Search

Within a conventional aircraft design process, the horizontal tail and vertical tail are generally sized via volume coefficient methods. In this manuscript, an improved method for conceptual aircraft tail design based on multidisciplinary design optimization (MDO) approach with stability and control constraints has been developed. To develop this method, first, the tail design requirements have been derived from the regulations and the fundamental functionalities of tail plans. Then, the empennage design is formulated as an MDO problem. Eventually the design optimization of horizontal and vertical tail is combined with the design optimization of the aircraft wing. A test case is presented for concurrent wing and tail plane design, which resulted in more than 9% reduction in aircraft block fuel weight and more than 3% reduction in aircraft maximal takeoff weight, which indicates a great potential for fuel burn and carbon reductions with empennage design optimization at conceptual aircraft design phase. [ABSTRACT FROM AUTHOR]

A realistic industrial conceptual design optimization problem for commercial transport airplanes was formulated with reasonable fidelity and comprehensiveness by selecting appropriate design parameters, constraints, and objectives, in order to provide a baseline to facilitate research on developing robust and efficient optimization methods for the industrial conceptual design of such airplanes. As a sample problem, a multiobjective simultaneous optimization of the design parameters for two types of civil passenger transport airplanes that constitute a family, with identical wing and tail geometries but different performance specifications, was performed using a genetic algorithm coupled with a constraint-handling technique. The results indicated that a realistic industrial conceptual design optimization of commercial transport airplanes, including simultaneous optimization of family airplanes, could be performed with the formulation. The findings from the sample optimization were also presented. [ABSTRACT FROM AUTHOR]

Conventional aircraft use discrete flight control surfaces to maneuver during flight. The gaps and discontinuities of these control surfaces generate drag, which degrades aerodynamic and power efficiencies. Morphing technology aims to replace conventional wings with advanced wings that can change their shape to control the aircraft with the minimum possible induced drag. This paper presents MataMorph-3, a fully morphing unmanned aerial vehicle (UAV) with camber-morphing wings and tail stabilizers. Although previous research has presented successful designs for camber-morphing wing core mechanisms, skin designs suffered from wrinkling, warping, or sagging problems that result in reduced reliability and aerodynamic efficiency. The wing and tail stabilizers of MataMorph-3 feature hybrid ribs with solid leading-edge sections that house servomotors, and compliant trailing-edge sections with integrated flexible ribbons that are connected to the servomotors to camber-morph the ribs. Thin laminated carbon fiber composite skin slides smoothly over the compliant rib sections upon morphing, guided by innovative trailing-edge sliders and skin-supporting linkage mechanisms strategically located between the ribs. Sample prototypes were built and tested to show the effectiveness of the proposed design solutions in enabling smooth camber-morphing. The proposed design provides a better alternative to stretchable skins in morphing airplane designs through the concept of skin sliding. [ABSTRACT FROM AUTHOR]

A multidisciplinary and multi-objective optimization approach that integrates the design of the control surfaces' sizes, active control system, and estimator for an aircraft's wing with three control surfaces is developed in this paper. Four objectives are considered: minimizing impacts of external gust loads, maximizing stability robustness, reducing control energy consumption, and minimizing the Frobenius norm of the estimator gains. For simulation purposes, a mathematical model of a flexible wing having three ailerons is used. The control system and observer are designed simultaneously. The optimization problem is formulated and solved by NSGA-II (non-dominated sorting genetic algorithm II). The solution of the optimization problem is called the Pareto set and the corresponding set of function evaluations is called Pareto front. The properties of the Pareto set and Pareto front; sensitives of the dominant poles of the open-loop system, closed-loop system, and estimator to the airspeed; and responses of the controlled, uncontrolled, and observer models at selected objective values are obtained. The results shows that the simultaneous design of the control and estimator algorithms, and the geometry of the ailerons in the multi-objective settings is very effective, the closed-loop control system can suppress the flutter and stabilize the system, and the estimator converges very quickly and always stable regardless of the air stream velocity. [ABSTRACT FROM AUTHOR]

Technology of photovoltaic cells and lithium batteries is developing rapidly. As a result, more and more attempts are made to build solar high altitude long endurance airplanes. Unfortunately, data on altitude impact on photovoltaic cells and batteries performance are not easily available. Moreover, acquisition cost of cells is still high. As a result, high altitude long endurance airplanes design is expensive and risky. Therefore, a tool for inexpensive testing of cells is needed. A small and very light unmanned aerial vehicle can be used for this purpose. It could fly as high as the envisaged high altitude long endurance airplane with a small number of cells and batteries, providing valuable information on them. The weight of such an experimental unmanned aerial vehicle could be minimized because long endurance would not be required, so heavy load of lithium batteries could be minimized, reducing also weights of other components. Wings of this unmanned aerial vehicle should enable installation of various types of photovoltaic cells including rigid ones. Therefore, it would be advantageous to apply an airfoil with a flat-upper-surface as large as possible. Unfortunately, flat-upper-surface airfoils are not popular in airfoils catalogs. Therefore, an attempt was undertaken to design an airfoil with 75% flat upper surface. The research focused on maximization of the lift-to-drag ratio and power factor assuming low Reynolds numbers conditions since it was designed for a small unmanned aerial vehicle for photovoltaic cells testing. This paper contains description of design methodology, design assumptions, and the obtained results. Moreover, the authors describe the experiment undertaken to verify the design. The wind tunnel and a semi-span model used for this experiment are presented together with the obtained results. The model has a similar structure to the envisaged structure of unmanned aerial vehicle, so flexibility of the wing is taken into account. [ABSTRACT FROM AUTHOR]

In this paper, the current state of the art of aircraft wing mass estimation methods is reviewed. The phases of aircraft design and the development process are discussed. The open literature on the subject of wing mass estimation methods and their applications in the aerospace industry is discussed, and relevant data are presented to provide the reader with background information on the field. Special attention is given to classifications of wing mass estimation methods. Current challenges and technological difficulties in wing mass estimation methods are identified, and perspectives are drawn and used to propose several key ideas for future research in the field. [ABSTRACT FROM AUTHOR]

This paper describes modeling design procedure of a flying wing UAV which is a tailless fixed wing aircraft and has no definite fuselage, with respect of the crew, payload and equipment being housed inside the main wing structure. In this thesis a 6 degree-of-freedom mathematical model describing the aircraft dynamics is first presented, then using these equations, the derivatives of the parameters and system identification of simplified, linear lateral and longitudinal models are estimated for the tailless aircraft. A detailed modeling procedure of flying wing UAV and stability analysis results using the linearized model at trim condition are represented. Finally, we have designed the flying wing UAV using classical control theory. [ABSTRACT FROM AUTHOR]

There is currently a large interest towards unconventional aircraft configurations. According to many experts, only revolutionary concepts can meet the demands of a sustainable aviation. Within this scenario, studies and investigations on innovative aircraft layouts have been flourished: Blended Wing Body, Joined Wings and Box Wings are relevant examples of configurations departing from the traditional monoplane. Joined Wings are characterized by strong geometric nonlinear effects, whose complexity is enhanced by the over-constrained layout and the intrinsic relevant bending-torsion coupling. Thus, the application of standard linear analysis tools of routine use in the aerospace industry is neither acceptable nor practical: the preliminary design phase needs a large number of simulations that have to be computationally fast and at the same time preserve an acceptable level of accuracy/fidelity. For this reason, reduced order models are particularly relevant for the design of these systems. However, several reduced order methods applied in the past to Joined Wings did not provide satisfactory accuracy and reliability. The poor performance is mainly due to the very early onset of nonlinear effects, even at very moderate deflections. This work provides an assessment, and an algorithm enhancement when needed, of several existing techniques aimed to build an effective reduction basis for model order reduction. In particular, Vibration Modes, Modal Derivatives, Ritz Vectors, Static Modes, Trial Solutions and Corrections Vectors are critically examined when applied to a representative JW configuration. The results indicate that Ritz Vectors and Modal Derivatives provide excellent accuracy of the reduced solution when compared with the full order solution. Specifically, a drastic increase in the performance of the reduced order model is noticed when ab-initio information regarding Modal Derivatives are included. This highlights and confirms the relevance of nonlinearities even at the early deformation stages. Hence, taking into account nonlinearities since the early design phases seems to be an unavoidable necessity for Joined Wings. As a consequence, efficient and reliable reduced order modeling might play an essential role in the design of such innovative configurations. [ABSTRACT FROM AUTHOR]

Purpose The aim of this study is to assess and describe possible consequences of a bird strike on a Polish-designed PZL-106 Kruk agricultural aircraft. Due to its susceptibility to such events, a wing slat has been chosen for analysis.Design/methodology/approach Smooth particle hydrodynamics (SPH) formulation has been used for generation of the bird finite element model. The simulations were performed by the LS-Dyna explicit finite element analysis software. Several test cases have been analysed with differing parameters such as impact velocity, initial velocity vector direction, place of impact and bird mass.Findings Results of this study reveal that the structure remains safe after an impact at the velocity of 25 m/s. The influence of bird mass on slat damage is clearly observable when the impact velocity rises to 60 m/s. Another important finding was that in each case where the part did not withstand the applied load, it was the lug where first failure occurred. Some of the analysed cases indicated the possibility a consequent wing box damage.Practical implications This finding provides the manufacturer an important insight into the behaviour of the slat and suggests that more detailed analysis of the current lug design might improve the safety of the structure.Originality/value Even though similar analyses have been performed, they tended to focus on large transport aircraft components. This investigation will enhance our understanding of structural response of small, low-speed aircraft to a bird impact, which is a realistic scenario for the chosen case of an agricultural plane. [ABSTRACT FROM AUTHOR]

Current economic constraints and environmental regulations call for design of more efficient aircraft configurations. An observed trend in aircraft design to reduce the lift induced drag and improve fuel consumption and emissions is to increase the wing aspect-ratio. However, a slender wing is more flexible and subject to higher deflections under the same operating conditions. This effect may lead to changes in dynamic behaviour and in aeroelastic response, potentially resulting in instabilities. Therefore, it is important to take into account geometric non-linearities in the design of high aspect-ratio wings, as well as having accurate computational codes that couple the aerodynamic and structural models in the presence of non-linearities. Here, a review on the state-of-the-art on non-linear aeroelasticity of high aspect-ratio wings is presented. The methodologies employed to analyse high aspect-ratio wings are presented and their applications discussed. Important observations from the state-of-the-art studies are drawn and the current challenges in the field are identified. [ABSTRACT FROM AUTHOR]

Many general aviation aircraft and commercial aviation aircraft are manufactured from high-polymer composite materials. The increasing use of polymer composite materials adds a time dimension to existing flutter analyses of an aircraft wing. Polymer composites exhibit viscoelastic material behaviors such as energy dissipation and a memory effect that influence the physical response of a wing. The proposed theory, derived from Lyapunov stability principles, predicts the time component of flutter, and provides a set of conditions for viscoelastic structural instabilities in general. The addition of the load history in the stability analysis of a structure is a significant change compared to elastic materials, and no standards exist for assessing the lifetime stability of a viscoelastic structure. Rigorous comparisons to the Goland wing, wings with and without stores, and a flying wing illustrate the practical applications of the theory during aircraft design. Further, the theory predicts the time to flutter of each wing for a range of representative viscoelastic materials and flow conditions. The existence and prediction of the time to flutter is the key, original contribution of the theory because no other method is available to predict the time to flutter. [ABSTRACT FROM AUTHOR]

Superplastic forming or gas forming is favorable method for the manufacture of sheet metal for making airplane skin panels. However, a major challenge in applying this method is controlling the thickness distribution in the formed product. Here, we explore a hot-bending-assisted process for gas forming of a deep trough with a complex shape for use in airplane wings. We propose the deliberate arrangement of wrinkles to achieve uniform thickness distribution. Because full-scale experiments are costly, sub-scale specimens are preferred for collecting experimental data. In linking sub-scale data to full-scale manufacturing, the size effect on wrinkle formation sites is important. The cause of the pre-form contour and larger-size specimen could be the greater variations in contour modes due to plastic buckling during the pre-forming stage. [ABSTRACT FROM AUTHOR]

In this paper, the experimental results of an unconventional joined-wing aircraft configuration are presented. The test model uses two different wings, forward and rear, both joined in tandem and forming diamond shapes both in plant and front views. The wings are joined in such a way that it is possible to change the rear wing dihedral angle values and the rear wing sweep angle values in 25 different positions that modify the relative distance and the relative height between the wings. To measure the system aerodynamic coefficients it is necessary to perform wind tunnel tests. The data presented corresponds to the lift, drag and induced drag aerodynamic coefficients, as well as the aerodynamic efficiency and the parameter for minimum required power, from the calculated values of the lift and drag time series measured by a 6-axis force and torque sensor. The results show the influence on the aerodynamic coefficients of the rear wing sweep and dihedral angles parameters. As a main result, it can be concluded that, in general terms, the lift and induced drag aerodynamic coefficients values decrease as both the distance and height between the wings increase, on the other hand, the total drag aerodynamic coefficient decreases if the distance between the wings increases, but nevertheless shows a slight tendency to increase if the height of the rear wing increases, whereas the aerodynamic efficiency and the parameter for minimum required power increase if the distance between the wings increases. [ABSTRACT FROM AUTHOR]

Purpose This paper aims to compare three closed non-planar wing configurations with a reference conventional wing-plus-horizontal tail aircraft, considering structural aspects, weights and aerodynamic characteristics, as well as operational issues, such as cruise performance.Design/methodology/approach A vortex lattice code is used and coupled with an in-house code for structural beam calculation subroutine to evaluate the configurations as a function of the four main parameters identified in the study.Findings The study concludes that the non-planar wing configurations have better performances than a conventional aircraft. Moreover, the joined-wing configuration seems to be better than the others, including the box-wing configuration, achieving an increase of 17 per cent in the range for maximum payload compared to the reference aircraft and a 3 per cent reduction of maximum take-off weight.Research limitations/implications In the study, characteristic tools for a conceptual design are used, and, thus, absolute results should be considered with caution. Nonetheless, as all the cases are studied in the same way, there is a good precision in comparative or relative results.Practical implications The work shows that the non-planar wing configurations can be used as an alternative to the conventional aircraft to meet the objectives for the future of the aviation industry.Social implications Non-planar wing configurations are able to reduce fuel consumption. Their use could lead to reductions in pollutant emissions and the impact on the environment of commercial aviation.Originality/value This study considers aerodynamic and structural aspects at the same time, as well as several non-planar wing configurations, making possible to obtain a more realistic comparison between them. [ABSTRACT FROM AUTHOR]

The goal of this paper is to provide a numerical flow analysis based on RANS equations in two directions: the study of augmented high-lift system for a cross-section airfoil of a wing up to transonic regime and the circulation control implemented by tangentially blowing jet over a highly curved surface due to Coanda effect on a rotor blade for a wind turbine. This study were analyzed the performance, sensitivities and limitations of the circulation control method based on blowing jet for a fixed wing as well as for one in rotation. Directions of future research are identified and discussed. [ABSTRACT FROM AUTHOR]

The strong variations in lift that occur with changes in forward speed lead to one of the fundamental difficulties in aircraft design: how to provide sufficient lift at landing and take-off without having oversized wings for cruise conditions. This fundamental problem is generally tackled using two approaches. First, by providing flaps and other high-lift devices, extra lift can be generated during landing and take-off, albeit at the cost of extra drag and complexity. Second, by using long and smooth runways, higher landing and take-off speeds can be tolerated, closing the gap between these speeds and those of operational flight. Even so, it is common for aircraft flying in the cruise condition to be operating with rather small main wing angles of attack (AoAs) compared with those at stall, implying that smaller wings would be desirable if acceptable landing and take-off could be achieved. A number of designers have attempted to tackle this problem with various forms of powered lift augmentation. This paper re-examines the idea of the Custer wing duct, also known as a channel wing, here applied to small unmanned air vehicles (UAVs). Such aircraft are generally not operated from long smooth runways and rarely have complex high-lift systems in their wings. It is shown that by using suitable ducts around the propellers, startlingly good take-off and landing performance can be achieved, and that suitable ducts can be readily incorporated into small UAVs with the use of 3D printing (selective laser sintering) for their manufacture. Computational fluid dynamics (CFD) analysis, wind tunnel tests, and flight trials of a Custer channel wing UAVare described. [ABSTRACT FROM AUTHOR]

To improve the efficiency of wing structural optimization in the aircraft preliminary design, an equivalent finite element model and a three-step optimization strategy were developed, taking the displacement, stiffness, buckling, and flutter requirements into account. The finite element model of a wing is reduced by the equivalent-strength-and-stiffness method instead of modeling the stiffener. It is beneficial to parameterize the modeling process for the wing primary structure. The composite wing of a regional aircraft is used to verify the equivalent finite element model with the static, buckling, and flutter analyses. The accuracy of this method is demonstrated through the comparison with the detailed finite element model. Then, the three-step optimization strategy was performed for the same case. The results show that the optimization strategy for the composite-wing design is efficient, and the equivalent finite element model in conjunction with the rapid modeling is suitable for the preliminary structural optimization with the static and dynamic constraints. [ABSTRACT FROM AUTHOR]

Purpose – The purpose of this paper is to present a structural design and optimization module for aircraft structures that can be used stand-alone or in a high-fidelity multidisciplinary design optimization (MDO) process. The module is capable of dealing with different design concepts and novel materials properly. The functionality of the module is also demonstrated. Design/methodology/approach – For fast sizing and optimization, linear static finite element (FE) models are used to obtain inner loads of the structural components. The inner loads and the geometry are passed to a software, where a comprehensive set of analytical failure criteria is applied for the design of the structure. In addition to conventional design processes, the objects of stiffened panels like skin and stringer are not optimized separately and discrete layups can be considered for composites. The module is connected to a design environment, where an automated steering of the overall process and the generation of the FE models is implemented. Findings – The exemplary application on a transport aircraft wing shows the functionality of the developed module. Originality/value – The weight benefit of not optimizing skin and stringer separately was shown. Furthermore, with the applied approach, a fast investigation of different aircraft configurations is possible without constraining too many design variables as it often occurs in other optimization processes. The flexibility of the module allows numerous investigations on influence of design concepts and failure criteria on the mass and layout of aircraft wings. [ABSTRACT FROM AUTHOR]

The aviation industry is dominated by the domain of heavier-than-air, fixed-wing, subsonic flight, and central to any design in this domain is the wing itself. One of the earliest debates in aviation still centres around the usefulness of the wing volume. On the one hand it is held that the wing, as an inevitable necessity, should provide the volume also for the payload. On the other, it is argued that more efficient wings do not even have sufficient volume for the entire wing structure. This work proposes precise definitions of the Wing Density and the Inflation Factor, two parameters that can quantitatively reflect the economic and technological trends in aviation. The wing volume of a hypothetical Ideal Wing is derived from the Operational Parameters of any given Flight Objective and compared to the volume requirement of that flight objective. We conclude that the dominant aircraft configuration of the future is likely to remain within the same family of the current dominant configuration, in conflict with some older predictions. [ABSTRACT FROM PUBLISHER]

Over the last few decades, the fuel efficiency of aircrafts was mainly improved by the application of refined aerodynamics, new materials, structural optimization and enhanced engine performance. A further potential for a greener aircraft is seen in the unconventional blended-wing-body (BWB) design configuration due to its advanced aerodynamic design and its reduced weight. For the certification of novel aircraft configurations, the airworthiness authorities require proof of at least equivalent safety standards compared to the existing conventional transport aircraft. In this context, the BWB configuration has to demonstrate sufficient crashworthiness for emergency landing on rigid surface and on water. Since structural modifications for improved safety should be applied in the early conceptual design phase of an aircraft, explicit simulations have been performed with the objective to estimate the crash behaviour of the BWB configuration. Based on the simulation results, design principles are derived for improved crashworthiness and ditching behaviour for the BWB aircraft configuration. [ABSTRACT FROM PUBLISHER]

An experimental investigation was conducted into the nature of the vortex shedding generated by truncated and nonflat serrated trailing edges of a NACA 0012 wing section as well as the aerodynamic performance of such trailing edges. In line with previous findings, the truncated trailing edge generates a significant amount of vortex shedding, while increasing both the maximum lift and drag coefficients, resulting in an overall reduction in the maximum lift-todrag ratio compared to a plain NACA 0012 wing section. By decreasing the chevron angle (p of the nonflat trailingedge serrations (i.e., by making them sharper), the energy of the vortex shedding significantly decreases, and the liftto- drag ratios increase compared to a plain NACA 0012 wing section. Fractal/multiscale patterns made of scaleddown repetitions of these serrations were also investigated with a view to further improve performance. It was found that the energy of the vortex shedding increases with an increasing fractal iteration if the chevron is broad (φ≈65 deg) but decreases for sharper chevrons (φ = 45 deg). It is believed that if is too big, the multiscale trailing edges are too far away from each other to interact and break down the vortex shedding mechanism. Fractal/multiscale trailing edges arc also able to improve aerodynamic performance compared to the NACA 0012 wing section. [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 a study which examines the effect of blockage on airfoils and wings of airplanes at high angles of attack. Topics discussed include the correction of the generated data for wall constraint, its comparison with those from other studies similarly corrected and the effect of aspect ratio. Also mentioned is the semi-empirical method used to estimate the lift coefficient at high angles of attack, based on the Rayleigh-Kirchhoff theory.

Numerical and experimental analyses of the aerodynamic performance of a triangular airfoil in low-Reynolds-number compressible flow are performed. This airfoil is one of the candidates for propeller blades on a possible future Martian air vehicle design. Based on past experimental studies conducted in the Mars Wind Tunnel at Tohoku University, this airfoil is known to exhibit nonlinear lift behavior. In the present study, direct numerical simulations of low-Reynolds-number compressible flow over a spanwise periodic triangular airfoil are conducted to identify the source of nonlinear lift. The numerical results reveal that the source of the nonlinear aerodynamic behavior is the lift enhancement provided by the large leading-edge vortex generated. For compressible low-Reynolds-number flow, the wake structure becomes elongated, causing the nonlinear lift enhancement to appear at higher angles of attack compared to the case of incompressible flow'. [ABSTRACT FROM AUTHOR]

The noise generated by a rod-airfoil configuration is investigated using unstructured Large-Eddy simulation coupled with a Ffowcs-Williams and Hawkings technique. The detailed experimental database and several numerical simulations available enable an extensive validation of the proposed methodology. Similar or improved results are obtained both in the near field (surface coefficient distributions and velocity profiles) and in the acoustic far field (power spectral densities obtained with both porous and solid surfaces, and directivities) compared with the best and most recent simulations. Dipolar noise radiation at the rod vortex-shedding frequency and its harmonics are found to be dominant. Some additional quadrupolar source contributions are seen at low and high frequencies. The rod is shown to have a significant contribution to the far-field noise at grazing angles. Constructive interferences occur normal to the rod-airfoil axis, whereas destructive interferences appear at grazing angles. [ABSTRACT FROM AUTHOR]

Wind-tunnel measurements were used to identify and characterize the low-frequency oscillation present in the flow field about an NACA 0012 airfoil with a horn-ice shape. This low-frequency mode was identified in the unsteady content of the airfoil lift coefficient and the corresponding Strouhal numbers compared well with those reported in the literature. To study the relationship between the ice-induced separation bubble dynamics and the airfoil circulation, the unsteady shear-layer reattachment location was determined using measurements from a surface-mounted hotfilm array. At the low-frequency mode, the lift coefficient led the reattachment location by a phase of n/2. Additional characterization of the relationship between the unsteady shear-layer reattachment location and the unsteady airfoil pressure distribution was also completed, along with an analysis of the convective qualities of the low-frequency oscillation throughout the airfoil pressure distribution. [ABSTRACT FROM AUTHOR]

A low-order phenomenological model relating dynamic wing motion to thrust production of micro flapping wings with different stiffness was examined. The motivation is toward closed-loop flight control of flapping micro air vehicles. The low-order model was based on a "black box" state-space description that used the intraperiod flappingwing motion as input to determine modeling coefficients that minimized the error between the predicted and actual intracycle thrust. The data time series used in the modeling were provided by a companion experiment with three flapping wings of different designs and stiffness. The wings were fabricated from laminated carbon-fiber strips that were covered by a thin polyvinylidene fluoride film sheet. The wings had a span of 12.5 cm and were attached to a commercial Cybird gearbox that could be operated at different flapping frequencies. Digital high-speed video was used to determine the dynamic displacement of the wing structure. The intraperiod wing displacement was quantified using digital image processing of each video frame. The low-order model successfully captured the dynamic thrust generation within a flapping cycle and produced the correct scaling of the peak thrust with flapping frequency for all three wing designs. The model mean thrust frequency scaling agreed with the experiments if the peak thrust was within approximately 40% of the cycle-averaged thrust. A modification of the model is suggested when that difference is too large. [ABSTRACT FROM AUTHOR]

In this paper, a droplet tracking method and a splashing model have been developed to calculate the droplet collection efficiency in a supercooled-large-droplets regime using the Lagrangian computational method. In the droplet tracking method, the droplet deformation and the droplet-wall effects (for example, splashing, bouncing, and reimpingement), which are the typical cases in a supercooled-large-droplets regime, are incorporated by introducing the mass residual ratio. The effects of the transition from the conventional-small-droplets impingement to a supercooled-large-droplets impingement, as well as the splashed secondary droplets, on the droplet collection are considered in the current splashing model. The performance and capacities of the droplet tracking method and the supercooled-large-droplets splashing model are validated against the alternative experimental reference data. The mass loss ratio and the mass back ratio are introduced in order to explore the distribution and the quantity of the mass loss and mass back caused by droplet splashing and reimpingement. The predicted results show that the quantity and the distribution range of the mass back ratio on airfoil surfaces are relatively lower than those of the mass loss ratio. A significant mass back is observed when the airfoil is contaminated with ice. No mass loss or mass back is observed beyond the impinging region for the given conditions. [ABSTRACT FROM AUTHOR]

Many current and future aircraft designs rely on the wing or other aircraft surfaces to shield observers on the ground from the engine noise. However, the available data showing how surfaces interact with a jet to shield and/or enhance the jet noise are currently limited. Therefore, far-field noise data and background-oriented schlieren images were acquired for a round jet, operating in the overexpanded, ideally expanded, and underexpanded supersonic flow regimes, near a planar surface to investigate how airframe surfaces might affect the shock-cell structure in the jet plume and the broadband shock noise produced. These data show that broadband shock noise is produced by the relatively weak shocks far downstream of the nozzle exit; consequently, a surface will be effective at reducing broadband shock noise only if it is long enough to shield the noise produced by shocks. Furthermore, the presence of a surface very near the edge of an underexpanded jet increases the shock-cell spacing, pushing the shock cells farther downstream. Conversely, the surface has a minimal affect on the shock cells in an overexpanded jet. [ABSTRACT FROM AUTHOR]

Boundary-layer stability analyses of mean flows extracted from unstructured-grid Navier-Stokes solutions have been performed. A procedure has been developed to extract mean flow profiles from the FUN3D unstructured-grid solutions for the purpose of stability analysis. Extensive code-to-code validations were performed by comparing the extracted mean flows as well as the corresponding stability characteristics to the predictions based on structured-grid mean flow solutions. Comparisons were made for a set of progressively complex geometric configurations ranging from a simple flat plate to a full aircraft configuration: a modified Gulfstream III with a natural laminar-flow glove. The results for the swept wing flow over the wing-glove assembly point to the need for stability analysis based on Navier-Stokes solutions or possibly fully three-dimensional boundary-layer codes when the underlying flow develops strong three-dimensionality. The effect of grid resolution, mean flow convergence, and low-order interpolation to a stability grid on metrics relevant to linear stability of the boundary-layer flow are also examined to provide guidelines for the use of both structured and unstructured grids in practical applications related to transition prediction for swept wing boundary layers. [ABSTRACT FROM AUTHOR]

This study demonstrates an autonomous formation flight of multiple flapping-wing flying vehicles for the first time. In order to extract the position and attitude information of each flying vehicle, an external motion capture system is applied. A path-following controller is designed so that each flapping-wing flying vehicle has a time-independent circular path with the predetermined radius, and the constant forward flight speed and altitude. Due to the variations in flight characteristics among four different flapping-wing flying vehicles, the gain of the directional control, which generates nonlinear rolling moment using the changes of the tension in flexible membrane wings, is individually tuned for each vehicle. The rotational speed of the circular path is kept constant by controlling the desired radius rather than the forward flight speed so that the formation angles between two closest flapping-wing flying vehicles are maintained at 90 degrees. The performances of the individual path-following controllers and the circular formation controller are statistically evaluated in terms of the mean and standard deviation of the flight state variables. [ABSTRACT FROM AUTHOR]

The article discusses a study which investigated the effect of a hot-wire probe and its traverse on flow development over a NACA 0018 airfoil for various positions and sizes of separation bubbles. Topics discussed include mean surface pressure distributions for all the cases investigated, causes of bubble bursting, and energy spectra from measurements with embedded microphones.

An advanced hybrid Reynolds-Averaged Navier-Stokesllarge eddy simulation (RANSI LES) turbulence model delayed detached eddy simulation (DDES) is conducted in thispaper to investigate the dynamic stall flows over 3D NACA0012 airfoil at 17 deg, 26deg, 45 deg, and 60 deg angle of attack (AOA). The spatially filtered unsteady 3D Navier-Stokes equations are solved using a fifth-order weighted essentially nonoscillatory (WENO) reconstruction with a low diffusion E-CUSP (LDE) scheme for the inviscid fluxes and a conservative fourth-order central differencing for the viscous terms. An implicit second-order time marching scheme with dual time stepping is employed to achieve high stability and convergency rate. A 3D flat plate is validated for the DDES model. For quantitative prediction of lift and drag of the stalled NACA0012 airfoil flows, the detached eddy simulation (DES) and DDES achieve much more accurate results than the Unsteady Reynolds-Averaged Navier-Stokes (URANS) simulation. In addition to the quantitative difference, the DESIDDES and URANS also obtain qualitatively very different unsteady stalled flows ofNACA00I2 airfoil with different vortical structures and frequencies. This may bring a significantly different prediction if those methods are used for fluid-structural interaction. For comparison purpose, a third-order WENO scheme with a second-order central differencing is also employed for the DDES stalled NACA0012 airfoil flows. Both the third- and fifth-order WENO schemes predict the stalled flow similarly for lift and drag at AOA less than 45 deg, while at AOA of 60 deg, the fifth-order WENO scheme shows better agreement with the experiment than the third-order WENO scheme. The high-order scheme of WENO 5 also resolves more small scales of flow structures than the second-order scheme. The prediction of the stalled airfoil flow using DDES with both the high-order scheme and second-order scheme is overall significantly more accurate than the URANS simulation. [ABSTRACT FROM AUTHOR]

This article presents a study which investigated analytical methods for predicting loads and moments on axisymmetric ring-wing ducts. Topics discussed include an approach for modeling the moments and center of pressure for all duct airfoil geometries, an approximate model for the force coefficient and the new moment/center-of-pressure model. Analytical models for guiding design optimization studies are explored.

Based on a hierarchical overset-grid strategy, enhanced implicit-hole-cutting method has been applied to the study of viscous flows around 30P30N multi-element airfoil and the wing-body configuration of the NASA common research model. Numerical simulations were performed on overlapping grids by an in-house computational fluid dynamics solver called Exstream. The present paper conducted the grid-convergence and buffet study of the NASA common research model, which was used in the Fifth AIAA Computational Fluid Dynamics Drag Prediction Workshop. The result of the 30P30N multi-element airfoil proves that the enhanced implicit-hole-cutting method and Exstream solver could obtain great robustness and accurateness. From the result of the grid-convergence study, the value of total drag follows a nearly linear trend and seems to converge to a given value under different grid refinements. Moreover, grid density has a nonnegligible impact on the skin-friction drag component. The consequence of buffet study reveals that there is a very interesting variation of lift coefficient that it decreases from 0.599 to 0.587 when the angle of attack increases from 3.00 to 3.25 deg. The streamlines and surface-pressure contours show that a flow separation bubble appears near the trailing edge of the wing-fuselage junction. Extrafine overlapping grid, matched structured grid, and different turbulence models have been tested for the flow separation bubble. [ABSTRACT FROM AUTHOR]

The article discusses the lifting-line theory applied to infinite airfoils and how a model of a high-aspect-ratio wing near stall has managed to predict the abrupt appearance of lift cells, in qualitative agreement with experiment and computational fluid dynamics (CFD) studies. Topics cited include the issue of the uniqueness of solutions to the Navier-Stokes Equations (NSEs) and an off-design condition for a wing for which an unsteady Reynolds-averaged Navier-Stokes URANS solution may exist.

A new bilevel optimization strategy for wing design is developed, in which the optimizations of the wing-planform and wing- airfoil shapes are decoupled from each other. The design of the wing-planform shape and the shape of the airfoils in several spanwise positions are considered as the goal of the optimization. In the new approach, the design problem is decomposed into a series of subproblems based on the design variables. The design variables defining the wing-planform shape are optimized in a top-level optimization, and the design variables defining the shape of airfoils in several spanwise positions are optimized in several sublevel optimizations. To take into account the influence of the airfoil shape in a specific spanwise position on the shape of the airfoils in other spanwise positions, a series of design variables are added to the design vector of the top-level optimization. The top-level optimizer is responsible for the consistency of the optimization. Using this approach, the number of design variables in the top-level optimization is reduced; the airfoils in several spanwise positions are optimized in parallel; and, instead of complex three-dimensional aerodynamic and structural solvers, much simpler and faster two-dimensional airfoil analysis tools can be used. [ABSTRACT FROM AUTHOR]

Supercooled large droplet (SLD) is beyond the envelope described in the Appendix C of FAR 25. When the droplets impact on the surface of airfoil, they will run downstream and freeze to shape of ridge. SLD runback ice may endanger the safety of the aircraft significantly, but the aerodynamic effect of icing on vehicle performance is not completely understand. The numerical simulation of SLD icing is carried out using Reynolds Stress Model (RSM) in this paper. The computational maximum lift coefficients and stalling angles are agree with Lee's experimental data on NACA 23012m, the applicability of RSM in analysis of flow separation resulting from runback ice is valialated. The impact degree of runback ice is quite different according to airfoil geometry, and the supercritical airfoils and tail airfoils are specially analyzed in large civil aircraft. The results demonstrate that runback ice influences the flow separation and significant deteriorates aerodynamic performance of supercritical airfoils. For the tail airfoils, the aerodynamic impact of SLD icing is not so severe, and the size of separation region is diminished at higher angles of attack. The contents of this paper have reference values in aircraft design and airworthiness certification for large civil 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 goal of this paper is to provide a numerical flow analysis based on RANS equations in two directions: the study of augmented high-lift system for a cross-section airfoil of a wing up to transonic regime and the circulation control implemented by tangentially blowing jet over a highly curved surface due to Coanda effect on a rotor blade for a wind turbine. This study were analyzed the performance, sensitivities and limitations of the circulation control method based on blowing jet for a fixed wing as well as for a rotating wing. Directions of future research are identified and discussed. [ABSTRACT FROM AUTHOR]