Aerodynamic Characteristics of an 11-Percent-Thick Symmetrical Supercritical Airfoil at Mach Numbers Between 0.30 and 0.85

An investigation was conducted in the Langley 8-foot transonic pressure tunnel over a Mach number range of 0.30 to 0.85 to determine the aerodynamic characteristics of an 11-percent-thick symmetrical supercritical airfoil. The Reynolds number of the tests, based on the airfoil chord, varied with Mach number over a range of 3.60 x 106 to 7.74 x 106. The geometric angle of attack varied from -0.5° to 10.5°. The results of the investigation indicate that the abrupt drag rise for the supercritical airfoil at zero-normal-force conditions occurs at a Mach number just above 0.82. The corresponding drag-rise Mach number for a conventional NACA 0012 airfoil is approximately 0.70. At zero-normal-force conditions, the level of supervelocity over the supercritical airfoil is considerably reduced from that for the NACA 0012 airfoil. Also, the shock wave for the supercritical airfoil is substantially weaker than that for the NACA 0012 airfoil. For a Mach number of 0.82 and zero normal force, the flow over the present airfoil is supercritical; however, there is no discernible shock wave in the flow, indicating near-isentropic recompression. At moderate-normal-force conditions, the supercritical airfoil has only a s light improvement over the conventional NACA 0012 airfoil in drag-rise Mach number.