Benedikt Konig

Numerical and Experimental Validation of Three-Dimensional Shock Control Bumps

Numerical and experimental studies have been performed to show the potential for drag reductions of an array of discrete three-dimensional shock control bumps. The bump contour investigated was specifically designed by means of CFD-based numerical optimization for wind tunnel testing on a modern transonic airfoil. The experimental investigations focused on turbulent flow at a Reynolds number of 5 million and were carried out at the

Shock Control Bumps on Flexible and Trimmed Transport Aircraft in Transonic Flow

Shock control bumps are a means to reduce wave drag that occurs at the upper limit of civil transport aircrafts flight envelope. An SCB was optimized for and applied to the rigid wing-body DLR F11 model. The effect of the SCB on the trim drag of the F11 configuration with an attached horizontal tail plane was investigated. Flexibility of the wing-body model was considered and the influence of aeroelasticity on the SCB performance was examined. RANS simulations with the DLR FLOWer code showed, that both the influence of trimming as well as of aeroelasticity is negligible for SCB design.

Validation of a Transonic Lattice-Boltzmann Method on the NASA Common Research Model

A novel transonic Lattice-Boltzmann Method is validated against the NASA Common Research Model. This model offers the most comprehensive experimental data base available but also suffers from a number of issues that make a comparison to CFD difficult. In order to achieve a rigorous validation of the new code, the major effects present in the wind tunnel experiments are included in the modeling. This work describes Lattice-Boltzmann simulations of the Common Research Model with updated wing twist distributions and including the wind tunnel support system. Preliminary results for both the baseline geometry and the updated model are presented and show a promising correlation to the experiments.

A Comparative Study of Simulated and Measured Main Landing Gear Noise for Large Civil Transports

Computational results for the NASA 26%-scale model of a six-wheel main landing gear with and without a toboggan-shaped noise reduction fairing are presented. The model is a high-fidelity representation of a Boeing 777-200 aircraft main landing gear. A lattice Boltzmann method was used to simulate the unsteady flow around the model in isolation. The computations were conducted in free-air at a Mach number of 0.17, matching a recent acoustic test of the same gear model in the Virginia Tech Stability Wind Tunnel in its anechoic configuration. Results obtained on a set of grids with successively finer spatial resolution demonstrate the challenge in resolving/capturing the flow field for the smaller components of the gear and their associated interactions, and the resulting effects on the high-frequency segment of the farfield noise spectrum. Farfield noise spectra were computed based on an FWH integral approach, with simulated pressures on the model solid surfaces or flowfield data extracted on a set of permeable surfaces enclosing the model as input. Comparison of these spectra with microphone array measurements obtained in the tunnel indicated that, for the present complex gear model, the permeable surfaces provide a more accurate representation of farfield noise, suggesting that volumetric effects are not negligible. The present study also demonstrates that good agreement between simulated and measured farfield noise can be achieved if consistent post-processing is applied to both physical and synthetic pressure records at array microphone locations.

Scale-Resolving Simulations Based on a Lattice-Boltzmann Method

This paper gives an overview on the scale-resolving capabilities of the Lattice-Boltzmann Method as implemented in the solver PowerFLOW. The basic concept of the approach is outlined, which comprises the turbulence modelling strategy. Following this are three examples ranging from fundamental to geometrically complex test cases: a shear layer, the flow over the NASA Hump and the flow over an iced airfoil. It is shown for all cases that transition from modelled to resolved turbulent fluctuations is achieved automatically once the flow separates with a flow based sensor if grid resolution is sufficient. Agreement with experimental reference data is good. As this paper is only intended to give a general overview, additional insight into each case is available in the corresponding reference papers.