F. Vogel, C. Breitsamter, N.A. Adams

The flow around a helicopter fuselage was investigated experimentally. For this purpose a scaled windtunnel model was instrumented with steady and unsteady pressure probes at the lower part of the fuselage tail section. The flow field in the wake area of the helicopter was measured with a Stereo-PIV system. A strong positive pressure gradient followed by stagnation at the lower tail section of the fuselage defines the flow separation. The pressure distribution for the base line case show some indications of vortex generation at the tail, which develops downstream and can also be detected by the flow field measurements. The influence of the support strut of the wind tunnel model was demonstrated by two model mount variations. The vortex strength is reduced by the disturbing effect of the support strut. A first estimation of the position of boundary layer transition was made. The unsteady effects of the flow separation at the lower tail section of the fuselage were investigated.

35th European Rotorcraft Forum 2009,

Conference Paper

englisch

21,0 x 29,7 cm, 9 Seiten

DGLR-Bericht, 2009, 2009-03, 35th European Rotorcraft Forum 2009 - Conference Proceedings; S.1-9; 2009; Deutsche Gesellschaft für Luft- und Raumfahrt - Lilienthal-Oberth e.V., Bonn

NA

in getr. Zählung;

flow distribution, fuselages, rotorcrafts

Bestellbar

2009