Showing papers in "Journal of Aircraft in 2010"

Almost 40 Years of Airframe Noise Research: What Did We Achieve?

Abstract: With the advent of low noise high bypass ratio turbofan engines airframe noise gained significant importance with respect to the overall aircraft noise impact around airports. Already around 1970 airframe noise, originating from flow around the landing gears and high-lift devices, was recognized as a potential “lower aircraft noise barrier” at approach and landing. Since then, the outcome of extensive acoustic flight tests and aeroacoustic wind tunnel experiments enabled a detailed description and ranking of the major airframe noise sources and the development of noise reduction means. In the last decade advances in numerical and experimental tools led to a better understanding of complex noise source mechanisms. Efficient noise reduction technologies were developed for landing gears while the benefits of high-lift noise reduction means were often compensated by a simultaneous degradation in aerodynamic performance. The focus of this paper is not on the historical sequence of airframe noise research but rather aims to provide a concise survey of the achievements in airframe noise source description and reduction over the last 40 years worldwide. Due to the vast amount of work focused on a variety of airframe noise problems, this review can only provide examples but does not claim to be complete.

Nonlinear Aeroelasticity of a Very Flexible Blended-Wing-Body Aircraft

Abstract: Blended-wing-body (BWB) aircraft with high-aspect-ratio wings is an important configuration for high-altitude long-endurance unmanned aerial vehicles (HALE UAV). Recently, Northrop Grumann created a wind tunnel model under the Air Force’s High Lift over Drag Active (HiLDA) Wing program to study the aeroelastic characteristics of blended-wing-body for a potential Sensorcraft concept. This paper presents a study on the coupled aeroelastic / flight dynamics stability and response of a BWB aircraft that is modified from the HiLDA experimental model. An effective method is used to model very flexible BWB vehicles based on a low-order aeroelastic formulation that is capable of capturing the important structural nonlinear effects and couplings with the flight dynamics degrees of freedom. A nonlinear strain-based beam finite element formulation is used. Finite-state unsteady subsonic aerodynamic loads are incorporated to be coupled with all lifting surfaces, including the flexible body. Based on the proposed model, body-freedom flutter is studied, and is compared with the flutter results with all or partial rigid-body degrees of freedom constrained. The applicability of wind tunnel aeroelastic results (where the rigid-body motion is limited) is discussed in view of the free flight conditions (with all 6 rigid-body degrees of freedom). Furthermore, effects of structural and aerodynamic nonlinearities as well as wing bending/torsion rigidity coupling on the stability and gust response are also studied in this paper.

Analytical Sensitivity Analysis of an Unsteady Vortex Lattice Method for Flapping Wing Optimization

Abstract: This work considers the design optimization of a flapping wing in forward flight with active shape morphing, aimed at maximizing propulsive efficiency under lift and thrust constraints. This is done with an inviscid three-dimensional unsteady vortex lattice method, whose lack of fidelity is offset by a relatively inexpensive computational cost. The design is performed with a gradient-based optimization, where gradients are computed with an analytical sensitivity analysis. Wake terms provide the only connection between the forces generated at disparate time steps, and must be included to compute the derivative of the aerodynamic state at a time step with respect to the wing shape at all previous steps. The cyclic wing morphing, superimposed upon the flapping motions, is defined by a series of spatial and temporal approximations. The generalized coordinates of a finite number of twisting and bending modes are approximated by cubic splines. The amplitudes at the control points provide design variables; increasing the number of variables (providing the wing morphing with a greater degree of spatial and temporal freedom) is seen to provide increasingly superior designs, with little increase in computational cost. I. Introduction HE design and optimization of artificial flapping wing flyers presents considerable difficulties in terms of computational cost: the complex physical phenomena associated with the flight (unsteady low Reynolds number vortical flows in conjunction with a nonlinear elastic wing surface undergoing large prescribed rotations and translations) may require a high-fidelity computational tool. Furthermore, the search optimization process typically requires many function evaluations to converge to a relevant optimum. Lower fidelity numerical tools may help alleviate the burden, either used during the search process in conjunction with a higher-fidelity model 1

Novel, Bidirectional, Variable-Camber Airfoil via Macro-Fiber Composite Actuators

Abstract: This study aims to enable solid-state aerodynamic force generation in high-dynamic-pressure airflow. A novel, high-load-output, bidirectional variable-camber airfoil employing a type of piezoceramic composite actuator known as a Macro-Fiber Composite is presented. The novel airfoil employs two active surfaces and a single four-bar (box) mechanism as the internal structure. The unique choice of boundary conditions allows variable and smooth deformation in both directions from a flat camber line. The paper focuses on actuation modeling and response characterization under aerodynamic loads. A parametric study of aerodynamic response is employed to optimize the kinematic parameters of the airfoil. The concept is fabricated by implementing eight Macro-Fiber Composite 8557-P1-type actuators in a bimorph configuration to construct the active surfaces. The box mechanism generates deflection and camber change as predicted. Wind-tunnel experiments are conducted on a 12.6% maximum thickness, 127 mm chord airfoil. Aerodynamic and structural performance results are presented for a flow rate of 15 m /s and a Reynolds number of 127,000. Nonlinear effects due to aerodynamic and piezoceramic hysteresis are identified and discussed. A lift coefficient change of 1.54 is observed purely due to voltage actuation. Results are compared with conventional, zero-camber NACA and other airfoils. A 72% increase in the lift-curve slope is achieved when compared with a NACA 0009 airfoil.

Computational investigation of micro-scale coaxial rotor aerodynamics in hover

Abstract: In this work, a compressible Reynolds-averaged Navier-Stokes solver is used to investigate the aerodynamics of a microscale coaxial-rotor configuration in hover, to evaluate the predictive capability of the computational approach and to characterize the unsteadiness in the aerodynamic flowfield of the microscale coaxial systems. The overall performance is well-predicted for a range of rpm and rotor spacing. As the rotor spacing increases, the top-rotor thrust increases and the bottom-rotor thrust decreases, while the total thrust remains fairly constant. The thrusts approach a constant value at very large rotor spacing. Top rotor contributes about 55 % of the total thrust at smaller rotor spacing, which increases to about 58% at the largest rotor separation. The interaction between the rotor systems is seen to generate significant impulses in the instantaneous thrust and power. Unsteadiness is mainly caused due to blade loading and wake effect. Additional high-frequency unsteadiness was also seen due to shedding near the trailing edge. The phasing of the top vortex impingement upon the bottom rotor plays a significant role in the amount of unsteadiness for the bottom rotor. Interaction of the top-rotor tip vortex and inboard sheet with the bottom rotor results in a highly three-dimensional shedding on the upper surface of the blade in the outboard region and a two-dimensional shedding on the lower surface at the inboard portion of the blade. The wake of the top rotor contracts faster compared with that of the bottom rotor because of the vortex-vortex interaction. Further, the top-rotor wake convects vertically down at a faster rate due to increased inflow.

Approximate Modeling of Unsteady Aerodynamics for Hypersonic Aeroelasticity

Abstract: DOI: 10.2514/1.C000190 Various approximations to unsteady aerodynamics are examined for the aeroelastic analysis of a thin doublewedge airfoil in hypersonic flow. Flutter boundaries are obtained using classical hypersonic unsteady aerodynamic theories: piston theory, Van Dyke’s second-order theory, Newtonian impact theory, and unsteady shock-expansion theory. The theories are evaluated by comparing the flutter boundaries with those predicted using computational fluid dynamics solutions to the unsteady Navier–Stokes equations. Inaddition, several alternative approaches to the classical approximations are also evaluated: two different viscous approximations based on effective shapes and combined approximate computational approaches that use steady-state computational-fluid-dynamics-based surrogatemodelsinconjunction withpistontheory.Theresultsindicatethat,with theexceptionof first-order piston theory and Newtonian impact theory, the approximate theories yield predictions between 3 and 17% of normalized root-mean-square error and between 7 and 40% of normalized maximum error of the unsteady Navier–Stokes predictions. Furthermore, the demonstrated accuracy of the combined steady-state computational fluid dynamics and piston theory approaches suggest that important nonlinearities in hypersonic flow are primarily due to steadystate effects. This implies that steady-state flow analysis may be an alternative to time-accurate Navier–Stokes solutions for capturing complex flow effects.

Span Efficiencies of Wings at Low Reynolds Numbers

Abstract: Elegant and inviscid analytical theory can predict the induced drag on lifting wings of finite span. The theoretical prediction is then often modified by multiplication with a dimensionless coefficient for which the departure from a value of 1 is used as a way to incorporate realistic and necessary departures from the idealized model. Unfortunately, there are conflicting definitions of these dimensionless coefficients, often known as span efficiencies, so that even if numerical values are assigned in a clear and transparent fashion, their application and validity remain unclear. Here, the differences between two commonly used definitions of span efficiency are identified and it is shown that for the case of airfoil sections and finite wings at chordwise Reynolds numbers less than 10 5 , neither one has values close to those commonly assumed in the aeronautics literature. The cause of these significant viscous modifications to inviscid theory is traced to the movement of separation points from the trailing edge of real airfoils. A modified nomenclature is suggested to reduce the likelihood of confusion, and appropriate formulations for the drag of streamlined bodies in viscous flows at moderate Reynolds number are considered, with application to small-scale flying devices, both natural and engineered.

Design Optimization of a Truss-Braced-Wing Transonic Transport Aircraft

Abstract: the conventional cantilever configuration. One comparison produces a reduction of 45% in the fuel consumption while decreasing the minimum takeoff gross weight by 15%. For a second comparison, the fuel weight is reduced by 33% with a decreased minimum takeoff gross weight of 19%. Very attractive vehicle performance can be achieved without the necessity of decreasing cruise Mach number. The results also indicate that a truss-braced wing has a greater potential for improved aerodynamic performance than other innovative aircraft configurations. Further studieswillconsidertheinclusionofmorecomplextrusstopologiesandotherinnovativetechnologiesthatarejudged to be synergistic with truss-braced-wing configurations.

Aerostructural Optimization of Nonplanar Lifting Surfaces

Abstract: to findoptimalnonplanarliftingsurfacesandtoexplainthevariousfactorsandtradeoffsatplayApanelmethodand anequivalentbeam finite-elementmodelareusedtoexplorenonplanarliftingsurfaces,whiletakingintoaccountthe coupling and design tradeoffs between aerodynamics and structures Both single-discipline aerodynamic optimization and multidisciplinary aerostructural optimization problems are investigated The design variables are chosen to give the lifting-surface arrangement as much freedom as possible This is accomplished by allowing a number of wing segments to vary their area, taper, twist, sweep, span, and dihedral, with the constraint that they must not intersect each other Because of the complexity of the resulting design space and the presence of multiple localminima, anaugmentedLagrangianparticle swarmoptimizer isusedto solvethe optimizationproblemsWhen only aerodynamics are considered, closed lifting-surface configurations, such as the box wing and joined wing, are found to be optimal When aerostructural optimization is performed, a winglet configuration is found to be optimal when the overall span is constrained, and a wing with a raked wingtip is optimal when there is no such constraint

Characteristics of Pitching and Plunging Airfoils Under Dynamic-Stall Conditions

Abstract: An experimental investigation into the dynamic-stall process of a pitching and plunging airfoil at low Reynolds numbers has been carried out using direct force measurements and smoke visualization in an Eiffel-type wind tunnel. The strong influence of reduced frequency (k = �fc/U∞) on the vortical wake of both pure-plunging and pure-pitching airfoils is revealed. Here a transition from a bluff-body to a mushroom-type wake has been observed at approximately k = 0.2. Some associated lift and moment hysteresis curves for combined pitching and plunging kinematics are then presented with an accompanying discussion on the nature of the dynamic-stall process. For these complex kinematics it is observed that both lift and moment phase lags grow with reduced frequency from k = 0.05 to k = 0.1. Despite substantial lift augmentation in the light- and deep-stall regimes, strong pitching-down moments are not avoided.

Damage of Carbon/Epoxy Composite Plates Subjected to Mechanical Impact and Simulated Lightning

Abstract: Damage is inflicted on carbon-fiber/epoxy composite plates using both simulated lightning strike and mechanical impact in the effort to understand the relative effect of the two damage mechanisms. A methodology is proposed to characterize the damage resistance and tolerance of unconfigured composite plates subjected to lightning strike in a fashion that is consistent with the extensive work previously done on low-velocity impact. Using current and voltage diagnostics, it is possible to extrapolate the amount of electromechanical energy absorbed by the plate during the strike and compare it to that absorbed during a mechanical impact. Damage resistance is characterized by means of ultrasonic C-scans and microscopy, whereas residual strength is measured by means of compression after impact testing. Results show that the energy dissipated in a specimen during the lightning strike is much greater than the strain energy introduced by mechanical impact, and hence a comparison based on energy is not recommended. However, based on the relative threat levels associated with the impact and the lightning strike events, the comparison yields insightful observations on both damage state and residual performance. In general, for the configurations tested, lightning strike damage seems to be less detrimental than the mechanical impact in terms of both damage area and residual strength.

Unsteady Lift Generation on Rotating Wings at Low Reynolds Numbers

Abstract: The rotating wing experiment is a fully three-dimensional simplification of the flapping-wing motion observed in nature. The spanwise velocity gradient and the wing starting and stopping acceleration that exist on an insectlike flapping wing are generated by the rotational motion of a finite-span wing. The flow development around a rotating wing at Re = 60, 000 has been studied using high-speed particle image velocimetry to capture the unsteady velocity field. Lift and drag forces have been measured for several different sets of wing kinematics and angles of attack. The lift curve shape was similar in all cases. A transient high lift peak, approximately 1.5 times the quasi-steady value, occurred in the first chord length of travel, and it was caused by the formation of a strong attached leading-edge vortex. This vortex then separated from the leading edge, resulting in a sharp drop in lift. As weaker leading-edge vortices continued to form and shed, lift values recovered to an intermediate value. The circulation of the leading-edge vortex has been measured and agrees well with the force data. Wing kinematics had only a small effect on the aerodynamic forces produced by the waving wing. In the early stages of the wing stroke, the velocity profiles with low accelerations affected the timing and the magnitude of the lift peak, but at higher accelerations, the velocity profile was insignificant.

Mechanism for Warp-Controlled Twist of a Morphing Wing

Abstract: A new concept for actively controlling wing twist is described. The concept relied on introducing warping deformation of the wing skin, which was split at the trailing edge to create an open-section airfoil. An internal screw mechanism was introduced near the trailing edge, so that the load-carrying capability of the wing was maintained while allowing the introduction of warping displacement between the lower and upper wing skins at the trailing edge. Simple structural modeling of the warping wing based on generalized thin-walled beam theory was performed. A demonstration wing was built based on a NACA 23012 airfoil section with a span of 0.68 m and a chord length of 0.235 m. A maximum peak-to-peak twist of 27 deg was demonstrated, with excellent correlation between theory and experiment. Wind-tunnel tests showed that warping could change the lift coefficient by as much as 0.7 at maximum peak-to-peak twist. Analytical and vortex-lattice models were demonstrated to give accurate predictions of the lift coefficient at smaller absolute twist angles. Furthermore, analytic modeling of the wing drag was shown to be in close correspondence with the drag measurements and showed that wing warping could be used to influence the lift induced drag. In general, it was demonstrated that at lower angles of attack, a more positive twist resulted in a higher lift-to-drag ratio. This study proved that a twist-active wing can have sufficient gain to control the rolling motion of an aircraft and to ensure that the lift-to-drag ratio is maximized at various flight conditions.

Ionospheric Threat Parameterization for Local Area Global-Positioning-System-Based Aircraft Landing Systems

Abstract: Observations of extreme spatial rates of change of ionospheric electron content and the characterization strategy for mitigation applied by the US local area augmentation system are shown During extreme ionospheric activity, the gradient suffered by a global navigation satellite system user a few kilometers away from a ground reference station may reach as high as 425 mm of delay (at the GPS L1frequency) per km of user separation The method of data analysis that produced these results is described, and a threat space that parameterizes these possible threats to user integrity is defined Certain configurations of user, reference station, global navigation satellite system satellite, and ionospheric storm-enhanced density may inhibit detection of the anomalous ionosphere by the reference station

Airfoil Optimization Using Practical Aerodynamic Design Requirements

Abstract: Practical aerodynamic design problems must balance the goal of performance optimization over a range of on-design operating conditions with the need to meet design constraints at various o-design operating conditions. Such design problems can be cast as multipoint optimization problems where the on-design and o-design operating conditions are represented as design points with corresponding objective/constraint functions. Two methods are presented for obtaining optimal airfoil designs that satisfy all design objectives and constraints. The first method uses an unconstrained optimization algorithm where the optimal design is achieved by minimizing a weighted sum of the objective functions at each of the operating conditions. To address the competing design objectives between on-design and o-design operating conditions, an automated procedure is used to eciently weight the o-design objective functions so as to limit their influence on the overall optimization while satisfying the design constraints. The second method uses the constrained optimization algorithm SNOPT, which allows the aerodynamic constraints imposed at the o-design operating conditions to be treated explicitly. Both methods are applied to the design of an airfoil for a hypothetical aircraft where the problem is formulated as an 18-point multipoint optimization.

Multidisciplinary Considerations in the Design of Wings and Wing Tip Devices

Abstract: Conceptual wing design analysis methods are combined with numerical optimization to find minimum-drag wings subject to constraints on lift, weight, pitching moment, and stall speed. Tip extensions and winglets designed for minimum drag achieve similar performance, with the optimal solution depending on the ratio of the maneuver lift coefficient to the cruise lift coefficient. The results highlight the importance of accounting for the depth of the wing structural box in the weight model and including constraints on stall speed. For tailless aircraft, C-wings show a slight performance advantage over wings with winglets when longitudinal trim and stability constraints are considered. This performance advantage is more significant for span-constrained or low-sweep designs. Finally, to demonstrate other possible applications of the method, planar wings with active load alleviation are optimized, showing drag savings on the order of 15%.

Airframe Design for Silent Fuel-Efficient Aircraft

Abstract: The noise goal of the Silent Aircraft Initiative, a collaborative effort between the University of Cambridge and Massachusetts Institute of Technology, demanded an airframe design with noise as a prime design variable and a design philosophy that cut across multiple disciplines. This paper discusses a novel design methodology synthesizing first-principles analysis and high-fidelity simulations, and it presents the conceptual design of an aircraft with a calculated noise level of 62 dBA at the airport perimeter. This is near the background noise in a well-populated area, making the aircraft imperceptible to the human ear on takeoff and landing. The all-lifting airframe of the conceptual aircraft design also has the potential for improved fuel efficiency, as compared with existing commercial aircraft. A key enabling technology in this conceptual design is the aerodynamic shaping of the airframe centerbody. Design requirements and challenges are identified, and the resulting aerodynamic design is discussed in depth. The paper concludes with suggestions for continued research on enabling technologies for quiet commercial aircraft.

Real Time Morphing Wing Optimization Validation Using Wind-Tunnel Tests

Abstract: In this paper, wind-tunnel results of a real time optimization of a morphing wing in the wind tunnel for delaying the transition toward the trailing edge are presented. A morphing rectangular finite aspect ratio wing, having a wind tunnel experimental airfoil reference airfoil cross section, was considered, with its upper surface made of a flexible composite material and instrumented with Kulite pressure sensors and two smart memory alloys actuators. Several wind-tunnel test runs for various Mach numbers, angles of attack, and Reynolds numbers were performed in the 6' x 9' wind tunnel at the Institute for Aerospace Research at the National Research Council Canada. Unsteady pressure signals were recorded and used as feedback in real time control while the morphing wing was requested to reproduce various optimized airfoils by changing automatically the two actuators' strokes. This paper shows the optimization method implemented into the control software code that allows the morphing wing to adjust its shape to an optimum configuration under the wind-tunnel airflow conditions.

Optimal Design of Tow-Placed Fuselage Panels for Maximum Strength with Buckling Considerations

Abstract: The introduction of advanced tow-placement machines has made it possible to fabricate novel variable-stiffness composite structures where the fiber orientation angle varies continuously within each ply and throughout the structure. This manufacturing capability allows designers of composites to use the fiber orientation angle as design variable in their analysis, not only for each ply as with conventional composites, but at each point within a ply. Consequently, beyond the improvements that can be accomplished with traditional composites with straight fibers, the directional material properties of composites can be fully exploited to improve the laminate performance. In this paper, design tailoring for the pressure pillowing problem of a fuselage panel bounded by two frames and two stringers is addressed using tow-placed steered fibers. The panel is modeled as a two-dimensional plate loaded by out-of-plane pressure and in-plane loads. A Python-ABAQUS script is developed to perform the linear and geometrically nonlinear finite element analyses of variable-stiffness panels. The design objective is to determine the optimal fiber paths within each ply of the laminate for maximum load carrying capacity and for maximum buckling capacity. Simulated-annealing algorithm is used to solve the optimization problems. Optimal designs are obtained for different loading cases and boundary conditions. As a basis ofcomparison, a practical constant-stiffness quasi-isotropic design is used. Numerical results indicate that by placing the fibers in their optimal spatial orientations within each ply, the load carrying capacity and the buckling load of the structure can be substantially improved compared with traditional straight fiber designs. It is shown that laminates optimized for maximum failure load have buckling loads that are higher than those for quasi-isotropic laminates. On the other hand, laminates optimized for maximum buckling load fail at load levels lower than laminates optimized for maximum failure load. However, the failure loads of those laminates may still be higher than those for their quasi-isotropic counterparts.

Multilevel Structural Optimization for Preliminary Wing-Box Weight Estimation

Abstract: This paper deals with the weight estimation of the wing box of a commercial aircraft by means of a procedure suitable for very large liners and/or unconventional configurations for which statistical data and empirical formulas may not be sufficiently reliable. Attention is focused on the need to account for aeroelastic interaction from a very preliminary stage of the design cycle. The procedure exploits the first of three levels of a multilevel structural optimization system conceived for the preliminary design of the wing primary structure and a simplified evaluation of the cross-sectional properties. The comparison between weight estimates obtained with the present procedure and predictions supplied by available literature shows a satisfactory agreement.

Data mining of Pareto-optimal transonic airfoil shapes using proper orthogonal decomposition

Abstract: A new approach to extract useful design information from Pareto-optimal solutions of optimization problems is proposed and applied to an aerodynamic transonic airfoil shape optimization. The proposed approach enables an analysis of line, face, or volume data of all Pareto-optimal solutions such as shape and flow field by decomposing the data into principal modes and corresponding base vectors using proper orthogonal decomposition (POD). Analysis of the shape and surface pressure data of the Pareto-optimal solutions of an aerodynamic transonic airfoil shape optimization problem showed that the optimized airfoils can be categorized into two families (low drag designs and high lift designs), where the lift is increased by changing the camber near the trailing edge among the low drag designs while the lift is increased by moving the lower surface upward among the high lift designs.

Nonlinear-Aerodynamics/Nonlinear-Structure Interaction Methodology for a High-Altitude Long-Endurance Wing

Abstract: Nonlinear aeroelastic analysis is essential for high-altitude long-endurance (HALE) aircraft. In the current paper, we have presented a computational aeroelastic tool for nonlinear-aerodynamics/nonlinear-structure interaction. Specifically, a consistent nonlinear time-domain aeroelastic methodology has been integrated via tightly coupling a geometrically exact nonlinear intrinsic beam model and the generalized unsteady vortex-lattice aerodynamic model with vortex roll-up and free wake. The effects of discrete gust as well as flow separation at various angles of attack from attached flow to the stall and poststall ranges are also included in the nonlinear aerodynamic model. A HALE-wing model is analyzed as a numerical example. The trim angle of attack is first found for the wing, and the results show that aeroelastic instability could occur at higher angles of attack. The HALE-wing model under the trim condition is then analyzed for various gust profiles to which it is subject. It is found that for certain gust levels, the elastic deformations of the HALE wing tend to become unstable: notably, the in-plane deflections become very significant. It is noted for the unstable solution of the HALE wing that the flow may be well beyond the stall range. An engineering approach with the use of the nonlinear sectional lift is attempted to consider such stall effects.

Modeling and testing of a morphing wing in open-loop architecture

Abstract: This paper presents the modeling and experimental testing of the aerodynamic performance of a morphing wing in open-loop architecture. We show the method used to acquire the pressure data from the external surface of the flexible wing skin, using incorporated Kulite pressure sensors and the instrumentation of the morphing controller. The acquired pressure data are analyzed through fast Fourier transforms to detect the magnitude of the noise in the surface airflow. Subsequently, the data are filtered by means of high-pass filters and processed by calculating the root mean square of the signal to obtain a plot diagram of the noise in the airflow. This signal processing is necessary to remove the inherent noise electronically induced from the Tollmien-Schlichting waves, which are responsible for triggering the transition from laminar to turbulent flow. The flexible skin is required to morph the shape of the airfoil through two actuation points to achieve an optimized airfoil shape based on the theoretical flow conditions similar to those tested in the wind tunnel. Two shape memory alloy actuators with a nonlinear behavior drive the displacement of the two control points of the flexible skin toward the optimized airfoil shape. Each of the shape memory actuators is activated by a power supply unit and controlled using the Simulink/MATLAB® software through a self-tuning fuzzy controller. The methodology and the results obtained during the wind-tunnel test proved that the concept and validity of the system in real time are discussed in this paper. Real-time acquisition and signal processing of pressure data are needed for further development of the closed-loop controller to obtain a fully automatic morphing wing system.

Validation Studies for Aeroelastic Trim and Stability of Highly Flexible Aircraft

Abstract: High-altitude, long-endurance (HALE) aircraft are highly flexible, requiring nonlinear aeroelastic analysis NATASHA (Nonlinear Aeroelastic Trim and Stability of HALE Aircraft) is a computer program developed to analyze the aeroelastic behavior of HALE aircraft The underlying formulation is based on the published geometrically exact, fully intrinsic beam equations This paper presents a wide range of results from NATASHA, comparing them with well-known solutions of beam stability and vibration problems, experimental data, or results from well-established computer codes Despite the simplicity of the underlying formulation, results obtained confirm the accuracy of the analysis and its suitability for conceptual and preliminary design of HALE aircraft

Closed-loop control validation of a morphing wing using wind tunnel tests

Abstract: In this paper a rectangular finite aspect ratio wing, having a wing trailing edge airfoil reference airfoil cross section, was considered. The wing upper surface was made of a flexible composite material and instrumented with Kulite pressure sensors and two smart memory alloys actuators. Unsteady pressure signals were recorded and visualized in real time while the morphing wing was being deformed to reproduce various airfoil shapes by controlling the two actuators displacements. The controlling procedure was performed using two methods which are described in the paper. Several wind-tunnel test runs were performed for various angles of attack and Reynolds numbers in the 6 × 9 foot wind tunnel at the Institute for Aerospace Research at the National Research Council Canada. The Mach number was varied from 0.2 to 0.3, the Reynolds numbers varied between 2.29 and 3.36 x 10 6 , and the angle-of-attack range was within -1 to 2 degrees. Wind-tunnel measurements are presented for airflow boundary layer transition detection using high sampling rate pressure sensors.

Improving the Aerodynamic Performance of Micro-Air-Vehicle-Scale Cycloidal Rotor: An Experimental Approach

Abstract: Performance and flowfield measurements were conducted on a small-scale cyclorotor for application to a micro air vehicle. Detailed parametric studies were conducted to determine the effects of the number of blades, rotational speed, and blade pitching amplitude. The results showed that power loading and rotor efficiency increased when using more blades; this observation was found over a wide range of blade pitching amplitudes. The results also showed that operating the cyclorotor at higher pitching amplitudes resulted in improved performance, independently of the number of blades. A momentum balance performed using the flowfield measurements helped to quantify the vertical and sideward forces produced by the cyclorotor; these results correlated well with the force measurements made using load balance. Increasing the number of blades increased the inclination of the resultant thrust vector with respect to the vertical because of the increasing contribution of the sideward force. The profile drag coefficient of the blade sections computed using a momentum deficit approach correlated well with typical values at these low chord Reynolds numbers. Particle image velocimetry measurements made inside the cage of the cyclorotor showed that there are rotational flows that, when combined with the influence of the upper wake on the lower half of the rotor, explain the relatively low efficiency of the cyclorotor.

Bistable Composite Flap for an Airfoil

Abstract: A study was conducted to address the challenges associated with investigating a bistable composite flap for an airfoil. A full-scale rotor blade section with a span of 2.114 m and a chord of 0.68 m, fitted with a 1 m span flap was wind-tunnel tested up to a speed of 60 m/s with the flap moving between two stable states for various angles of attack. The blade was approximated as a NACA 24016 section with a 20% chord trailing-edge flap to simplify the analysis. The trailing-edge flap was designed to change between its stable geometries between hover and forward flight conditions for aerodynamic performance improvements. The flap was driven by an electromechanical actuator that was mounted inside the blade D-spar at the leading edge. All of the rotor blade structure remote from this bistable flap region was unmodified and assumed to be completely rigid during wind-tunnel testing.

Effect of High-Fidelity Ice-Accretion Simulations on Full-Scale Airfoil Performance

Abstract: The simulation of ice accretion on a wing or other surface is often required for aerodynamic evaluation, particularly at small scale or low Reynolds number. Although there are commonly accepted practices for ice simulation, there are no established and validated guidelines. The purpose of this paper is to report the results of an experimental study establishing a high-fidelity, full-scale, iced-airfoil aerodynamic performance database. This research was conducted as a part of a larger program with the goal of developing subscale aerodynamic simulation methods for iced airfoils. Airfoil performance testing was carried out at the ONERA F1 pressurized wind tunnel using a 72 in. (1828.8 mm) chord NACA 23012 airfoil over a Reynolds number range of 4.5 x 10 6 to 16.0 × 10 6 and a Mach number range of 0.10 to 0.28. The high-fidelity ice-casting simulations had a significant impact on the aerodynamic performance. A spanwise-ridge ice shape resulted in a maximum lift coefficient of 0.56 compared with the clean value of 1.85 at Re = 15.9 x 10 6 and M = 0.20. Two roughness and streamwise shapes yielded maximum lift values in the range of 1.09 to 1.28, which was a relatively small variation compared with the differences in the ice geometry. The stalling characteristics of the two roughness ice simulations and one streamwise ice simulation maintained the abrupt leading-edge stall type of the clean NACA 23012 airfoil, despite the significant decrease in maximum lift. Changes in Reynolds and Mach numbers over the large range tested had little effect on the iced-airfoil performance.

Neural-network-based flush air data sensing system demonstrated on a mini air vehicle

Abstract: Flush air data sensing systems have been widely applied to large (manned) aircraft, where pressure orifices are typically located at the nosetip. This paper investigates the feasibility of a flush air data sensing system designed to estimate the air data states of a small unmanned air vehicle flown at speeds as low as Mach 0.07. Furthermore, due to the presence of a nose propeller, the pressure orifices are located at the wing leading edge. The motivation behind this project is the fact that traditional air data booms are physically impractical for small unmanned air vehicles. Overall, an 80 and 97% reduction in instrumentation weight and cost, respectively, were achieved. Both parametric and multilayer perceptron neural network models have been previously applied in the literature to model the aerodynamic relationship between aircraft surface pressure and the air data states. In this paper, an extended minimum resource allocating network radial basis function neural network is used as the flush air data sensing system model, due to its good generalization capabilities and compact structure. Computational fluid dynamic simulations are implemented to identify the ideal pressure port locations, and wind-tunnel tests are carried out to train and test the extended minimum resource allocating network radial basis function neural network.

Advanced Experimental and Numerical Validation and Analysis of Propeller Slipstream Flows

Abstract: A comparison of velocity(-derived) quantities is made in the slipstream of a turboprop equipped transport aircraft between phase-locked experimental Particle Image Velocimetry (PIV) data and unsteady Reynolds Averaged Navier-Stokes (uRANS) calculations. Velocity results indicate a high level of agreement between the numerical and experimental results. In addition to velocity vector data, also derived quantities, such as vorticity, allow a verification and validation of the modeling in CFD. Although a good agreement is achieved, future needs for the comparison and validation of computational (CFD) and experimental (PIV) data are highlighted.