Cm*0.5*rho*V_local**2*S_w*MAC +\
Cmh*0.5*rho*V_local**2*S_h*MAC_h
u_dot = Fx / mtow - q * w
w_dot = Fz / mtow + q * u
q_dot = My / I_yy
u += u_dot * dt
w += w_dot * dt
q += q_dot * dt
theta += q * dt
alpha_nose = np.arctan(w / u)
V_local = np.sqrt(u**2 + w**2)
alst.append(alpha_nose.to('degree').magnitude)
qlst.append(q.to('degree /s').magnitude)
qdotlst.append(q_dot.to('degree/s**2').magnitude)
thetalst.append(theta.to('degree').magnitude)
Vlst.append(V_local.magnitude)
print('finished')
plt.figure()
plt.plot(tlst, qlst, label='pitch rate', color='black')
plt.plot(tlst, thetalst, label='theta [deg]', color='black', ls='--')
plt.plot(tlst, alst, label='alpha [deg]', color='black', ls='-.')
plt.xlabel("time [s]")
plt.ylabel("pitch rate [deg/s]")
#plt.plot(tlst,qdotlst,label='q dot')
plt.legend()
import sys
sys.path.append(
'../'
) # This makes sure the parent directory gets added to the system path
from Misc import ureg, Q_
import math as m
import numpy as np
import scipy as sp
from Geometry import Geometry
from pint import UnitRegistry
import Aeroprops
#Input variables, to be filled in by user
AoA = Q_("5 deg") #fill in angle of attack flight condition drag to be known
AoA = AoA.to(ureg.rad)
Rwb = 1.05 #Read in DATCOM p1164, wing body interference factor
Rhtb = 1.05
Rvtb = 1.05
CL_trim = 0 #lift coefficient needed to trim aircraft
height = Q_('500 m') # height in m
velocity = Q_('94 m/s') #velocity in m/s
loc_max_tc_wing = 0 # 0 if t/c max is < 30%c, 1 if t/c max is >= 30%c
loc_max_tc_ht = 0 # 0 if t/c max is < 30%c, 1 if t/c max is >= 30%c
loc_max_tc_vt = 0 # 0 if t/c max is < 30%c, 1 if t/c max is >= 30%c
Sref = Geometry.Wing.S
CL_wing = Aeroprops.CL_alpha_wing * AoA #max lift coefficient
AR_wing = Geometry.Wing.A
ih = Geometry.H_tail.i_h #incidende angle HT
CL_ht = Aeroprops.CL_alpha_ht * AoA.to(ureg.rad) - Aeroprops.de_da * \
disc_wing_h[i, 7] = y_i.magnitude
# Moments and Forces
sum_Fn_w = sum(disc_wing_w[:, 1]) * Q_("N")
sum_Fn_h = sum(disc_wing_h[:, 1]) * Q_("N")
My = dx_w * sum_Fn_w + dx_h * Fn_h
Mx = (sum(disc_wing_w[:, 3]) + sum(disc_wing_h[:, 3])) * Q_("N*m")
Mz = sum(disc_wing_w[:, 4]) * Q_("N*m")
p_dot = Mx / I_xx
q_dot = My / I_yy
p += p_dot * dt
q += q_dot * dt
plst.append(p.to('degree/s').magnitude)
pdotlst.append(p_dot.to('degree/s**2').magnitude)
Mzlst.append(Mz.magnitude)
plt.plot(tlst, plst, label='roll rate', c='black')
#plt.plot(tlst,Mzlst,label='moment due to roll')
#plt.plot(tlst,pdotlst,label='p dot')
plt.xlabel("Time [s]")
plt.ylabel("Roll rate [" r'$^\circ$' "/s]")
plt.legend()
plt.show()
"============================================================================="
#### Pitch
####alpha_nose, de, Thrust = trim()
##
#V_local = V_inf