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What is the parasite drag or CD,0 for the Airbus A320? How do we calculate CD? Is it correct to use CD = CD,0 + ( CL2)/$\pi$AR

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  • $\begingroup$ A good aircraft design has a $c_{D0}$ of 0.012, an average one 0.02, and bad ones exceed 0.03. If the gear is not retractable or the wing is held by bracing wires, the $c_{D0}$ can even reach 0.05. Gliders have a $c_{D0}$ of below 0.005 at high speed. Calculation is hard: You need to bookkeep the contribution of every component. $\endgroup$ Jun 1, 2015 at 16:33

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For numerical usage (in some question) you can use the above formula, but for exact drag (and lift) calculation, the industry standard is to use Computer Solvers to compute this drag because of the following reasons:

  1. The formula used above is correct but the value of Cd varies for every point on the wing, as things like wing twist, sweep back, wing tip sharklets contribute to a different Cd throughout the wing.
  2. Other effects like drag due to fuselage etc can not be calculated by using a generalized equation for the whole plane, as usually local Cd for every point on the airplane surface are known.
  3. Application of force on the surface of an airplane (Lift/Drag force) causes deformation in the geometry of the surface and this leads to recursive change in the value of Cd/Cl, which hence needs further set of calculation to again arrive at a value of Cd from the Solver.

For approximate usage though an averaged out Cd, Cd0 and AR if known, then the above formula can be applied to arrive at the results.

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