Me 163B polars

The polar as presented in 'Ein Dreieck fliegt; die Entwicklung der Delta-Flugzeuge bis 1945' by Alexander Lippisch and Fritz Trenkle. The exact source of the data is not mentioned. The dotted line represents the actual measurements, a table of which can be found at the bottom of the page. The solid line reads 'Gleichgewichtspolare' which I freely translate as 'polar in the trimmed condition' (not sure though!). The graph caption remarks that the drag is decreased by control surface deflection in high speed flight. I interpret this as a trim surface deflection, trailing edge down in high speed flight. The value of dCm/dCa is positive, which means the pitching moment increases to a more nose-up value as the lift coefficient increases. The presence of the 'Gleichgewichtspolare' line and the dCm/dCa value suggest that the measurements are wind tunnel (as opposed to in-flight) measurements.

Max L/D of the dotted line occurs around the Cl=0.3 data point, and gives 15.8. The solid line achieves an L/D of 19.2 at a Cl of about 0.27. 19.2 is very high for an aircraft with an aspect ratio of only 5.

A big question mark is which reference wing area was used to calculate the coefficients of this graph. According to current conventions, the wing area is 17.3 square meters. However, 17.6 square meters and 19.6 square meters are also used in German wartime documents. There is one clue for this graph. RAF flight tests after the war (see 'AVIA 6/10072' file in the National Record Office) report a maximum lift coefficient of 1.0 for a stall speed of 83 mph at 3800 lbs. This translates into a wing area of calculated with a wing area of 20.0 square meters, which is close to the 19.6 square meters used by the Germans. Since the maximum lift coefficient of this graph is 0.95, it appears that 19.6 square meters is the reference area of this polar.

Drag coefficient Lift coefficient
0.0201 -0.200
0.0123 -0.036
0.0115 0.000
0.0104 0.048
0.0109 0.093
0.0123 0.145
0.0140 0.194
0.0162 0.241
0.0190 0.301
0.0266 0.408
0.0392 0.581
0.0539 0.709
0.0782 0.873
0.1003 0.948
0.1164 0.894

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