1) The document provides aerodynamic calculations for a wing and tail at a cruise speed of Mach 0.30257 at 30,000 feet, including drag coefficients for the wing, horizontal tail, vertical tail, and fuselage. 2) Calculations are shown for the wing wetted area, reference area, aspect ratio, taper ratio, and Oswald efficiency factor. Drag coefficients are calculated based on laminar flow percentages and Reynolds numbers. 3) Tail drag coefficient calculations are also provided for the horizontal and vertical tails, accounting for laminar flow percentages and Reynolds numbers at the root and tip. 4) Fuselage drag coefficient and wetted area calculations are presented based on the fu

1. 0
The Drag Polar at Maximum cruisespeed of Mach 0.30257at30,000ft.
From Design Data Sheets:
 = 0.0027717 slug/ft3,  = 3.1324010-7 lb.sec/ft2, a = 1116.45 ft/s, then
the airspeed is = 337.8 ft/s
1. Wing contribution:
Assume35% laminar flow on the root and 45% at the tip, Cr = 14.07 ft,
given Ct = 4.11
Rer = 30993843.18
Then 0.02975
Then Cfr = 1.753*10-9
 At tip:
Ret = 9053638.626
Then 0.05523
Then Cft = 5.469*10-9
Cfavg = 3.611*10-9
Exposed wing area:
, SEXPO = 397.0931 ft2
(t/c) = 0.12, λ = 0.25, τ= 1.4, Swet =825.63 ft2
Refu = 2.0045*108

2. 1
Fromdata design sheets:
M Rwf
0.8 0.97
0.78 X
0.7 0.99 , then X = Rwf = 0.986
AR = 7.8, Ʌn = 21.72o
, n =0.25, m = 0.37, λ = 0.25, then Ʌm = 19.8o
at
max
cos Ʌm = 0.94
Fromdata design sheets:
M RLS
0.6 1.15
0.78 x
0.8 1.26
Then at 0.78, RLS = 1.249
L = 1.2
,Then (CDo)w = 1.175*10-8
2). Tail:
Tail; horizontal: (assuming at root45% laminar, at tip 60%)
At root: (Cr = 6.575 ft)
Ret = 14483618.97, Ret =14.48*106

3. 2
0.046, then Cfr = 1.808*10-3
At tip (Ct = 4.11)
Ret = 9.0536*106
0.0661, then Cft = 1.631*10-3
Cfavg = 1.72*10-3
, Sh =120.6 ft2
, from the Previous equation then;
Swet = 250.217 ft2
.
AR = 5.0793, Ʌn =20o
, n = 0, m = 0.37, λ = 0.625, then Ʌm = 16.52o
at
max
Then cos(Ʌm) = 0.95o
M RLs
0.6 1.15
0.78 X
0.8 1.27 then RLs = 1.258, L = 1.2
(CDo
)H =0.001144463
Tail; vertical: (assuming 50% laminar flow at root and 35% at the tip)
At tip: (Ct = 6.49ft)
Ret = 14.29*106
, 0.00397,then Cft = 2.037*10-3
, L =1.2,

4. 3
At root: (Cr =11.73 ft)
Re = 25.839*106
0.0404, Cfr =1.489*10-3
Cavg = 1.763*10-3
, Sexpov = 77.5 ft2
,
(t/c) = 0.12, λ = 0.553, τ= 1.4, then Swetv = 160.84 ft2
AR = 1.6392, Ʌn =40o
, n = 1, m = 0.37, λ = 0.554, then Ʌm = 30.112o
at
max
Cos(Ʌm) =0.86
M RLs
0.6 1.14
0.78 X
0.8 1/24 ,then RLs = 1.23
(CD0)v = 0.000737271
3). Fuselage contribution: (assuming 10% laminar flow)
L = 91.633 ft (correcting value by WebPlotDigitizer),
Refu = 2.15*108, 0.00657,
Then Cffus = 1.47456*10-3,

5. 4
Fuselage wetted Area:
Sfus = (πD2/4), then Sfus = 43.93 ft2
Then Df = 7.38
Assuming (L1 = 12.4 ft, L2 = 68.30 ft, L3 = 10.933 ft):
Then fuselage wetted area = 171.75 + 1810.9 + 155.367 = 2138.017 ft2
dp = 0.95 ft, (CDf)B = 0.00606 ,(CD)b = 0.00005448211
then (CD0)B = 0.00660821
4). the span (Oswald) efficiency factor:
Assuming leading edge radius = 1.505%
then MAC = 11.403
LLER = 11.403*1.505%, then LLER = 0.1716 ft, Re = 3.78*105
ɅLE = 25.42o
P1 = 2.16

6. 5
P2 = 2.4*105, then P2 > 1.3*105
(Assume CLα = 2π)
Then e = 0.993407387