The Aeroelastic Prediction Workshop brought together an international community of computational fluid dynamicists as a step in defining the state of the art in computational aeroelasticity. This workshop's technical focus was prediction of unsteady pressure distributions resulting from forced motion, benchmarking the results first using unforced system data. The most challenging aspects of the physics were identified as capturing oscillatory shock behavior, dynamic shock-induced separated flow and tunnel wall boundary layer influences. The majority of the participants used unsteady Reynolds-averaged Navier Stokes codes. These codes were exercised at transonic Mach numbers for three configurations and comparisons were made with existing experimental data. Substantial variations were observed among the computational solutions as well as differences relative to the experimental data. Contributing issues to these differences include wall effects and wall modeling, nonstandardized convergence criteria, inclusion of static aeroelastic deflection, methodology for oscillatory solutions, post-processing methods. Contributing issues pertaining principally to the experimental data sets include the position of the model relative to the tunnel wall, splitter plate size, wind tunnel expansion slot configuration, spacing and location of pressure instrumentation, and data processing methods.

Reduced-order modelling (ROM) methods are applied to the Computational Fluid Dynamics (CFD)-based aeroelastic analysis of the AGARD 445.6 wing in order to gain insight regarding well-known discrepancies between the aeroelastic analyses and the experimental results. The results presented include aeroelastic solutions using the inviscid Computational Aeroelasticity Programme-Transonic Small Disturbance (CAP-TSD) code and the FUN3D code (Euler and Navier-Stokes). Full CFD aeroelastic solutions and ROM aeroelastic solutions, computed at several Mach numbers, are presented in the form of root locus plots in order to better reveal the aeroelastic root migrations with increasing dynamic pressure. Important conclusions are drawn from these results including the ability of the linear CAP-TSD code to accurately predict the entire experimental flutter boundary (repeat of analyses performed in the 1980s), that the Euler solutions at supersonic conditions indicate that the third mode is always unstable, and that the FUN3D Navier-Stokes solutions stabilize the unstable third mode seen in the Euler solutions. IntroductionClassical linear aeroelastic analyses typically produce velocity-damping-frequency (V-g-f) plots and/or root locus plots. The use of these plots has enabled the aeroelastician to view the nature of the flutter mechanism(s) in addition to identifying the condition(s) at which flutter occurs. The rapid creation of these plots was facilitated by the use of linear unsteady aerodynamics and linear aeroelastic equations of motion (Adams and Hoadley 1993).During the last few years, higher order CFD-based methods have become an important method for the computation of nonlinear unsteady aerodynamics for use in aeroelastic analyses. The use of these higher order methods provides valuable insight regarding complex flow physics at conditions where linear methods are not theoretically valid. However, the increased computational cost associated with the computation of unsteady aerodynamics and aeroelastic responses using higher order methods has resulted in a subtle change in the manner in which the aeroelastician evaluates and interprets these analyses. First, the increased computational cost of these analyses has tended to dictate a 'snapshot' approach to aeroelastic analyses whereby the aeroelastic response at a handful of dynamic pressures is all that is computed. This 'snapshot' approach is used to identify the flutter dynamic pressure but the actual flutter mechanism is not easily discernible. Second, due to the complexity of the computational methods, methods that could

Static and dynamic aeroelastic analyses have been performed for the Ares I crew launch vehicle during atmospheric ascent. It is shown that, through the transonic speed range, there is a rapid change in the static aeroelastic center-of-pressure increment with increasing Mach number. The greatest sensitivity to grid resolution is observed through the transonic range. Dynamic aeroelastic analyses are also performed to assess the aeroelastic stability of the launch vehicle. Flexible dynamic linearized quasi-steady analyses using steady rigid line loads are compared with fully coupled aeroelastic time-marching computational fluid dynamic analyses. There are significant differences between the methods through the transonic Mach number range. The largest difference is at Mach 1. At that Mach number, the linearized quasi-steady method produces strong damping in modes 1 and 2. The unsteady computational aeroelastic method indicates that the first mode is significantly undamped, while mode 2 is strongly damped. The cause of the disparity in damping between modes 1 and 2 is also investigated. A vehicle with no protuberances other than rings produced damping values in modes 1 and 2 that were nearly identical. It is shown that the disparity in damping of modes one and two is due to asymmetric placement of protuberances around the vehicle circumference.

This paper summarizes the plans for the second AIAA Aeroelastic Prediction Workshop. The workshop is designed to assess the state-of-the-art of computational methods for predicting unsteady flow fields and aeroelastic response. The goals are to provide an impartial forum to evaluate the effectiveness of existing computer codes and modeling techniques, and to identify computational and experimental areas needing additional research and development. This paper provides guidelines and instructions for participants including the computational aerodynamic model, the structural dynamic properties, the experimental comparison data and the expected output data from simulations. The Benchmark Supercritical Wing (BSCW) has been chosen as the configuration for this workshop. The analyses to be performed will include aeroelastic flutter solutions of the wing mounted on a pitch-and-plunge apparatus. Nomenclature Roman Symbols a speed of sound, ft/sec b wing span, in c, c re f chord length, chord reference length, in C p coefficient of pressure k reduced frequency M Mach number H, p total pressure, static pressure, psf Pr Prandtl number q dynamic pressure, psf T , T stat total temperature, static temperature, • F V free stream velocity, ft/sec f , f * , f f frequency, excitation frequency, flutter frequency, Hz

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