def Calc_skin_inertia_Ixx(Spar1, Spar2):
n = 100 # number of sections
dx = ((Spar2 - Spar1) / n)
x = Spar1
Ixx = 0
for i in range(n):
x = x + dx
dxlength = dx * Wing.Chordlength
y = ((Wing.airfoilordinate(x - dx) + Wing.airfoilordinate(x)) /
2) * Wing.Chordlength
dy = abs(Wing.airfoilordinate(x - dx) -
Wing.airfoilordinate(x)) * Wing.Chordlength
length = np.sqrt(dxlength**2 + dy**2)
dIxx = length * Wing.ThSkin * (y**2)
Ixx = Ixx + dIxx
Ixx = Ixx * 2
return Ixx
def Calc_skin_inertia_Iyy(Spar1, Spar2):
n = 100 # number of sections
dx = ((Spar2 - Spar1) / n)
x = Spar1
Iyy = 0
for i in range(n):
x = x + dx
xlength = x * Wing.Chordlength
dxlength = dx * Wing.Chordlength
y = ((Wing.airfoilordinate(x - dx) + Wing.airfoilordinate(x)) /
2) * Wing.Chordlength
dy = abs(Wing.airfoilordinate(x - dx) -
Wing.airfoilordinate(x)) * Wing.Chordlength
length = np.sqrt(dxlength**2 + dy**2)
dIyy = length * Wing.ThSkin * (
(abs(xlength - (Wing.centroid * Wing.Chordlength)))**2)
Iyy = Iyy + dIyy
Iyy = Iyy * 2
return Iyy
I_XX_TOT += loc_I_xx + stringer_area * (y_coords[i])**2
I_YY_TOT += loc_I_yy + stringer_area * (
x_coords[i] - Wing.centroid * Wing.Chordlength)**2
I_XY_TOT += loc_I_xy + stringer_area * (
x_coords[i] - Wing.centroid * Wing.Chordlength) * (y_coords[i] -
Y_CEN)
return I_XX_TOT, I_YY_TOT, I_XY_TOT
# returns stiffener x,y locations and rotation
# return z_y_angle_coords # [(stringer0 z,y,rot),(stringer1 x,y,rot)] m,m,rad
#Ixx and Iyy of the clamps
ClampsIxx = ((Wing.AreaClamps / 2) *
((Wing.airfoilordinate(Wing.ChSpar1) * Wing.Chordlength)**2)) * 2
ClampsIyy = ((Wing.AreaClamps / 2) *
((abs(Wing.ChSpar1 - Wing.centroid) * Wing.Chordlength)**2)) * 2
I_XX_TOT_str, I_YY_TOT_str, I_XY_TOT_str = calc_total_stringer_inertia(
Wing.get_coord_from_perim(Wing.N_stringers / 2, Wing.ChSpar1, Wing.ChSpar2,
Wing.Chordlength)[0],
calc_stringer_inertia(Wing.h_str, Wing.w_str, Wing.t_str))
I_XX_Spar1, I_YY_Spar1 = Calc_spar_inertia(Wing.HSpar1, Wing.ThSpar1,
Wing.ChSpar1, Wing.z)
I_XX_Spar2, I_YY_Spar2 = Calc_spar_inertia(Wing.HSpar2, Wing.ThSpar2,
Wing.ChSpar2, Wing.z)
I_XX_Skin = Calc_skin_inertia_Ixx(Wing.ChSpar1, Wing.ChSpar2)
I_YY_Skin = Calc_skin_inertia_Iyy(Wing.ChSpar1, Wing.ChSpar2)
# plt.plot(zlist, L_momentlist)
# plt.show()
def Normal_stress_due_to_bending(x, y): # Normal stress due to bending
denominator_inertia_term = Inertia.Ixx_wb * Inertia.Iyy_wb - Inertia.Ixy_wb**2
inertia_term_1 = (Inertia.Iyy_wb * y * Wing.Chordlength - Inertia.Ixy_wb *
x * Wing.Chordlength) / denominator_inertia_term
inertia_term_2 = (Inertia.Ixx_wb * x * Wing.Chordlength - Inertia.Ixy_wb *
y * Wing.Chordlength) / denominator_inertia_term
sigma_zs = D_moment * inertia_term_1 + L_moment * inertia_term_2
strain = sigma_zs / youngs_modulus
return sigma_zs, strain #Gives the normal stress function for a given span zs, and x- and y- coordinate
NS = Normal_stress_due_to_bending(0.18, Wing.airfoilordinate(0.18))[0]
# print('sigma_zs', Normal_stress_due_to_bending(0.15, Wing.airfoilordinate(Wing.c)))
# def Shear_wb(zs, dL, dD, dM):
# #section 01
# qtorque = dM/(2*Wing.Area_cell())
# n = 100
# ds = Wing.HSpar1/n
# qs1 = np.array([])
# s1 = np.array([])
# qs1 = np.append(qs1, 0)
# s1 = np.append(s1, 0)
# y = Q_("0 m")
# s = Q_("0 m")
# for i in range(n):