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Composite materials are widely used in aerospace applications due to their good mechanical properties such as low density and good resistance. Due to their versatility it is possible to design different configurations making it difficult to find the optimal option. The objective of this work is to compare different laminates on the wing surface, the advantage of using structural elements such as spars and ribs, also to see the influence on the results of different load simplifications. In the study, the finite element analysis technique is used to find the best characteristics regarding resistance, deformation, and failure modes under the established conditions. The analysis showed that the best load simplification is a decreasing distributed load along the pressure center, the best surface laminate is [45/0/0/45] with carbon fiber cloth with a maximum displacement of 15.27 mm, the spar presented a 29.6% decrease in maximum displacement and the ribs did not represent a significant change in the structural response. [ABSTRACT FROM AUTHOR]

Abstract In this paper, the dynamic modeling and vibration control of a two-blade propeller system of an aircraft are investigated in the presence of input magnitude and rate constraints and external disturbances. To avoid discretizing the original system, a partial differential equation (PDE) model preserving all system modes is developed. Based on the proposed PDE model, boundary control laws are derived to restrict the bending and twist deformations of the two-blade propeller under disturbances. To meet the input magnitude and rate constraints, a backstepping controller is designed. The closed-loop system stability is proved by using the Lyapunov's direct method. Numerical simulations are presented to demonstrate the effectiveness of the proposed control method. [ABSTRACT FROM AUTHOR]

This Special Issue highlights the latest enhancements in the abatement of noise and vibrations of aerospace and automotive systems. The reduction of acoustic emissions and the improvement of cabin interior comfort are on the path of all major transportation industries, having a direct impact on customer satisfaction and, consequently, the commercial success of new products. Topics covered in this Special Issue deal with computational, instrumentation and data analysis of noise and vibrations of fixed wing aircrafts, satellites, spacecrafts, automotives and trains, ranging from aerodynamically generated noise to engine noise, sound absorption, cabin acoustic treatments, duct acoustics and vibroacoustic properties of materials. The focus of this Special Issue is also related to industrial aspects, e.g.,: numerical and experimental studies have been performed for an existing and commercialized engine to enable design improvements aimed at reducing noise and vibrations; moreover, an optimization is provided for the design of low vibroacoustic volute centrifugal compressors and fans whose fluids should be strictly kept in the system without any leakage. Existing procedures and algorithms useful to reach the abovementioned objectives in the most efficient way are illustrated in the collected papers. [ABSTRACT FROM AUTHOR]

The objective of this study is to ascertain the steady-state aerodynamic characteristics of a thin circular arc airfoil with a large-camber by comparing its camber with various cambers at Reynolds numbers ranging from 1x10³ to 10x10³, similar to that of insect flight. A large camber of over 20% chord is seen in the elytra of beetles (Coleoptera), many of which fly by opening the elytra. A low-pressure wind tunnel that we uniquely designed was used to measure the very small forces and moment acting on the wing. As a result, a high maximum lift coefficient was obtained for a large-camber circular arc airfoil, even at Reynolds numbers less than 3x10³. In addition, the pitching moment of the airfoil was found to be zero in a wide range of angles of attack when the moment center was set to the appropriate position. The large-camber circular arc airfoil will be useful in designing very-low Reynolds number aircraft such as insect-sized micro-air vehicles (MAVs). [ABSTRACT FROM AUTHOR]

The Wilcox 2006 stress-m model, a Reynolds stress model, implemented in both the NASA Langley codes FUN3D and CF L3D, has been used to study a number of 2-D and 3-D cases. This study continues the assessments of the stress- w model by simulating the flow over two wings: the DPW-W1 and the DLR-F11 wings. Using FUN3D, which uses unstructured grids, and CFL3D, which uses structured grids, the results were compared to solvers employing oneand two-equation turbulence models and experimental data. In general, in situations where experimental data is available, the stress-co model performs as well or better than one- and two-equation models. [ABSTRACT FROM AUTHOR]

The wing-body junction flow is investigated parametrically using computational fluid dynamics in an attempt to understand the effects of junction flow on aircraft drag, with a focus on application to large business jet or commercial transport aircraft. A Reynolds-averaged Navier-Stokes computational fluid dynamics methodology is validated against detailed experimental data for a junction flow. Computational fluid dynamics results for a wing with a leading-edge strake (an aerodynamic surface designed to reduce flow separation, thereby reducing aircraft drag) are presented, and the effects of scaling this strake are explored. The effectiveness of the strake on a swept wing is compared to the same for a straight wing. Finally, the results from this parametric study are successfully applied to sizing a leading-edge strake for a commercial transport aircraft. It is demonstrated that a systematic approach, starting with a simple, validated model and building up to a realistic aircraft application, can build confidence in computational fluid dynamics results. [ABSTRACT FROM AUTHOR]

The German Biirgernahes Flugzeug (Citizen-Oriented Airplane) research program advances technologies for high-lift systems using circulation control Coanda-type flaps. The interaction of these high-lift systems with a propeller engine is the subject of the present work. A large and complex wind-tunnel model equipped with an active high-lift system and a powered propeller was designed, manufactured and instrumented. The wind-tunnel data show the potentials of the active high-lift system in three-dimensional wing applications. The results characterize and quantify the interactions between a propeller and a high-lift wing by using systematic variations of blowing momentum and propeller thrust. The results represent a unique database for future efforts to improve the efficiency of active lift and for validation of numerical simulation methods. [ABSTRACT FROM AUTHOR]

Manufacturing tolerances play an important role in designing low-cost and low-weight aircraft structures. A probabilistic cost model is developed that explicitly models performance cost along with quality and manufacturing costs to optimize manufacturing tolerances for damage-tolerant aircraft structures. The tolerance optimization is demonstrated for the design of a wing spar with fastener holes. The advantage of including performance cost is illustrated by comparing to an optimum found by ignoring the performance cost, showing substantial change in the optimum tolerance and total cost. The random aspect of quality cost is illustrated using measured manufacturing error data collected from the wing spar assemblies of a business jet. The paper also quantifies the uncertainties in the optimum due to finite samples of measured data used for modeling manufacturing error distributions, which were found to be small due to a large number of samples. [ABSTRACT FROM AUTHOR]

The formation and accretion of ice on the leading edge of a wing can be detrimental to airplane performance. Complicating this reality is the fact that even a small amount of uncertainty in the shape of the accreted ice may result in a large amount of uncertainty in aerodynamic performance metrics (e.g., the stall angle of attack). The main focus of this work concerns using the techniques of polynomial chaos expansions to quantify icing uncertainty much more quickly than traditional methods (e.g., Monte Carlo). First, we present a brief survey of the literature concerning the physics of wing icing, with the intention of giving a certain amount of intuition for the physical process. Next, we give a brief overview of the background theory of polynomial chaos expansions. Finally, we compare the results of Monte Carlo simulations to polynomial-chaos-expansion-based uncertainty quantification for several different airfoil-icing scenarios. The results are in good agreement and confirm that polynomial chaos expansion methods are much more efficient for the canonical airfoil icing uncertainty quantification problem than Monte Carlo methods. [ABSTRACT FROM AUTHOR]

The article focuses on study which examines a rigid paraglider reference wing. Topics discussed include the external flow and the influence of the spanwise arched shaped on stationary longitudinal and lateral aerodynamic characteristics, full-scale wing dimensions and model wing geometry data at panel's ends. Also mentioned are the lifting curve at zero sideslip, the polar curve at zero sideslip and the pitching moment curve.

Hover-capable flapping-wing micro air vehicles are well suited for missions in confined spaces. The best design practices for flapping wings and their kinematics are largely unknown, especially for flexible wings. To address this issue, numerical optimization is applied to the design of the kinematics and structural sizing of a flapping wing using a surrogate-based approach. The surrogates are generated using kriging interpolation of the time-averaged thrust generated and power required by the wings. The thrust and power data are computed using a nonlinear approximate aeroelastic model developed in previous studies by the authors. A numerical optimization algorithm is used to identify designs that produce the desired combination of thrust and power. The design variables consist of parameters describing the flap-pitch kinematics and the stiffness of the flexible wings. A trend study of thrust and power indicate that the phase angle between flap and pitch motions significantly affects the wing performance when the stroke amplitudes and the frequency are fixed. Smaller amounts of pitch actuation produced peak thrust in flexible wings when compared to the rigid wings. Several flexible configurations produce higher thrust when compared to the best thrust-producing rigid configuration. However, rigid wings have higher propulsive efficiency when compared to flexible wings for the same amount of generated thrust. Thus, the actual design of a flapping wing will depend on the relative importance given to thrust production and propulsive efficiency. [ABSTRACT FROM AUTHOR]

Aircraft wings carrying stores are susceptible to nonlinear aeroelastic limit-cycle oscillations, which can lead to reduced flight and mission performance. Limit-cycle-oscillation dynamics and control simulation studies, although of great importance, are often based on simplified typical-section airfoil models. In this work, the more accurate beam-rod representation is used to capture the spanwise varying displacement of a wing with store. The spanwise variation of wing geometry and structural properties, as well as the presence of multiple stores on rigid or flexible mounts, is efficiently modeled by the primitive-modes approach. Aeroelastic limit-cycle oscillation due to structural nonlinearity is demonstrated via time-marching simulations, as well as the computationally more efficient harmonic-balance method. Some novel forms of limit-cycle-oscillation behavior are observed as the model parameters are varied, and these are explained in terms of the flutter/divergence properties of the base linear aeroelastic system. [ABSTRACT FROM AUTHOR]

High-aspect-ratio wings present in very flexible aircraft can undergo large deformations, which results in significant changes in natural frequencies as well as in static and dynamic aeroelastic response. This geometric nonlinear behavior becomes an integral part of any aeroelastic analysis to be conducted in such class of vehicles. Aeroelastic scaling is an important way to study the aeroelastic behavior of aircraft, and it is an integral part in risk mitigation for aircraft development. However, the current aeroelastic scaling methodologies have focused on geometrically linear structures. This paper demonstrates a methodology for geometrically nonlinear aeroelastic scaling of very flexible aircraft. The known linear scaling factors and similarity rules are extended to address geometrically nonlinear aeroelastic scaling. A high-aspect-ratio flying wing in free flight is taken as an example to verify the new scaling procedure, and numerical studies are conducted using the University of Michigan's Nonlinear Aeroelastic Simulation Toolbox. Numerical results support the new approach for aeroelastic scaling of very flexible aircraft. [ABSTRACT FROM AUTHOR]

In preparation for wind-tunnel acoustic experiments of a hybrid wing-body vehicle with two jet engine simulator units, a series of twin jet aeroacoustic investigations was conducted leading to increased understanding and risk mitigation. An existing twin jet nozzle system and a fabricated hybrid wing-body aft deck fuselage are combined for a 1.9 % model scale study of jet nozzle spacing and jet cant angle effects, elevon deflection into the jet plume, and acoustic shielding by the fuselage body. Linear and phased array microphone measurements are made, and data processing includes the use of the deconvolution approach for the mapping of acoustic sources. Closely spaced twin jets with a 5 deg inward cant angle exhibit reduced noise levels compared with their parallel flow counterparts at similar and larger nozzle spacings. A 40 deg elevon deflection into the twin jet plume, which is required for hybrid wing-body ground rotation, can significantly increase upstream noise levels (more than 5 dB overall sound pressure level) with only minimal increases in the downstream direction. Lastly, deconvolution approach for the mapping of acoustic sources processing can resolve the noise source distribution of multiple shielded jet sources within the limits of the incoherent source assumption. [ABSTRACT FROM AUTHOR]

Under the hypotheses of linear potential flow and rigid wake aligned with the freestream, a configuration-invariant analytical formulation for the induced drag minimization of single-wing nonplanar systems is presented. Following a variational approach, the resulting Euler-Lagrange integral equation in the unknown circulation distribution is obtained. The kernel presents a singularity of the first order, and an efficient computational method, ideal for the early conceptual phases of the design, is proposed. Munk's theorem on the normalwash and its relation with the geometry of the wing under optimal conditions is naturally obtained with the present method. Moreover, Munk's constant of proportionality, not provided in his original work, is demonstrated to be the ratio between the freestream velocity and the optimal aerodynamic efficiency. The augmented Munk's minimum induced drag theorem is then formulated. Additional induced drag theorems are demonstrated following the derivations of this invariant procedure. Several nonplanar wing systems are proposed and analyzed, and the optimal induced drag and circulation are provided. The conjecture regarding the equality of the optimum induced drag of a quasi-closed C-wing with the induced drag of the corresponding closed system is also verified for several configurations including the box wing. [ABSTRACT FROM AUTHOR]

Flight experiments carried out with a motorized glider yield insights into the aerodynamic processes acting on a laminar wing section under varying on-coming flow conditions. Measurement data are obtained on a laminar wing glove equipped with pressure transducers and wall microphones to detect transition and investigate the boundary-layer instabilities. The on-coming flow, including the instantaneous angle of attack and the turbulence characteristics, is measured with hot-wire sensors. The flight tests were carried out under calm and moderately turbulent conditions. Characteristic results for both on-coming flow conditions are presented and differences are analyzed. Selected comparisons of the experimental results with steady boundary-layer and linear stability computations are provided. The effects observed in the experiments cannot be described solely by quasi-steady theory. Unsteady changes of the pressure distribution are identified as one source for the modification of the transition development. A conclusion on the observed superimposed effect of increased levels of atmospheric small-scale turbulence on boundary-layer receptivity cannot yet be drawn from the present experiments. [ABSTRACT FROM AUTHOR]

In this paper, a morphing-wing concept based on compliant ribs is considered as a replacement for conventional ailerons and its performance is investigated. The roll performance of the three-dimensional morphing-wing solution for different design speeds and under structural and strength constraints is analyzed. The design approach uses optimization techniques and considers the three-dimensional aerostructural behavior. The results show the possibility of producing sufficient roll control authority with a morphing solution, thus replacing conventional ailerons up to a design speed of 250 km /h. In a further step, the weight of the system as a function of the produced rolling moment is compared with the one of a conventional system. The obtained relations offer an estimation of the weight penalties associated with morphing. This estimation can be useful in the preliminary design of morphing aircraft. [ABSTRACT FROM AUTHOR]

The term supercirculation refers to an active, fluidic circulation control method used for a variety of aeronautical applications. In particular, this paper refers to supercirculation of entrainment wings. Using experimental data as well as some basic analytical equations, a mathematical expression is used to define supercirculation; subsequently an optimization calculation is provided for the minimal blowing pressure ratio needed for achieving supercirculation. Also, based on the proposed definition, a screening equation is determined in order to discern whether or not a certain airfoil has a geometry suited for supercirculation. [ABSTRACT FROM AUTHOR]

The article presents an analysis of a propeller wake interacting with an aircraft wing using high-resolution numerical methods to resolve the important flow features within the propeller wake. The unsteadiness of the propeller wake-wing interaction is described. It outlines the computation for the interaction of a time-averaged propeller wake with a wing and compares the results of the fully resolved interaction.

This study focuses on an increase in efficiency in the construction of flight procedures considering wind effect in protection areas. In Turkish airspace, the International Civil Aviation Organization wind model is used in the construction of flight procedures. This model estimates extremely high wind-speed values compared to actual wind speed, and causes inefficient airspace utilization in procedure design. Better wind-speed estimations reduce the size of the protection areas. Because the protection areas are not allowed to overlap, this reduction provides flexibility for capacity enhancement plans and high-terrain intersections for flight procedures. In this regard, considering long-term meteorological statistical data, a unique wind-speed model has been developed for use in Turkish airspace. The wind effect of the Turkish model is discussed over the wind effect found using the International Civil Aviation Organization wind model, in which the reduction in the protection areas for 90 deg course changes is found to be 0.2 n miles (450 m) for ground and 0.7 n miles (1332 m) for an altitude of 24,000 ft. For holding maneuvers at 10,000 ft, at certain flight configurations for instance, the reduction in the protection areas can be up to 1.6 n miles (3000 m). [ABSTRACT FROM AUTHOR]

This paper addresses the aerodynamics of membrane flapping wings. In a series of wind-tunnel experiments on the 25 and 74-cm-wingspan models, the stroke-averaged lift and horizontal force were measured at an angle of attack that varied from 0 deg (horizontal) to 90 deg (hovering position). The flapping frequency was held constant, and freestream velocity was varied with the advance ratio range 0-1.2. For high angles of attack, flapping wings do not exhibit a typical abrupt stall seen with fixed wings. This feature in flapping wings allows a smooth transition from a level flight to hovering and back. The angle of attack, corresponding to a balance of forces in the horizontal direction, decreases exponentially with advance ratio. Aerodynamic coefficients are defined using a reference velocity as a vector sum of a freestream velocity and a stroke-averaged wing-tip flapping velocity. The lift and drag coefficients obtained using this procedure collapse well for the studied advance ratios, especially at lower angles of attack. Polynomial approximations for aerodynamic coefficients are proposed that can be used in flight-performance analysis. [ABSTRACT FROM AUTHOR]

This work extends advanced beam models to carry out a more accurate free-vibration analysis of conventional (straight, or with sweep/dihedral angles) and joined wings. The beam models are obtained by assuming higher-order (up to fourth) expansions for the unknown displacement variables over the cross-section. Higher-order terms permit bending/torsion modes to be coupled and capture any other vibration modes that require in-plane and warping deformation of the beam sections to be detected. Classical beam analyses, based on the Euler-Bernoulli and on Timoshenko beam theories, are obtained as particular cases. Numerical solutions are obtained by using the finite element (FE) method, which permits various boundary conditions and different wing/section geometries to be handled with ease. A comparison with other shell/solid FE solutions is given to examine the beam model. The capability of the beam model to detect bending, torsion, mixed and other vibration modes is shown by considering conventional and joined wings with different beam axis geometries as well as with various sections (compact, plate-type, thin-walled airfoil-type). The accuracy and the limitations of classical beam theories have been highlighted for a number of problems. It has been concluded that the proposed beam model could lead to quasi-three-dimensional dynamic responses of classical and nonclassical beam geometries. It provides better results than classical beam approaches, and it is much more computationally efficient than shell/solid modeling approaches. [ABSTRACT FROM AUTHOR]

Some fixed-wing vertical-takeoff-and-landing aircraft can transition between two flight modes: forward flight and near-hover. This study was conducted to improve the transition performance of such vehicles. The experimental model consists of a rigid Zimmerman wing and a propulsion system with contra-rotating propellers. It was rapidly pitched about its aerodynamic center at an average freestream Reynolds number of 86,000. Five nondimensional pitching rates were used, along with two elevator deflections and three propulsive settings. The model was tested statically, and several observations were made. At constant throttle setting, from 20 to 70 deg, both advance ratio and thrust coefficient increase linearly with angle of attack. Higher throttle setting results in greater stall delay, causing the maximum lift coefficient to increase. When throttle setting increases, lift and drag coefficients increase throughout the tested angle-of-attack domain. Rapid-pitching tests showed that nose-up pitching delays stall and nose-down pitching hastens it. Lift and drag coefficients generally increase with positive pitching rate and decrease with negative pitching rate. Wing aerodynamic efficiency is virtually independent of throttle setting, elevator deflection, and pitching rate between 30 and 70 deg angle of attack. The propulsion system is not sensitive to rapid pitching with regard to thrust, normal force, or propulsive moment. [ABSTRACT FROM AUTHOR]

The article discusses a study which identified the rigid body derivatives and the aerolastic derivatives representing the elastic effects from flight data, using an integrated model structure. Estimates of rigid body and aeroelastic derivatives were made in the longitudinal axis for the first two elastic modes. The two structural modes included in the flight simulation characterize fuselage bending with the wing participating in phase and wing bending with fuselage participating out of phase.

Recent studies on multidisciplinary design optimization of truss-braced-wing airplane configurations have indicated that these airplanes offer significant potential for performance improvements in terms of fuel efficiency. The typical truss-braced-wing configurations obtained from these multidisciplinary design optimization studies have wing half-span on the order of 170 ft with low wing-chord and thickness-to-chord ratios, which makes the wings very flexible in nature. This paper presents a set of parametric studies of truss-braced-wing configurations to understand the influence of the wing geometry parameters on the wing structural and aeroelastic characteristics. The primary parameters considered here are the wing half-span, strut sweep, spanwise location of wing-strut joint, and number of truss members in the wing configuration. Each truss-braced-wing parametric configuration is sized based on strength considerations and studied for aeroelastic behavior. The results indicate strong influence of all the parameters considered here. For most cases, increasing the half-span monotonically increases the wing weight and reduces both the natural frequencies and the flutter speed. A larger difference between the wing- and strut sweep angles is seen to increase wing weight, but with a positive influence on flutter speed for various truss-braced-wing configurations. The spanwise intersection location has distinct optima for wing weight and flutter speed, which typically lie in between 55 and 70%. These results are expected to provide guidance for future multidisciplinary design optimization studies for truss-braced-wing configurations. [ABSTRACT FROM AUTHOR]

We conducted large eddy simulations of the control of separated flow over an airfoil using body forces and discuss the role of a three-dimensional vortex structure in separation control. Two types of cases are examined: (1) the body force is distributed in a spanwise uniform layout and (2) the body force is distributed in a spanwise intermittent layout, with three-dimensional vortices being expected to be generated in the latter cases. The flow fields in the latter cases have a shorter separation bubble than those in the former cases although the total momentum of the body force in the latter cases is the same as or half of the former cases. In the flow fields of the latter type, the three-dimensional vortices, which are not observed in the former cases, are generated by the body force downstream of the body force distributed. Thus, three-dimensional vortices are considered to be effective in controlling the separated flow. [ABSTRACT FROM AUTHOR]

For searching the stiffness structure under the constraints of given material volume fraction, the subset simulation-based topology optimization method is established on the combination of the advanced sensitivity filtering technique with the subset simulation, an efficient reliability analysis method. The subset simulation-based topology optimization is convergent, checkerboard free, and mesh independent. After one example is used to testify these advantages of the established method, the optimal design of the leading-edge rib of an aircraft wing is searched by the subset simulation-based topology optimization method, and the results demonstrate that the established method is an effective and rational tool for the topology optimization of the continuum structure. [ABSTRACT FROM AUTHOR]

Shape optimization with a nmltiobjective optimization technique and tradeoff analysis of the three-dimensional wings-in-ground effect have been performed. Unlike an airplane out-of-ground effect, a wing-in-ground vehicle has to satisfy static height stability in order to cruise steadily in ground effect. However, it is difficult to simultaneously satisfy both high aerodynamic performance and static height stability because of the strong tradeoffs between the two. A multiobjective optimization technique can address this problem. In multiobjective optimization, the optimum is not unique; rather, it is a set of nondominated potential solutions called Pareto optima. In this study, after 19 evolutions of a genetic algorithm, 74 nondominated Pareto optima are obtained and four different groups can be observed from the Pareto optima. Each group represents a set of the potential global solutions for a single objective or combined objectives. The analysis of the Pareto optima suggests that the lift is increased because of increased pressure on the lower wing surface, but the influence of pressure on the upper surface is not significant. The wing thickness and curvature increase the lift further, but they are not favorable for static height stability and the lift-drag ratio; a swept-forward wing can improve the static height stability. [ABSTRACT FROM AUTHOR]

The cruise or loiter performance of an aircraft is intimately tied to its wing loading and its thrust-to-weight ratio. Paradoxically, mission performance is often not considered when these fundamental aircraft parameters are determined in conceptual design. In this paper, the traditional constraint diagram is extended to include contours of range or endurance parameter. These performance metrics represent the mission-performance capability of the aircraft without sizing the aircraft to a particular mission. This gives the designer an immediate and intuitive understanding of the tradeoff between the point and mission performance of the aircraft. The potential freedom for the designer to choose the operating condition of the aircraft is also considered. This improved constraint diagram presents the designer with a more complete basis for understanding and weighing the consequences of decisions. Through improved insight, the secondary impact of technologies can be better understood early in the design process. [ABSTRACT FROM AUTHOR]

For inverted wings in close ground proximity, such as race car configurations, the aero-dynamic ground effect can produce local velocities significantly greater than the free-stream and the effects of compressibility may occur sooner than would be expected for a wing that is not close to a ground plane. A three-dimensional computational fluid dynamics study was conducted, involving a modified NASA GA(W)-2 LS [1]-0413 MOD inverted wing with an endplate, to investigate the onset and significance of compressibility for low subsonic Mach numbers. With the wing angle of incidence fixed, Mach numbers from 0.088 to 0.4 were investigated, at ground clearances ranging from infinite (free flight) to a height-to-chord clearance of 0.067. The freestream Mach number at which flow compressibility significantly affects the predicted aerodynamic coefficients was identified to be as low as 0.15. Beyond this point, as the compressible flow conditions around the wing result in changed pressure distribution and separation behavior, treating the flow as incompressible becomes inappropriate and leads to consistent underprediction of lift and drag. The influence on primary vortex behavior of density changes around the wing was found to be relatively inconsequential even at the higher end of the Mach scale investigated. By a freestream Mach number of 0.4 and at low clearances, local supersonic flow regions were established close to the suction peak of the lower wing surface in compressible simulations; the formation of a normal shock wave between the wing and the ground was shown to result in significant increases in separation and therefore overall drag, as well as a distinct loss of downforce. [ABSTRACT FROM AUTHOR]

Radiography testing is one of the most important nondestructive testing (NDT) techniques for aircraft wing inspection. Traditionally, radiography inspection is time and manpower consuming work. In addition, human inspection of cracks and damage based on film radiography is very subjective, inconsistent, and sometimes biased. Digital radiography allows all the benefits of producing significant increases in magnification and image clarity, as well as efficient management of digital images. Therefore, it is desirable to develop a computer aided system to assist interpretation of radiographic images to increase the objectivity, accuracy, and efficiency of radiographic inspection. In this article, an automated crack detection algorithm using region of interest (ROI) minimization and improved Canny edge detection techniques is developed for aircraft wing inspection. Cracks of up to 2 mm long on the aircraft wing around the fastener holes are detected automatically. [ABSTRACT FROM AUTHOR]

The article presents a study which aims to pinpoint the optimal location of the leading-edge sweep break point. It states that after the range of the potential optimum has been bracketed, an additional point was examined. It found that the strake which has the leading-edge break point located at 37.5% of the stake root chord achieved the best performance.

Two approaches are examined for finding the best stacking sequence of laminated composite wing structures with blending and manufacturing constraints: smeared-stiffness-based method and lamination-parameter-based method. In the first method, the material volume is the objective function at the global level, and the stack shuffling to satisfy blending and manufacturing constraints is performed at the local level. The other method introduced in this paper is to use lamination parameters and numbers of plies of the predefined angles (0, 90, 45, and -45 deg) as design variables with buckling, strain, and ply percentage constraints while minimizing the material volume in the top-level optimization run. Given lamination parameters from the top-level optimization as targets for the local level, an optimal stacking sequence is determined to satisfy the global blending requirements. On a benchmark problem of an 18-panel wing box, the results from these two approaches are compared to published results to demonstrate their potential. [ABSTRACT FROM AUTHOR]

The article investigates the twisteron and circulation-controlled (CC) wing technologies which could allow an aircraft to fly efficiently in both the high- and low-speed regimes. It describes a method for reducing the induced drag generated on a CC wing by varying the blowing along the span of the wing. Computational fluid dynamics (CFD) simulations of flows over CC airfoils and wings were performed.

The mechanisms responsible for the wakelike and jetlike axial flows of a tip vortex were investigated experimentally at Re = 3.07 x 106. Both square and round tips were tested. Along the wing tip, a pocket of a higher- than-freestream, or jetlike, fluid was continually present, regardless of the tip condition and airfoil angle of attack α. Its maximum velocity increased with α and always exhibited a peak value around the trailing edge. In the near field, this jetlike fluid pocket was entrained by the shear layers and the wing wake, and it resulted in a wakelike axial flow for smaller α. For higher α, the jetlike fluid pocket was, however, surrounded by the shear layers that, in turn, protected it from the destructive effects of the wing wake and remained jetlike. The switchover angle at which the axial flow switched from wakelike to jetlike appeared around 7 deg, regardless of the tip condition, which also coincided with the angle at which the lift-to-drag ratio was a maximum. Finally, the round-tip-produced vortex was more concentrated and had a higher peak tangential velocity and core circulation but a smaller core radius compared with those of the square tip. [ABSTRACT FROM AUTHOR]

The pitching and plunging motions of airfoils have received a lot of attention recently, due to the increased interest in the design of micro air vehicles. The use of combined pitch-plunge motion with phase difference between them has often been used for the generation of thrust and lift. These aerodynamic forces could be significantly enhanced under similar operating conditions by using generalized pitch motion with variable center of wing rotation. The current study investigated the flowfield and aerodynamic forces for this generalized pitching motion. Two-dimensional rigid airfoils were taken as prototypes of micro air vehicle wings. First, the computational results were compared with the available measurements for an SD 7003 airfoil in pure-pitch and pure-plunge motions at Re = 10, 000. Good agreement was observed between the numerical computations and the experimental results in terms of streamwise velocity, location of the vorticity contours, and wake profiles. Next, the pure-pitch case was considered with the stationary centers of rotation located at different positions along the chord of the airfoil, it was found that the maximum value of the computed average coefficient of lift was obtained when the pitching axis was positioned at either the leading edge or the trailing edge. The generalized pitching motion computations were performed next. It was observed that a phase difference of 90 deg between the pitching motion and the motion of the axis caused a twofold increase of the mean coefficient of lift compared to the pitching about leading edge and combined pitch- plunge motion with a 90 deg phase difference. The stability of the leading-edge vortex was found to be responsible for the enhancement of lift by reduction in pressure at the upper surface of the airfoil. However, thrust force was not generated by applying the generalized pitching alone, whereas it was generated by the combined pitch-plunge motion. Finally, a generalized pitching motion combined with a superimposed plunging motion was studied. It was found that for this motion, thrust was generated and the generated lift was higher than that for the generalized pitching motion. This result may help in the use of superposition of kinematic motions of wings to produce the desired amount of lift and thrust. [ABSTRACT FROM AUTHOR]

Inverse design methods are elegant and efficient tools for the design of aerodynamic shapes like airfoils, wings, and propellers. In the case of propeller design, only a few general parameters have to be prescribed and the design procedure returns the propeller geometry in terms of chord-length and blade-angle distributions along the radius. The common design procedures require the prescription of the lift coefficient over the blade radius in order to obtain the unknown chord length. In this paper, a modified design method is presented that allows for designing optimum propellers having a prescribed chord-length distribution. Now the radial distribution of the lift coefficient becomes the unknown, which is determined by the design procedure. [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]

This paper presents the results of recent investigations into the aerodynamics of simulated runback ice accretion on airfoils. Aerodynamic testing was performed on a full-scale, 72-in.-chord (1828.8-mm-chord), NACA 23012 airfoil model over a Reynolds number range of 4.7 × 106 to 16.0 × 106 and a Mach number range of 0.10 to 0.28. A high-fidelity ice-casting simulation of a runback ice accretion was attached to the model leading edge. For Re = 16.0 × 106 and M = 0.20, the artificial ice shape decreased the maximum lift coefficient from 1.82 to 1.51 and decreased the stalling angle of attack from 18.1 to 15.0 deg. In general, the iced-airfoil performance was insensitive to Reynolds and Mach number changes over the range tested. Aerodynamic testing was also conducted on a quarter-scale NACA 23012 model [18 in. (457.2 mm) chord] at Re = 1.8 × 106 and M = 0.18, using low-fidelity geometrically scaled simulations of the full-scale casting. It was found that simple two-dimensional simulations of the upper- and lower-surface runback ridges provided the best representation of the full-scale, high-Reynolds-number, iced-airfoil aerodynamics. Higher-fidelity simulations of the runback ice accretion that included geometrically scaled three-dimensional features resulted in larger performance degradations than those measured on the full-scale model. Based upon this research, a new subclassification of spanwise-ridge ice is proposed that distinguishes between short and tall ridges. This distinction is made in terms of the fundamental aerodynamic characteristics as described in this paper. [ABSTRACT FROM AUTHOR]