Control of Separation Induced on the Wing of Transonic Speed Aircraft by Means of Micro Vortex Generator

Abstract

This experimental study provides an assessment of the potential of applying a control device to mitigate undesired
separation formed on the wing of an aircraft causing instability of the airliner. They are used in a wide range of
aircrafts and yet needs to be improved for wider applications. The study was conducted in a transonic wind
tunnel. The concept of passive control, led to applying a pair of micro vortex generator at a certain distance from
the location of normal shock on the wing. Detailed measurements of a fully developed flat plate turbulent
boundary layer were used to assess the performance of the control device. The performance of vortex generator
was assessed by comparing flow before separation and after the reattachment. Mean flow data such as static and
Preston pressure distributions, boundary layer total pressure and velocity profile and also Schlieren method were
used in evaluating the performance of the control device. Micro vortex generator has shown a reduction of 30%
on displacement thickness at the area of separation in comparison with undisturbed flow at shock location and an
outstanding result declared a reduction of about 60% in shape factor of the device at the shock location. The
result promises a considerable reduction of separation which provide more manoeuvrability and stability of the
aircraft.
Keywords: Boundary Layer, Micro Vortex Generator, Separation, Shock Wave, Transonic Flow, Wind Tunnel