def example1():
# run trim analysis
import aircraft
ac = aircraft.load('V0510')
fc = flightConditions()
CG = numpy.array([1.7,0,0])
I = numpy.array([500., 1000, 1000])
M = 600.
V = 50.
density = 1.2255
Cd0 = ac.get_drag(V,0.0)
fc.addTrimmedFlightCondition('cruise1',M,CG,I,V,density,CmTrim=0.0,Cd0=Cd0)
a3d = aero3d_VLM(ac)
rslt = a3d.runVLM(fc,False)
print 'GENERAL'
rslt.results[0].display()
print 'PHUGOID'
rslt.results[0].dynamicStability.phugoid.display()
print 'SHORT PERIOD'
rslt.results[0].dynamicStability.shortPeriod.display()
print 'DUTCH ROLL'
rslt.results[0].dynamicStability.dutchRoll.display()
print 'SPIRAL'
rslt.results[0].dynamicStability.spiral.display()
print 'ROLL'
rslt.results[0].dynamicStability.roll.display()
def calc_hinges():
import aircraft
ac = aircraft.load('V0510')
fc = flightConditions()
CG = [1.824,0.0,-0.125]
I = ac.get_inertia()
m = 620.0
V = 75.0
rho = 1.2255
CD0 = ac.get_drag(V,0.0)
n = 1.0
fc.add_flight_condition('hinge',m,CG,I,V,rho,n,CD0,alpha=20.0,elevator=30.0)
aero3d = aero3d_VLM(ac)
rslt = aero3d.runVLM(fc)
rslt.results[0].hingeMoments.display()
def run_test():
import aircraft
ac = aircraft.load('V0510')
fc = flightConditions()
CG = ac.get_CG(False)
I = ac.get_inertia(False)
M = ac.get_mass_total(False)
V = 1.0
rho = 1.2255
Cd0 = ac.get_drag(V,0.0)
fc.addTrimmedFlightCondition('trim',M, CG, I, V, rho)
fc.flightConditionList[0].setRudder(0.)
aero3D = aero3d_VLM(ac)
rslt = aero3D.runVLM(fc)
rslt.results[0].display()
def add_flight_condition(self,name,mass,CG,inertia,velocity,density,
loadFactor=1.0,CD0=0.0,alpha=0.0,beta=0.0,flap=0.0
,elevator=0.0,aileron=0.0,rudder=0.0):
"""
Non trimmed flight conditions
"""
newFC =self._flightCondition()
newFC.pitchTrim =False
newFC.CG =CG
newFC.name =name
newFC.mass =float(mass)
newFC.inertia =inertia
newFC.velocity =float(velocity)
newFC.density =float(density)
newFC.densityAltitude =fc.get_density_altitude(newFC.density)
atm =fc.ISAtmosphere(newFC.densityAltitude,0.0)
newFC.soundSpeed =atm.soundSpeed
newFC.pressure =atm.pressure
newFC.temperature =atm.temperature
newFC.Mach =newFC.velocity/newFC.soundSpeed
newFC.loadFactor =loadFactor
newFC.Cd0 =CD0
newFC.alpha =alpha
newFC.beta =beta
newFC.flap =flap
newFC.elevator =elevator
newFC.aileron =aileron
newFC.rudder =rudder
self.flightConditionList.append(newFC)
def run_climbRate(self,airspeed,powerSetting):
"""
Calculates the climb rate at a given airspeed and power setting.
Supports descent calculations which will appear as negative climb
rates. Does not utilize small angle approximations.
"""
basicInput = self.bi
thrustModule = self.tm
V =airspeed
rho =basicInput.density
Cd0 =basicInput.Cd0
g =basicInput.g
k =basicInput.k
S =basicInput.refArea
W =basicInput.mass*g
Q =0.5*rho*V*V
tol =1.0e-4
change =tol+1.0
CA =0.0 #climb angle, rad
iMax =100
i =0
if powerSetting>=5.0:
tm=thrustModule.runAnalysis
else:
tm=thrustModule._zeroThrustModule
while change>tol and i<=iMax:
tm(V,rho,powerSetting)
T =thrustModule.thrust
L =W*np.cos(CA)-T*np.sin(CA)
Cl =L/Q/S
Cd =Cd0+k*Cl*Cl
D =Q*S*Cd
PR =D*V
PA =T*V
RC =(PA-PR)/W
ca =CA
CA =np.arcsin(RC/V)
change=np.abs(ca-CA)
i+=1
self.__init__(self.bi, self.tm)
self.velocity =V
self.equivalentAirspeed=self.EAS(V,basicInput.density)
self.densityAltitude =fc.get_density_altitude(basicInput.density)
self.density =basicInput.density
self.climbRate =RC
self.climbAngle =CA
self.fuelFlow =thrustModule.fuelFlow
self.RPM =thrustModule.RPM
self.powerSetting =powerSetting
self.thrust =T
self.power =PA
self.propEfficiency =thrustModule.propEfficiency
self.beta =thrustModule.beta
self.SAR_km_per_kg =thrustModule.SAR_km_per_kg
self.lift =L
self.drag =D
self.L_D =L/D
self.turnRadius =0.0
self.bankAngle =0.0
self.loadFactor =1.0
def addTrimmedFlightCondition(self,name,mass,CG,inertia,velocity,density,
loadFactor=1.0,Cd0=0.0,CmTrim=0.0,flapDefl=0.0,
aileron=0.0,rudder=0.0):
"""
Sets and stores a list of flight conditions to be analyzed by the
aerodynamics solvers.
Parameters
----------
name : string
name of the flight condition case
mass : float, kg
total mass of the aircraft
CG : float array, m
array of CG coordinates [x,y,z]
inertia : float array
array of moments of inertia [Ixx, Iyy, Izz]
velocity : float, m/sec
aircraft TAS
density : float, kg/m**3
atmospheric density
loadFactor : float
load factor for steady pull-out flight conditions
Cd0 : float
zero lift drag coefficient of aircraft
CmTrim : float
trim pitch moment coefficient. Set CmTrim=None if non-trimmed flight
condition is required
flapDefl : float, deg
plain flap deflection angle
"""
newFC =self._flightCondition()
newFC.pitchTrim =True
newFC.CG =CG
newFC.name =name
newFC.mass =float(mass)
newFC.inertia =inertia
newFC.velocity =float(velocity)
newFC.density =float(density)
newFC.densityAltitude =fc.get_density_altitude(newFC.density)
atm =fc.ISAtmosphere(newFC.densityAltitude,0.0)
newFC.soundSpeed =atm.soundSpeed
newFC.pressure =atm.pressure
newFC.temperature =atm.temperature
newFC.Mach =newFC.velocity/newFC.soundSpeed
newFC.loadFactor =loadFactor
newFC.Cd0 =Cd0
newFC.CmTrim =CmTrim
newFC.flap =flapDefl
newFC.aileron =aileron
newFC.rudder =rudder
self.flightConditionList.append(newFC)
def run_test3():
import aircraft
import matplotlib.pyplot as plt
ac = aircraft.load('V0510')
fc = flightConditions()
CG = ac.get_CG()
I = ac.get_inertia()
m = ac.get_mass_total()
V = 50.0
rho = 1.2255
CD0 = ac.get_drag(V,0.0)
ail = [-10,0,2,5,7,10]
ailDisp = list()
r = 15
for a in ail:
fc.add_flight_condition('untrimmed',m,CG,I,V,rho,CD0=CD0,aileron=a,rudder=r)
ailDisp.append(a)
aero3d = aero3d_VLM(ac)
rslt = aero3d.runVLM(fc)
cl = [rslt.results[i].coefficients.Cl for i in range(len(ail))]
plt.plot(ailDisp,cl,'ro-')
plt.show()
def update_notrim(self,velocity,density,loadFactor=1.0,flap=0.0,aileron=0.0,
elevator=0.0,rudder=0.0,alpha=0.0,beta=0.0,mass=None,
CG=None,inertia=None):
"""
Performs aerodynamic analysis at specified flight conditions.
Parameters
----------
velocity : float, m/s
density : float, kg/m**3
loadFactor : float
flap : float,deg
flap deflection angle. Note that plain flap is analyzed, slotted flap
analysis is not available
aileron : +float, deg
aileron deflection angle. Deflection is mirrored.
elevator : float, deg
elevator deflection angle
rudder : float, deg
rudder deflection angle
alpha : float, deg
angle of attack. Note that solver may fail on high angles of attack.
Use values [-5;5].
beta : float, deg
sideslip angle.
mass : float, kg
if mass is not defined, than it will be obtained from current
aircraft configuration
CG : float vector, m
aircraft center of gravity in format [x,y,z].
If CG is not defined it will be obtained from current aircraft configuration
inertia : float vector, kg*m**2
aircraft moment of inertia in format [Ixx,Iyy,Izz].
"""
if mass==None:
mass = self.aircraft.get_mass_total()
if CG==None:
CG = self.aircraft.get_CG()
if inertia==None:
inertia = self.aircraft.get_inertia()
altitude = fc.get_density_altitude(density)
CD0 = self.aircraft.get_drag(velocity,altitude)
flightCond = flightConditions()
flightCond.add_flight_condition('untrimmed',mass,CG,inertia,velocity,density,
loadFactor,CD0,alpha,beta,flap,elevator,
aileron,rudder)
vlm = aero3d_VLM(self.aircraft)
out = vlm.runVLM(flightCond)
return out.results[0]
def getCd0(self,V,rho):
"""
Executes AeroCD0 skin friction / form factor drag analysis code
Returns parasite drag (Cd0)
Parameters
----------
V : float, m/sec
velocity
rho : float, kg/m**3
density
"""
alt = fc.get_density_altitude(rho)
Cd0= self.aircraft.get_drag(V,alt)
return Cd0
def _setAttributes(self,velocity):
drag=self.requiredThrust(velocity)
self.__init__(self.bi, self.tm)
self.tm.matchRequiredThrust(velocity,self.bi.density,drag)
self.velocity =velocity
self.equivalentAirspeed=self.get_EAS(velocity,self.bi.density)
self.densityAltitude =fc.get_density_altitude(self.bi.density)
self.density =self.bi.density
self.fuelFlow =self.tm.fuelFlow
self.RPM =self.tm.RPM
self.powerSetting =self.tm.powerSetting
self.thrust =self.tm.thrust
self.power =self.tm.power
self.propEfficiency =self.tm.propEfficiency
self.beta =self.tm.beta
self.SAR_km_per_kg =self.tm.SAR_km_per_kg
self.drag =drag
self.lift =self.bi.wt
self.L_D =self.lift/self.drag
def update(self,velocity,density,loadFactor=1.0,CmTrim=0.0,flapDefl=0.0,
mass=None,CG=None,inertia=None,name='default'):
"""
Updates aerodynamics results in self.results
Parameters
----------
velocity : float, m/sec
aircraft TAS
density : float, kg/m**3
air density
loadFactor : float
load factor
CmTrim : float
trim pitch moment coefficient
flapDefl : float, deg
plain flap deflection
"""
self.flightConditions =flightConditions()
densityAltitude =fc.get_density_altitude(density)
if mass==None:
mass =self.aircraft.get_mass_total()
if CG==None:
CG =self.aircraft.get_CG()
if inertia==None:
inertia =self.aircraft.get_inertia()
Cd0 =self.aircraft.get_drag(velocity,densityAltitude)
self.flightConditions.addTrimmedFlightCondition(name,mass,CG,inertia,velocity,density,loadFactor,Cd0,CmTrim,flapDefl)
vlm =aero3d_VLM(self.aircraft)
out =vlm.runVLM(self.flightConditions)
self.results =out.results[0]
self.results.name = name
NP = self.results.xNP
self.results.SM = (NP-CG[0])/self.aircraft.wing.MAC
self.results.mass = mass
self.results.CG = CG
self.results.inertia = inertia
self.results.velocity = velocity
self.results.density = density
def _setAttributes(self,n,airspeed,powerSetting):
V =airspeed
rho =self.bi.density
Cd0 =self.bi.Cd0
g =self.bi.g
k =self.bi.k
S =self.bi.refArea
W =self.bi.mass*g
Q =0.5*rho*V**2
self.tm.runAnalysis(V,rho,powerSetting)
phi =np.arccos(1./n)
turnRadius =V**2/(g*(n**2-1.)**.5)
turnRate =g*(n**2-1)**.5/V
L =n*W
Cl =L/Q/S
Cd =Cd0+k*Cl**2
D =Q*S*Cd
self.tm.matchRequiredThrust(V,self.bi.density,D)
self.__init__(self.bi,self.tm)
self.velocity =V
self.equivalentAirspeed=self.EAS(V,self.bi.density)
self.densityAltitude =fc.get_density_altitude(self.bi.density)
self.density =self.bi.density
self.fuelFlow =self.tm.fuelFlow
self.RPM =self.tm.RPM
self.powerSetting =powerSetting
self.thrust =self.tm.thrust
self.power =self.tm.power
self.propEfficiency =self.tm.propEfficiency
self.beta =self.tm.beta
self.SAR_km_per_kg =self.tm.SAR_km_per_kg
self.lift =L
self.drag =D
self.L_D =L/D
self.turnRadius =turnRadius
self.bankAngle =phi
self.loadFactor =n
self.turnRate =turnRate
def analyze_prop(self,betaSet,N,V,rho):
"""
Performs propeller analysis using BEMT. In case of poor convergence
(required J cannot be achieved at given rpm and airspeed)
function returns advance ratio and CP=None, CT=None, effy=None.
Method returns thrust, power required, propeller advance ratio *J*,
thrust coefficient *CT*, power coefficient *CP* and efficiency *effy*
at given pitch angle, rpm, velocity and density
Parameters
----------
betaSet : float, degree
pitch angle setting at 75% of diameter
N : float, rpm
propeller rotation speed
V : float, m/sec
true airspeed of aircraft (propeller)
rho : float, kg/m^3
air density
Returns
-------
J : float
advance ratio
CT : float
thrust coefficient
CP : float
power coefficient
effy : float
propeller efficiency
"""
if N*V*rho==0.0:
return zeros(6)
if betaSet!=None:
self.set_beta(betaSet)
Jcorr = 0.97
n = N/60.0
d = self.diameter
B = self.numBlades
dT = list()
dQ = list()
altitude = FlightConditions.get_density_altitude(rho)
soundSpeed = FlightConditions.ISAtmosphere(altitude).soundSpeed
Jreq = Jcorr * V / (n*d)
def master_func(alpha,beta,x,r,B,soundSpeed,d,af,c,n):
phi = radians(beta - alpha)
# compute Prandtl loss factor
sinPhiL1 = (1.0 + 1.0/(x*tan(phi))**2)**0.5
f = B * (1.0/x-1.0) * sinPhiL1/2.0
F = 2.0 * arccos(exp(-f)) / pi
vLocal = pi*x*n*d / cos(phi)
mach = vLocal / soundSpeed
cl = float(af.polar.clAlpha(mach,alpha))
cd = float(af.polar.cdAlpha(mach,alpha))
sinPhi = sin(phi)
cosPhi = cos(phi)
lambdaT = cl*cosPhi - cd*sinPhi
lambdaP = cl*sinPhi + cd*cosPhi
sigma = c*B / (2.0*pi*r)
J = pi*x*(4.0*F*sinPhi**2 - lambdaT*sigma) / (4.0*F*sinPhi*cosPhi + lambdaP*sigma)
return J,vLocal, lambdaT,lambdaP
def get_J(alpha,Jreq,beta,x,r,B,soundSpeed,d,af,c,n):
rslt = master_func(alpha,beta,x,r,B,soundSpeed,d,af,c,n)
return rslt[0] - Jreq
for i in range(self.numStations):
x = self.x[i]
c = self.chord[i]
beta = self.beta[i]
r = self.r[i]
alpha0List = [-20,-10,-5,0,5,10,20,30,40,50]
af = self.airfoil[i]
for alpha0 in alpha0List:
try:
alpha = newton(get_J,alpha0,args=(Jreq,beta,x,r,B,soundSpeed,d,af,c,n),tol=0.01)
break
except RuntimeError:
return 0.0,0.0,Jreq, 0.0, 0.0, 0.0
J,vLocal,lambdaT,lambdaP = master_func(alpha,beta,x,r,B,soundSpeed,d,af,c,n)
dT.append( rho*vLocal**2*lambdaT*c*B / 2.0 )
dQ.append( rho*vLocal**2*lambdaP*c*B*r / 2.0 )
dT.insert(0,0.0)
dQ.insert(0,0.0)
dT.append(0.0)
dQ.append(0.0)
r_ = copy(self.r)
radius = self.diameter/2.0
r_ = hstack([0.0,r_,radius])
T = simps(dT,r_)
Q = simps(dQ,r_)
CT = T / (rho*n**2*d**4)
CQ = Q / (rho*n**2*d**5)
effy = CT/CQ*Jreq/(2.0*pi)
effy = effy
CP = Jreq*CT/effy
P = CP*rho*n**3*d**5
if CT<=0.0 or CT>1.0:
CT=0.0
effy = 0.0
if CQ<=0.0 or CQ>1.0:
CQ=0.0
effy = 0.0
if CP<=0.0:
CP=0.0
if not effy==None:
effy = effy*0.95
return T,P,Jreq/Jcorr,CT,CP,effy