Elizabeth M. Lee-Rausch

Wing flutter boundary prediction using unsteady Euler aerodynamic method

Modifications to an existing three-dimensional, implicit, upwind Euler/Reynolds-averaged Navier-Stokes code (CFL3D Version 2.1) for the aeroelastic analysis of wings are described. These modifications, which were previously added to CFL3D Version 1.0, include the incorporation of a deforming mesh algorithm and the addition of the structural equations of motion for their simultaneous time-integration with the governing flow equations. This article gives a brief description of these modifications and presents unsteady calculations that check the modifications to the code. Euler flutter results for an isolated 45-deg swept-back wing are compared with experimental data for seven freestream Mach numbers that define the flutter boundary over a range of Mach number from 0.499 to 1.14. These comparisons show good agreement in flutter characteristics for freestream Mach numbers below unity. For freestream Mach numbers above unity, the computed aeroelastic results predict a premature rise in the flutter boundary as compared with the experimental boundary. Steady and unsteady contours of surface Mach number and pressure are included to illustrate the basic flow characteristics of the time-marching flutter calculations and to aid in identifying possible causes for the premature rise in the computational flutter boundary.

Application of Parallel Adjoint-Based Error Estimation and Anisotropic Grid Adaptation for Three-Dimensional Aerospace Configurations

This paper demonstrates the extension of error estimation and adaptation methods to parallel computations enabling larger, more realistic aerospace applications and the quantification of discretization errors for complex 3-D solutions. Results were shown for an inviscid sonic-boom prediction about a double-cone configuration and a wing/body segmented leading edge (SLE) configuration where the output function of the adjoint was pressure integrated over a part of the cylinder in the near field. After multiple cycles of error estimation and surface/field adaptation, a significant improvement in the inviscid solution for the sonic boom signature of the double cone was observed. Although the double-cone adaptation was initiated from a very coarse mesh, the near-field pressure signature from the final adapted mesh compared very well with the wind-tunnel data which illustrates that the adjoint-based error estimation and adaptation process requires no a priori refinement of the mesh. Similarly, the near-field pressure signature for the SLE wing/body sonic boom configuration showed a significant improvement from the initial coarse mesh to the final adapted mesh in comparison with the wind tunnel results. Error estimation and field adaptation results were also presented for the viscous transonic drag prediction of the DLR-F6 wing/body configuration, and results were compared to a series of globally refined meshes. Two of these globally refined meshes were used as a starting point for the error estimation and field-adaptation process where the output function for the adjoint was the total drag. The field-adapted results showed an improvement in the prediction of the drag in comparison with the finest globally refined mesh and a reduction in the estimate of the remaining drag error. The adjoint-based adaptation parameter showed a need for increased resolution in the surface of the wing/body as well as a need for wake resolution downstream of the fuselage and wing trailing edge in order to achieve the requested drag tolerance. Although further adaptation was required to meet the requested tolerance, no further cycles were computed in order to avoid large discrepancies between the surface mesh spacing and the refined field spacing.

Rotor Airloads Prediction Using Unstructured Meshes and Loose CFD/CSD Coupling

The FUN3D unsteady Reynolds-averaged Navier-Stokes solver for unstructured grids has been modified to allow prediction of trimmed rotorcraft airloads. The trim of the rotorcraft and the aeroelastic deformation of the rotor blades are accounted for via loose coupling with the CAMRAD II rotorcraft computational structural dynamics code. The set of codes is used to analyze the HART-II Baseline, Minimum Noise and Minimum Vibration test conditions. The loose coupling approach is found to be stable and convergent for the cases considered. Comparison of the resulting airloads and structural deformations with experimentally measured data is presented. The effect of grid resolution and temporal accuracy is examined. Rotorcraft airloads prediction presents a very substantial challenge for Computational Fluid Dynamics (CFD). Not only must the unsteady nature of the flow be accurately modeled, but since most rotorcraft blades are not structurally stiff, an accurate simulation must account for the blade structural dynamics. In addition, trim of the rotorcraft to desired thrust and moment targets depends on both aerodynamic loads and structural deformation, and vice versa. Further, interaction of the fuselage with the rotor flow field can be important, so that relative motion between the blades and the fuselage must be accommodated. Thus a complete simulation requires coupled aerodynamics, structures and trim, with the ability to model geometrically complex configurations. NASA has recently initiated a Subsonic Rotary Wing (SRW) Project under the overall Fundamental Aeronautics Program. Within the context of SRW are efforts aimed at furthering the state of the art of high-fidelity rotorcraft flow simulations, using both structured and unstructured meshes. Structured-mesh solvers have an advantage in computation speed, but even though remarkably complex configurations may be accommodated using the overset grid approach, generation of complex structured-mesh systems can require months to set up. As a result, many rotorcraft simulations using structured-grid CFD neglect the fuselage. On the other hand, unstructured-mesh solvers are easily able to handle complex geometries, but suffer from slower execution speed. However, advances in both computer hardware and CFD algorithms have made previously state-of-the-art computations routine for unstructured-mesh solvers, so that rotorcraft simulations using unstructured grids are now viable. The aim of the present work is to develop a first principles rotorcraft simulation tool based on an unstructured CFD solver.

Adjoint-Based Design of Rotors in a Noninertial Reference Frame

Optimization of rotorcraft flowfields using an adjoint method generally requires a time-dependent implementation of the equations. The current study examines an intermediate approach in which a subset of rotor flowfields are cast as steady problems in a noninertial reference frame. This technique permits the use of an existing steady-state adjoint formulation with minor modifications to perform sensitivity analyses. The formulation is valid for isolated rigid rotors in hover or where the freestream velocity is aligned with the axis of rotation. Discrete consistency of the implementation is demonstrated by using comparisons with a complex-variable technique, and a number of single- and multipoint optimizations for the rotorcraft figure of merit function are shown for varying blade collective angles. Design trends are shown to remain consistent as the grid is refined.

FUN3D and CFL3D Computations for the First High Lift Prediction Workshop

Workshop, held in Chicago in June 2010. The unstructured-grid code FUN3D and the structured-grid code CFL3D were applied to several different grid systems. The effects of code, grid system, turbulence model, viscous term treatment, and brackets were studied. The SST model on this configuration predicted lower lift than the Spalart-Allmaras model at high angles of attack; the Spalart-Allmaras model agreed better with experiment. Neglecting viscous cross-derivative terms caused poorer prediction in the wing tip vortex region. Output-based grid adaptation was applied to the unstructured-grid solutions. The adapted grids better resolved wake structures and reduced flap flow separation, which was also observed in uniform grid refinement studies. Limitations of the adaptation method as well as areas for future improvement were identified.

CFD Sensitivity Analysis of a Drag Prediction Workshop Wing/Body Transport Configuration

The current work re-visits calculations for the First AIAA Drag Prediction Workshop (DPW-I) configuration and uses a grid convergence study to evaluate the quantitative effects of discretization error on the code-tocode variation of forces and moments. Four CFD codes commonly used at NASA Langley Research Center are used in the study: CFL3D and OVERFLOW are structured grid codes, and NSU3D and FUN3D are unstructured grid codes. Although the drag variation reported in the summary of DPW-I results was for the constantlift cruise condition, the focus of the current grid convergence study is a constant angle-of-attack condition (α = 0◦) near the same cruise lift in order to maintain identical boundary conditions for all of the CFD codes. Forces and moments were computed on the standard DPW-I structured overset and node-based unstructured grids, and the results were compared for the required transonic drag polar case. The range in total drag predicted using the workshop standard grids at α = 0◦ was 14 counts. The variation of drag in terms of standard deviation was 6 counts. Additional calculations at α = 0◦ were performed on the two families of structured and unstructured grids to evaluate the variation in forces and moments with grid refinement. The structured grid refinement study was inconclusive because of difficulties computing on the fine grid. The grid refinement study for the unstructured grid codes showed an increase in variation of forces and moments with grid refinement. However, all of the unstructured grid results were not definitively in the range of asymptotic grid convergence. The study indicated that certain numerical schemes (central vs. upwind, thin-layer vs. full viscous) or other code-to-code differences may have a larger effect than previously thought on grid sizes considered to be “medium” or “fine” by current standards. ∗Member AIAA, Research Engineer NASA Langley Research Center(LaRC), Hampton, Virginia. †Associate Fellow AIAA, Senior Research Scientist NASA LaRC. ‡Associate Fellow AIAA, Research Fellow National Institute of Aerospace, Hampton, Virginia. §Senior Member AIAA, Research Scientist NASA LaRC. ¶Member AIAA, Research Engineer NASA LaRC. ‖Member AIAA, Research Scientist NASA LaRC. ∗∗Associate Fellow AIAA, Senior Research Scientist NASA LaRC. This material is declared a work of the U.S. Government and is not subject to copyright protection in the United States. Introduction The AIAA Applied Aerodynamics Technical Committee conducted a Drag Prediction Workshop (DPW-I) in the summer of 2001 to evaluate CFD transonic cruise drag predictions for subsonic transports. Workshop participants were required to calculate the lift, drag and pitching moment for the DLR-F4 wing-body configuration at the cruise condition (Mach = 0.75, CL = 0.5), as well as the Mach = 0.75 drag polar. The participants were given a required grid to run and were encouraged to develop their own grid. The DLR-F4 wing-body was chosen since it had been tested in multiple wind tunnels. 1 A total of 35 solutions were computed with 14 different CFD codes; multiple turbulence models were used; structured and unstructured grids were used; 21 solutions were submitted on the required grids and an additional 14 solutions were provided on grids developed by the participants. In Ref. 2, Levy et al. provided a description of the workshop requirements and summary of the data submitted by the workshop participants. Hemsch 3 analyzed all of the solutions using a statistical framework. The variation in the drag from all 35 solutions at the cruise condition as measured by an estimate of the population standard deviation was 0.0021. The variation in the drag from the experiment was 0.0004. Thus, the computational drag variation was over 5 times the variation between wind tunnels. Designers typically state that they require drag prediction within one count (one count = 0.0001). Thus, the wind tunnel variation was 4 times the designer’s requirement, and the CFD variation was 21 times the designer’s requirement. Roache4 stated that multiple grids must always be used in order to verify a CFD solution. The design of the first DPW-I did not require that the participants provide solutions on multiple grids. Hence, the solutions were evaluated in the original study without the benefit of a quantitative measure of grid convergence. Each participant was free to choose whichever turbulence model and numerical scheme that they preferred for their calculations. Additionally, in order to accommodate the maximum number of CFD codes possible, the transition was specified at the leading edge of the vehicle, i.e. fully turbulent, rather than matching the experimentally determined transition pattern. Also, although aeroelastic deformations were incorporated into the geometry, they

NASA Trapezoidal Wing Computations Including Transition and Advanced Turbulence Modeling

Flow about the NASA trapezoidal wing is computed with several turbulence models by using grids from the first high-lift prediction workshop in an effort to advance understanding of computational fluid dynamics modeling for this type of flowfield. Transition is accounted for in many of the computations. In particular, a recently developed four-equation transition model is used and works well overall. Accounting for transition tends to increase lift and decrease moment, which improves agreement with the experiment. Upper surface flap separation is reduced, and agreement with experimental surface pressures and velocity profiles is improved. The predicted shape of wakes from upstream elements is strongly influenced by grid resolution in regions above the main and flap elements. Turbulence model enhancements to account for rotation and curvature have the general effect of increasing lift and improving the resolution of the wing-tip vortex as it convects downstream. However, none of the models improve the predi...

Three-dimensional effects in multi-element high lift computations

Abstract In an effort to discover the causes for disagreement between previous two-dimensional (2-D) computations and nominally 2-D experiment for flow over the three-element McDonnell Douglas 30P-30N airfoil configuration at high lift, a combined experimental/CFD investigation is described. The experiment explores several different side-wall boundary layer control venting patterns, documents venting mass flow rates, and looks at corner surface flow patterns. The experimental angle of attack at maximum lift is found to be sensitive to the side-wall venting pattern: a particular pattern increases the angle of attack at maximum lift by at least 2°. A significant amount of spanwise pressure variation is present at angles of attack near maximum lift. A CFD study using three-dimensional (3-D) structured-grid computations, which includes the modeling of side-wall venting, is employed to investigate 3-D effects on the flow. Side-wall suction strength is found to affect the angle at which maximum lift is predicted. Maximum lift in the CFD is shown to be limited by the growth of an off-body corner flow vortex and consequent increase in spanwise pressure variation and decrease in circulation. The 3-D computations with and without wall venting predict similar trends to experiment at low angles of attack, but either stall too early or else overpredict lift levels near maximum lift by as much as 5%. Unstructured-grid computations demonstrate that mounting brackets lower the lift levels near maximum lift conditions.

Grid-Adapted FUN3D Computations for the Second High Lift Prediction Workshop (Invited)

Contributions of the unstructured Reynolds-averaged Navier-Stokes code FUN3D to the 2 nd AIAA CFD High Lift Prediction Workshop are described, and detailed comparisons are made with experimental data. Using workshop-supplied grids, results for the clean wing conguration are compared with results from the structured code CFL3D Using the same turbulence model, both codes compare reasonably well in terms of total forces and moments, and the maximum lift is similarly over-predicted for both codes compared to experiment. By including more representative geometry features such as slat and ap brackets and slat pressure tube bundles, FUN3D captures the general eects of the Reynolds number variation, but under-predicts maximum lift on workshop-supplied grids in comparison with the experimental data, due to excessive separation. However, when output-based, o-body grid adaptation in FUN3D is employed, results improve considerably. In particular, when the geometry includes both brackets and the pressure tube bundles, grid adaptation results in a more accurate prediction of lift near stall in comparison with the wind-tunnel data. Furthermore, a rotation-corrected turbulence model shows improved pressure predictions on the outboard span when using adapted grids.

Computational Analysis of the G-III Laminar Flow Glove

Under NASA’s Environmentally Responsible Aviation Project, flight experiments are planned with the primary objective of demonstrating the Discrete Roughness Elements (DRE) technology for passive laminar flow control at chord Reynolds numbers relevant to transport aircraft. In this paper, we present a preliminary computational assessment of the Gulfstream-III (G-III) aircraft wing-glove designed to attain natural laminar flow for the leading-edge sweep angle of 34.6 o . Analysis for a flight Mach number of 0.75 shows that it should be possible to achieve natural laminar flow for twice the transition Reynolds number ever achieved at this sweep angle. However, the wing-glove needs to be redesigned to effectively demonstrate passive laminar flow control using DREs. As a by-product of the computational assessment, effect of surface curvature on stationary crossflow disturbances is found to be strongly stabilizing for the current design, and it is suggested that convex surface curvature could be used as a control parameter for natural laminar flow design, provided transition occurs via stationary crossflow disturbances .