def runTest_thrustAnalysis():
import aircraft
print '\n Testing Single Engine Variable Pitch'
ac=aircraft.load("sampleInput1")
th=thrustAnalysis(ac)
th.runAnalysis(60.,1.225,100.)
th.printResults()
print '\n Testing Thrust Matching - Single Engine Variable Pitch'
th.matchRequiredThrust(40.,1.225,600)
th.printResults()
print '\n Testing - Twin Engine Variable Pitch'
ac=aircraft.load("sampleInput2")
th=thrustAnalysis(ac)
th.runAnalysis(60.,1.225,100.)
th.printResults()
print '\n Testing Thrust Matching - Twin Engine Variable Pitch'
th.matchRequiredThrust(40.,1.225,1200)
th.printResults()
print '\n Testing Twin Engine Fixed Pitch'
ac=aircraft.load("fixedPitchSample")
th=thrustAnalysis(ac)
th.runAnalysis(60.,1.225,100.)
th.printResults()
print '\n Testing Thrust Matching - Twin Engine Fixed Pitch'
th.matchRequiredThrust(40.,1.225,1200)
th.printResults()
def run_clean_configuration():
ac = aircraft.load('V0510')
CG = ac.get_CG()
CG = [1.711,0.0,-0.104]
I = ac.get_inertia()
m = ac.get_mass_total()
V = 34.0
rho = 1.2255
CD0 = ac.get_drag(V,0.0)
aero3d = aerodynamics.aero3d_VLM(ac)
case = [9,10,11]
alpha = array([-2,0.0,2,4,6,8,10,12])
beta = array([0.0,10.0,20.0])
wtRslt = list()
avlRslt = list()
for i in range(len(case)):
fc = aerodynamics.flightConditions()
b = beta[i]
for a in alpha:
fc.add_flight_condition('wt validation',m,CG,I,V,rho,CD0=CD0,alpha=a,beta=b)
wtPath = find_file_by_number(case[i])
wtRslt.append(WTresult(wtPath))
_avlrslt = aero3d.runVLM(fc).results
_avlRslt = WTresult()
_avlRslt._read_vlm_results(_avlrslt)
avlRslt.append(_avlRslt)
compare_results(wtRslt[i],avlRslt[i])
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 runTest2():
import aircraft
AC =aircraft.load("Xtail_sample")
FC =flightConditions()
CG =numpy.array([0.65,0,0])
I =numpy.array([500.0,1000.0,1000.0])
M =75.0
V =50.0
rho=1.225
CD0=AC.drag.total.get_total_drag()
FC.addTrimmedFlightCondition('cruise1',M,CG,I,V,rho,CmTrim=0.0,Cd0=CD0)
FC.flightConditionList[0].flap = 0.0
#FC.addTrimmedFlightCondition('cruise2',M/2,CG,I,V/2,rho,Cd0=CD0)
a3d=aero3d_VLM(AC)
R=a3d.runVLM(FC,False)
R.results[0].display()
print_header('PHUGOID')
R.results[0].dynamicStability.phugoid.display()
print_header('SHORT PERIOD')
R.results[0].dynamicStability.shortPeriod.display()
print_header('DUTCH ROLL')
R.results[0].dynamicStability.dutchRoll.display()
print_header('SPIRAL')
R.results[0].dynamicStability.spiral.display()
print_header('ROLL')
R.results[0].dynamicStability.roll.display()
AC.display()
def get_drag_polar():
rsltFile = 'drag_polar.txt'
ac = aircraft.load('V0510')
fc = aero.flightConditions()
CG = ac.get_CG(False)
I = ac.get_inertia(False)
M = ac.get_mass_total(False)
V = 34.0
rho = 1.2255
CD0 = ac.get_drag(V,0.0)
fc.addTrimmedFlightCondition('trim',M, CG, I, V, rho)
aero3D = aero.aero3d_VLM(ac)
rslt = aero3D.runVLM(fc)
alpha = np.linspace(-10.,20.,20)
CLexp, CDexp, alphaExp = experimental_data()
CLa = rslt.results[0].a * (np.pi/180.0)
CL0 = rslt.results[0].CL0
k = rslt.results[0].k
CL = CL0 + alpha*CLa
CD = CD0 + CL*CL*k
fid = open(rsltFile,'wt')
fid.write('Velocity %.2f m/sec\n'%V)
fid.write('Density %.2f kg/m3\n'%rho)
fid.write('CD0 %.4f\n'%CD0)
fid.write('alpha[deg]\tCL\tCD\n')
for i,a in enumerate(alpha):
fid.write('%.1f\t%.4f\t%.4f\n'%(a,CL[i],CD[i]))
fid.close()
def aero_valid():
ac = aircraft.load('V0510')
fc = aero.flightConditions()
CG = ac.get_CG(False)
I = ac.get_inertia(False)
M = ac.get_mass_total(False)
V = 34.0
rho = 1.2255
CD0 = ac.get_drag(V,0.0)
fc.addTrimmedFlightCondition('trim',M, CG, I, V, rho)
aero3D = aero.aero3d_VLM(ac)
rslt = aero3D.runVLM(fc)
alpha = np.linspace(-10.,20.,20)
CLexp, CDexp, alphaExp = experimental_data()
CLa = rslt.results[0].a * (np.pi/180.0)
CL0 = rslt.results[0].CL0
k = rslt.results[0].k
rslt.results[0].display()
CL = CL0 + alpha*CLa
CD = CD0 + CL*CL*k
plt = mpl.pyplot
font = {'family' : 'normal',
'weight' : 'normal',
'size' : 20}
mpl.rc('font',**font)
plt.figure(1)
plt.hold(True)
plt.grid(True)
plt.plot(CDexp,CLexp,'bo-')
plt.plot(CD,CL,'r-')
plt.ylabel('lift coefficient')
plt.xlabel('drag coefficient')
plt.axis([0,0.3,-1.0,1.5])
plt.legend(['wind tunnel','calculated'],'lower right')
plt.figure(2)
plt.hold(True)
plt.grid(True)
plt.plot(alphaExp,CDexp,'bo-')
plt.plot(alpha,CD,'r-')
plt.xlabel('angle of attack,deg')
plt.ylabel('drag coefficient')
plt.legend(['wind tunnel','calculated'],'upper left')
plt.figure(3)
plt.hold(True)
plt.grid(True)
plt.plot(alphaExp,CLexp,'bo-')
plt.plot(alpha,CL,'r-')
plt.xlabel('angle of attack,deg')
plt.ylabel('lift coefficient')
plt.legend(['wind tunnel','calculated'],'lower right')
plt.axis([-10,25,-0.5,1.5])
plt.show()
def run_test2():
import aircraft
ac = aircraft.load('V0510')
aero = aerodynamics(ac)
aero.update(65.,0.8,3.0)
aero.results.display()
ac.display()
def kla100_mass():
import aircraft
ac = aircraft.load('V0510')
print 'Aft CG\n======'
ac.mass.update_item_mass('Payload','baggage',70.)
ac.mass.display()
ac.mass.update_item_mass('Fuel','Fuel Tank Right',47.5/2.)
ac.mass.update_item_mass('Fuel','Fuel Tank Left',47.5/2.)
ac.mass.total.display()
ac.mass.update_item_mass('Fuel','Fuel Tank Right',2.5)
ac.mass.update_item_mass('Fuel','Fuel Tank Left',2.5)
ac.mass.total.display()
print 'front CG\n========'
ac.mass.update_item_mass('Payload','passenger',86.)
ac.mass.update_item_mass('Payload','pilot',86.)
ac.mass.update_item_mass('Payload','baggage',0.0)
ac.mass.update_item_mass('Fuel','Fuel Tank Right',47.5)
ac.mass.update_item_mass('Fuel','Fuel Tank Left',47.5)
ac.mass.total.display()
ac.mass.update_item_mass('Fuel','Fuel Tank Right',47.5/2)
ac.mass.update_item_mass('Fuel','Fuel Tank Left',47.5/2)
ac.mass.total.display()
ac.mass.update_item_mass('Fuel','Fuel Tank Right',2.5)
ac.mass.update_item_mass('Fuel','Fuel Tank Left',2.5)
ac.mass.total.display()
ac.display()
def example2():
# run trim using modified function
import aircraft
ac = aircraft.load('V0510')
aero = aerodynamics(ac)
aero.update(65.,0.8,3.0)
aero.results.display()
ac.display()
def debug1():
import aircraft
ac = aircraft.load('V0510')
ac.mass.fuel.set_fuel_mass(30,'total')
ac.mass.display()
ac.mass.fuel.set_fuel_mass_burned(1.5,'total')
ac.mass.display()
baggage = MassComponent('baggage', 70.0, [2.85,0,-0.125])
baggage.display(True,True)
def test_flight_mechanics():
ac = aircraft.load('V0510')
bi = BasicInput(ac,1.6,2000.)
tm = ac.analysis.thrust
fm = FlightMechanics(bi,tm)
print fm.get_EAS(50,1.05)
print fm.get_LDmax()
print fm.get_requiredThrust(65)
print fm.get_Vstall(1.2)
def new_report():
ac = aircraft.load('V0510')
sr = StabilityReport(ac)
sr.init_report()
sr.add_cg([1.656,0.0,0.0],726.0,1.0,'current FWD')
sr.add_cg([1.800,0.0,0.0],726.0,1.0,'current AFT')
sr.add_flight_condition(0.0,[20.0,75.0,15])
sr.reportStaticMargin = True
sr.set_required_param('derivatives',['Clb','Cnb','Cma'],['Cl_beta','Cn_beta','Cm_alpha'])
sr.set_required_param('basic',['elevator','xNP'],['elev deflection,deg','neutral point'])
sr.generate(14)
def run_takeoff():
name = 'V0510-MDO'
velocity =50.0
altitude =2000.
ac =aircraft.load(name)
mass =ac.mass.totalMass
rslt = ac.analyze_aero_trim(velocity,altitude)
ac.analysis.aerodynamics.update(velocity,1.225)
Cd0 =ac.get_drag(velocity,altitude)
tm =ac.analysis.thrust
TO = Takeoff(ac,0.0,tm)
print TO.analyze(1.5,800.0)
def run_test1():
ac = aircraft.load('V0510-MDO')
alt = 2000.0
vel = ny.linspace(10,100.0,20)
Cd0 = list()
for v in vel:
Cd0.append(ac.get_drag(v,alt))
plt.figure(2)
plt.plot(vel,Cd0,'b-')
plt.hold(True)
plt.grid(True)
plt.xlabel('velocity, m/sec')
plt.ylabel('CD0')
plt.show()
def generate_stability_report():
ac = aircraft.load('V0510')
names,massList,cgList = get_mass_data()
velocity = 60.0
density = 1.0581
loadFactor = 1.0
aero = aerodynamics.aerodynamics(ac)
out = ''
show = [True,False,False,False]
for name,m,cg,s in zip(names,massList,cgList,show):
aero.update(velocity,density,loadFactor,name=name,mass=m,CG=cg)
sr = rt.StabilityTextReport(aero.results,velocity,m,cg)
out += sr._report(s,dynamic=True) + '\n'
print out
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 runTest_fixPitch():
import aircraft
import matplotlib.pyplot as plt
ac =aircraft.load("sampleInput1")
thrust =thrustPerEngine(ac.engine.engineList[0],ac.propeller.propellerList[0])
N=10
Blist =frange(15.,45.,N) # pitch angle list
propRPMlist =frange(500.,3000.,N)
Vlist =frange(1,70,N)
thrust.buildPropTable(Blist,propRPMlist,Vlist)
V =30.
rho =1.2
beta =20.0
#pwrList =np.array(frange(50.,100.,20))
pwrList = np.array([50.,75.,100.])
Vlist =np.array(frange(1., 70.,20))
t1 = time.time()
for i in range(0,100,1):thrust.thrust_fixedPitch(V,rho,100,beta,True)
t2=time.time()
for i in range(0,100,1):thrust.thrust_fixedPitch(V,rho,100,beta,False)
t3 = time.time()
print 'exact time = '+str(t2-t1)
print 'approx time = '+str(t3-t2)
T =np.zeros([len(pwrList),len(Vlist)])
P =np.zeros([len(pwrList),len(Vlist)])
N =np.zeros([len(pwrList),len(Vlist)])
RPM=np.zeros([len(pwrList),len(Vlist)])
VV =np.zeros([len(pwrList),len(Vlist)])
beta =25.
plt.figure(1)
plt.hold(True)
for i,pwr in enumerate(pwrList):
for j,V in enumerate(Vlist):
T[i,j],P[i,j],FF,RPM[i,j],N[i,j],beta=thrust.thrust_fixedPitch(V,rho,pwr,beta,1)
#aprx=thrust.thrust_fixedPitch(V,rho,pwr,beta,0)
VV[i,j]=V
plt.plot(VV[i,:],T[i,:],'o-')
plt.xlabel('Velocity, m/sec')
plt.ylabel('Thrust, N')
plt.grid(True)
plt.legend(['50% throttle','75% throttle','full throttle'])
plt.show()
def run_performance_test(inputSheet='V0510'):
from report_tools import header
velocity =50.
altitude =2000.
Clmax =1.4
nMax =4.5
ac =aircraft.load(inputSheet)
tm =ac.analysis.thrust
bi =BasicInput(ac, Clmax, altitude)
slf = SteadyLevelFlight(bi,tm)
clm = ClimbAndDescendingFlight(bi,tm)
trn = TurningFlight(bi,tm)
print header('TURN')
trn.run_turnRate(60.*np.pi/180.,velocity,nMax,100)
trn.display()
print header('CLIMB RATE')
clm.run_climbRate(velocity,100)
clm.display()
print header('MAXIMUM CLIMB ANGLE')
clm.run_maximumClimbAngle(100.)
clm.display()
timer.start()
print header('MAXIMUM CLIMB RATE')
clm.run_maximumClimbRate(100.)
clm.display()
timer.finish()
print header('MINIMUM GLIDE ANGLE')
clm.run_minGlideAngle()
clm.display()
print header('MINIMUM GLIDE SINK RATE')
clm.run_minGlideSinkRate()
clm.display()
print header('MAXIMUM VELOCITY')
slf.run_velocity_maxTAS(75.)
slf.display()
timer.start()
print header('MAXIMUM SAR')
slf.run_velocity_maxSAR()
slf.display()
timer.finish()
print header('MINIMUM FUEL')
slf.run_velocity_minFuel()
slf.display()
def run_test1():
import matplotlib.pyplot as plt
name = 'V0510-MDO'
velocity =50.0
altitude =2000.
Clmax =1.4
ac =aircraft.load(name)
rslt = ac.analyze_aero_trim(velocity,altitude)
ac.analysis.aerodynamics.update(velocity,1.225)
tm =ac.analysis.thrust
bi = BasicInput(ac,Clmax,altitude)
print_header('Cruise')
crs = Cruise(bi,tm)
rslt = crs.get_range_at_power_setting(80.,10.,2000.0,75.0,6)
rslt.display()
crs.display()
def new_report():
ac = aircraft.load('V0510')
sr = StabilityReport(ac)
sr.init_report()
sr.plot = True
sr.report = False
sr.add_cg([1.711,-0.040,-0.124],593.1,1.0,'aft load')
sr.add_cg([1.824,-0.0,-0.103],583.1,1.0,'fwd load')
sr.add_flight_condition(1500.0,[30.0,70.0,7])
#sr.set_alpha_sweep(0.0,70.0,aileron=20.0)
sr.reportStaticMargin = True
# sr.set_required_param('basic',['xNP','alpha','elevator'],['neutral point','trim AOA','trim elevator defl.'])
# sr.set_required_param('coefficients',['CL','CD'],['lift coef.','drag coef.'])
# sr.set_required_param('derivatives',['Clb','Cnb','Cma'],['Cl beta','Cn beta','Cm alpha'])
# sr.set_required_param('short period',['Wn','Z'],['Short period nat freq.','Short period damping ratio'])
# sr.set_required_param('dutch roll',['Wn','Z'],['Dutch roll nat freq.','Dutch roll damping ratio'])
sr.set_required_param('hinge moment',['Mrudder','Maileron','aileron','Melevator'],['rudder, N*m','aileron, N*m','aileron coef','elevator, N*m'])
sr.generate(14)
def run_test3():
ac = aircraft.load('V0510-MDO')
altitude = 0.0
dT = 0
velocity = 50.0
mass = ac.mass.totalMass
rslt = ac.analyze_aero_trim(velocity,altitude)
k =rslt.k
Cd0 =ac.get_drag(velocity,altitude)
S =ac.wing.area
Clmax = 1.8
tm =ac.analysis.thrust
bi =basicInput(altitude,dT,Cd0,k,S,mass,Clmax)
climb = Climb()
climb._init_climb(ac,bi,tm)
print climb.climb_simulation_integr(0,2000)
print climb.climb_simulation_max(0,2000)
def runTest_propTable():
import aircraft
import matplotlib.pyplot as plt
N=20
ac =aircraft.load("sampleInput1")
prop =ac.propeller.propellerList[0]
Blist =frange(15.,45.,N)
propRPMlist =frange(500.,3000.,N)
Vlist =frange(1,70,N)
thrust =thrustPerEngine(ac.engine.engineList[0],ac.propeller.propellerList[0])
thrust.buildPropTable(Blist,propRPMlist,Vlist)
exactMode =thrust.thrust_fixedPitch(40.,1.22,100,25.,True)
approxMode=thrust.thrust_fixedPitch(40.,1.22,100,25.,False)
print exactMode
print approxMode
#propTbl =propTable(engine,prop,Blist,propRPMlist,Vlist)
propTbl =thrust.propTable
beta =25.
rho =1.225
propRPM =1500.
V =35.
print 'CP = '+str(propTbl.getCP(beta,propRPM,V))
print 'Power = '+str(propTbl.getPropPower(beta,propRPM,V,rho)/1000.)+' kw'
print 'Thrust = '+str(propTbl.getPropThrust(beta,propRPM,V,rho))+' N'
T,P,Jreq,CT,CP,effy=prop.analyze_prop(beta,propRPM,V,1.225)
print 'CP = '+str(CP)
print 'Power = '+str(P/1000.)+' kw'
print 'Thrust = '+str(T)+' N'
T=np.zeros([len(Vlist)])
plt.figure(1)
plt.hold(True)
for i,RPM in enumerate(propRPMlist):
for j,V in enumerate(Vlist):
#T[j]=CPtab.getPropThrust(beta,RPM,V,rho)
T[j]=propTbl.getCT(beta,RPM,V)
plt.plot(Vlist,T,'o-')
plt.show()
def v200_stability():
# report for KLA100 CDR 2014-03-31
ac = aircraft.load('V204-ext_aileron')
#ac.display()
sr = StabilityReport(ac)
sr.init_report()
sr.plot = True
wingLoc = ac.wing.aapex[0]
wingMAC = ac.wing.MAC
sr.add_cg([0.2513*wingMAC+wingLoc, 0,0],620,1.0,'maximum mass,fw CG')
sr.add_cg([0.3691*wingMAC+wingLoc, 0,0],620.,1.0,'maximum mass, aft CG')
sr.add_cg([0.1649*wingMAC+wingLoc, 0,0],403.5,1.0,'minimum mass, abs fw CG')
sr.add_flight_condition(1500.0,[30,80,10])
sr.reportStaticMargin = True
sr.set_required_param('basic',['xNP','alpha','elevator'],['neutral point','trim AOA','trim elevator defl.'])
sr.set_required_param('coefficients',['CL','CD'],['lift coef.','drag coef.'])
sr.set_required_param('derivatives',['Clb','Cnb','Cma'],['Cl beta','Cn beta','Cm alpha'])
sr.set_required_param('short period',['Wn','Z'],['Short period nat freq.','Short period damping ratio'])
sr.set_required_param('dutch roll',['Wn','Z'],['Dutch roll nat freq.','Dutch roll damping ratio'])
sr.generate()
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 test_report():
ac = aircraft.load('V0510')
print ac.wing.MAC
print ac.vStab.MAC
print ac.hStab.MAC
sr = StabilityReport(ac)
sr.init_report()
CG = generate_payload_mass_list()
sr.add_flight_condition(0.0,[20.0,30.0,10],False,flap=15.0)
for cg in CG:
sr.add_payload_mass_list(cg)
print cg.display()
# sr.reportStaticMargin = True
# sr.reprtHingeMoments = True
# sr.set_required_param('basic',['alpha','elevator'],['Trim AOA, deg','elev. deflection,deg'])
# sr.set_required_param('coefficients',['CL','CD'],['lift coefficient','drag coefficient'])
# sr.set_required_param('derivatives',['Cma','Cnb','Clb'],['Cm-alpha','Cn-beta','Cl-beta'])
sr.set_required_param('short period',['w','Z'],['short period nat.freq.','short period damp.ratio'])
sr.set_required_param('dutch roll',['w','Z'],['dutch roll nat.freq.','dutch roll damp.ratio'])
# sr.set_required_param('hinge moment',['elevator','flap'],['elev hinge moment coeficient','flap hinge moment coefficient'])
sr.generate()
def calculate_part_inertia():
ac = aircraft.load('V0510')
name = ['main wing','hStab','vStab','fuselage']
Ixx, Iyy, Izz = zeros(len(name)),zeros(len(name)),zeros(len(name))
i = 0
m = 56.30
a = ac.wing.MAC
b = ac.wing.thicknessRatio*a
l = ac.wing.span
Ixx[i],Iyy[i],Izz[i] = rect_inertia(m,a,b,l)
i = 1
m = 10.40
a = ac.hStab.MAC
b = ac.hStab.thicknessRatio*a
l = ac.hStab.span
Ixx[i],Iyy[i],Izz[i] = rect_inertia(m,a,b,l)
i = 2
#m = ac.mass.airframe.get_item_mass_by_name('vStab')
m = 3.8
a = ac.vStab.MAC
b = ac.vStab.span
l = ac.vStab.thicknessRatio*a
Ixx[i],Iyy[i],Izz[i] = rect_inertia(m,a,b,l)
i = 3
m = 66.20+16.80-1.9+66.9
r = ac.fuselage.diameter / 2.0
l = ac.fuselage.length
Ixx[i],Iyy[i],Izz[i] = circ_inertia(m,r,l)
print 'unit: kg*m^2'
out = '{0:15}|{1:^12}|{2:^12}|{3:^12}|'.format('name','Ixx','Iyy','Izz')
print out + '\n' + '='*len(out)
for i,n in enumerate(name):
print '{0:15} {1:12.4f} {2:12.4f} {3:12.4f}'.format(n, Ixx[i],Iyy[i],Izz[i])
def runTest_varPitch():
import aircraft
import matplotlib.pyplot as plt
import matplotlib
ac=aircraft.load("V0510")
T=thrustPerEngine(ac.engine.engineList[0],ac.propeller.propellerList[0])
V =76.
rho =1.2
pctThrottle =75.
# N=20
# Blist =frange(15.,45.,N)
# propRPMlist =frange(500.,3000.,N)
# Vlist =frange(1,70,N)
#T.buildPropTable(Blist,propRPMlist,Vlist)
Vlist =np.array(range(30,70,2))
thrust1=np.zeros(len(Vlist))
effy1 =np.zeros(len(Vlist))
beta1 =np.zeros(len(Vlist))
thrust2=np.zeros(len(Vlist))
effy2 =np.zeros(len(Vlist))
beta2 =np.zeros(len(Vlist))
t1 = time.time()
for i in range(0,100,1):
T.thrust_variablePitch(V,rho,pctThrottle,mode='max_thrust')
t2 = time.time()
print 'Time = '+str((t2-t1)/100)+ ' seconds per run'
for i,V in enumerate(Vlist):
thrust1[i],powerAchieved1,fuelFlow1,RPM,effy1[i],beta1[i]=T.thrust_variablePitch(V,rho,pctThrottle,mode='max_thrust')
thrust2[i],powerAchieved2,fuelFlow2,RPM,effy2[i],beta2[i]=T.thrust_variablePitch(V,rho,pctThrottle,mode='max_efficiency')
print str(thrust1[i])+' '+str(thrust2[i])
font = {'family' : 'sans-serif',
'style' : 'normal',
'size' : 16}
matplotlib.rc('font', **font)
plt.figure(1)
plt.hold(True)
plt.grid(True)
l1,=plt.plot(Vlist,effy1,'b-')
l2,=plt.plot(Vlist,effy2,'r-')
plt.legend([l1,l2],['Max Thrust Mode','Max Efficiency Mode'],loc=2)
plt.title('Velocity vs. Efficiency')
plt.xlabel('Velocity (m/s)')
plt.ylabel('Efficiency')
plt.figure(2)
plt.hold(True)
plt.grid(True)
l1,=plt.plot(Vlist,thrust1,'b-')
l2,=plt.plot(Vlist,thrust2,'r-')
plt.legend([l1,l2],['Max Thrust Mode','Max Efficiency Mode'],loc=2)
plt.title('Velocity vs. Thrust')
plt.xlabel('Velocity (m/s)')
plt.ylabel('Thrust (N)')
plt.figure(3)
plt.hold(True)
plt.grid(True)
l1,=plt.plot(Vlist,beta1,'b-')
l2,=plt.plot(Vlist,beta2,'r-')
plt.legend([l1,l2],['Max Thrust Mode','Max Efficiency Mode'],loc=2)
plt.title('Velocity vs. Beta')
plt.xlabel('Velocity (m/s)')
plt.ylabel('Beta (deg)')
plt.show()
def cross_wind_landing_manual_input():
"""
Performs crosswind takeoff and landing simulation.
"""
Vcross = 10.0 #kt
rho = 1.2255
Vstall = 20.6
VTo = 1.1*Vstall
aileron = 2.0
alphaGround=0.0
steering = 0.0
g = 9.81
n = 25
Cbeta = 0.025
mu = 0.03
mass = 620.0
#CG = np.array([1.925,0,-0.144])
#CG = np.array([1.813,0,-0.125])
Vcross = Vcross*0.5144 # kt to m/s
aileron = np.radians(aileron)
steering = np.radians(steering)
ac = aircraft.load('V204')
xCG = 1.4+1.116*0.2554 #ac.wing.get_fs_on_mac(40.0)
CG = np.array([xCG,0,0])
#aero = aerodynamics.aerodynamics(ac)
#aeroData = aero.update_notrim(0.5*Vstall,rho,flap=flap,aileron=0.0,rudder=0.0,alpha=alphaGround,beta=beta)
T = 0
cref = 1.166
bref = 10.311
Sref = 11.607
Xmlg = ac.landingGear.groundContact_X[2]
Xnlg = ac.landingGear.groundContact_X[0]
Ymlg = ac.landingGear.groundContact_Y[2]
Zmlg = ac.landingGear.groundContact_Z[2]
Znlg = ac.landingGear.groundContact_Z[0]
zmg = CG[2] - Zmlg
zng = CG[2] - Znlg
ymg = abs(Ymlg)
V = np.linspace(0.8*VTo,VTo,n)
Png = np.zeros(n)
Pmg = np.zeros(n)
betaCross = np.zeros(n)
rudder = np.zeros(n)
PmgL = np.zeros(n)
PmgR = np.zeros(n)
PsiA = np.zeros(n)
Ytot = np.zeros(n)
Yng = np.zeros(n)
YmgL = np.zeros(n)
YmgR = np.zeros(n)
CYb = -0.541
CYda = 0.063
CYdr = 0.097
Clb = -0.175
Clda = 0.1356
Cldr = 0.0096
Cnb = 0.075
Cnda = -0.0156
Cndr = -0.0471
Cm = -0.1925
CD = 0.22
CL = 1.65
Mweight = mass*g*(CG[0] - Xmlg)
A = np.ones([3,3])
A[1,2] = 0.0
A[2,0] = 0.0
A[0,2] = mass/mu
A[2,2] = mass*(CG[2]-Zmlg)
b = np.zeros(3)
A1 = np.ones([3,3])
b1 = np.ones(3)
for i,v in enumerate(V):
betaCross[i] = np.arctan(Vcross/v)
qS = rho*v*v*Sref / 2.0
qSb = qS*bref
Maero = Cm*qS*cref
Mthrust = T*(-ac.landingGear.groundContact_Z[2])
D = qS*CD
L = qS*CL
b[0] = (D-T*np.cos(np.radians(alphaGround)))/mu
b[1] = mass*g-L-T*np.sin(np.radians(alphaGround))
b[2] = Maero - Mthrust - Mweight
x = np.linalg.solve(A,b)
Pmg[i] = x[0]
Png[i] = x[1]
A1[0,0] = CYb*qS + Cbeta*(Png[i] + Pmg[i])
A1[0,1] = 0.0
A1[0,2] = CYdr*qS
b1[0] = -( (CYb*betaCross[i] + CYda*aileron)*qS + Cbeta*Png[i]*steering )
A1[1,0] = Clb*qSb - Cbeta*(Png[i]*zng + Pmg[i]*zmg)
A1[1,1] = 2.0*ymg
A1[1,2] = Cldr*qSb
b1[1] = - ((Clb*betaCross[i] + Clda*aileron)*qSb - Pmg[i]*ymg - Cbeta*steering*Png[i]*zng)
A1[2,0] = Cnb*qSb + Cbeta*(Png[i]*(Xnlg-CG[0]) - Pmg[i]*(Xmlg-CG[0]))
A1[2,1] = 0.0
A1[2,2] = Cndr*qSb
b1[2] = - ( (Cnb*betaCross[i]+Cnda*aileron)*qSb - Cbeta*steering*Png[i]*(Xnlg-CG[0]) )
x1 = np.linalg.solve(A1,b1)
betaCross[i] = np.degrees(betaCross[i])
PsiA[i] = np.degrees(x1[0])
PmgL[i] = x1[1]
PmgR[i] = Pmg[i] - PmgL[i]
Yng[i] = Cbeta*Png[i]
YmgL[i] = Cbeta*PmgL[i]
YmgR[i] = Cbeta*PmgR[i]
Ytot[i] = Yng[i]+YmgL[i]+YmgR[i]
rudder[i] = np.degrees(x1[2])
#print '%.4f\t%.4f\t%.4f\t%.4f'%(np.degrees(betaCross[i]),np.degrees(x1[0]), x1[1], np.degrees(x1[2]))
print '%.1f\t%.1f\t%.1f\t%.1f\t%.1f\t%.1f\t%.1f'%(Yng[-1],PmgL[-1],PmgR[-1],Yng[-1],YmgL[-1],YmgR[-1],Ytot[-1])
fig1 = plt.figure(1)
ax1 = fig1.add_subplot(211)
ax1.hold(True)
ax1.grid(True)
ax1.plot(V,betaCross,'bo-')
ax1.plot(V,PsiA,'ro-')
ax1.legend(['beta cross','crab angle'],loc='upper left')
ax1.set_ylim([-30,30])
ax1.set_ylabel('deg')
ax1.set_xlim([0.85*V[0],V[-1]])
ax2 = fig1.add_subplot(212)
ax2.hold(True)
ax2.grid(True)
ax2.plot(V,rudder,'bo-')
ax2.set_ylabel('rudder defl, deg')
ax2.set_ylim([-45,45])
ax2.set_xlabel('Velocity,m/s')
ax2.set_xlim([0.85*V[0],V[-1]])
fig2 = plt.figure(2)
ax3 = fig2.add_subplot(211)
ax3.hold(True)
ax3.grid(True)
ax3.plot(V,Pmg/1e3,'bo-')
ax3.plot(V,PmgL/1e3,'ro-')
ax3.plot(V,PmgR/1e3,'go-')
ax3.plot(V,Png/1e3,'ko-')
ax3.legend(['main total','main left','main right','nose'],loc='upper left')
ax3.set_ylabel('Landing gear reaction, kN')
#ax3.set_ylim([-4,8])
ax3.set_xlim([0.85*V[0],V[-1]])
ax4 = fig2.add_subplot(212)
ax4.hold(True)
ax4.grid(True)
ax4.set_ylabel('Side force, kN')
ax4.plot(V,Ytot/1e3,'bo-')
ax4.plot(V,YmgL/1e3,'ro-')
ax4.plot(V,YmgR/1e3,'go-')
ax4.plot(V,Yng/1e3,'ko-')
ax4.legend(['total','main left','main right','nose'],loc='upper left')
#ax4.set_ylim([-0.05,Cbeta*mass*0.01])
ax4.set_xlim([0.85*V[0],V[-1]])
ax4.set_xlabel('Velocity,m/s')
plt.show()
def run_test3():
ac = aircraft.load('V05')
alt = 2000.0
ac.drag.total.display()