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This paper examines the application of Gauss's principle of least constraint to the classical problem of the lift of a two-dimensional airfoil in ideal, incompressible fluid flow, including the role of the empirical Kutta condition that determines the airfoil circulation in classical theory. Gauss's principle demonstrates that the fluid pressure field consists of two components: a constraint pressure field that enforces the continuity and nonpenetration constraints and an impressed pressure field of arbitrary strength that satisfies but does not influence any constraints. The analysis shows that a flow solution that minimizes the difference between the acceleration and the impressed pressure gradient is a valid solution of Euler's equation. Furthermore, Gauss's principle proves that ideal potential flow airfoil theory is incomplete and that a first principles closure condition cannot exist within potential flow. This reaffirms the view that an external, empirical condition is needed to complete classical airfoil theory, i.e., the Kutta condition. The paper also investigates and disproves a new variational theory of lift developed from Hertz's principle of minimum curvature that is claimed to provide a closure condition based on first principles to replace the empirical Kutta condition of classical airfoil theory. [ABSTRACT FROM AUTHOR]

The efficient design of air intake is crucial for high-speed airbreathing propulsion systems. This paper presents a novel design method that integrates the recently introduced internal conical flow M (ICFM) basic flowfield (Musa et al., "New Parent Flowfield for Streamline-Traced Intakes," AIAA Journal, Vol. 61, No. 7, 2023, pp. 2906-2921) with the pressure-corrected osculating axisymmetric flows and streamline tracing techniques. The new design method takes into account the effects of lateral pressure gradients by incorporating pressure corrections into the original design method using crossflow velocity information from the ICFM flowfield. Additionally, new entrance and throat profiles are introduced for designing two hypersonic internal waverider intakes with shape transition. One intake is designed using the new method, and the other using the original method. Viscous simulations have been performed at design (i.e., Mach 6.0) and off-design (i.e., Mach 4.0 and 3.5) conditions. The results indicate that the pressure-corrected intake exhibits superior performance in terms of total pressure recovery and reduced flow distortion. This reveals that the new design method can achieve a smooth shape and efficient aerodynamic transitions between the intake entrance and throat. The new entrance and throat profiles realized better performance than the typical ones. The startability analysis reveals that the Van Wie curve controls the considered intakes. Thus, combining the pressure-corrected osculating axisymmetric flows with the ICFM flowfield enhances the performance of hypersonic intakes, making this approach a promising avenue for future hypersonic vehicles. [ABSTRACT FROM AUTHOR]

Unsteady-flow wind tunnels typically employ control louvers installed in the wind-tunnel circuit to produce rapid changes in the test-section flow velocity. For these wind tunnels, an open-return circuit is advantageous because of the lower mass of air that must be accelerated by the louver forcing. However, in addition to inertial effects, the wind-tunnel behavior is also affected by the interaction of the pressure disturbances from the louvers and their reflection from the open boundary conditions of the wind tunnel, resulting in a windspeed response that is nonintuitively related to the louver motion. In this paper, an experimental investigation into the unsteady behavior of a model-scale, open-return wind tunnel is described. An incompressible method of characteristics model for the wind tunnel is shown to successfully predict the wind-tunnel dynamic behavior, and an approach to produce controlled windspeed histories that compensates for the wind-tunnel response is demonstrated. [ABSTRACT FROM AUTHOR]

A wire mesh casing treatment is proposed and numerically tested on the transonic axial compressor (Rotor 37). A theoretical stall inception prediction model was developed to account for the wire mesh casing treatment, which was modeled as an impedance boundary condition. This model was validated on a low-speed compressor, replicating the experimental trend with reliable precision. The impedance boundary condition effectively attenuates disturbances and confines the distribution range of perturbations compared to the solid wall boundary condition. Furthermore, the installation of the wire mesh casing treatment induced a downstream shift in the passage shock wave and enhanced the flow capacity in the blade tip region. The effects of the casing treatment on different installation locations were investigated, and the results indicated that the maximum stall margin improvement occurred when the casing treatment was positioned at the midchord location. The integrated design of the casing treatment and blade profile, along with the established prediction model, is also discussed. This model demonstrates reliable precision when compared to the experimental results and addresses the need for rapid design iteration during the initial design phase. [ABSTRACT FROM AUTHOR]

Mode transition is essential for the combined-cycle engine inlet. During actual inlet operation, mode transition may occur in the off-design state. An experiment was performed with a generic over/under combined-cycle engine inlet mode to investigate the characteristics of the inlet mode transition in the off-design state. The transition Mach number was set to Mach 3.5 by design, while it was 2.9 during the experiment. Oscillatory flow is obtained during both down- and up-rotation processes of the splitter. Results show that the size and location of the separation bubble vary with the splitter position. The dominant frequency of the flowfield oscillation is always below 200 Hz. Combining the analysis of schlieren and dynamic pressure data, it is found that the main factor affecting the flowfield oscillation features is the self-excited oscillation of the shock train, while the acoustic resonance frequency is also reflected near the unstart point. Comparison of the pressure curves of the two processes shows that the hysteresis interval accounts for 17.8% of the entire rotation time of the splitter and that the hysteresis stems from the presence of the separation bubble on the splitter during the up-rotation of the splitter. The analysis of the unstart and restart points shows that they deviate from the isentropic and Kantrowitz limits, respectively. The new unstart and restart boundaries should consider the combined effect of the separation bubble at the duct entrance, the internal wave system, and the throat Mach number. [ABSTRACT FROM AUTHOR]

A new design of fluidic oscillator, in which the oscillator is curved along its primary flow direction, allows for a sweeping jet to be placed near the leading edge of a wing as an active flow control device. Previous studies of fluidic oscillator applications have shown their promise as an active flow control device by increasing the lift over a wing and preventing flow separation. However, the planar extent of the flat fluidic oscillators creates challenges when trying to optimally implement them in a confined geometry. The application of these new curved oscillators is experimentally investigated on the leading edge of an upswept NACA0018 airfoil full-span wing. Global load cell measurements of lift, drag, and pitching moment are presented and analyzed using suction surface static pressure taps, and particle image velocimetry data are presented. The effects of active flow control on lift, drag, and stall at a Reynolds number of 100,000 are documented for a variety of oscillator chordwise locations near the leading edge and compared to traditional vortex-generating jets. While further aft-positioned oscillators produced a larger increase in maximum lift, more forward locations resulted in better poststall performance and delayed flow separation over the wing. [ABSTRACT FROM AUTHOR]

A machine learning model is developed to establish wake patterns behind oscillating foils for energy harvesting. The role of the wake structure is particularly important for array deployments of oscillating foils since the unsteady wake highly influences the performance of downstream foils. This work explores 46 oscillating foil kinematics, with the goal of parameterizing the wake based on the input kinematic variables and grouping vortex wakes through image analysis of vorticity fields. A combination of a convolutional neural network with long short-term memory units is developed to classify the wakes into three classes. To fully verify the physical wake differences among foil kinematics, a convolutional autoencoder combined with k -means++ clustering is used to reveal four wake patterns via an unsupervised method. Future work can use these patterns to predict the performance of foils placed in the wake and build optimal foil arrangements for tidal energy harvesting. [ABSTRACT FROM AUTHOR]

We carry out a parametric study of hypersonic impinging shock-wave/turbulent-boundary-layer interactions (SBLIs) at hypersonic conditions (Mach number 6.0) by means of numerical simulation of the Reynolds-averaged Navier-Stokes (RANS) equations, with the eventual goal of establishing wall temperature and Reynolds number effects. Comparison with available direct numerical simulation and experimental data shows that RANS is capable of predicting the main features of SBLI, namely, typical size and distribution of the wall properties. A large number of flow cases were computed to examine the scaling of the global flow parameters over a wide range of wall temperatures and Reynolds numbers. As expected, the interaction zone is reduced as the wall is cooled, and as the Reynolds number is increased. A modified form of the scaling originally introduced by Souverein et al. [Journal of Fluid Mechanics, Vol. 714, 2013, pp. 505-535] is here considered to account for nonadiabatic wall state, which is found to successfully collapse the data to the distributions obtained for adiabatic flow, with some remaining deviations associated with low-Reynolds-number effects. [ABSTRACT FROM AUTHOR]

We present three-dimensional scaling relations for the thrust production and power consumption of combined heaving and pitching hydrofoils by extending the three-dimensional pitching scaling laws introduced by Ayancik et al. ("Scaling Laws for the Propulsive Performance of Three-Dimensional Pitching Propulsors," Journal of Fluid Mechanics, Vol. 871, July 2019, pp. 1117-1138). Self-propelled inviscid simulations and previously published experimental data are used to validate the scaling laws over a wide range of motion amplitudes, Strouhal numbers, heave ratios, aspect ratios, and pitch axis locations. The scaling laws are shown to predict inviscid numerical and experimental data well, within ±25% and ±16% of the thrust and power data, respectively. It reveals that both the circulatory and added mass forces are important when considering a wide range of motion amplitudes and that nonlinear corrections to the classic linear theory are essential to modeling the power performance across a wide amplitude and aspect ratio range. By using the scaling laws as a tool, it is obtained that peak efficiency occurs when dimensionless amplitude A*>1 and for these large-amplitude motions there is an optimal nondimensional heave ratio h*, where the efficiency maximizes in the narrow range of 0.75

The prediction of flowfield evolution can provide valuable reference information for the development of hypersonic technology. Flowfield prediction with the introduction of deep learning techniques is a promising method to provide future flowfield evolution in scramjet isolators. A multipath flowfield prediction model has been proposed to achieve flowfield prediction based on wall pressure sequence. The prediction model is mainly constructed with a convolutional neural network. An experimental dataset was built with supersonic experimental data under different evolution laws in an isolator. The flowfield prediction model is trained and validated using independent experimental data. The proposed model's prediction performance under different prediction spans is discussed in depth. The results demonstrate that the predicted flowfield is in good agreement with the ground truth, and the background wave and shock train structure are basically restored, even when the shock train leading edge changes intermittently. The influence of pressure sequence length on the proposed model's prediction performance is also analyzed. [ABSTRACT FROM AUTHOR]

A partially lagging one-equation eddy viscosity model, based on the transformation of the k-ε turbulence closure through Bradshaw et al. ("Calculation of Boundary Layer Development Using the Turbulent Energy Equation," Journal of Fluid Mechanics, Vol. 23, No. 3, 1967, pp. 593-616), is proposed. The model attempts to account for the turbulence history effects by spatially lagging the destruction/diffusion terms. The amount of relaxation is based on the von Kármán length scale. The present model does not assume equal diffusion coefficients of the k and ε equations; therefore, third-order velocity gradients emerge. The performance of the present model is assessed against the Spalart-Allmaras model ("A One-Equation Turbulence Model for Aerodynamic Flows," AIAA Paper 1992-0439, 1992) and the one-equation model proposed by Menter ("Eddy Viscosity Transport Equations and Their Relation to the k-ε Model," Journal of Fluids Engineering, Vol. 119, 1997, pp. 876-884), which is based on Bradshaw et al. ("Calculation of Boundary Layer Development Using the Turbulent Energy Equation," Journal of Fluid Mechanics, Vol. 23, No. 3, 1967, pp. 593-616), and the assumption of equal diffusion coefficients. The behavior of the one-equation transformed model without the presence of third-order velocity gradients is also examined. The proposed model is validated through a comparison to a wide class of internal and external turbulent flows. [ABSTRACT FROM AUTHOR]

The asymmetric vortex regime of a von Karman ogive with a fineness ratio of 3.5 is experimentally studied at a Reynolds number of 156,000. The wake of an axisymmetric bluff body is an ideal candidate for active feedback flow control because minute fluidic disturbances and geometry perturbations near the tip of the ogive get amplified through the flow's convective instability. The resulting disturbance interacts with the quasi-steady vortex location and produces a deterministic port or starboard asymmetric vortex state (i.e., side force). Accurate control or manipulation of this asymmetric vortex state holds the potential for increased maneuverability and stability characteristics of slender flight vehicles. For implementation of an active feedback flow-control system, plasma actuators at the tip of the ogive are used as the flow effector, and surface-mounted pressure sensors are used to estimate the vortex configuration in real time. A linear time invariant model developed from open-loop experimental tests and a proportional-integral control law are used to close the loop in the experimental setting. Closed-loop experimentation shows the ability to arbitrarily track a side force set point while also suppressing low-frequency fluctuations. Thus, the adopted model-based feedback flow-control approach is validated experimentally for a complex, three-dimensional flow. [ABSTRACT FROM AUTHOR]

An in-depth investigation of the Hartmann tube (HT) was carried out to gain a better understanding of and the effects of various parameters on its output flow characteristics. The HT used in this work consisted of an underexpanded jet directed into a close-ended cylindrical tube of the same diameter. The effects of tube depth, separation distance between the tube and the nozzle, and the jet Mach number were explored. Experiments were also performed on an HT fluidic actuator (HTFA). The HTFA operates under the same principles as the HT, except a major portion of the area between the nozzle and tube is shielded off, so that it can be used as a pulsating injector for flow control. Dynamic pressure measurements in the near field and microphone measurements in the far field provided temporal and spectral data. Instantaneous and phase-averaged images of the flow were obtained to explore the nature of turbulence structures within the flow. Limited hot-wire measurements were performed to characterize the velocity fluctuations at the exit of the HTFA. Striking similarities were observed in the frequency content and amplitude of tonal frequencies in near-field pressure and far-field acoustic measurements, as well as the flow results. Time traces of pressure in the tube were used to explain the major differences between the primary frequency of HT and the quarter-wave frequency. Flow visualization results showed a pulsating flow very rich in vortical structures. These images also revealed that the separation distance between the nozzle and the tube has a significant effect on the direction of the pulsating jet out of HT. The results illustrate many similarities, but also some differences between the HT and the HTFA.

A Reynolds stress relaxation model, specifically the lag Reynolds stress transport model, is applied to a wingtip vortex flow, and its performance is assessed and compared with other aerospace standard turbulence models. A Reynolds stress relaxation model allows for Reynolds stress history effects due to streamline curvature, which are seen to play an important role in the nondiffusive nature of turbulent vortices. This study shows that the lag Reynolds stress transport turbulence model is capable of predicting mean flow results as accurately as those of the well-performing Spalart-Allmaras model with correction for streamline curvature and system rotation. Furthermore, in this wingtip vortex flow, the lag Reynolds stress transport model predicts turbulence quantities more accurately than the rotation/curvature-corrected Spalart-Allmaras model. Although the lag Reynolds stress transport model well predicts this flow, it is more computationally intensive to solve than the rotation/curvature-corrected Spalart-AIlmaras model, and it has some deficiencies, such as an inability to independently control the Reynolds stress magnitude and relaxation amount. [ABSTRACT FROM AUTHOR]

This paper describes a method to detect three-dimensional, steady inviscid shock waves from computational fluid dynamics data. The method is based on the principle that the collision of the characteristics of the same family causes the generation of a shock wave in a two-dimensional flow. In a three-dimensional flowfield, however, there is an infinite number of characteristics. Therefore, the plane on which the streamline can be treated locally as a planar curve is introduced, and the characteristic as an intersection between the Mach cone and this plane is defined. As a result, the shock waves in a three-dimensional, steady flowfield can be treated by a similar methodology as that in a two-dimensional flowfield. [ABSTRACT FROM AUTHOR]

In a previous paper, a novel approach to axisymmetric actuator-disk modeling was presented. In the present paper, this modeling is extended to include the case of a skewed flow, which is asymmetric. Similar to the axisymmetric case, it is assumed that the pressure difference between both sides of each point of the disk is a time average of the pressure difference between both sides of the blade elements that pass through the same point. The influence of the loaded disk on the flowfield is modeled by a distribution of sinks/sources over the disk plane. Closed-form integral expressions are obtained for the induced velocity through and upstream of the disk. By using Bernoulli's equation, which relates between the flow far upstream and the flow just before crossing the disk plane, an additional nonlinear algebraic equation for the induced-velocity distribution over the disk is obtained. Discretization of the integral equations results in a system of nonlinear algebraic equations. An iterative procedure leads to the final solution. The model is efficient, stable, easy to apply, and solve. Comparisons with results of another method from the literature exhibit a very good agreement. [ABSTRACT FROM AUTHOR]

Far-field drag-prediction and decomposition methods are powerful tools that increase the accuracy of the drag coefficient computed from computational fluid dynamics results by removing the spurious drag caused by numerical procedures. Furthermore, these methods allow a physical decomposition of the drag in terms of viscous, wave, and induced drag. However, they are currently limited to steady flows. This paper presents a generalization of the commonly used drag-prediction and decomposition method to unsteady flows. This generalized method, designed for three-dimensional viscous, subsonic, and transonic flows, is defined for both inertial and noninertial coordinate systems and allows drag decomposition to be performed on either static or moving/rotating meshes. This generalization also allows the drag caused by the unsteady fluctuations of the flow to be identified. [ABSTRACT FROM AUTHOR]

This contribution describes the development and evaluation of a new statistical modeling approach for active vortex generator jets. Previous experiments from the Technisehe Universität Braunsehweig and their subsequent evaluation by the present authors showed that the induced flowfield can be reasonably well represented by two- dimensional Lamb--Oseen vortices. Based on that, an analytical expression for the Lamb-Oseen-vortex-model maximum circulation Γmax was derived in terms of the freestream velocity U∞, the jet-to-freestream velocity ratio λ, and the jet skew angle β, as well as the actuator diameter ΦVGJ. Based on the parameterized results, universal values for the Lamb--Oseen-vortex-model parameters at the actuator position were determined for the development of the statistical vortex-generator-jet model. The idea behind the statistical modeling approach is that the vortices are represented by their spanwise-averaged velocity correlations, or vortex stresses, that are added to the turbulence stresses in a Reynolds stress turbulence model. The spanwise-averaged vortex stresses are derived by computing the spanwise-averaged second-order statistics of the vortex flowfield. These vortex-generator-jet model results were compared to the spanwise-averaged vortex stresses from experiments and from fully resolved computational fluid dynamics investigations, and reasonable qualitative as well as quantitative agreement was found. [ABSTRACT FROM AUTHOR]

Non-unique solutions of the Euler equations were originally discussed by Jameson in 1991 ("Airfoil Admitting Non-Unique Solutions to the Euler Equations," AIAA Paper 1991-1625, June 1991) for several highly cambered airfoils which were the result of aggressive shape optimization. In 1999 Hafez and Guo ("Nonuniqueness of Transonic Flows," Acta Mechanica, Vol. 138, Nos. 3-4, 1999, pp. 177-184) found non-unique solutions for a symmetric parallel sided airfoil, and subsequently Kuz'min and Ivanova ("The Structural Instability of Transonic Flow Associated with Amalgamation/Splitting of Supersonic Regions," Theoretical and Computational Fluid Dynamics, Vol. 18, No. 5, 2004, pp. 335-344) have discovered some fully convex symmetric airfoils that provide non-unique solutions. In this article four new symmetric airfoils, all of which exhibit non-unique solutions in a narrow band of transonic Mach numbers, are studied. The first, NU4 was the result of shape optimization. The second, JF1 is an extremely simple parallel sided airfoil. The third JB1, is also parallel sided but has continuous curvature over the entire profile. The fourth, JC6, is convex and C ∞ continuous. CL -- α plots of these airfoils exhibit three branches at zero angle of attack, the P-, Z- and N-branches with positive, zero and negative lift respectively. At some Mach numbers no stable Z-branch could be found. When the P-branch is continued to negative α in some cases there is a transition to the Z-branch, while in other cases there is a direct transition from the P- to the N-branch. [ABSTRACT FROM AUTHOR]

In computational fluid dynamics, for problems with periodic flow solutions, the computational cost of spectral methods is significantly lower than that of full, unsteady computations. As is the case for regular steady-flow problems, there are various interesting periodic problems, such as those involving helicopter rotor blades, wind turbines, or oscillating wings, that can be analyzed with spectral methods. When conducting gradient-based numerical optimization for these types of problems, efficient sensitivity analysis is essential. The authors developed an accurate and efficient sensitivity analysis for time-spectral computational fluid dynamics. By combining the cost advantage of the spectral-solution methodology with an efficient gradient computation, the total cost of optimizing periodic unsteady problems can be significantly reduced. The efficient gradient computation takes the form of an automatic differentiation discrete adjoint method, which combines the efficiency of an adjoint method with the accuracy and rapid implementation of automatic differentiation. To demonstrate the method, the authors computed sensitivities for an oscillating ONERA M6 wing. The sensitivities are shown to be accurate to 8-12 digits, and the computational cost of the adjoint computations is shown to scale well up to problems of more than 41 million state variables. [ABSTRACT FROM AUTHOR]

The disruption of simulated fuel droplets in supersonic flow is examined experimentally in a drawdown supersonic wind tunnel. The droplets are accelerated in the supersonic flow, achieving supersonic velocities relative to the surrounding air with relative Mach numbers as high as 1.8 and Weber numbers as high as 300. Monodisperse 100-µm-diam fluid droplets are generated using a droplet-on-demand generator upstream of the tunnel entrance. The droplets are imaged by direct close-up single- and multiple-exposure imaging. Three test liquids were employed: 2-propanol and tetraethylene glycol dimethyl ether as nonvolatile fluids, and a more volatile 50/50 hexanol-pentane mixture. The decreased static pressure in the supersonic stream had the potential to give rise to superheating of the droplet fluid, as in some cases, the static pressure became significantly lower than the vapor pressure of the droplet liquid. Droplet lifetimes for the hexanol/pentane mixture appear to be shorter due to accelerated vaporization consistent with superheating, although little impact is observed on the droplet velocity and relative Mach number. Droplet-disruption patterns for these supersonic flow conditions can be classified into four different flow regions by considering the changes in the Weber number with downstream distance as the droplets accelerate. The drag coefficients associated with the droplet disruption under locally supersonic conditions are generally higher than those expected for solid spheres, largely due to the cross-sectional area change associated with droplet deformation/ breakup. [ABSTRACT FROM AUTHOR]

This paper presents the results of insertion loss measurements and numerical impedance eduction of three different liner samples. An overview of the test rig and methodology is given, and preprocessed results in terms of reflection and transmission coefficients as well as the energy dissipation are discussed. These coefficients are calculated for discrete frequencies within the investigated frequency range. Subsequently, a numerical postprocessing is performed in the time domain, and the educed impedance function for each sample and flow Mach number is presented. This postprocessing in the time domain uses an impedance model based on the extended Helmholtz resonator with five free parameters. The parameters of the model are fitted via an optimization, which determines the whole frequency response in one optimization process. The comparison of measured and numerically evaluated energy coefficients proves the reliability of the tools for impedance evaluation under flow conditions. Finally, the impedance results of the different samples are discussed, including a comparative study with Aermacchi data of the National Aerospace Laboratory (The Netherlands) flow tube and Aermacchi impedance tube experiments. [ABSTRACT FROM AUTHOR]

The article offers information on supersonic convex-corner flows. It is said that an interaction law can be developed in an explicit form that would relate the displacement effect of the boundary layer to the pressure. The pressure is induced in the inviscid pan of subsonic flows. Studies that attributed the flow properties near the corner to the viscous-inviscid interaction are scarce.

The thermoacoustic coupling caused by dynamic flow/flame interactions was investigated in a gas-turbine model combustor using high-repetition-rate measurements of the three-component velocity field, OH laser-induced fluorescence, and OH* chemiluminescence. Three fuel-lean, swirl-stabilized flames were investigated, each of which underwent self-excited thermoacoustic pulsations. The most energetic flow structure at each condition was a helical vortex core that circumscribed the combustor at a frequency that was independent of the acoustics. Resolving the measurement sequence with respect to both the phase in the thermoacoustic cycle and the azimuthal position of the helix allowed quantification of the oscillatory flow and flame dynamics. Periodic vortex/flame interactions caused by deformation of the helices generated local heat-release oscillations having spatially complex phase distributions relative to the acoustics. The local thermoacoustic coupling, determined by statistically solving the Rayleigh integral, showed intertwined regions of positive and negative coupling due to these vortices. In the quietest flame, the helical vortex created a large region of negative coupling that helped damp the oscillations. In the louder flames, the shapes of the oscillating vortices and flames were such that large regions of positive coupling were generated, driving the instability. From these observations, flame/vortex configurations that promote stability are identified. [ABSTRACT FROM AUTHOR]

The aeroelastic response of a biologically inspired flapping membrane wing to unsteady inertial and aerodynamic forces is strongly predicated upon the topological distribution of the underlying skeletal reinforcement (venation patterns). Proper exploitation of this relationship may enable the development of effective flapping wing micro air vehicles. The skeletal topology is parameterized into a developmental program which, when compiled and executed, evolves the wing topology in stages. A genetic algorithm can optimize the details of this program, rather than explicitly considering the topology itself, thus removing the link between design variables and topological geometry resolution. This cellular division tool is coupled to an unsteady aeroelastic representation of a flapping wing in forward flight in order to obtain the Pareto tradeoff curves between thrust generation, lift generation, and input power requirements. The topologies obtained along the front provide an understanding of the key relationships between skeletal topology, aeroelastic behavior, and performance metrics during flapping flight. [ABSTRACT FROM AUTHOR]

The fundamental vibration frequencies of thin-walled elliptical composite cylinders for which the fiber orientations in some or all of the layers vary linearly with circumferential location are discussed. The study is motivated by previous work that demonstrated that a circumferentially varying fiber orientation scheme can be used to improve the axial buckling capacity of an elliptical composite cylinder. At issue then is the degree to which fundamental frequencies are influenced by the circumferentially varying fiber orientation. Circular cylinders are also considered, so an additional finding of the study is the influence of cylinder cross-sectional geometry on the fundamental vibration frequency, with all other cylinder geometric and material parameters remaining the same. The results, which are based on finite element models of the elliptical and circular cylinders, indicate that, even for what is considered a significant range of fiber orientations that vary linearly with circumferential location, the fundamental frequencies of both elliptical and circular cylinders are not significantly influenced. On the other hand, the circumferential wave numbers are influenced. The frequency results for circular cylinders are substantiated by the work of other researchers. [ABSTRACT FROM AUTHOR]

This work presents the buckling optimization of composite stiffened panels loaded in compression and shear. The procedure relies on the coupled use of a semi-analytical formulation for the structural analysis together with an optimization based on genetic algorithms. The semi-analytical formulation allows the assessment of the local buckling behavior of panels with blade, J, T, and hat-stiffener cross sections. The out-of-plane buckling deflections of the skin and of the stiffener are represented by trigonometric shape functions, and the governing equations are derived applying the principle of the minimum potential energy and the method of Ritz. The formulation is used to obtain optimal configurations in terms of skin and stiffener lay-ups, stiffener cross sections, and geometry. Constraints can be imposed on the structural response in terms of buckling load and prebuckling stiffness as well as on technological requirements. The optimal design is presented for the buckling load maximization of a flat hat-stiffened panel loaded in compression and shear and for the minimum weight under buckling constraints of a curved panel with open-section stringers loaded in compression. The results obtained from the semi-analytical formulation are then compared with finite element analyses. [ABSTRACT FROM AUTHOR]

A workshop on the computational fluid dynamics (CFD) prediction of shock boundary-layer interactions (SBLIs) was held at the 48th AIAA Aerospace Sciences Meeting. As part of the workshop, numerous CFD analysts submitted solutions to four experimentally measured SBLIs. This paper describes the assessment of the CFD predictions. The assessment includes an uncertainty, analysis of the experimental data, the definition of an error metric, and the application of that metric to the CFD solutions. The CFD solutions provided very similar levels of error and, in general, it was difficult to discern clear trends in the data. For the Reynolds-averaged Navier-Stokes (RANS) methods, the choice of turbulence model appeared to be the largest factor in solution accuracy. Scale-resolving methods, such as large-eddy simulation (LES), hybrid RANS/LES, and direct numerical simulation, produced error levels similar to RANS methods but provided superior predictions of normal stresses. [ABSTRACT FROM AUTHOR]

A study of shock-buffet onset and instability mechanism via Reynolds-averaged Navier-Stokes simulations on several airfoils is presented. The numerical setup and the Spalart-Allmaras turbulence closure are validated based on wind-tunnel data from NACA 0012 and RA16SCI airfoils. The paper presents simulations of the flow past three airfoils: the subsonic NACA 0012, the supercritical RA 16SC1, and the thin, transonic/supersonic NACA 64A204, at pre- and postbuffet conditions, and within a cycle of developed shock buffet. The shock-buffet cycle is found to be similar in nature for all airfoils, originating in unstable interaction of the shock and the separation bubble. Simulation results support the notion that buffet onset is not related to the bursting of the separation bubble behind the shock. Shock-buffet categorizing is posited as a transonic prestall instability phenomenon that depends on the shock strength and location. Shock-buffet onset conditions occur when the shock position is behind and sufficiently close to the upper-surface maximum curvature location. Additionally, it is suggested that offset conditions are when the shock is at an upstream location and the flow aft of it is fully separated. [ABSTRACT FROM AUTHOR]

A computational fluid-dynamics investigation, including passive vortex generators (VGs) that generate streamwise counter-rotating vortex structures, usually requires a grid with fully resolved VG geometries and vortex structures with a corresponding large number of grid points to obtain an accurate solution. An efficient way to avoid such a setup and time-consuming process in turbulent shear-layer flows is to introduce statistics-based vortex-generator modeling. The second-order statistics of the initial vortices are computed by using a vortex model in combination with the lifting-line theory. The statistics are added as additional turbulence stress terms to the equations within a differential Reynolds stress-turbulence model. In this investigation, results from statistical VG model computations for zero- and adverse-pressure-gradient flat-plate boundary-layer flows, as well as for the flow in a plane asymmetric diffuser, are evaluated against results from fully resolved VG computations and experiments. It could be shown that the initial near-field forcing is too weak for the proposed VG model. An improved VG model description removes some drawbacks by adding additional statistical forcing terms. Results become more comparable, resulting in improved predictions when compared to experiments and fully resolved computations. [ABSTRACT FROM AUTHOR]

This paper presents analytical and experimental studies of a modified semiactive hydraulic damper operating in periodic working regimes. This work was completed as a part of the Rotor Embedded Actuator Control Technology project sponsored by the Technology Strategy Board in the United Kingdom. The damper tested is based on modifications to an industrially employed helicopter damper. The work presented covers the relevant aspects of the model development, damper modification, test planning, and analysis, as well as a model-simulation correlation study. The model of the damper directly reflects the actual hydraulic modification, enabling a semiactive mode of operation. Pressure-flow representations used in the damper modeling are based on a novel testing methodology using triangular piston excitation waveforms. A specific test structure based on varying the relative phase difference between two harmonic input signals is used to assess the damper properties in periodic working regimes. The test configuration with a base harmonic piston excitation combined with a harmonic modulation at a frequency 3 times higher than the base frequency leads to significant changes in the second and fourth harmonic components of the resulting periodic damper forces. A successful model-simulation correlation study suggests that a simple one-state dynamic model of this damper features good predictive properties, equally applicable in the extended simulation contexts. [ABSTRACT FROM AUTHOR]

The aerodynamic steering forces and moments on an axisymmetric bluff body are controlled by exploiting an induced, asymmetric, segmented attachment of the base flow to a Coanda surface at the body tail end. Control is effected by four individually addressable, aft-facing, synthetic jet actuators. The jets emanate from azimuthally segmented slots, equally distributed around the perimeter of the tail. The model is suspended in the wind tunnel by eight thin wires for minimal support interference with the wake. Each wire is instrumented with a miniature strain-gage transducer for direct dynamic force measurements. The aerodynamic effects associated with continuous and transitory asymmetric activation of the Coanda effect are characterized using force and hot-wire measurements. Transitory modulation of the actuation waveform of multiple actuators leads to the generation of significant dynamic radial forces of controlled magnitude and direction with potential utility for flight stabilization and fast maneuvering. [ABSTRACT FROM AUTHOR]