- The Section Geometry For A Naca 4 Digit Series Airfoil Is Completely Fixed By The Maximum Camber The Location Of Maximu 1 (240.47 KiB) Viewed 34 times
of attack and the quarter chord moment related to the value of
maximum thickness.
b. Considering the NACA 2410 airfoil, determine the lift
coefficient and moment coefficient at the location of 25% of the
chord. Given that the geometric angle of attack is 6 degrees.
c. The NACA 2410 is designed to have a trailing–edge flap
with a length equal to 20% of the chord. Flaps are fastened to the
main wing. Assume that the flaps are designed to have maximum
performance. Calculate the lift coefficient of the wing with flaps.
Given that the geometric angle of attack is 6 degrees and the flap
deflection angle is 5 degrees.
The section geometry for a NACA 4-digit series airfoil is completely fixed by the maximum camber, the location of maximum camber, the maximum thickness, and the chord length. The first digit indicates the maximum camber in percent of chord. The second digit indicates the distance from the leading edge to the point of maximum camber in tenths of the chord. The last two digits indicate the maximum thickness in percent of chord. The camber line is given by X Ymc 2 0SX SXmc Xmc Xmc y(x) = |---2-01 [() C-X Ymc C-X C-Xmc Xmc sxsc C-Xmc where yme is the maximum camber and Xose is the position of maximum camber. The thickness about the camber line, measured perpendicular to the camber line itself, is given by t(x) im 2.969 1 03 - 2004 - 1.00(9)-3.50()*288(11015(9 (9 –8)* * X +2.843 where in is the maximum thickness.