# All the quantities are expressed in SI units
from math import pi,sqrt
AR = 8.; # Aspect ratio of the wing
alpha = 5.*pi/180.; # Angle of attack experienced by the wing
a0 = 2.*pi # airfoil lift curve slope
alpha_L0 = 0; # zero lift angle of attack is zero since airfoil is symmetric
# from fig. 5.20, for AR = 8 and taper ratio of 0.8
delta = 0.055;
tow = delta; # given assumption
# thus the lift curve slope for wing is given by
a = a0/(1.+(a0/pi/AR/(1.+tow)));
# thus C_l can be calculated as
C_l = a*alpha;
# from eq.(5.61)
C_Di = C_l**2./pi/AR*(1.+delta);
print"Cl =",round(C_l,2)
print"CDi =",round(C_Di,2)
# All the quantities are expressed in SI units
from math import sqrt,pi
CDi1 = 0.01; # induced drag coefficient for first wing
delta = 0.055; # induced drag factor for both wings
tow = delta;
alpha_L0 = -2.*pi/180.; # zero lift angle of attack
alpha = 3.4*pi/180.; # angle of attack
AR1 = 6.; # Aspect ratio of the first wing
AR2 = 10.; # Aspect ratio of the second wing
# from eq.(5.61), lift coefficient can be calculated as
C_l1 = sqrt(pi*AR1*CDi1/(1.+delta));
# the lift slope for the first wing can be calculated as
a1 = C_l1/(alpha-alpha_L0);
# the airfoil lift coefficient can be given as
a0 = a1/(1.-(a1/pi/AR1*(1.+tow)));
# thus the list coefficient for the second wing which has the same airfoil is given by
a2 = a0/(1.+(a0/pi/AR2*(1.+tow)));
C_l2 = a2*(alpha-alpha_L0);
CDi2 = C_l2**2./pi/AR2*(1.+delta);
print"The induced drag coefficient of the second wing is CD,i =",CDi2
# all the quantities are expressed in SI units
from math import pi
a0 = 0.1*180./pi; # airfoil lift curve slope
AR = 7.96; # Wing aspect ratio
alpha_L0 = -2.*pi/180.; # zero lift angle of attack
tow = 0.04; # lift efficiency factor
C_l = 0.21; # lift coefficient of the wing
# the lift curve slope of the wing is given by
a = a0/(1+(a0/pi/AR/(1.+tow)));
# thus angle of attack can be calculated as
alpha = C_l/a + alpha_L0;
print"alpha =",alpha*180./pi,"degrees\n"
# All the qunatities are expressed in SI units
from math import pi,sqrt
alpha_L0 = -1.*pi/180.; # zero lift angle of attack
alpha1 = 7.*pi/180.; # reference angle of attack
C_l1 = 0.9; # wing lift coefficient at alpha1
alpha2 = 4.*pi/180.;
AR = 7.61; # aspect ratio of the wing
taper = 0.45; # taper ratio of the wing
delta = 0.01; # delta as calculated from fig. 5.20
tow = delta;
# the lift curve slope of the wing/airfoil can be calculated as
a0 = C_l1/(alpha1-alpha_L0);
e = 1./(1.+delta);
# from eq. (5.70)
a = a0/(1.+(a0/pi/AR/(1.+tow)));
# lift coefficient at alpha2 is given as
C_l2 = a*(alpha2 - alpha_L0);
# from eq.(5.42), the induced angle of attack can be calculated as
alpha_i = C_l2/pi/AR;
# which gives the effective angle of attack as
alpha_eff = alpha2 - alpha_i;
# Thus the airfoil lift coefficient is given as
c_l = a0*(alpha_eff-alpha_L0);
c_d = 0.0065; # section drag coefficient for calculated c_l as seen from fig. 5.2b
# Thus the wing drag coefficient can be calculated as
C_D = c_d + ((C_l2**2.)/pi/e/AR);
print"The drag coefficient of the wing is C_D =",C_D