The following are all hand-calcs or short scripts. Compute the parasitic drag for a large commercial airliner (we will compute just the wing and fuselage). The flight conditions are Mach = 0.8, altitude = 35,000 ft. At this Reynolds number the boundary layer will be mostly turbulent so for all components we’ll assume fully turbulent to be conservative. You should compute the drag of each component in isolation to keep things simple for this problem. In other words, don’t try to subtract out the area of the wing that will be buried in the fuselage.
Open flightlab and use the Airfoils tab. Compare the performance of the following three airfoils: NACA 2412, E212, ESA40. You can find the last one, and a bunch more, at the following: database. The files for these airfoils are fine as is, but for future reference some files in the database are in slightly different formats and need to be adjusted to the standard format. Analyze the airfoils at a Reynolds number of 100,000 and a Mach number of 0. Start with natural transition (x/c=1.0). Download the CSV files to enable you to to plot the following graphs.
In the first set of graphs plot all 3 airfoils on one graph using natural transition: a lift curve (\(c_l\) vs \(\alpha\)), drag polar (\(c_l\) vs \(c_d\)), lift to drag ratio (\(l/d\) vs \(\alpha\)), and a moment curve (\(c_m\) vs \(\alpha\)). Briefly comment (a sentence or two) on the differences observed.
In the second graph plot the drag polar for just the NACA 2412 but with two curves: one with natural transition and a second curve using a forced transition at 5% chord (x/c=0.05).
Be sure that you analyze a broad enough angle of attack range to include the zero-lift angle of attack on the low end, and stall (or near-stall) at the high end. Going much past stall is not helpful as the assumptions of the method will be violated. When you plot the drag polars set an appropriate upper limit on the drag axis. Otherwise, if the drag shoots off to high values (past stall) you can barely see the part of the curve that you actually care about (it just looks flat). In contrast, moment cofficient only varies across a narrow range so the default axes may make it look much less flat than it really is. Often it is plotted on the same range as lift coefficient.