class HALEweight(object):
def __init__(self,aircraft):
self.ac = aircraft
self.emptyMass = MassList(self.ac.name)
self.fuelProp = constants.load('fuel_density')
self.constMass = constants.load('mass')
def analyze(self):
self._mass_wing()
def _mass_wing(self):
m = 100.
cgX = 5.0
self._add_mass1('wing',m, [cgX, 0, 0],False)
def _add_mass1(self,name,mass,cg=None,lb=True):
""" NOTE: CG is in meters"""
if lb:
mass = convert.lb_to_kg(mass)
self.emptyMass.add_item(name, mass, cg)
def __init__(self,aircraft):
self.ac = aircraft
self.emptyMass = MassList(self.ac.name)
self.fuelProp = constants.load('fuel_density')
self.constMass = constants.load('mass')
class BlendedWingBodyMass(object):
def __init__(self,aircraft):
self.ac = aircraft
self.emptyMass = MassList(self.ac.name)
self.fuelProp = constants.load('fuel_density')
self.constMass = constants.load('mass')
def analyze(self):
self._get_data()
self._mass_wing(self.ac.wing)
self._mass_fuselage()
self._mass_mlg()
self._mass_nlg()
self._mass_engine()
self._mass_engine_mount()
self._mass_engine_section()
self._mass_engine_cooling()
self._mass_airintake()
self._mass_oil_cooling()
self._mass_engine_controls()
self._mass_starter_pneumatic()
self._mass_fuel_system()
self._mass_flight_controls()
self._mass_hydraulics()
self._mass_electrical()
self._mass_avionics()
def _get_coefficients(self):
self.Kdw = 1.0 # 0.786 for delta wing, otherwise Kdw=1.0
self.Kvs = 1.0 # 1.19 for variable sweep, otherwise 1.0
self.Kdwf = 1.0 # 0.774 for delta wing, otherwise 1.0
self.Kcb = 1.0 # 2.25 for cross beam landing gear, otherwise 1.0
self.Ktpg = 1.0 # 0.826 for tripod landing gear, otherwise 1.0
def _get_data(self):
self._get_coefficients()
self.Nz = 1.5 * self.ac.designGoals.loadFactor # ultimate load factor
self.Wdg = convert.kg_to_lb(self.ac.designGoals.grossMass)
self.Nl = self.ac.designGoals.loadFactorLanding
self.Wt = self.Wdg
self.Wl = self.Wdg
self.Ne = self.ac.propulsion.numberOfEngines
self.T = convert.kgf_to_lbf(self.ac.propulsion.totalThrust)
self.mUav = convert.kg_to_lb(self.ac.designGoals.avionicsMass)
Vi = self.ac.designGoals.fuelMass/self.fuelProp['kerosene']
self.Vi = convert.cubm_to_gal(Vi)
def _add_mass1(self,name,mass,cg=None,lb=True):
""" NOTE: CG is in meters"""
if lb:
mass = convert.lb_to_kg(mass)
self.emptyMass.add_item(name, mass, cg)
def _mass_engine(self):
mKg = self.ac.propulsion.totalMass
CG = self.ac.propulsion.CG
self._add_mass1('engine',mKg,CG,False)
def _mass_wing(self,wing):
# by Raymer - fighter weights
Sw = convert.sqm_to_sqft(wing.area)
A = wing.aspectRatio
tcRoot = wing.airfoils[1].thickness #FIXME: assume thickness of 2nd section
# --- own corrections ---
eqArea = (wing.chords[0]+wing.chords[-1])*wing.span/2.0
corr = wing.area/eqArea
Kms = 1.1 # correction for multisegment wings
TR = wing.taperRatio
sweepC4 = wing.equivSweepC4rad
Scsw = wing.csArea
m = 0.0103*self.Kdw*self.Kvs* (self.Wdg*self.Nz)**0.5* Sw**0.622* A**0.785 *tcRoot**(-0.4)* (1+TR)**0.5*np.cos(sweepC4)**(-1.0)* Scsw**0.04
m = corr*Kms*m
wingCGratio = self.constMass['wingCGratio']
xCG = self.ac.wing.MAClocation[0] + self.ac.wing.MAC*wingCGratio
zCG = self.ac.wing.secApex[0,2]
if wing.material=='composite':
m = m*self.constMass['compositeWing']
self._add_mass1('wing',m,np.array([xCG,0.0,zCG]))
def _mass_fuselage(self):
# Raymer - fighter
L = convert.m_to_ft(self.ac.wing.chords[0])
D = convert.m_to_ft(self.ac.wing.airfoils[0].thickness)
W = convert.m_to_ft(self.ac.fusWidth)
m = 0.499*self.Kdwf*self.Wdg**0.35 *self.Nz**0.25 *L**0.5 *D**0.849 *W**0.685
xCG = self.ac.wing.chords[0] * self.constMass['fuseCGratio'][1]
zCG = self.ac.wing.secApex[0,2]
if self.ac.wing.material=='composite':
m = m*self.constMass['compositeBody']
self._add_mass1('fuselage',m,np.array([xCG,0.0,zCG]))
def _mass_mlg(self):
# Lm = convert.m_to_ft(self.ac.landingGear.strutLength[1]) #TODO: check value
# m = self.Kcb*self.Ktpg*(self.Wt*self.Nl)**0.25 * Lm**0.973
# Gundlach - UAV design
m = self.Wt * 0.04*0.85
xCG = self.ac.landingGear.groundContactX[1]
self._add_mass1('main landing gear',m,np.array([xCG,0,0]))
def _mass_nlg(self):
# Ln = convert.m_to_ft(self.ac.landingGear.strutLength[0])
# Nnw = 1.0 # number of nose wheels
# m = (self.Wl*self.Nl)**0.29 *Ln**0.5 *Nnw**0.525
# Gundlach - UAV design
m = self.Wt * 0.04*0.15
xCG = self.ac.landingGear.groundContactX[0]
self._add_mass1('nose landing gear',m,np.array([xCG,0,0]))
def _mass_engine_mount(self):
cg = self.emptyMass.get_item_cg_by_name('engine')
m = 0.013*self.Ne**0.795*self.T**0.579 *self.Nz
self._add_mass1('engine mount',m,cg)
def _mass_engine_section(self):
cg = self.emptyMass.get_item_cg_by_name('engine')
We = convert.kg_to_lb(self.ac.propulsion.engine.mass)
m = 0.01*We**0.717*self.Ne*self.Nz
self._add_mass1('engine section',m,cg)
def _mass_airintake(self):
Kvg = 1.0 # 1.62 for variable geometry
Kd = 3.43 # duct constant. see fig 15.2 Raymer
Ld = self.ac.wing.chords[0]-self.ac.propulsion.engine.length
Ld = convert.m_to_ft(Ld)
Ls = Ld
De = convert.m_to_ft(self.ac.propulsion.engine.diameter)
mAir = 13.29*Kvg * Ld**0.643* Kd**0.182 *self.Ne**1.498 * (Ls/Ld)**(-.373)*De
self._add_mass1('air intake',mAir)
def _mass_engine_cooling(self):
#FIXME: engine shroud length is assumed 0.15*total length
cg = self.emptyMass.get_item_cg_by_name('engine')
Lsh = convert.m_to_ft(self.ac.propulsion.engine.length)*0.15
De = convert.m_to_ft(self.ac.propulsion.engine.diameter)
m = 4.55*De*Lsh*self.Ne
self._add_mass1('engine cooling',m,cg)
def _mass_oil_cooling(self):
m = 37.82*self.Ne**1.023
self._add_mass1('oil cooling',m)
def _mass_engine_controls(self):
#FIXME: assumed that engine controls are close to engine since it is UAV
cg = self.emptyMass.get_item_cg_by_name('engine')
Lec = 0.25*convert.m_to_ft(self.ac.propulsion.engine.length)
m = 10.5*self.Ne**1.008*Lec**0.222
self._add_mass1('engine controls',m,cg)
def _mass_starter_pneumatic(self):
Te = convert.kgf_to_lbf(self.ac.propulsion.engine.thrustMC)
m = 0.025*Te**0.760*self.Ne**0.72
self._add_mass1('engine starter',m)
def _mass_fuel_system(self):
#FIXME: assumption that all tanks are protected and pressurized
Vi = 0.5*self.Vi
Vt = self.Vi
Vp = 0*self.Vi
sfc = self.ac.propulsion.engine.sfcMC
Nt = self.ac.propulsion.numberOfTanks
m = 7.45*Vt**0.47*(1.+Vi/Vt)**(-0.095)*(1.+Vp/Vt)*Nt**0.066*self.Ne**0.052*(self.T*sfc/1e3)**0.249
self._add_mass1('fuel system',m)
def _mass_flight_controls(self):
M = self.ac.designGoals.cruiseMach
Scs = self.ac.wing.csArea
Nc = 1.0 #FIXME: number of crew is set to 1 for UAV
Ns = 2.0
m = 36.28*M**0.003*Scs**0.489*Ns**0.484*Nc**0.127
self._add_mass1('flight controls',m)
def _mass_hydraulics(self):
Kvsh = 1.0 #1.425 for variable sweep wing, otherwise 1.0
Nu = 10.0 # number of hydraulic utility functions (typically 5-15)
m = 37.23*Kvsh*Nu**0.664
self._add_mass1('hydraulics',m)
def _mass_electrical(self):
Rkva = 120.0 # system electrical rating
Kmc = 1.45
Nc = 1.0
La = convert.m_to_ft(self.ac.wing.chords[0])
Ngen = self.Ne
m = 172.2*Kmc*Rkva**0.152*Nc**0.1*La**0.1*Ngen**0.0091
self._add_mass1('electric system',m)
def _mass_avionics(self):
#mUav = 600.#FIXME: uninstalled avionics weight is assumed 1000lb
mUav = self.mUav
m = 2.117*mUav**0.933
self._add_mass1('avionics',m)
