Experimental Study of a Transonic Flow over a NACA0012 Airfoil by Measurement of Pressure Fluctuations

In this research، shock wave motion، shock/boundary layer interaction، and flow separation in a transonic flow over a NACA0012 airfoil were investigated experimentally. Thirteen Kulite sensors were utilized for measurement of pressure fluctuations over the airfoil. Experiments were performed for angles of attack between -4 to 4 degrees and flow Mach numbers of M= 0. 39 to 0. 82. The Kulite sensors were distributed on one side of the airfoil surface from 13% to 63% of the chord from the leading edge. Data acquisitioning were performed using a 10KHz sample rate and a resolution of 0. 05 KPa/bit. A Schlieren system was also used for visualization of shock lines. The results show that the transition from laminar to turbulent cause increase in amplitude of pressure fluctuations during the process. If boundary layer separation occurs due to interaction of boundary layer with shock waves، amplitude of pressure fluctuations reduces compared to upstream of the separated region. The results also show that as Mach number increases، the onset location of laminar to turbulent transition moves toward the leading edge and the shock lines move toward the trailing edge. Another significant phenomenon was occurrence of a phase difference between pressure signals of two positions namely: after shock wave and inside separation bubble.