def initial(self,V0,y0,ydot0,R0,q0,p0,r0,a0,b0,m0,rho0):
self.V0=V0
self.y0=y0
self.ydot0=ydot0
self.R0=R0
self.p0=p0
self.q0=q0
self.r0=r0
self.a0=a0
self.b0=b0
self.m0=m0
self.a=util.scale_a()
self.M0=self.V0/self.a
self.rho0=rho0
self.q=0.5*self.rho0*self.V0**2 # dynamic pressure
def update(self):
self.a=util.scale_a()
self.M0=self.V0/self.a
if self.diag:
print "V: ",self.V0
if self.diag:
print "Mach: ", self.M0
self.shell.update(self.M0)
CT,CN,CM=shell.analyse(self.a0,mode="rad")
self.CL=CN*np.cos(self.a0)-CT*np.sin(self.a0)
self.CD=CT*np.cos(self.a0)+CN*np.sin(self.a0)
if self.para and self.T_sim>=self.para_t:
self.CD+=self.para_CD
if self.diag:
if self.para and self.T_sim>=self.para_t:
print "Para Deployed: True"
print "CL, CD: ",self.CL,self.CD
self.q=0.5*util.scale_height(self.R0-Re)[2]*self.V0**2 # dynamic pressure
if self.diag:
print "Dyn P: ",self.q
if self.para and self.T_sim>=self.para_t:
self.D0=self.q * self.S * (self.CD-self.para_CD) + self.q * self.para_S * self.para_CD
else:
self.D0=self.q * self.S * self.CD # aerodynamic force
self.L0=self.q * self.S * self.CL # aerodynamic force
if self.diag:
print "Lift, Drag : ",self.L0,self.D0
self.g0=-self.grav(self.R0-Re,0,0)
self.ydot0 = 0# -(self.V0/self.R0 - self.g0/self.V0)*np.cos(self.y0) + self.L0/(self.m*self.V0)
#print self.ydot0
self.q_data.append(self.q)
self.rho_data.append(util.scale_height(self.R0-Re)[2])
self.ydot_data.append(self.ydot0)
def norm_moi(Ixx,Iyy,Izz,Ixz,c,b,mass):
KX2 = Ixx/(b**2*mass)
KY2 = Iyy/(c**2*mass)
KZ2 = Izz/(b**2*mass)
JXZ = Ixz/(b**2*mass)
return KX2,KY2,KZ2,JXZ
if __name__=="__main__":
stabMargin=0.1
configuration="t"
ratio=0
velocity = 100
mach = velocity/util.scale_a()
canard,main,tail,vert=dummyWings()
xac = main.root/4.
xnp = xac + stabMargin*main.root
r_fus=0.6
t_fus=4./1000
width_fus=r_fus*2
l_fus = 5.
mtv = 0.1 # perpendicular distance between zero lift line and tail
rati=np.arange(0,3,0.1)
mass_r=[]
