunnel testing was used to characterize the propeller-motor efficiency. The drag polar was constructed using data from flight tests. The proposed methods were implemented for a small UAV. A drag polar was found for this aircraft with CDo equal to 0.021, K1 equal to 0.229, and K2 equal to -0.056. This drag polar was then used to find the following performance parameters; maximum velocity, minimum velocity, velocity for maximum range, velocity for maximum endurance, maximum rate of climb, maximum climb angle, minimum turn radius, maximum turn rate, and maximum bank angle. Applications in UAV control and mission planning are also proposed.

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