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A theoretical aeroelastic study for a flexibly suspended high-aspect ratio wing aeroelastic model excited by gust loads is presented along with a companion wind-tunnel test. A flexible support system at the wing root has been constructed to simulate a rigid body mode. Structural equations of motion based on a nonlinear beam theory are combined with the ONERA aerodynamic model. Also, a dynamic perturbation analysis about a nonlinear static equilibrium is used to determine the small perturbation flutter boundary and structural natural frequencies. The effects of the flexible support system (bi-beam system) on the wing structural dynamics, the static aeroelastic displacement at root, nonlinear flutter, and gust response to a harmonic or a frequency sweep excitation are discussed. Also the effect of gust distribution along the span on the response is described based upon computations. The fair to good quantitative agreement between theory and experiment demonstrates that the present analysis method has reasonable accuracy. DOI: 10.2514/1.J050309

This study addresses the aerodynamics of elastic membrane flapping wings. Several applications of the actuator disk theory to the flapping wings of insects and birds are reviewed. In previous studies, to account for spatial and temporal variance in the wake behind the flapping wings, empirical corrections were proposed for the induced velocity and power. In the present paper, a new procedure for determination of the correction factor is proposed, using membrane-type flapping-wing devices. Wind-tunnel experiments were conducted and the stroke-averaged propulsive thrust was measured on 25-cm-wingspan (flat and 9% camber) and 74-cm-wingspan flapping-wing models. Either flapping frequency or input power was held constant during the tests. Obtained thrust forces were compared to theoretical values predicted by the actuator disk theory. Empirical correction factors to the actuator disk theory were determined, providing a best fit to the experimental data when the flapping axis aligned with freestream velocity. It is noteworthy that the numerical value for the correction factor for the 25 cm cambered wing agrees with the results obtained on large insects. The theoretical corrections for angle of attack of the flapping wing give satisfactory agreements with the experimental data only for relatively low forward speeds. DOI: 10.2514/1.J050139

Scale model approach in thermal testing of aircraft components has been proposed in this study. The idea of thermal scaling has been validated experimentally. Quite often, it may not be possible to carry out experimental validation of design analysis at high temperatures on prototype structures at the design stage since exact simulation of thermal environments may not be possible. Under such conditions, the designer may resort to study the structural behavior using scale model testing at a different temperature where the material does not change phase at the scaled temperature. DOI: 10.1061/(ASCE)0893-1321(2010)23:1(24) CE Database subject headings: Dimensional analysis; Thermal factors; Finite element method; Tests; Aircraft. Author keywords: Similitude; Dimensional analysis; Thermal scaling.

A methodology for improving the quality of high-lift-system performance prediction within a multidisciplinary conceptual design process is presented. The high-lift-system geometry is explicitly modeled and a multiple-lifting-line method is used to compute its aerodynamic characteristics. Computation times are acceptable for use in a conceptual design process. The results for several test cases show good agreement with wind-tunnel and/or high-fidelity numerical data. In addition, the method allows for further enhancement by using nonlinear airfoil polars for interpolation, improving drag prediction, and introducing some degree of nonlinear aerodynamic behavior. DOI: 10.2514/1.42845

Following the program to test a hybrid actuation system for high-agility aircraft buffeting load alleviation on the full-scale F/A-18 vertical fin structure, an investigation has been performed to understand the aerodynamic effects of high-speed vortical flows on the dynamic characteristics of vertical fin structures. Extensive wind-tunnel tests have been conducted on a sealed model fin integrated with piezoelectric actuators and accelerometers to measure the aft-tip vibration responses under various freestream and vortical airflow conditions. Test results demonstrated that the airflow induced considerable damping to the fin structure, which generally increased with airflow speed as well as the vertical fin angle of attack relative to the airflow direction. Moreover, it was observed that at the angle of attack of 10 deg, the high-speed airflow introduced large deflection to the smart fin structure and caused significant frequency shift to the vibration modes due to nonlinear geometrical coupling of bending and torsional modes. These aerodynamic effects may adversely affect the performance and robustness of the closed-loop control laws developed based on vertical fin dynamic model identified without considering the varying aerodynamic effects. To explore this problem, the structured singular values synthesis technique was adopted to develop robust control law using smart fin model identified without aerodynamic excitations, and the aerodynamic effects on the fin structure were assumed as smart fin parametric and dynamic uncertainties. The effectiveness and robust performance of the control law was demonstrated through extensive closed-loop wind-tunnel tests using various airflow conditions. This provided a verified control law design strategy for future flight tests of the full-scale aircraft buffeting load alleviation system. DOI: 10.2514/1.42489

A series of experiments was performed at Wichita State University's water tunnel on a 70-degree-sweep delta wing using a towing mount. A video camera captured dye-flow visualization images of the vortex burst that were subsequently analyzed using a computer-assisted image analysis software. The aim was to better understand the relationship between the freestream velocity and the time constants involved in the movement of the vortex-burst point. Experiments indicated that a change in the freestream velocity changed the forward progression of the vortex burst. Under pitch-up conditions, deceleration resulted in a momentary retardation in the forward progression of the burst, whereas acceleration resulted in a faster progression toward the apex.

The application of active flow control for high-lift multi-element wing sections is investigated. A computational fluid dynamics procedure is used to simulate the interactive flow in conjunction with distributed flow control. Favorable nonlinear interactions using flow control on a conventional wing section and an advanced airfoil lead to major aerodynamic performance improvements at representative takeoff and landing conditions. This study demonstrates that judicious application of flow control at select locations on the individual wing elements may lead to near inviscid lift levels in the linear lift range and substantial gains in maximum lift. DOI: 10.2514/1.41236

The behavior of the tip vortex behind a square NACA0015 wing was numerically investigated. The problems studied include the stationary and the oscillating wings at static and dynamic stall conditions. Reynolds-averaged Navier--Stokes and detached-eddy simulation schemes were implemented. Vortex structures predicted by Reynolds-averaged Navier--Stokes were mainly diffused while detached-eddy simulation was able to produce qualitatively and quantitatively better results as compared to the experimental data. The breakup of the tip vortex, which started at the end of the upstroke and continued to the middle of the downstroke over an oscillation cycle, was observed in detached-eddy simulation data. DOI: 10.2514/1.40685

Wind-tunnel tests have been carried out on a battle-damaged NACA [64.sub.1]-412 half-wing aspect ratio of 8.2. The simulated gunfire damage had a diameter of 0.2 wing chord and was located at midchord and at one of two spanwise locations. Tests were carried out at a Reynolds number of 5.5 x [10.sup.5]. Compared with an undamaged wing, the damage resulted in reduced lift, increased drag and a positive increase in pitching moment at zero lift. Moving the damage to near the tip reduced the magnitude of these effects. Using the static pressure difference between the upper and lower surfaces of the undamaged wing allowed the data from the present study to be successfully compared with previously published drag and lift data for a two-dimensional damaged airfoil. Tests on wings with aspect ratios of 6.2 and 10.3 produced similar trends in the aerodynamic characteristics and showed that the use of static pressure difference was equally effective in allowing comparisons with two-dimensional data. DOI: 10.2514/1.39839

This paper focuses on the drag characteristics of optimally span-loaded planar, wingletted, and C wings. The span load is optimized resulting in minimum induced or total drag. The wing-root bending moment is kept constant for all analyzed wings to ascertain that different wings have comparable weight. The optimum span loadings for the different types of wing are calculated using a fast and simple numerical method. The wings are analyzed in the Trefftz plane, infinitely far behind the wing. Lagrange multipliers are used to calculate the optimum span loading resulting in minimum induced or total drag, with the wing-root bending moment and/or the lift coefficient as constraint. The induced drag can be calculated using the optimum span loading. The profile drag is assumed to be a function of the local lift coefficient. The results indicate that the C wing does not have real aerodynamic performance advantages compared to a wingletted wing. For wings with span and/or aspect ratio constraints, a winglet offers a drag reduction relative to a planar wing. DOI: 10.2514/1.39426

Shimmy of aircraft nose-wheel landing gear is a coupled lateral and torsional dynamic instability that may occur during take off, taxi, and landing. The potential for shimmy is often discovered when the aircraft undergoes taxi (runway) tests. At this stage the options available to the designer are limited, often to introducing a damper in the steering degree of freedom. This paper presents modeling and analysis of shimmy of a 3 degrees-of-freedom linear nose-wheel landing gear model with a closed-loop shimmy damper mechanism attached to its hydraulic steering. A typical control law is considered for representing the closed-loop hydraulic control in the steering rack and pinion system. The stability of the system is studied using root locus plots. The results for critical velocity obtained for different values of gain in the closed-loop shimmy damper show that it can be simulated by an equivalent open-loop shimmy damper with appropriate value of damping. This suggests that an equivalent open-loop shimmy damper model can be used in place of the closed-loop shimmy damper to simplify the modeling of integrated airframe--landing gear dynamics.

A numerical investigation of a fully three-dimensional swept-wing geometry featuring separation from a curved leading edge is presented. The implicit large-eddy-simulation strategy based on a third-order high-resolution method for discretizing the advective fluxes and a second-order Runge-Kutta time-stepping scheme with an extended stability region have been employed. No attempt to incorporate a wall model has been made. Instead, the boundary layer is fully resolved over the majority of the wing. Qualitative and quantitative comparisons with experimental oil-film visualizations and three-dimensional laser Doppler anemometry measurements show very good agreement between the experiment and the numerically predicted flow structures, as well as velocity and stress profiles near the wing. Furthermore, data from a hybrid Reynolds-averaged Navier-Stokes and large-eddy simulation have been included for comparison.

Load-path-based topology optimization is used to synthesize a compliant adaptive aircraft wing leading edge, which deforms in a prescribed way when subject to a single point internal actuation. The load-path-based optimization method requires the specification of a parent lattice. Increasing the complexity of this lattice means the number of parameters required for a complete representation of the structure in the topology optimization becomes prohibitive, although it is desirable to enable a full exploration of the design space. A new method based on graph theory and network analysis is proposed, which enables a substantial reduction in the required number of parameters to represent the parent lattice. The results from this load-path-based approach are compared with those obtained from the better-known density-based topology optimization method.

In flutter testing, it is often implicitly assumed that the most critical case at any given Mach number will occur at sea level conditions. This is a reasonable expectation, as the highest dynamic pressure at a fixed Mach number will be encountered at sea level, where the air density is highest. In the present paper, a counterexample involving a generic swept wing representative of transport aircraft is presented in which transonic limit cycle flutter is predicted to occur at altitude rather than at sea level. The calculations are based on a time-accurate Eulerian-Lagrangian finite element scheme that fully accounts for both structural and aerodynamic nonlinearities arising from large deflections and shock motion. Flutter onset is around Mach 0.84 and the limit cycle amplitudes grow until Mach 0.95 is reached, at which point they start decreasing, vanishing abruptly around Mach 0.97. The limit cycle flutter persists down to very low air densities, corresponding to altitudes above 75,000 ft, and the maximum limit-cycle-oscillation flutter amplitude does not occur at sea level, but at altitude. If sufficient structural damping is added, the wing will be flutter-free at lower altitudes but not at higher (cruise) altitudes. At Mach 0.95, the wing is stable below about 15,000 ft, even in the absence of structural damping, but experiences strong limit cycle flutter at higher altitudes.

Boeing Co. is to conduct flight test of 12 new and derivative airplane models in 2008 and the company's new Test Operations Center for centralized test control will be handling concerns on the tests and address pre-position requirements. Other details about TOC's operations are presented.

Experiments were conducted on a high-fidelity, model-scale, main landing gear in the Virginia Polytechnic Institute and State University Stability Wind Tunnel. This wind tunnel was recently upgraded to an anechoic facility, which allowed aeroacoustic measurements to be carried out in the far field and in an environment with significantly less reflections. The model was a very faithful replica of the full-scale landing gear, designed to address the issues associated with low-fidelity models. A 63-element microphone phased array was used to locate the noise-source components of the landing gear from different streamwise positions, both in the near and far fields. The same landing-gear model was previously tested in the original hard-walled configuration of the tunnel with the same phased array mounted on the wall of the test section (i.e., near-field position). The new anechoic configuration of the Virginia Polytechnic Institute and State University wind tunnel offered a unique opportunity to directly compare data collected in hard-walled and semi-anechoic test sections, using the same landing-gear model and phased-array instrumentation. Through these tests, some of the limitations associated with testing in hard-walled wind tunnels were addressed.

A systematic series of wind-tunnel tests was conducted on an ornithopter configuration consisting of two sets of symmetrically flapping wings, located one behind the other in tandem. It was discovered that the tandem arrangement can give thrust and efficiency increases over a single set of flapping wings for certain relative phase angles and longitudinal spacing between the wing sets. In particular, close spacing on the order of 1 chord length is generally best, and phase angles of approximately 0 [+ or -] 50 deg give the highest thrusts and propulsive efficiencies. Asymmetrical flapping was also studied, which consists of the two sets of wings rocking relative to one another 180 deg out of phase. It was found that the performance of such an arrangement is poor, relative to the best performing symmetrical tandem flapping. DOI: 10.2514/1.28160

The design and development of a shape memory alloy based smart landing gear for aerospace vehicles is based on a novel design approach. The smart landing gear comprises a landing beam, an arch, and a superelasfic nickel-titanium shape memory alloy element. This design is of a generic nature and is applicable to a certain class of light aerospace vehicles. In this paper a specific case of the shape memory alloy based smart landing gear design and development applicable to a radio controlled semirigid airship (radio controlled blimp) of 320 [m.sup.3] volume is presented. A judicious combination of carbon fiber reinforced plastic for the landing beam, cane (naturally occurring plant product) wrapped with carbon fiber reinforced plastic for the arch, and superelastic shape memory alloy is used in the development. An appropriate sizing of the arch and landing beam is arrived at to meet the dual requirement of low weight and high-energy dissipation while undergoing 'large elastic' (large nonlinear recoverable elastic strain) deformations to ensure soft landings when the airship impacts the ground. The soft landing is required to ensure that shock and vibration are minimized (to protect the sensitive payload). The inherently large energy-dissipating character of the superelastic shape memory alloy element in the tensile mode of deformation and the superior elastic bounce back features of the landing gear provide the ideal solution. A nonlinear analysis based on the classical and finite element method approach is followed to analyze the structure. Necessary experiments and tests have been conducted to check the veracity of the design Good correlation has been found between the analyses and testing. This exercise is intended to provide an alternate method of developing an efficient landing gear with satisfactory geometry for a 'certain class of light aerospace vehicles' such as airships, rotorcraft, and other light unmanned air vehicles. DOI: 10.2514/1.26811

This paper describes experimental investigations of possibilities to destabilize the vortex wake behind a wing by using winglets with integrated rudders. These rudders can be deflected statically, as well as oscillated periodically around their static deflection. The oscillation is intended to excite instabilities inherent in the vortex system to provoke an accelerated decay of the vortices. The used model also comprises deflectable ailerons and was equipped with a pair of trailing edge flaps to modify the vortex system. Quantitative wake investigations by particle image velocimetry have been conducted up to 43 spans behind the wing model in a water towing tank. Results show a clear earlier decay of the vortex system and abatement of induced rolling moments in the wake for certain rudder configurations in comparison to the clean wing configuration. However, the oscillation of the winglet rudders did not lead to a further acceleration of vortex decay compared to the statical case in the frame of the experiments conducted.

Trailing vortex structure of low aspect ratio wings was studied in a water tunnel at Reynolds numbers of 8000 and 24,000 using dye injection and digital particle image velocimetry in cross-flow planes in the near wake, for rectangular, semi-elliptical, and delta-wing planforms. The velocity data were used to calculate lift via circulation and effective span, and the results were compared with force balance measurements and classical inviscid theory. The objectives of the study were to assess how low-Reynolds number effects might affect the measurement of lift coefficient from tip-vortex circulation, how well the measurements fit the various theoretical models of lift curve slope for low aspect ratio wings, and the extent to which planform shape affects lift coefficient while aspect ratio and planform area are kept constant. All models were thin flat plates with square edges. For the conditions studied, apparent viscous effects on vorticity integrations in the wake were small, resulting in good agreement between lift coefficient values inferred from the measurements, the force balance data, and the classical inviscid formulas. Whereas the three wings exhibited distinctly different wake structures, lift curve slopes were similar.

The flowfield downstream of a blade interacted normally with a winglet-generated double vortex was investigated experimentally. Similar to the baseline (with no winglet) interaction, the passage of the double vortex over the blade significantly altered the blade circulation distribution, causing negative vorticity to be shed into its wake, which, together with the turbulent wake motions, acted to trigger the turbulent decay of the vortex core, resulting in an increase in its size and loss of circulation. For a junction vortex interacted directly with the blade leading edge, a pair of two counter-rotating vortices, similar to the baseline interaction case, was observed at 1.5 chords downstream of the blade trailing edge. Qualitatively, the variation in double vortex-blade separation distance produced large variations in the vortex core parameters associated with the junction vortex, whereas minor changes for the tip vortex.

Active control of separated flows has become an attractive approach enabling the designer to meet increased performance demands for various systems. Whereas a significant part of the work published so far in the area of closed-loop flow control is based on simulation studies, this paper presents an example of a successful application of a multivariable controller in wind tunnel experiments. A robust [H.sub.[infinity]] controller is used to control the spanwise variable reattachment length downstream of a benchmark problem, the backward-facing step. To reduce the conservatism of this approach, a nonlinear static precompensation is included, first. The synthesis of the controller is based on a family of identified linear black-box models, which describe the compensated input/output behavior of the plant. Tracking performance and disturbance rejection of the controller are tested in wind tunnel experiments and shown in this paper.

In this study, flow computations are performed on three-dimensional high-lift configurations on multiblock structured and unstructured meshes with two turbulence models, Spalart-Allmaras model and Menter's shear stress transport model. Two kinds of three-element trapezoidal wings with full-span flap or part-span flap are computed. The objectives of this paper are to investigate the mesh dependency and the effect of turbulence models and improve the reliability in simulating the flow around high-lift devices. Importance of the mesh resolution to resolve the separated corner flows near the wing-fuselage junction is shown to improve the accuracy of computational results. It is also shown that the maximum lift and the angle at which it occurs are very sensitive to turbulence model and Menter's model gives better results at a higher angle of attack in the present computations.

Limit-cycle oscillation was simulated for a rectangular wing referred to as the [Goland.sup.+] wing. It was found that the aerodynamic nonlinearity responsible for limit-cycle oscillation in the [Goland.sup.+] wing was shock motion and the periodic appearance/disappearance of shocks. The [Goland.sup.+] structural model was such that in the transonic flutter dip region, the primary bending and twisting modes were in phase and coupled to produce a single-degree-of-freedom, torsional flutter mode about a point located ahead of the leading edge of the wing. It was determined that the combination of strong trailing-edge and lambda shocks which periodically appear/disappear, limited the energy flow into the structure. This mechanism quenched the growth of the flutter, resulting in a steady limit-cycle oscillation. Underwing and tip stores were added to the [Goland.sup.+] wing to determine how they affected limit-cycle oscillation. It was found that the aerodynamic forces on the store transferred additional energy into the structure increasing the amplitude of the limit-cycle oscillation. However, it was also found that the underwing store interfered with the airflow on the bottom of the wing, which limited the amplitude of the limit-cycle oscillation.

The purpose of the study is to investigate the probability density function of turbulent boundary layer fluctuating pressures measured on the outer sidewall of a supersonic transport aircraft and to approximate these probability density functions by analytical models. Experimental flight results show that the fluctuating pressure probability density functions differ from the Gaussian distribution even for standard smooth surface conditions. The probability density function tails are wider and longer than those of the Gaussian model. For pressure fluctuations in front of forward-facing step discontinuities, deviations from the Gaussian model are more significant and the probability density functions become asymmetrical: There is a certain spatial pattern of the skewness and kurtosis behavior depending on the distance upstream from the step. All characteristics related to non-Gaussian behavior are highly dependent upon the distance from the step and the step height, less dependent on aircraft speed, and not dependent on the fuselage location. A Hermite polynomial transform model and a piecewise-Gaussian model fit the flight data well both for the smooth and stepped conditions. The piecewise-Gaussian approximation can be additionally regarded for convenience in usage after the model is constructed. DOI: 10.2514/1.18294

A high resolution computational methodology is developed for the solution of the compressible Reynolds-averaged Navier-Stokes equations. This methodology is used to study the effect of spanwise blowing as a method of tip vortex control. The numerical error is reduced by using high order accurate schemes on appropriately refined meshes. For vortex evolution problems, the equations are solved on multiple overset grids that ensure adequate resolution in an efficient manner. Reliable validation of the mean flowfield for the baseline and control configurations is obtained by adding a simple correction to the production term in the Spalart-Allmaras turbulence model. A detailed study of the underlying physics of the effects of spanwise blowing is presented.

Frequency response functions (FRF) measured by piezoelectric macrofiber composite (MFC) actuators/sensors are used to detect subsurface delamination in a composite plate. The plate is impacted to seed damage in the form of ply delamination. Then, the MFC-based active-sensing system exerts an excitation into the plate and measures the subsequent responses. Traditional piezoceramic materials are also mounted in comparable locations on the plate to compare their performances. FRF and damage indicator features are derived from the measured signals and used to assess the condition of the plate. Validation of the delamination is completed using an ultrasonic C-scan method. The effective area of observed damage is well correlated to the damage indicator feature.

A zonal hybrid Reynolds-averaged Navier-Stokes large-eddy simulation (RANS/LES) approach, called zonal-DES, used to handle a two-dimensional high-lift configuration with deployed slat and flap is presented. This method allows to reduce significantly the cost of an accurate numerical prediction of the unsteady flow around wings compared to a complete LES. Some issues concerning grid generation as well as the use of zonal-detached-eddy simulation for a multi-element airfoil are discussed. The basic planar grid has 250,000 points and the finest spanwise grid has 31 points with [DELTA]z/c = 0.002. The effort is geared toward detailed comparison of the numerical results with the Europiv2 experimental particle image velocimetry data including both mean and fluctuating properties of the velocity field (Arnott, A., Neitzke, K. P., Agocs, J., Sammer, G., Schneider, G., and Schroeder, A., 'Detailed Characterisation Using PIV of the Flow Around an Aerofoil in High Lift Configuration,' EUROPIV2 Workshop on Particle Image Velocimetry, Springer, Berlin, 2003). The results also provide an insight into the real unsteady nature of the flow around a three-element airfoil that cannot be reproduced by classical RANS models. The current calculation displays extremely complex flow dynamics in the slat and flap coves like the ejection process through the gaps of several vortices issued from the impingement of the free shear layers on the lower walls of the different elements.

A study of the transverse normal strain effect on the static thermoelastic response of homogeneous and multilayered plates is presented. Numerical evaluations have been given for classical, refined, and advanced zig-zag plate theories. Constant, linear, and higher-order forms of temperature profile in the plate thickness direction have been accounted for. Basic assumptions of the considered theories are quoted. The related governing equations are not given because these were derived from a unified formulation that was described elsewhere. Closed-form solutions are discussed by addressing three plate problems: a homogeneous plate made of an isotropic layer; a two-layered plate consisting of two layers made of different isotropic materials; a multilayered composite plate made of three cross-ply layers. It has been confirmed that any refinements of classical models are generally meaningless, unless the effects of transverse shear and normal strains are both taken into account in a plate theory. Furthermore, it has been found that transverse normal strains cannot be discarded even though thin plates are considered and the accurate description of the temperature profile in the plate thickness direction could result to be meaningless, unless transverse normal strains are taken into account, and vice versa.

Shape control of a local part of a structure rather than the entire structure is often encountered in engineering practice using as less control energy as possible. Energy optimization for local shape control of structures integrated with nonlinear piezoelectric actuators is investigated. The local shape is maintained by assigning the generalized displacements to the desired ones at some given points. The distortion of the residual shape is also limited to a given range. This optimal control voltage distribution is solved by using elimination and the Lagrange multiplier method. A unique solution scheme based on Cholesky decomposition and eigenvalue problem is given to reduce the computational load for finding the Lagrange multipliers from an algebraic equation. The final optimal control voltage distribution for the nonlinear system is obtained by an iteration procedure. Finally, an example is given to illustrate the present method, and the results show that proper relaxation of the error tolerance of the residual shape can significantly reduce the control energy requested for the local shape control of the designated area.

The flow structure of a tip vortex in the near field of a NACA 0015 wing with an effective aspect ratio of 5.04 undergoing a deep-stall oscillation with [alpha](t) = 18 deg + 6 deg sin [omega]t at Re = 1.86 x [10.sup.5] was investigated. The wing oscillation imposed a strong discrepancy in contour shapes and magnitudes between the pitch-up and pitch-down phases of the oscillation cycle. The vortex was more organized and nearly axisymmetric for x/c > 0.5 during pitch-up than during pitch-down. The peak tangential velocity and vorticity and the strength and size of the vortex increased with [alpha](t), except in the vicinity of [[alpha].sub.max], and had higher values during pitch-up than during pitch-down. The axial flow was always wake like and the velocity deficit decreased with [alpha](t), while exhibiting a sharp increase and decrease on the upstroke and downstroke in the vicinity of [[alpha].sub.max], respectively. The tangential velocity decreased slightly with the downstream distance. The vortex size increased rather significantly with x/c during pitch-down while remaining virtually unchanged during pitch-up. The inner region of the vortex exhibited a self-similar structure, similar to that of a stationary wing. The induced drag increased with [alpha](t) and had a local maximum at 20 deg during pitch-up.

Within a parametric study on brush-type trailing-edge extensions, the noise reduction potential of several design concepts was determined. The obtained database represents the first phase of an ongoing project with the long-term objective to develop scaling laws for a future application of such devices as add-on solutions for today's aircraft components. The experiments comprised both acoustic and aerodynamic measurements on a zero-lift generic plate model (Re=2.1 x [10.sup.6] to 7.9 x [10.sup.6]) in DLR's open jet Aeroaconstic Wind Tunnel Braunschweig. Noise data were taken by means of a directional microphone system. Measurement results indicate a significant source noise reduction potential in excess of 10 dB, depending on the configuration. Two relevant noise reduction mechanisms were identified: 1) the suppression of narrowband bluntness noise, as well as 2) the reduction of broadband turbulent boundary-layer trailing-edge noise.

The Abrupt-Wing-Stall (AWS) Program has addressed the problem of uncommanded lateral motions, such as wing drop and wing rock, at transonic speeds. The genesis of this program was the experience of the F/A-18E/F program in the late 1990s, when wing drop was discovered in the heart of the maneuver envelope for the preproduction aircraft. Although the F/A-18E/F problem was subsequently corrected by a leading-edge flap scheduling change and the addition of a porous door to the wing fold fairing, the AWS program was initiated as a national response to the lack of technology readiness at the time of the F/A-18E/F development program. The AWS program objectives were to define causal factors for the F/A-18E/F experience, to gain insights into the flow physics associated with wing drop, and to develop methods and analytical tools so that future programs could identify this type of problem before going to flight test. The major goals of the AWS Program, the status of the technology before the program began, the program objectives, the accomplishments, and the impacts are reviewed. Lessons learned are presented for the benefit of programs that must assess whether a future vehicle will have uncommanded lateral motions before going to flight test.

During a developmental flight test of the F/A-18E/F Super Hornet, uncommanded rolling motion was experienced while conducting maneuvers at transonic speeds. Abrupt, large-amplitude lateral events, referred to as wing drop, impaired the pilot's ability to effectively perform air-to-air tracking tasks at these flight conditions. A modified leading-edge flap schedule and the addition of porosity in the wing fold area eliminated the problem. However, finding the solution was slowed by the lack of a validated figure of merit to correlate with computational fluid dynamics, wind-tunnel, or flight-test analysis. A method has been developed to assess lateral activity using flight-test data. The lateral activity assessment involves both qualitative and quantitative evaluations. The pilot performs the qualitative evaluation, and the quantitative evaluation is based on analysis of roll rate, roll acceleration, lateral stick position, and bank angle. This technique provides a reliable approach to assessing lateral activity levels from flight test. An excellent correlation was attained between the pilots qualitative evaluation and the quantitative flight-test data assessment.

Numerical simulation of three-dimensional dynamic stall has been undertaken using computational fluid dynamics. The full Navier-Stokes equations, coupled with a two-equation turbulence model, where appropriate, have been solved on multiblock strucured grids in a time-accurate fashion. Results have been obtained for wings of square planform and of NACA 0012 section. Efforts have been devoted to the accurate modeling of the flow near the wing tips, which, for this case, were sharp without tip caps. The obtained results revealed the time evolution of the dynamic stall vortex, which, for this case, takes the shape of a capital omega [OMEGA] spanning the wing. The obtained results compare well against experimental data both for the surface pressure distribution on the wing and the flow topology. Of significant importance is the interaction between the three-dimensional dynamic stall vortex and the tip vortex. The present results indicate that once the two vortices are formed both appear to originate from the same region, which is located near the leading edge of the tip. The overall configuration of the developed vortical system takes a II-[OMEGA] form. To our knowledge, this is the first detailed numerical study of three-dimensional dynamic stall appearing in the literature.

There is a wide body of evidence that suggests that the turbulent mixing noise from high-speed jets consists of two components. These two components are generated by sources/mechanisms associated with the fine-scale turbulence and the large-scale structures of the jet plume. The noise generated by the large-scale structures radiates predominantly to the aft quadrant and typically peaks at angles close to the jet axis. The noise from the fine-scale turbulence is dominant in the forward quadrant and at near-normal angles to the jet axis. The initial objective of this study is to find a means to alter this basic radiation process in a high-speed jet. The ultimate goal of this study, however, is the achievement of noise reduction by accomplishing the preceding objective. A simple concept, a beveled nozzle, is proposed, and detailed aeroacoustic measurements are performed on two nozzles of different bevel angles. Noise measurements, over a wide range of polar angles, are made at several azimuthal angles to map the azimuthal variations. The performances of the beveled nozzles are assessed against a reference round nozzle. The beveled nozzles introduce significant azimuthal variations in the spectra, resulting in major differences in the polar directivities of the overall sound pressure levels at different azimuthal angles; these differences become pronounced when the jet velocity is increased. It is demonstrated that significant noise reduction is achieved in the azimuthal directions below the longer lip of the beveled nozzle, principally in the polar angular range of ~110 to ~140 deg. Furthermore, this reduction is observed at all frequencies, with a low performance penalty of ~ 1%. The magnitude of the noise reduction is a strong function of the jet velocity, with progressively higher reductions as the jet velocity is increased. With proper orientation of the nozzle, the noise footprint on the ground can be reduced. It is shown that the noise reduction is caused by the modification of the noise generated by the large-scale structures in the jet plume. Thus, noise reduction is achieved through the manipulation of the generation mechanisms.

A rather novel approach to predict jet noise is the stochastic noise generation and radiation (SNGR) method. The SNGR method uses the linearized Euler equations as an acoustic analogy together with source terms that are modeled. Previously, source terms for the linearized Euler equations have been derived and presented by Goldstein (Goldstein, M. E., 'A Unified Approach to Some Recent Developments in Jet Noise Theory,' International Journal of Aeroacoustics, Vol. 1, No. 1, 2002, pp. 1-16), as well as Bailly et al. (Bailly, C., Lafon, P., and Candel, S. M., 'A Stochastic Approach to Compute Noise Generation and Radiation of Free Turbulent Flows,' AIAA Paper 95-092, June 1995). Two different nonconservative formulations of the inhomogeneous linearized Euler equations are presented in these previous studies. Because the linearized Euler equations in conservative form are used in the present work, new source terms have to be derived for the conservative set of equations. A formal derivation of source terms for the linearized Euler equations in conservative form is presented, and the system of equations is compared to those of Goldstein and Bailly et al. Simplified versions of the derived source terms are also developed. To test the derived source terms, a direct simulation of a forced two-dimensional mixing layer is carried out. The direct simulation provides the mean flow and source field for the inhomogeneous linearized Euler equations. The solutions to the linearized Euler equations with source terms are compared to the solution of the direct simulation and show good agreement. All simulations are performed using Tam and Webb's fourth-order dispersion-relation-preserving scheme and a four-step fourth-order Runge-Kutta time-marching technique. Artificial selective damping introduced through the numerical scheme is used to avoid spurious waves. Absorbing boundary conditions based on characteristic variables are used at the free boundaries and a buffer layer is added at the outflow.

A numerical investigation of the adiabatic and condensing flowfield around the ONERA M6 and F-16 Falcon wing is performed. For the ONERA M6 wing, three different mean aerodynamic chord [c.sub.mac] lengths, 0.5, 1.0, and 2.0 m, three angles of attack, 1.07, 3.06, and 6.06 deg, and three different number densities of particles, [10.sup.6], [10.sup.12], and [10.sup.16] [m.sup.-3], are analyzed. At [10.sup.6] [m.sup.-3], the flow is dominated by homogeneous condensation, and due to the large chord length, the effects of condensation are small. As the particle density increases, a mixture of homogeneous/heterogeneous condensation occurs at [10.sup.12] [m.sup.-3], and at [10.sup.16] [m.sup.-3], only heterogeneous condensation occurs. For the F-16 Falcon wing, the size of the wing and the number density of particles [10.sup.12] [m.sup.-3] are kept constant. The angle of attack and freestream Mach number are changed, where a decrease in the lift-to-drag ratio reached 20%. In general, if the wing is experiencing separation, condensation can improve the lift-to-drag ratio, whereas for attached flow with small flight angles, condensation reduces the lift-to-drag ratio.

A higher order assumed modes (Ritz) method is developed for rectangular plate analysis using two-dimensional plate mode shape functions. These two-dimensional plate mode shape functions were determined using the extended Kantorovich-Krylov method. The rectangular plate is assumed isotropic and uniform, and has each edge of its span (root and tip) damped and remaining edges free (damped--free--clamped--free boundary conditions). Natural frequencies, mode shape functions, and frequency responses of the plate were calculated and compared to the results of traditional analysis using one-dimensional beam modes to approximate plate modes in both x and y directions. The updated two-dimensional plate mode shape functions substantially reduce the computational cost compared to the ease of using one-dimensional beam mode shape functions. Fewer modes are needed in the proposed method to yield the same accuracy. Experiments were conducted to validate our predictions and the data agree with our predictions.

In axial flight, rotors with three or more blades result in constant coefficient equations of motion, but the polar asymmetry of a two-bladed rotor leads to equations of motion with periodic coefficients. Although numerous studies on whirl flutter of three-bladed rotors are available in the literature, only a few of these address the issue of two-bladed rotors. The aeroelastic stability of two-bladed proprotor/pylon/wing combinations is examined using an elastic-blade analysis. Several parametric studies have been conducted by varying the natural frequencies of the system. Results of these studies have been compared with the results of a simple rigid-blade analysis developed previously. It is shown that the rigid-blade analysis captures the important interactions between the various rotor and wing modes. However, significant differences are observed in the behavior of the whirl flutter speeds, as obtained from the two analyses, when the natural frequencies of the rotor are varied. The parametric studies show that blade elasticity affects the wing chord mode damping at flight speeds close to the flutter speed, leading to differences in the critical flight speed of the wing chord mode as compared to the rigid-blade analysis. Also, the wing beam mode, which is always stable as per the rigid-blade analysis, is found to become unstable if the wing torsional stiffness is reduced from its baseline value.

A study is presented of the effects of spanwise lift distribution on aerodynamic efficiency for a blended wing body (BWB) configuration of a given baseline planform. The baseline geometry is initially assessed by a high-fidelity aerodynamic model based on a multiblock structured grid Reynolds-averaged Navier-Stokes (RANS) solution. The accuracy of the simulation is investigated by a grid sensitivity study regarding total drag and its pressure drag and skin-friction drag components. Excessive outer wing loading with associated shock wave, hence, wave drag, has been revealed to be the major factor degrading the aerodynamic performance of the baseline BWB model. To relieve the outer wing, an efficient low-fidelity panel method aerodynamic model is used for the inverse design to achieve the target lift distributions through the variation of twist distribution along the span, shifting the load inboard. For the given BWB geometry, the baseline model is retwisted to achieve an elliptic, a triangular, and an averaged elliptic/triangular spanwise loading distribution. The designs are then analyzed using the high-fidelity RANS aerodynamic model. The wave drag component of the total drag for different span loadings is extracted from the flowfield solution to gain insight into the drag reduction provided by the new twist designs. At the high subsonic speed of Maeh 0.85, a fine balance of the vortex drag and the wave drag is found to be essential to achieve the best aerodynamic performance. The elliptic lift distribution no longer gives the minimum drag if the nonlinear compressibility effects, that is, shock waves, feature in the flowfield of the designed aircraft. Although the concern is primarily with the transonic aerodynamic performance, the effects of the various lift distributions on pitching moments, therefore, trim and their implication on structural weight, are also discussed.

The morphing concept for unmanned aerial vehicles is a topic of current research interest in aerospace engineering. One concept of morphing is to change the wing configuration during flight to allow for multiple flight regimes. A particular approach to planform morphing is a variable-span morphing wing to increase wingspan to reduce induced drag and increase range and endurance. The wing area and the aspect ratio of the variable-span morphing wing increase as the wingspan increases. This means that the total lift increases while the induced drag is reduced, whereas the wing-root bending moment increases, thus, requiring a larger bending stiffness of the wing structure. Therefore, a study of the variable-span morphing wing requires not only aerodynamic analysis but also an investigation of the aeroelastic characteristics of the wing. The aerodynamic characteristics of the variable-span morphing wing are investigated, and a static aeroelastic analysis is performed.

Trends in the environmental impact of military aviation between 1960 and 2000 are articulated. The focus is on community noise, local air quality, and global climate impacts and the discussion is restricted to fixed-wing aircraft. Comparisons are made to trends within the commercial air transport industry. The unique features of military aircraft technology and operations responsible for the differences in environmental impacts are described. The discussion also considers the effects of environmental restrictions on the deployment and combat readiness of military aviation services. Regulations designed to mitigate environmental impacts from military and civil aviation are also reviewed. The analysis shows that military aviation has been responsible for a small and decreasing fraction of total fossil fuel use in the United States. Furthermore, when averaged nationally, contributions to local air-quality impacts and community noise have decreased over the period considered. These trends are a result of historical reductions in fleet sizes and number of operations. However, because base closures were largely responsible for these reductions, the impacts at any given installation may not reflect overall trends. Community noise and air quality are expected to be a growing concern for military aviation due to increasing urbanization, increasing public and regulatory attention, and use of training spaces for larger, multiservice operations involving longer range, higher speed weapon systems.

The computational techniques required for flight-test simulation in which the deflection angle of a tail wing is changed are discussed. To treat the computational grid for a moving control surface, an unstructured dynamic mesh method with surface-grid movement is used. This method is applied to the numerical simulation of the control-surface response for an experimental supersonic airplane at the National Aerospace Laboratory of Japan. Its high capability is demonstrated in both steady and unsteady simulations.

A linear model for analyzing the performance of pulsed-blowing actuator systems commonly used in many flow-control applications is presented and compared with experiments. The pulsed-blowing system consists of a regulated air supply, an oscillatory valve, transmission tubing, and an actuator. The elements of the actuator are a cavity and slot at the location in the flow to be controlled. The typical design objective is to achieve the largest possible velocity-fluctuation levels at the actuator-slot exit for a given fluctuating pressure input. However, the performance of the system is strongly dependent on component geometry, flow rate, and frequency of pulsation. Lumped-element models are useful for estimating the performance of the actuator cavity and slot, but fail to account for the effects of the transmission tubing. A distributed model for the tubing combined with a lumped-element model for the actuator provides estimates of the system resonant frequencies and amplitudes. Comparisons with experiments are made for a wide range of tubing lengths, slot widths, mean flow velocities, and forcing frequencies. The resonances of the pulsed blowing system and the changes in resonant behavior are characterized by considering the reflection coefficient of the system. Unexpected behavior such as reversed flow at the slot exit and reduced fluctuation levels with increasing pressure are demonstrated and explained with the system model.

The time-dependent acoustic field generated by the mutual interaction of a pair of coaxial viscous vortex rings advecting along a common axis of symmetry is simulated using a two-step numerical procedure. In the first step, the flowfield is predicted by solving the incompressible Navier-Stokes equations. In the second step, Powell's vortex sound theory is used to predict the far-field acoustic field associated with the flow unsteadiness. This study considers the effect of the initial toroidal ring radii on the sound radiation. The acoustic predictions are compared with those obtained using the method of matched asymptotic expansions. Our results indicate that the acoustic signature for the Mach-number range simulated shows an axisymmetric lateral quadrupole pattern. In addition, the time variation of the far-field acoustic signal displays significant effects caused by vortex dynamics associated with the finite-sized cores and viscous diffusion. These effects are more pronounced for rings with a smaller initial toroidal radius. The influence of Mach number on the far-field directivity (e.g., location of polar extinction angles) and the multipole content of the acoustic signal is examined through a multipole decomposition of the predicted acoustic fields.

A method for generating a reduced-order model of a flapping airfoil is presented. Using proper orthogonal decomposition, sets of orthogonal basis functions are generated from snapshots produced from previous computational (Navier--Stokes) results. With proper orthogonal decomposition, most of the energy in the flow is concentrated in just a few basis functions. These functions are used for a Galerkin projection of the Navier-Stokes equations, reducing them to a small set of coupled, nonlinear ordinary differential equations. The reduced-order model is used to simulate heaving motions that are both similar to and different from the motion(s) used to generate the basis functions, and the errors in the model are quantified. Several methods are used to generate mode sets that can be used over a range of heaving parameters, including snapshots from one, two, and multiple Navier-Stokes simulations. As snapshots from additional simulations are added to the decomposition, the mode sets become 'richer' and can simulate a wider range of parameter space. This richness, however, requires retaining more modes in the reduced-order model, adding to the computational complexity. The suitability of this approach for controlling a flapping wing over a broad range of parameters is discussed.

Reynolds averaged Navier-Stokes simulations of a static and pitching delta wing within three wind tunnels have been performed. These simulations have been compared with the ease of the wing in free air to ascertain the various influences of the walls on the vortical flow. The presence of tunnel walls has been found to promote vortex breakdown, with side wall proximity being the dominant factor. Roof and floor proximity has been seen to have a negligible effect on vortex breakdown. During pitching motion, side wall proximity delays vortex reformation after breakdown has reached its most upstream location, during cyclic pitching motion. This delay is recovered on the upstroke of the motion. These results confirm previous work with Euler simulations of tunnel interference.