Consider the wing planform shown in figure Q3. It is an unswept wing of span b = 12m and root chord cs = at = 0.5, and outer portion linear taper profile from 2 Ct = 1m. The wing has constant aerofoil section with lift curve slope ko 27 per radian and has a drag bucket profile drag coefficient 0.002. Wing performance is to be estimated from lifting line theory limiting the calculation to the first two non-zero A coefficients. i) Calculate the wing lift curve slope. ii) For the untwisted wing calculate the wing drag polar in the form Cp Cpo + KC. %3D 1.6m with inner portion linear taper profile down to chord length c = = 1.2m 0.5 to the tip with tip chord %3D %3D %3D
Consider the wing planform shown in figure Q3. It is an unswept wing of span b = 12m and root chord cs = at = 0.5, and outer portion linear taper profile from 2 Ct = 1m. The wing has constant aerofoil section with lift curve slope ko 27 per radian and has a drag bucket profile drag coefficient 0.002. Wing performance is to be estimated from lifting line theory limiting the calculation to the first two non-zero A coefficients. i) Calculate the wing lift curve slope. ii) For the untwisted wing calculate the wing drag polar in the form Cp Cpo + KC. %3D 1.6m with inner portion linear taper profile down to chord length c = = 1.2m 0.5 to the tip with tip chord %3D %3D %3D
Advanced Engineering Mathematics
10th Edition
ISBN:9780470458365
Author:Erwin Kreyszig
Publisher:Erwin Kreyszig
Chapter2: Second-order Linear Odes
Section: Chapter Questions
Problem 1RQ
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Transcribed Image Text:inner
Cs =1.6m
outer
6/4
Ct =1.0m
trailing edge
b/2
Figure Q3: Wing planform for Q3
..mid-span
1.2m
Transcribed Image Text:b) Consider the wing planform shown in figure Q3. It is an unswept wing of span b = 12m and
root chord Cs =
1.6m with inner portion linear taper profile down to chord length c
1.2m
at i = 0.5, and outer portion linear taper profile from
Ct = 1m. The wing has constant aerofoil section with lift curve slope ko = 2TT per radian and has
a drag bucket profile drag coefficient 0.002. Wing performance is to be estimated from lifting
line theory limiting the calculation to the first two non-zero A coefficients.
i) Calculate the wing lift curve slope.
ii) For the untwisted wing calculate the wing drag polar in the form Cp CD, + KC.
= 0.5 to the tip with tip chord
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