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The paper presents the design of a symmetric, active, gust load alleviation system for a regional transport aircraft, based on a static output feedback with a constrained structure. The design is carried out on a comprehensive finite state aeroservoelastic model, including sensor units and actuator transfer functions, and verified by taking into account saturated control positions, rates, and hinge moments. The controller is designed within a quadratic optimal framework, through a second-order Hessian-based optimization algorithm, exploiting block diagonal Schur transformations of the closed-loop state equations and performance weightings. An accurately chosen worst discrete gust and a reference flight condition provide a baseline design, which is significantly effective in alleviating continuous turbulence loads. Such a reference design proves itself robust enough to alleviate atmospheric loads over the complete flight envelope and is eventually further improved and robustified through a simple bilinear q∞- M∞ algebraic scheduling. [ABSTRACT FROM AUTHOR]

A concept for lift modification on a conventional aircraft wing flow roll control at low angle of attack with dielectric barrier discharge plasma actuators is proposed and assessed through computational fluid dynamics simulations and preliminary wind-tunnel experiments. The concept consists of placing plasma actuators around the wing tip to add momentum in the direction opposite to that of the flow forming the tip vortex. Because of the limited strength of existing plasma actuators, the assessment is carried out for a relatively small two-dimensional wing (NACA 4418) with a rounded tip at zero angle of attack and 15 m/s for a Reynolds number in the range of 1.5 x 105. Computational fluid dynamics simulations show a significant alteration of the vorticity field downstream of the trailing edge characterized by a more diffused vortex surrounded by zones of negative vorticity induced by the actuators and, but not necessarily, outboard displacement of the tip vortex. This leads to a reduced downwash that results in a change in lift of up to almost 20% for actuator strength levels that should be achievable in the short term with a new generation of dielectric barrier discharge actuators. The actuator placed on the suction side contributes the most to the lift increase, with its induced jet blocking the flow around the wind tip at the origin of the formation of the tip vortex. Wind-tunnel experimental results support the computational fluid dynamics predictions in both magnitude and trend. Furthermore, preliminary computational fluid dynamics simulations are carried out for a symmetric nonlifting wing (NACA 0018), representative of aircraft tail surfaces at zero angle of attack to generate lift for pitch and yaw control. Results indicate lift generation that increases and becomes larger than drag at higher actuator strengths. These promising results show a potential for the proposed concept to replace movable flight control surfaces on future aircraft wings and empennages. [ABSTRACT FROM AUTHOR]

The objective of this study is to demonstrate the concept of active trailing-edge flaps as primary flight control and vibration reduction devices for a typical full-scale helicopter. A comprehensive rotorcraft analysis based on UMARC was developed to analyze the swashplateless rotor. A parametric study of various key design variables involved in the trailing-edge flap design was carried out. An optimal design of a trailing-edge flap system that provides effective control authority within the complete range of advance ratios as well as minimum actuation requirements was achieved. Trailing-edge flaps demonstrated the capability of performing both primary flight control and active vibration control functions. At a high forward speed (advance ratio of 0.32), the 4/rev vertical force and roll and pitch moments at hub ate successfully eliminated (by 90%), and the 4/rev in-plane hub forces are reduced by more than 40%. The half peak-to-peak value of the trailing-edge flap deflection for primary flight control is 7.1 deg, and an additional 4.7 deg is required for active vibration control. [ABSTRACT FROM AUTHOR]

A rational and robust scheme for obtaining recovery controls for nonlinear flight conditions, such as spins, is presented. In a spin, the loss of control due to the large angular rates and high angle of attack that the airplane experiences could lead to disastrous situations in general aviation. In this study, to arrest rotation and at the same time recover from a spin, the directions and magnitudes of linear and angular momentum vectors are effected to restore their prespin conditions. The problem of computation of the required controls is formulated as a nonlinear programming problem. Simulated time histories are presented and compared to the actual flight-test data for a light, single-engine general aviation aircraft. The proposed method produced effective control time histories for recovery from both a steep and a flat spin, demonstrating the effectiveness of the proposed approach. [ABSTRACT FROM AUTHOR]

Weight and longitudinal center-of-gravity position of an aircraft are valuable information for flight monitoring, automatic flight control, and fault detection and identification algorithms. In trimmed symmetric flight (as cruise, climb, or descent) the aircraft weight is counteracted mainly by the aerodynamic lift, whereas the associated couple is counterbalanced by a moment generated by an appropriate elevator deflection. Because, for given flight conditions, the aerodynamic lift and pitching moment are functions of the angle of attack and elevator deflection only, the last two variables carry the information about the aircraft weight and its longitudinal center-of-gravity position. It is demonstrated that this information can be effectively recovered in flight using nonparametric artificial neural networks (NNs), pretrained on flight-test data. To enhance the accuracy and speed up the training process, basic flight mechanics relations are incorporated into the NN. It is also demonstrated that the NN-based lift and pitching moment functions can be used a posteriori to estimate the aircraft longitudinal neutral point. [ABSTRACT FROM AUTHOR]

A perspective on the practice, issues, and potential of aeroelasticity is presented. The issues of cycle time reduction and improvements in the fidelity of aeroelastic analysis need to be emphasized for aeroelasticity to support adequately the airplane configuration development and configuration decisions. The introduction of fly-by-wire control systems has increased the complexity as well as the fidelity requirements for the aeroelastic analysis. The business cost of inaccuracies in loads and dynamics predictions can be seriously high. Therefore, innovative approaches should be encouraged to address these issues. Aeroelasticity has an opportunity to contribute to airplane design by evolving to become a more integrating discipline. Although many of the methods and processes in different subdisciplines are in place already, others need to be developed further for improved usability, efficiency, fidelity, and integration. The most challenging problems in realizing the full potential of aeroelasticity are not just technical but are also organizational in nature. The organizational issues are related to management of multidisciplinary teams in a lean engineering environment. Thus, to realize the multidisciplinary integrating potential of aeroelasticity and its contributions to the configuration decisions, technical and management foresight, leadership, and vision are required. [ABSTRACT FROM AUTHOR]

This paper reviews concepts for flow control based on jet spoilers and presents a new device based on the jet principle for application to fixed-wing aircraft. The new concept relies on the naturally blown inverse spoiler. The jet is blown parallel to the surface and points against the freestream. The theoretical basis of the device is presented. Calculations are performed with a Navier-Stokes program to investigate the key design parameters. These parameters include the spoiler deflection (equal or differential), the position of the hinge points, the shape of the bypass channel, the effects of backward- and forward-facing steps in a partial deflection, and the nature and magnitude of the control forces. A comparison is done with conventional inlay spoilers. Finally, a multidimensional optimization has been carried out to study the combined effects of hinge position, deflection, and spoiler length. Optimal configurations with respect to jet momentum and drag coefficient have been analyzed. [ABSTRACT FROM AUTHOR]