Reid Melville

Accuracy and Coupling Issues of Aeroelastic Navier-Stokes Solutions on Deforming Meshes

An implicit time-accurate approach to aeroelastic simulation was developed with particular attention paid to the issues of time accuracy, structural coupling, grid-deformation strategy, and geometric conservation. A Beam ‐Warming, approximate-factored algorithm, modie ed to include Newton-like subiterations was coupled with a structural model, also in subiteration form. With a sufe cient number of subiterations, this approach becomes a fully implicit, e rst- or second-order-accurate aeroelastic solver. The solver was used to compute time-accurate solutions of an elastically mounted cylinder. The fully implicit coupling allowed the overall scheme to become second-order accurate in time, signie cantly reducing the workload for a given accuracy. A new algebraic grid deformation strategy was developed that preserves grid orthogonality near the surface under large deformations. Finally, the oscillatory behavior of an elastically mounted cylinder was reproduced accurately by the present approach, and results compared favorably to previous experiments and simulations.

Implementation of a fully-implicit, aeroelastic Navier-Stokes solver

An implicit time-accurate approach to aeroelastic simulation is developed using a Beam-Warming, approximate factored algorithm, modified to include Newton-like subiterations and coupled with a structural model. With a sufficient number of subiterations, this approach becomes a fully implicit, firstor second-order accurate aeroelastic solver. The solver was used to compute solutions of an elastically mounted cylinder, a static aeroelastic wing deformation, and a dynamic wing flutter condition. All three showed agreement with other solutions and experiments. For the cylinder case, the second-order, implicit formulation showed significant improvements in solution accuracy. For the wing flutter case, there was little difference between the current formulation and a standard lagged, first-order method.

Transonic flutter simulations using an implicit aeroelastic solver

Flutter computations are presented for the AGARD 445.6 standard aeroelastic wing configuration using a fully implicit, aeroelastic Navier-Stokes solver coupled to a general, linear, second-order structural solver. This solution technique realizes implicit coupling between the fluids and structures using a subiteration approach. Results are presented for two Mach numbers, M∞ = 0.96 and 1.141. The computed flutter predictions are compared with experimental data and with previous Navier-Stokes computations for the same case. Predictions of the flutter point for the M∞ = 0.96 case agree well with experimental data. At the higher Mach number, M∞ = 1.141, the present computations overpredict the flutter point but are consistent with other computations for the same case. The sensitivity of computed solutions to grid resolution, the number of modes used in the structural solver, and transition location is investigated. A comparison of computations using a standard second-order accurate central-difference scheme and a third-order upwind-biased scheme is also made.

Fully implicit aeroelasticity on overset grid systems

An implicit time-accurate approach to aeroelastic simulation is developed using a Beam-Warming, approximate factored algorithm, coupled with a linear, second-order structural model. With subiteration, this approach becomes a fully implicit, second-order accurate aeroelastic solver. The flow domain is decomposed into an overset grid system for parallel implementation, and a grid deformation methodology is introduced to accommodate the deflection of aeroelastic surfaces. Several strategies for coupling the fluid and structural solvers are assessed for accuracy and efficiency.

Physical mechanisms for limit-cycle oscillations of a cropped delta wing

This paper presents simulations of limit-cycle oscillations of a cropped delta wing using a computational technique that implicitly couples a full Navier-Stokes solver to a linear structural solver. Computational results are compared with existing experimental measurements of the limit-cycle response of the delta wing. The limit-cycle oscillation is shown to consist primarily of a response of the first bending mode and the first torsional mode. Aerodynamic features of the limit cycle identified include a leading-edge vortex and a double shock structure. The leading edge vortex acts like an aerodynamic spring and provides the physical mechanism for the evolution of the limit cycle on the delta wing. Computations are repeated using the Euler equations to determine if the inviscid flow equations adequately simulate the observed limit cycle response.

Numerical Simulation of Limit-cycle Oscillations of a Cropped Delta Wing Using the Full Navier-Stokes Equations

This paper presents simulations of limit-cycle oscillations of a cropped delta wing using a computational technique that implicitly couples a full Navier-Stokes solver to a linear structural solver. Computational results are compared with existing experimental measurements of the limit-cycle response of the delta wing. The limit-cycle oscillation is shown to consist primarily of a response of the first bending mode and the first torsional mode. Aerodynamic features of the limit cycle identified include a leading-edge vortex and a double shock structure. The leading edge vortex acts like an aerodynamic spring and provides the physical mechanism for the evolution of the limit cycle on the delta wing. Computations are repeated using the Euler equations to determine if the inviscid flow equations adequately simulate the observed limit cycle response.

Numerical simulation of the jet produced by an internal aircraft explosion

Steady flowfields about a generic aircraft fuselage were simulated numerically by integration of the NavierStokes equations, including a two-equation (k-s) turbulence model. A steady, sonic, underexpanded jet issuing from a small square aperture in the fuselage surface was used to model the results of an internal explosion that ruptured the aircrafts skin. The computed solutions include simulations of a wind-tunnel test, and of flight at cruise conditions typical of a large transport aircraft. In each case, both the jet-on and jet-off situations were considered. Details of the computations are presented, and features of the flowfield are discussed. Comparisons were made with experimental data in terms of surface static pressure distributions and total pressure loss profiles, and found to be acceptable for engineering purposes.

Numerical simulation of large amplitude aeroelastic wing response

A dynamic aeroelastic solver is used to simulate large amplitude limit cycle oscillation on a thin plate wing. A Beam-Warming, approximate factored algorithm, coupled with a linear, second-order structural model via subiteration, becomes a fully implicit, second-order accurate aeroelastic solver. The flow domain is decomposed into an overset grid system for parallel implementation. Qualitative agreement with experimental data is achieved for a sustained nonlinear, aeroelastic oscillation. The response is seen to be a classical flutter that is limited by the interaction of the secondary bending mode of the wing. Several highly nonlinear flow features are observed on the wing over the course of the oscillation.

Accuracy issues for transonic flutter using 3-D Navier-Stokes

This paper presents flutter computations for the AGARD 445.6 standard aeroelastic wing configuration using a fully implicit, aeroelastic Navier-Stokes solver coupled to a general, linear, second-order structural solver. This solution technique realizes implicit coupling between the fluids and structures using a subiteration approach. Results are presented for two Mach numbers, MXJ = 0.96, where no clearly resolved shocks are present on the wing and Mx — 1.141 where a shock is resolved on the aft portion of the wing. The computed flutter predictions are compared with experimental data and with previous Navier-Stokes computations for the same case. Predictions of the flutter point for the MOO = 0.96 case agree well with experimental data. At the higher Mach number, M^ = 1.141, the present computations overpredict the flutter point but are consistent with other computations for the same case. The sensitivity of computed solutions to grid resolution and the number of modes used in the structural solver is investigated. A comparison of computations using a standard second-order accurate central-difference scheme and a third-order upwind-biased scheme is also made.

Numerical simulation of jet expulsion from an aircraft fuselage

Steady flowfields about a generic aircraft fuselage were simulated numerically by integration of the time-dependent, three-dimensional, compressible, massaveraged, Navier-Stokes equations. Effects of fine scale turbulence were accounted for by means of a twoequation (k E ) turbulence model which included a generalized formulation, low-Reynolds number terms, and a compressibility correction. A steady, sonic, underexpanded jet issuing from a small square aperture in the fuselage surface was used to model the results of an internal explosion which ruptured the aircrafts skin. The computed solutions include simulations of a wind tunnel test, and of flight a t cruise conditions typical of a large transport aircraft. In each case, both the jeton and jet-off situations were considered. Details of the computations are presented, and features of the flowfield are discussed. Comparisons are made with experimental data in terms of surface static pressure distributions and total pressure loss profiles.