Abstract
In this contribution, three methodologies based on temperature-sensitive paint (TSP) data were further developed and applied for the optical determination of the critical locations of flow separation and reattachment in compressible, high Reynolds number flows. The methodologies rely on skin-friction extraction approaches developed for low-speed flows, which were adapted in this work to study flow separation and reattachment in the presence of shock-wave/boundary-layer interaction. In a first approach, skin-friction topological maps were obtained from time-averaged surface temperature distributions, thus enabling the identification of the critical lines as converging and diverging skin-friction lines. In the other two approaches, the critical lines were identified from the maps of the propagation celerity of temperature perturbations, which were determined from time-resolved TSP data. The experiments were conducted at a freestream Mach number of 0.72 and a chord Reynolds number of 9.7 million in the Transonic Wind Tunnel Göttingen on a VA-2 supercritical airfoil model, which was equipped with two exchangeable TSP modules specifically designed for transonic, high Reynolds number tests. The separation and reattachment lines identified via the three different TSP-based approaches were shown to be in mutual agreement, and were also found to be in agreement with reference experimental and numerical data.
Highlights
Published: 28 July 2021Improvements in the aerodynamics of commercial aircraft, and in particular in the reduction of drag [1,2], are needed in order to reach the targets in polluting emission reduction set by the European Commission [3] and by NASA [4]
This complex phenomenon [11,12,13] is likely to occur at transonic flow conditions on aircraft wings designed with laminar flow technology, where the laminar boundary layer over the suction side of the wing reaches supersonic speeds; the supersonic flow region is typically terminated by a shock, which interacts with the boundary layer and induces flow separation [1,14,15]
After a strong acceleration over the leading-edge region up to locally supersonic flow conditions, the boundary layer underwent a deceleration at approximately x/c > 10%, culminating in a very strong adverse pressure gradient at approximately 20.5% ≤ x/c ≤ 22.5% related to a shock wave, which terminated the supersonic flow region
Summary
Published: 28 July 2021Improvements in the aerodynamics of commercial aircraft, and in particular in the reduction of drag [1,2], are needed in order to reach the targets in polluting emission reduction set by the European Commission [3] and by NASA [4]. Flow separation may have a negative impact on the aerodynamic performance of the aircraft surfaces, especially when induced by a Shock-Wave/Boundary-Layer Interaction (SWBLI) [7,8,9,10,11] This complex phenomenon [11,12,13] is likely to occur at transonic flow conditions on aircraft wings designed with laminar flow technology, where the laminar boundary layer over the suction side of the wing reaches supersonic speeds; the supersonic flow region is typically terminated by a (normal) shock, which interacts with the boundary layer and induces flow separation [1,14,15].
Sign-in/Register to view highlights & summary Sign-in/Register to access full text options 
Talk to us
Join us for a 30 min session where you can share your feedback and ask us any queries you have
Schedule a call
