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Therefore, the Drag coefficient on a supersonic airfoil is described by the following expression: C D = C D,friction + C D,thickness + C D,lift. Experimental data allow us to reduce this expression to: C D = C D,O + KC L 2 Where C DO is the sum of C (D,friction) and C D,thickness, and k for supersonic flow is a function of the Mach number. [3]
For example, the NACA 2412 airfoil has a maximum camber of 2% located 40% (0.4 chords) from the leading edge with a maximum thickness of 12% of the chord. The NACA 0015 airfoil is symmetrical, the 00 indicating that it has no camber. The 15 indicates that the airfoil has a 15% thickness to chord length ratio: it is 15% as thick as it is long.
In February 1976, work commenced to automate the methods contained in the USAF Stability and Control DATCOM, specifically those contained in sections 4, 5, 6 and 7.The work was performed by the McDonnell Douglas Corporation under contract with the United States Air Force in conjunction with engineers at the Air Force Flight Dynamics Laboratory in Wright-Patterson Air Force Base.
The supersonic flow over a supercritical airfoil terminates in a weaker shock, thereby postponing shock-induced boundary layer separation. A supercritical aerofoil (supercritical airfoil in American English) is an airfoil designed primarily to delay the onset of wave drag in the transonic speed range.
The result was a long string of fundamental breakthroughs, including "thin airfoil theory" (1920s), "NACA engine cowl" (1930s), the "NACA airfoil" series (1940s), and the "area rule" for supersonic aircraft (1950s).
Thin airfoil theory assumes the air is an inviscid fluid so does not account for the stall of the airfoil, which usually occurs at an angle of attack between 10° and 15° for typical airfoils. [20] In the mid-late 2000s, however, a theory predicting the onset of leading-edge stall was proposed by Wallace J. Morris II in his doctoral thesis. [ 21 ]
Generally, the drag coefficient peaks at Mach 1.0 and begins to decrease again after the transition into the supersonic regime above approximately Mach 1.2. The large increase in drag is caused by the formation of a shock wave on the upper surface of the airfoil, which can induce flow separation and adverse pressure gradients on the aft portion ...
It is used for near-supersonic flight and produces a higher lift-to-drag ratio at near supersonic flight than traditional airfoils. Supercritical airfoils employ a flattened upper surface, highly cambered (curved) aft section, and greater leading-edge radius as compared to traditional airfoil shapes. These changes delay the onset of wave drag.