Alexander Allen

Transport Aircraft Wake Influenced by Oscillating Winglet Flaps

The wake vortex development in the near field and extended near field behind a four engined large transport aircraft model fitted with an active winglet is presented. A detailed wind-tunnel investigation is conducted using a half-model focusing on the high-lift case of a typical approach configuration at a Reynolds number of 0.5 x 10 6 based on the wing mean aerodynamic chord and an angle of attack of 6.5 deg. The flowfield is observed using advanced hot-wire anemometry mainly focusing on the crossflow plane at 5.6 spans downstream of the model. Based on the time-dependent velocity components the wake flowfield is analyzed by distributions of mean vorticity, turbulence intensities, and spectral densities. Seven main vortical structures dominating the near-field wing vortex sheet roll up and merge to form the remaining trailing vortex in the extended near field. By the use of oscillating winglet flaps the velocity fluctuations at the core region of the remaining vortex are significantly influenced. Distinct narrowband concentrations of turbulent kinetic energy can be found for the farthest downstream plane documenting the presence of the disturbances generated by the winglet flaps which may result in an amplification of inherent far-field instabilities. The frequencies at which these narrowband energy concentrations occur are, on the one hand, dependent on the oscillation frequency of the winglet flaps; on the other hand, there are also independent energy concentrations within the frequency bands associated with wake instabilities.

Transport Aircraft Wake Influenced by a Large Winglet and Winglet Flaps

Detailed flowfields of a wind-tunnel investigation are discussed presenting the wake vortex development and evolution in the near field and extended near field behind a four-engined large transport aircraft model fitted with a large winglet. The tests use a half-model of 1:32 scale focusing on the high-lift case of a typical approach configuration at a Reynolds number of 0.5 x 10 6 based on the wing mean aerodynamic chord and at an angle of attack of 7 deg. Flowfields are carefully inspected by advanced hot-wire anemometry at seven crossflow planes up to 5.6 spans downstream of the model. Based on the measured time-dependent velocity components, the wake flowfield is analyzed by distributions of mean velocity and vorticity, turbulence intensities, and spectral densities. The near-field wing vortex sheet is dominated by seven main vortical structures, namely the winglet vortex, the wing tip vortex, the outboard flap vortex, the outboard and inboard nacelle vortices, and the vortices shed at the wing-body junction and the horizontal tail plane. In the extended near field, these vortices roll up and merge to form the remaining rolled-up vortex. The deflection of winglet flaps produce additional vortices influencing the wing tip near field and enhancing the overall merging process. Especially for the cases with asymmetrical flap deflection, the remaining rolled-up vortex shows a significant narrowband concentration of turbulent kinetic energy which may result in an amplification of inherent far-field instabilities.

Landing Gear Influence on the Wake Vortex of a Large Transport Aircraft

An experimental investigation on the wake vortex formation and evolution of a four-engine large transport aircraft model in the near field and extended near field has been conducted by means of hot-wire anemometry in a wind tunnel. The half-model used has a scale of 1:19.25 and the tests focus on the high-lift case of a typical landing configuration at a Reynolds number of 0.52 x 10 6 based on the wing mean aerodynamic chord. The flowfield is investigated up to 4.7 spans downstream of the model, which is investigated with and without landing gear. Based on the measured time-dependent velocity components, the wake flowfield is analyzed by distributions of vorticity, turbulence intensities, and spectral densities. The landing gear consists of a stay and four wheels, which create a bluff-body wake that interferes with the shear layer caused by the wing and wing junction. The landing gear creates a low-energy vortex system and a highly turbulent wake. Due to the landing gear, the main vortex centers are located slightly further outboard and the roll-up process is retarded, especially the inboard movement of the horizontal tail plane vortex. The velocity spectra of the landing gear wake show broadband characteristics with some moderate narrowband energy concentrations in the close near field, which do not enhance inherent wake instabilities. Although the turbulence caused by the landing gear is still present in the extended near field, only a small influence on the overall vortex system of the entire aircraft can be found. The landing gear influence is therefore negligible in the context of reducing minimum aircraft separation distances.

Computation of Delta Wing Flap Oscillations with a Reynolds-Averaged Navier-Stokes Solver

A selection of steady and unsteady validation results for the Reynolds-averaged Navier-Stokes solver FLM-NS with respect to an experimental delta wing test case is presented and compared to other numerical results. Having put the numerical method’s validity into evidence, the unsteady flow induced by an oscillating flap is investigated for a Fighter Type Delta Wing.

Flap efficiency of a delta wing with an external store using an Euler code for small disturbances

The results of a numerical investigation on flap efficiency of a delta wing with an external store attached are discussed. The wing has an inboard and an outboard flap. The store is attached to the wing on the lower surface in the area of the outboard flap. For this geometry a structured multiblock mesh has been generated and elliptically smoothed. Numerical investigations are carried out with an Euler code using a small disturbance solver (FLM-SDEu) and an unsteady solver (FLM-Eu). The first assumes that the unsteady disturbances are small in comparison to the overall steady mean flow [1] [2], the latter is only used for comparison to FLM-SDEu. The small disturbance code reduces the calculation time significantly and the results obtained show good conformity with the results from the unsteady solver [3] [4].