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Droplet deformation and breakup in the continuously accelerated flowfield generated by an incoming airfoil have been studied. The upper limit of droplet deformation and the minimum distance to the ...

In the present work, LEWICE-based ice-shape prediction results are presented when coupled to a computational fluid dynamics (CFD) model using a discrete-element roughness method (DERM) prediction o...

A new breakup criterion is proposed in this paper for droplets subject to the flowfield generated by an incoming airfoil (that is, the criterion should be applied only to this type of aerodynamics ...

This paper presents the development of a new framework of multiblock structured elliptic grid smoothing for full-aircraft configurations. The framework is fully automated and applies different solv...

A method is presented for using large-eddy simulation to generate urban wake fields for use in studying the effects on the autonomous flight performance of a small quadrotor. The flowfield is solved around a single square building using OpenFOAM and stored in a database accessed by a MATLAB/Simulink flight simulator. Four flight missions are evaluated to compare the difference in performance between wake fields generated by Reynolds-averaged Navier–Stokes and large-eddy simulation solutions. The results of the holding position in a constant freestream wind show both methods produce similar results and can hold position in all three directions within approximately ±1.5 body lengths. When the quadrotor is in or on the boundary of the building wake, the maximum deviation volumes, as calculated when using a Reynolds-averaged Navier–Stokes or large-eddy simulation air wake, can differ by two orders of magnitude. Additionally, the large-eddy simulation air wakes can cause skewed deviations by as much as five to...

Information concerning droplets impinging on airfoils, including their drag coefficient, that depends on the droplet shape is needed to feed computational codes used to simulate the ice accretion process on wing surfaces. A new model is presented to predict the droplet ratio of deformation for a given flowfield. This model is based on the analogy of a distorted droplet and a damped spring-mass system. It differs from known Taylor analogy breakup model, Clark’s model, and droplet deformation and breakup model in some departing assumptions. Comparisons of the model results with experimental data in the stagnation line show excellent agreement. The new model proposed here will aid to determine the drag coefficient evolution, being able to be included in ice accretion codes.

Adaptive morphing trailing-edge wings have the potential to reduce the fuel burn of transport aircraft. However, to take full advantage of this technology and to quantify its benefits, design studies are required. To address this need, the aerodynamic performance benefits of a morphing trailing-edge wing are quantified using aerodynamic design optimization. The aerodynamic model solves the Reynolds-averaged Navier–Stokes equations with a Spalart–Allmaras turbulence model. A gradient-based optimization algorithm is used in conjunction with an adjoint method that computes the required derivatives. The baseline geometry is optimized using a multipoint formulation and 192 shape design variables. The average drag coefficient is minimized subject to lift, pitching moment, geometric constraints, and a 2.5g maneuver bending moment constraint. The trailing edge of the wing is optimized based on the multipoint optimized wing. The trailing-edge morphing is parameterized using 90 design variables that are optimized i...

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

Cd = drag coefficient of the airship model D = hull diameter of the airship model, m L = hull length of the airship model, m ReV = Reynolds number based on the cubic root of thevolume of the model U∞ = freestream velocity, m∕s urms = streamwise velocity fluctuation intensity, m∕s Xfr∕L = nondimensional axial length coordinate of sublimation front position Xtr∕L = nondimensional axial length coordinate at transition x = coordinate in freestream direction from hull nose, m y = coordinate in wall normal direction, mm z = coordinate in circumferential direction, mm α = angle of attack of the model, deg δ = boundary-layer thickness, mm

This document is the Accepted Manuscript version of the following article: C. Wang, S. Chang, and H. Wu, "Lagrangian Approach for Simulating Supercooled Large Droplets’ Impingement Effect", Journal of Aircraft, Vol. 52, No. 2 (2015), pp. 524-537. The Version of Record is available online at doi: https://doi.org/10.2514/1.C032765.

The goal of this study is to show the efficiency of the wake-integral technique coupled with vorticity confinement method for the prediction of induced drag by inviscid numerical modeling. While the vorticity confinement method has previously been used to capture trailing vortices, its quantitative effect on improving wake-integral drag predictions has not been investigated. In the present study, parameters of vorticity confinement are tuned and the advantages of using a variable confinement parameter are demonstrated. A combination of vorticity confinement with the wake-integral technique indicates that the computational fluid dynamics-predicted induced drag for wing with and without winglet becomes insensitive to wake-integration plane location, the computed drag approaches its theoretical lifting-line value, and the spurious entropy drag is suppressed.

Results presented at the Fifth Drag Prediction Workshop using CFL3D, FUN3D, and NSU3D are described. These are calculations on the workshop provided grids and drag adapted grids. The NSU3D results have been updated to reflect an improvement to skin friction calculation on skewed grids. FUN3D results generated after the workshop are included for custom participant generated grids and a grid from a previous workshop. Uniform grid refinement at the design condition shows a tight grouping in calculated drag, where the variation in the pressure component of drag is larger than the skin friction component. At this design condition, A fine-grid drag value was predicted with a smaller drag adjoint adapted grid via tetrahedral adaption to a metric and mixed-element subdivision. The buffet study produced larger variation than the design case, which is attributed to large differences in the predicted side-of-body separation extent. Various modeling and discretization approaches had a strong impact on predicted side-of-body separation. This large wing root separation bubble was not observed in wind tunnel tests indicating that more work is necessary in modeling wing root juncture flows to predict experiments.

A far-field drag prediction and decomposition method has been applied to the results of the AIAA Drag Prediction Workshop 5 held in Louisiana during the summer of 2012. The method has two principal advantages: it allows the removal of spurious drag inherent to computational fluid dynamics solutions, and it allows the decomposition of drag into viscous, wave, and induced physical drag components. This research shows that accurate drag coefficients can be predicted on coarse grids when the spurious drag is extracted with the far-field method and that these results are closer to experimental values than drag coefficients computed on finer meshes when spurious drag is not extracted. The research also investigated the reasons behind the lift and drag losses found by some participants in the workshop. It is shown that the lift loss is caused by the boundary-layer separation at the wing root, inducing a reduction of 20% of the shock wave drag and a significant change in the wing loading. The initiation of buffet...

CFS Engineering and RUAG Aviation participated in the 4th Drag Prediction Workshop organized by AIAA in June 2009. Calculations were made for the Common Research Model configuration using the Navier–Stokes multiblock solver on the grids generated at CFS Engineering. After the workshop, the polars were computed on the medium multiblock structured grids provided by other workshop participants for the Common Research Model configuration with 0 deg horizontal tail deflection. All these results were processed by a far-field drag extraction tool developed jointly by RUAG Aviation and CFS Engineering. This paper first summarizes the theory behind the drag extraction tool. The results of the different calculations are presented and a detailed analysis of the drag breakdown on the different grids is given. This shows that the spread in drag coefficients obtained on different grids is much lower when using the effective drag computed by the drag extraction tool than when using the near-field drag.

The measurement and prediction of the aerodynamic performance of airfoils and wings at chord Reynolds numbers below 105 are both difficult and increasingly important in the application to small-scale aircraft. Not only are the aerodynamics strongly affected by the dynamics of the unstable laminar boundary layer, but the flow is decreasingly likely to be two-dimensional as the Reynolds number decreases. The spanwise variation of the nominally two-dimensional flow along a two-dimensional geometry is often held to be responsible for the large variations in the measured profile drag coefficient cd at this scale. Here, local two-dimensional drag coefficients are measured along a finite wing using nonintrusive particle image velocimetry methods. Variations in cd(y) can be related to local flow variations on the wing itself. Integrated values can then be compared with direct force balance data, and the dynamical significance of spanwise variability will be reevaluated.

A low-speed experimental investigation detailing the design and evaluation of a self-contained circulation-control wing is presented. A moderate aspect-ratio wing (3.34) utilizing a low-Reynolds-number S8036 profile was tested at Re=170,000. The trailing-edge jet was produced using two internal impellers drawing air off the wing tips. The effect of trailing-edge bluntness was examined as was the jet’s quiescent characteristics. Achievable jet-momentum coefficients were low, limiting circulation control to the boundary-layer control regime. Force-balance testing indicated that the circulation-control system could achieve a moderate zero-lift angle-of-attack shift; however, the primary effect of the jet was flow control at high incidence, in which stall was delayed significantly. The blunt trailing edge (Coanda surface) caused a significant increase in the minimum drag coefficient, reducing both the range and endurance parameters.

The previous works of Prandtl, Jones, and Klein and Viswanathan addressed the problem of determining the lift distribution that minimizes induced drag for a given lift and specified bending moment. In these formulations, bending moment is considered to be a surrogate for wing weight. These classical methods require the bending constraints to be imposed at the same lift coefficient at which drag is minimized. In practice, however, it is commonly desired to minimize drag at a representative cruise lift coefficient while imposing the bending constraints at a limiting structural load condition, such as a maneuver lift coefficient. This paper presents an approach to extend the classical methods by allowing the bending constraints to be imposed at different lift coefficients than that at which induced drag is minimized. An example for a wing planform similar to that of a Boeing 737 shows that the penalty for optimizing induced drag at the maneuver lift coefficient as implied in the classical methods results in ...

In this work, the performance of the assumed-transition-point criterion, in conjunction with Reynolds-averaged Navier–Stokes simulations is evaluated over the hull of a lighter-than-air vehicle. The evaluation is performed by comparing the Reynolds-averaged Navier–Stokes computations of drag, lift, and moment coefficients against those measured in experiments where 1) transition is forced to occur (forced-transition case) and 2) transition is allowed to occur naturally (free-transition case). It was found that by using the assumed-transition-point criterion Reynolds-averaged Navier–Stokes computations satisfactorily recover the drag coefficient both at zero angle of attack and nonzero angles of attack up to five degrees when compared against the forced-transition case. However, irrespective of the location of the assumed transition point, Reynolds-averaged Navier–Stokes fails to recover the experimentally observed values of the force and moment coefficients for the free-transition case. At this point, it ...

A series of wind-tunnel experiments were conducted in which the drag characteristics and inflated geometry of model parachute canopies with rectangular parallelepiped geometries (polyhedron) were examined. The model-canopy layouts were the same as cross canopies with the adjacent sides completely attached together. All models had a base dimension of 0.2 m, and aspect ratios ranged from 0.2 to 1.2. The models did not have a central vent or any other geometric porosity. The data show the inflated geometry of the canopy differs from the constructed geometry with the smallest change occurring at a constructed aspect of 0.8 and the variation becomes larger for increasing or decreasing constructed aspect ratios. The data also indicate the aerodynamic drag coefficient, based on the projected area, has a maximum value of approximately one for the constructed aspect ratio of 0.3 corresponding to an inflated aspect ratio of 0.53. The drag coefficient is less for smaller and larger aspect-ratio models. If scaled by ...

Multidisciplinary Design Optimization (MDO) studies show the Strut/Truss Braced Wing (SBW/TBW) concept has the potential to save a significant amount of fuel over conventional designs. For the SBW/TBW concept to achieve these reductions, the interference drag at the wing strut juncture must be small compared to other drag sources. Computational Fluid Dynamics (CFD) studies have indicated that the interference drag is small enough to be manageable. However, the RANS formulation and turbulence models used in these studies have not been validated for high Reynolds number transonic junction flows. This study assesses turbulence models by comparing flow separation characteristics obtained from experiment and CFD. The test model used is a NACA 0012 wing of aspect ratio 2 at Mach number of 0.76 and a Reynolds number of 6 million with varying angle of attack. The CFD study involved an 18.8 million cell structured grid of the wind tunnel test section using the ANSYS Fluent 12.0 solver. The k-ω SST turbulence model was the main turbulence model employed. Experiments were conducted in a high Reynolds number transonic Ludwieg tunnel. The wing was tested at different Mach numbers and inlet conditions to account for some of the experimental variations. Porous walls eliminate shock reflection across the tunnel. Surface oil flow visualization is used to indicate the interference flow patterns. The assessment shows CFD overpredicts separation and therefore interference drag, likely due to deficiencies in the turbulence model. Nomenclature

0.73 and0.78.Measurements are acquiredat several radial locations inphase-locked mode, encompassing the blade length;specificdatapostprocessingisaimedatdeterminingtheaerodynamicforces,namely,thesectionalthrustand torque at each cross section. The evaluation of the pressure field is based on the integration of the Navier–Stokes equations with the experimental velocity fields expressed in a frame moving with the propeller blade. The approach returns the evaluation of the three-dimensional pressure field, in particular over the blade surface, considerably simplifyingthe flowvisualizationandanalysisinraisingpropellerstudies.Thevelocityandpressuredataarefurther integrated by means of a contour approach to yield the propeller thrust and torque, and compared to data directly derived from a multicomponent balance and from the engine power consumption. Nomenclature AS = cross-sectional area of model support, m 2 CP, CD = pressure and drag coefficients c = chord, m cmax = maximum chord, m D, D0 = drag of the propeller and of the model support, N dE, lS = diameter and length of the engine cowling, m dp = propeller diameter, m ds =i nfinitesimal surface element, m 2

An experimental study is presented investigating the effect of ground proximity on an airfoil at a low Reynolds number. A S8036 airfoil was evaluated atRe 120; 000 and 180,000. The groundwas simulated using a thin splitter plate. Ground clearance was varied from 1 to 0.1 of the chord, measured from the airfoil’s quarter-chord. Force balance as well as surface pressure and ground plane pressures were recorded. Two semiempirical prediction methods were developed to estimate the two-dimensional ground effect. Results indicate that ground proximity increases the lift coefficientwith a peak occurring at a gap of 0.2 of the chord.Basedupon lift coefficient, ground effect may reduce drag, while for a given airfoil incidence, it increases the drag coefficient. Surface pressure evaluation at 0 deg incidence indicated that ground effect increases suction on both airfoil surfaces. The theoretical methods showed close accord with the experiment and predicted the location of the lift increment maxima. Theoretical decomposition suggests that ground effect has the net impact of increasing airfoil camber and reducing the effective freestream.

As part of the NATO Applied Vehicle Technology 161 technical group, a study of the aerodynamic behavior of the stability and control configuration wind-tunnel model is presented. Sharp and round leading-edge versions of the model are computed. A validation of Reynolds-averaged Navier–Stokes predictions obtained using two block structured codes are made. Static cases are analyzed and compared with wind-tunnel measurements. The vortical flow features are described in detail for a range of angles of attack. The predictions are in good agreement with the experiments at low angles of attack, whereas for higher angles of incidence (alpha > 15), some discrepancies are seen. A dual vortex structure is present in this region for both leading-edge configurations, resulting in highly nonlinear aerodynamic behavior.

Nomenclature A = wing area, m Cdp = pressure (form) drag coefficient Cl = lift coefficient Cn = normal force coefficient Ct = tangential force coefficient C = momentum coefficient [ _ mjet jet =qA] Do = initial injection width, mm e = 2.71828. . . (Euler’s number) _ m = mass flow rate, kg=s q = dynamic pressure, N=m Xe = distance over which injection width/velocity increases by a factor of e, mm = angle of attack, deg jet = velocity of jet, m=s o = initial injection velocity, m=s 1 = freestream velocity, m=s

TO IMPROVE aircraft performance, innovative technology should be pursued to drastically reduce the weight and fuel consumption of the aircraft and to significantly increase aircraft mission payload and stall margin. Both military and commercial aircraft will benefit from the same technology. Flow control is the most promising route to bring significant performance improvement to aircraft [1–7]. Recently, Zha et al. [8–10] andWang et al. [11] have developed a promising airfoil flow control technique using a coflow jet, which significantly increases lift, stall margin, and drag reduction. The coflow jet airfoil is designedwith an injection slot near leading edge and a suction slot near trailing edge on the airfoil suction surface. The slots are opened by translating a large portion of the suction surface downward. A high-energy jet is injected tangentially near the leading edge and the same amount of mass flow is drawn in near the trailing edge. The turbulent shear layer between the main flow and the jet causes strong turbulence diffusion and mixing under the severe adverse pressure gradient, which enhances lateral transport of energy from the jet to the main flow and allows the main flow to overcome the severe adverse pressure gradient and remain attached at high angles of attack (AOA). The high-energy jet induces high circulation and hence generates high lift. The energized main flow fills the wake and therefore reduces drag. The coflow jet (CFJ) airfoil achieves net-zero mass-flux flow control and can significantly reduce the penalty to the propulsion system bydumping zero jetmass flow. In [11], the CFJ airfoils with three different slot sizes are computed using Spalart–Allmaras (S-A) model. It is observed that at lowAOA, the computed results agree well with the experiment. But at high AOA, the deviation of the lift and drag is large because a Reynoldsaveraged Navier–Stokes (RANS) model is not able to simulate the flow separation accurately. In particular, the predicted stall AOA by RANS is significantly smaller than the measured one. Recently, in [12], the DES predicts the stalled airfoil flow at high AOA significantly better than the RANS model. To overcome the disadvantages of RANSmodels and to avoid the intensive CPU requirement for LES, Spalart et al. developed the detached-eddy simulation (DES) strategy [13], which is a hybrid RANS and LES method. Near the solid surface within the wall boundary layer, the unsteady RANSmodel is realized in DES. Away from thewall surface, themodel automatically converts to large-eddy simulation (LES). By using the RANSmodel near thewall, the mesh size as well as the CPU time can be tremendously reduced. By using the LESmodel away from thewall, the regions of massive separation and other free shear flows are properly simulated. Its application for turbulence simulation has already achieved encouraging success, as shown in thework of Tarvin et al. [14], Spalart [15,16], Forsythe et al. [17], Viswanathan et al. [18], Squires et al. [19,20], Hansen and Forsythe [21], Subbareddy and Candler [22],Wang and Zha [23,24], and Im and Zha [12]. These flows using DES include those for airfoils, cylinders, forebodies, baseflows, etc. The results are qualitatively and quantitatively better than the solutions using RANS. The objective of this paper is to simulate the flowfield of a CFJ airfoil at high AOA using DES in order to better capture the jet mixing and flow separation. A low-diffusion E-CUSP scheme [25,26] with fifth-order weighted essentially nonoscillatory (WENO) scheme [27,28] is employed for the inviscid fluxes. The computed lifts atAOA 30, 37, and 39 agree better with the experiment than those using S-A one-equation RANS. The computed drag usingDES is also improved greatly, even though the deviation remains very large. The computed stall AOA using DES is 37 , which agrees very well with the experiment and is much better than the one using S-A one-equation RANS. The computed results using DES show significantly improvements at high AOA compared with those using S-A one-equation RANS.

Thisworkpresents the results of an experimental study on droplet deformation and breakupnear the leading edge of an airfoil. The experiment was conducted in the rotating rig test cell at the National Institute of Aerospace Technology (INTA) inMadrid, Spain. An airfoil model placed at the end of the rotating armwas moved at speeds of 50 to 90 m=s. Amonosized droplet generatorwas employed to produce droplets that were allowed to fall fromabove, perpendicular to the path of the airfoil at a given location. High-speed imaging was employed to observe the interaction between the droplets and the airfoil. The high-speed imaging allowed observation of droplet deformation and breakup as the droplet approached the airfoil near the stagnation line. A tracking software programwas used to measure, from the high-speed movies, the horizontal and vertical displacement of the droplet against time. The velocity, acceleration,Weber number, Bond number, Reynolds number, and drag coefficients were calculated along the path of a given droplet from the beginning of the deformation to the breakup and/or hitting of the airfoil. Results are presented for droplets with a diameter of 490 m at airfoil speeds of 50, 60, 70, 80, and 90 m=s.

This paper presents the computational studies done at ONERA in the context of the 4th AIAA CFD Drag PredictionWorkshop (DPW4). Furthermore, it gives a detailed description of the far-field methods developed in the Applied Aerodynamics Department. Concerning the DPW4 configuration, a grid convergence study and a downwash study are proposed. Then the effects due toMach andReynolds numbers variations are quantified. All the multiblock structured grids used in this work have been provided by The Boeing Company to the drag prediction community. All the Reynolds-averaged Navier–Stokes computations are performed by using the ONERA-elsA solver with the Spalart–Allmaras turbulence model, and the solutions are postprocessed with the ONERA-ffd72 far-field drag-extraction tool. Concerning drag predictions, a very good agreement has been observed between ONERA-elsA results and the near-field drag coefficients (pressure and friction) computed by other DPW4 participants such asBoeing orAirbus.Moreover, the far-field software ffd72 givesONERA the singular capability to determine the values of the different physical drag components (viscous, wave, and lift-induced productions). Concerning the pitching moment, ONERA results are very close to Boeing, Airbus, or DLR, German Aerospace Center predictions.

This paper presents results from a Navier-Stokes finite volume flow solver simulating the flowfields around a platform and cargo configuration representative of platforms used for military parachute airdrops. The platform and cargo configuration consists of a flat plate model with an aspect ratio (width/length) of 0.56, upon which a box representing cargo is placed. This combination is simulated in conditions approximating the fall of a container prior to, and after parachute deployment and inflation, including a full 360 ◦ angle of attack range at a Reynolds number of 2.94 ·10 6 (freestream velocity of 30ft/s). The static simulations approximate those cases where the tumbling of the container (mostly prior to parachute deployment) is slow enough to approximate near-steady state flow conditions. Results include lift, drag, and moment coefficients over the range of flow conditions, as well as pressure contours to help elucidate relevant flow physics around the pallet-cargo configuration. Results show the flow orientation (into the nose bumper or flat side first) significantly affects the drag and moment behavior, but not the lift coefficient. Lift curve slopes match well to previously published data on pallet and cargo geometries as well as flat plates with similar aspect ratios. Drag and moment coefficients were significantly different between flow orientations; drag coefficients exhibit asymmetry between positive and negative angles of attack while moment coefficients illustrate that flow into the nose bumper may be more stable during extraction but much less stable during steady descent.

The paper describes a study of the effects of several key aerodynamic considerations on the conceptual design of minimum fuel / emissions, long range transport, truss-braced wing aircraft configurations. This unconventional configuration has a large benefit over conventional cantilever wing configurations. The truss system enables an increased aspect ratio with lower sweep, thickness ratio, and chords, thus exploiting natural laminar flow. The design problem is solved by an MDO process, which takes into account both aerodynamic and structural considerations. The paper contains several studies, each of which investigates the dependency of the design space on a specific aerodynamic parameter such as the extent of laminar flow on the wing, cruise Mach number, maximum cruise twodimensional lift coefficient, the supercritical characteristics of the airfoil, winglet influence, and intersection fairing design. In addition, various fuselage drag reduction technologies are investigated: fuselage relaminarization, surface riblets, tailless arrangements and Goldschmied apparatus. All of these studies illustrate large potential of the truss-braced wing along with additional drag reduction technologies, which may substantially decrease the fuel weight and vehicle emissions. The paper emphasizes the importance of appropriate wing section airfoils, which can satisfy various contradicting criteria of natural laminar flow, supercritical characteristics, high lift coefficient, and low drag coefficient. Finally, these studies illustrate the importance of fuselage drag reduction for a low induced drag, NLF aircraft like a truss-braced wing configuration.

Nomenclature AR = aspect ratio, b=S b = span, mm CD = configuration drag coefficient CL = configuration lift coefficient c = chord, mm ds = angle of strake setting relative to wing chord plane (positive, strake leading edge up), q = freestream dynamic pressure, 1 2 1V 2 1, kPa Re = Reynolds number based on wing model chord, 1V1c= 1 S = reference area, m V1 = freestream velocity, m=s x, y, z = aerodynamic axes = angle of attack, deg = absolute viscosity, N s=m = air density, kg=m

To improve aerodynamic performance, the active and passive flow control parameters are optimized in different combinations. The flow is simulated by solving the Navier–Stokes equations and a turbulence model. The active flow control is applied using flow suction or blowing on an airfoil, and its design variables are set to be the location and the angle of the suction/blowing port and the mass flow coefficient at this port. Airfoil surfaces are parameterized by using Bezier curves to optimize an airfoil shape in a passive way. To validate the computational methodology, successful comparisons have beenmadewith available data. A reel coded genetic algorithm is used in all optimization processes. For the demonstration problems considered herein, the best result is provided by the sequential optimization process, which results in 575.8% increment in aerodynamic performance, 183.76% increment in lift coefficient, and 68% decrease in drag coefficient.