David M. Schuster

Computational Aeroelasticity: Success, Progress, Challenge

This article asserts a much broader dee nition of the term, one in whichCAEencompassesalllevelsofaeroelasticanalysis.Aeroelastic tools based on both linear unsteady aerodynamics and nonlinear CFD methods have been developed and successfully applied. We refer to both of these methodologies as components of CAE. Likewisestructuralmodelingassimpleasbeamtheorytostate-of-the-art e nite element modeling (FEM) have been incorporated into aeroelastic tools, and these techniques should also be included under the CAE heading. It is not the intention of this article to provide an exhaustive history of the development and application of CAE over the past 70 years. Rather, the subject will be examined in the context of two primary themes: 1) aeroelastic problems requiring theoretical

Static aeroelastic analysis of fighter aircraft using a three-dimensional Navier-Stokes algorithm

An aeroelastic analysis method for fighter aircraft operating at extreme flight conditions has been developed and tested. The method involves the use of state-of-the-art zonal grid generation methods, three-dimensional Reynoldsaveraged Navier-Stokes analysis, and linear structures to analyze the flow over complex, flexible aircraft. The main objective of this effort is to develop a method capable of analyzing aircraft operating at flight conditions where vortices, strong shock waves, separated flow, and even highly unsteady flow may be present. The present application focuses on the static aeroelastic analysis of fighter aircraft operating at high angle of attack and high transonic Mach number. The developed method has been compared against static aeroelastic wind-tunnel data on an aeroelastically tailored wing/fuselage configuration, and the results are very encouraging.

Overview of the Aeroelastic Prediction Workshop

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 workshops 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, non-standardized 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.

Transonic Unsteady Aerodynamics of the F/A-18E at Conditions Promoting Abrupt Wing Stall (Invited)

A transonic wind tunnel test of an 8% F/A-18E model was conducted in the NASA Langley Research Center (LaRC) 16-Foot Transonic Tunnel (16-Ft TT) to investigate the Abrupt Wing Stall (AWS) characteristics of this aircraft. During this test, both steady and unsteady measurements of balance loads, wing surface pressures, wing root bending moments, and outer wing accelerations were performed. The test was conducted with a wide range of model configurations and test conditions in an attempt to reproduce behavior indicative of the AWS phenomenon experienced on full-scale aircraft during flight tests. This paper focuses on the analysis of the unsteady data acquired during this test. Though the test apparatus was designed to be effectively rigid, model motions due to sting and balance flexibility were observed during the testing, particularly when the model was operating in the AWS flight regime. Correlation between observed aerodynamic frequencies and model structural frequencies are analyzed and presented. Significant shock motion and separated flow is observed as the aircraft pitches through the AWS region. A shock tracking strategy has been formulated to observe this phenomenon. Using this technique, the range of shock motion is readily determined as the aircraft encounters AWS conditions. Spectral analysis of the shock motion shows the frequencies at which the shock oscillates in the AWS region, and probability density function analysis of the shock location shows the propensity of the shock to take on a bi-stable and even tri-stable character in the AWS flight regime.

Development of an Euler/Navier-Stokes aeroelastic method for three-dimensional vehicles with multiple flexible surfaces

In 1989, an aeroelastic analysis method coupling high-level Computational Fluid Dynamics (CFD) methods with linear structural models was developed to analyze structurally flexible fighter aircraft This method, known as the Euler/Navier-Stokes 3Dimensional Aeroelastic (ENS3DAE) method was written to analyze both the steady and unsteady aerodynamic and aeroelastic performance of aircraft Initially, the code was validated for steady rigid sad structurally flexible configurations operating at a wide range of flight conditions. Since 1989, many enhancements to the aerodynamic and aeroelastic modules of the code have been added, including aeroelastic analysis of shell structures and vertical tails, aeroelastic analysis of multiple flexible surfaces, enhanced dissipation models, and real-gas capability. However, the program has not been validated for these options or for unsteady flows. The analyses described in this document are a portion of the effort to accomplish this validation.

Transonic Unsteady Aerodynamics of the F/A-18E Under Conditions Promoting Abrupt Wing Stall

A transonic wind-tunnel test of an 8% F/A-18E model was conducted in the NASA Langley Research Center 16-Foot Transonic Tunnel to investigate the abrupt wing stall characteristics of this aircraft. During this test, both steady and unsteady measurements of balance loads, wing surface pressures, wing-root bending moments, and outer-wing accelerations were performed. The test was conducted with a wide range of model configurations and test conditions in an attempt to reproduce behavior indicative of the abrupt wing stall phenomenon experienced in full-scale aircraft during flight tests. This study focuses on the analysis of the unsteady data acquired during this test. Though the test apparatus was designed to be effectively rigid, model motions due to sting and balance flexibility were observed during the testing, particularly when the tunnel was operated under conditions representative of those where wing drop was experienced in flight. The correlation between observed aerodynamic frequencies and model structural frequencies is analyzed and presented

Navier-Stokes simulation of burst vortex flowfields for fighter aircraft at high incidence

A computational study has been conducted to investigate vortex breakdown on a generic fighter conf iguration operating at high incidence. The flow simulations are based on solutions of the full Reynolds-averaged Navier-Stokes equations using an implicit finite-difference algorithm. Results of the analysis have been correlated with an experimental data base obtained using a 3-D Laser Velocimeter system. Good correlation has been obtained between the analysis and the data base in terms of the vortex core path, cross-flow velocity fields and the approximate vortex breakdown onset location.

Plans and Example Results for the 2nd AIAA Aeroelastic Prediction Workshop

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.

Analysis of Test Case Computations and Experiments for the First Aeroelastic Prediction Workshop

This paper compares computational and experimental data from the Aeroelastic Prediction Workshop (AePW) held in April 2012. This workshop was designed as a series of technical interchange meetings to assess the state of the art of computational methods for predicting unsteady flowfields and static and dynamic aeroelastic response. The goals are to provide an impartial forum to evaluate the effectiveness of existing computer codes and modeling techniques to simulate aeroelastic problems and to identify computational and experimental areas needing additional research and development. Three subject configurations were chosen from existing wind-tunnel data sets where there is pertinent experimental data available for comparison. Participant researchers analyzed one or more of the subject configurations, and results from all of these computations were compared at the workshop.

Transport Wing Flutter Model Transonic Limit Cycle Oscillation Test

The model for aeroelastic validation research involving computation semispan wind-tunnel model, a transport wing-fuselage flutter model, was tested in NASA Langleys Transonic Dynamics Tunnel with the goal of obtaining experimental limit cycle oscillation behavior data at transonic separation onset conditions. This research model is notable for its inexpensive construction and instrumentation installation procedures. Unsteady pressures and wing responses were obtained for three wing-tip configurations: clean, tip store, and winglet. Traditional flutter boundaries were measured over the range of M = 0.6-0.9, and maps of limit cycle oscillation behavior were made in the range of M = 0.85-0.95. The effects of dynamic pressure and angle of attack were measured. Testing in both R134a heavy gas and air provided unique data on the Reynolds number, transition effects, and the effect of speed of sound on limit cycle oscillation behavior. This report gives an overview of the test results, including experimental flutter boundaries, and the conditions involving shock-induced transonic flow separation onset at low wing angles, including maps of limit cycle oscillation behavior.