def Calc_spar_inertia(
HSpar, TSpar, ChSpar, zs
): # Input height spar (m), thickness spar (m), location centroid w.r.t. chord and chordwise location (-) spar respectively (-)
Ixx = (1 / 12) * TSpar * (HSpar**3) # Calculation of Ixx
Iyy = (1 / 12) * HSpar * (TSpar**3) # Caclulation of Iyy w/o steiner term
Iyysteiner = TSpar * HSpar * (abs((Wing.centroid * Wing.length_chord(zs)) -
(ChSpar * Wing.length_chord(zs)))**2
) # Calculation of steiner term Iyy
Iyy = Iyy + Iyysteiner # Adding both Iyy moments of inertia together
return Ixx, Iyy
def computeloads(z):
Sectioncenters = np.array([])
dLlist = np.array([])
dDlist = np.array([])
L = 0
D = 0
M = 0
L_moment = 0
D_moment = 0
Llist = np.array([])
Dlist = np.array([])
Lmomentlist = np.array([])
Dmomentlist = np.array([])
L_momentlist = np.array([])
zs = b - (sectionlength / 2)
dL = Q_("0 N")
dD = Q_("0 N")
dM = Q_("0 N*m")
while zs > z: #zs is measured is m from
Areaofsection = sectionlength * Wing.length_chord(zs)
Sectioncenters = np.append(Sectioncenters, zs)
'''Lift drag and moment for the section'''
dL = (cl * 0.5 * rho * (V**2) *
Areaofsection).magnitude #lift of the section
dD = (cd * 0.5 * rho * (V**2) *
Areaofsection).magnitude #drag of the section
dM = (cm * 0.5 * rho * (V**2) * Areaofsection *
Wing.length_chord(zs)).magnitude #moment of the section
dL *= Q_('kg * m / s**2')
dD *= Q_('kg * m / s**2')
dM *= Q_('kg * m**2 / s**2')
if zs < Geometry.Fuselage.b_f * 0:
dL = 0
dD = 0
dM = 0
dLlist = np.append(dLlist, dL)
dDlist = np.append(dDlist, dD)
'''Total lift, drag and moment for the wing'''
L = L + dL # Total lift for one wing
D = D + dD # Total drag for one wing
M = M + dM # Total moment for one wing
Llist = np.append(Llist,
L) # put the values in a list so we can plot them
Dlist = np.append(Dlist, D)
zs = zs - sectionlength # Select other section for the next loop
for i in range(0, len(Sectioncenters)):
arm = (Sectioncenters[i] - z.magnitude)
dLmoment = (arm * dLlist[i])
dDmoment = (arm * dDlist[i])
L_moment = L_moment + dLmoment
D_moment = D_moment + dDmoment
Lmomentlist = np.append(Lmomentlist, L_moment)
Dmomentlist = np.append(Dmomentlist, D_moment)
'''For the 20G manoeuver'''
MTOW = Geometry.Masses.W_MTOW
Max_20G_N = MTOW * 9.81 * 20
Tot_L = 2 * totallift
if Tot_L.magnitude > 0.:
fac_20G = Max_20G_N / Tot_L
fac_20G = fac_20G.magnitude
else:
fac_20G = 0
L_moment = L_moment * fac_20G
D_moment = D_moment * fac_20G
L = L * fac_20G
D = D * fac_20G
M = M * fac_20G
return L, D, M, L_moment, D_moment, dL, dD, dM
Lmomentlist = np.array([])
Dmomentlist = np.array([])
dLlist = np.array([])
dDlist = np.array([])
Llist = np.array([])
Dlist = np.array([])
Sectioncenters = np.array([])
sectionlength = (b.magnitude -
0 * Geometry.Fuselage.b_f.magnitude) / n * ureg.meter
zs = b - (sectionlength / 2)
while zs.magnitude > 0: # zs is measured is m from
Areaofsection = sectionlength * Wing.length_chord(zs)
Sectioncenters = np.append(Sectioncenters, zs)
'''Lift drag and moment for the section'''
dL = cl * 0.5 * rho * (V**2) * Areaofsection.magnitude * Q_(
"m**2") # lift of the section
if zs < Geometry.Fuselage.b_f * 0:
dL = Q_('0 kg * m / s**2')
'''Total lift, drag and moment for the wing'''
L = L + dL # Total lift for one wing
zs = zs - sectionlength # Select other section for the next loop
totallift = L