Shown here is the polar as presented in 'Der Streng Geheime Vogel Me 163' by Wolfgang Späte, Appendix C, page 309-310. It is a combination of actual measurements (source yet unknown) and calculations. The calculations for Mach 0.5 to 0.8 were made according to a method presented in 'Aircraft performance, stability and control' by Perkins and Hage, Wiley, New York, 1963. The graph identifies the polars as approximations. The graph identifies CL=0.57 as the lift-off value.
According to the accompanying text, the basic polar (Mach below 0.5) is given by the standard equation CD = CD0 + CL^2 / (pi*A*e). CD0 is 0.013, aspect ratio 4.41, Oswald correction factor 0.786. The aspect ratio figure is based on wing surface figure of 19.6 m^2, while I keep measuring 17.4 m^2, using the standard definition of wing area of extending the leading and trailing edges to the centerline. Possibly the gross area (wing plus fuselage) is 19.6 m^2. Using a wing area figure of 17.4 m^2, I calculate an aspect ratio of 5.0.
Measurement of the data points and graphs below.
The data points were measured digitally.
|Drag coefficient||Lift coefficient|
The curves were measured and the equations below were the calculated as best fit equations.
|M<0.5||CD = 0.0130 + 0.0953 * CL^2|
|M=0.6||CD = 0.0141 + 0.1037 * CL^2|
|M=0.7||CD = 0.0154 + 0.1015 * CL^2|
|M=0.8||CD = 0.0283 + 0.4473 * CL^2|