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_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_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_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_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 get_parasite_drag_fw(ac,altitude=0.0):
path1 = pth.aeroCD0in
path2 = pth.aeroCD0inWave
W = ac.designGoals.grossMass*9.81
rho = ac.designGoals.fc.atm.density
V = ac.designGoals.cruiseSpeed
Sref = ac.wing.area
CL=W/(0.5*rho*V*V*Sref)
# --- friction + form drag input ---
fid1 = open(path1,'wt')
#h = ac.designGoals.cruiseAltitude
#h = convert.m_to_ft(h)
h = float(altitude)
Sref = convert.sqm_to_sqft(Sref)
Abase = 0.095
Dexit = 0.0
Ewdd = 1.8
prleak = 10.0
fid1.write('%d %.0f %.0f %.4f %.1f %.1f %.1f\n'%(1,h,Sref,Abase, Dexit, Ewdd, prleak))
ale = ac.wing.equivSweepLEdeg
aqc = ac.wing.equivSweepC4deg
ahc = ac.wing.equivSweepC2deg
a4 = ac.wing.span
a4 = convert.m_to_ft(a4)
# ---
t1 = ac.wing.airfoils[0].thickness*ac.wing.chords[0]
t2 = ac.wing.airfoils[1].thickness*ac.wing.chords[1]
Amax = (t1+t2)*ac.wing.segSpans[0]
# ---
ttc = ac.wing.equivThickness
fc = ac.wing.equivCamber
cld = CL
xa = 1.02
fid1.write('%.4f %.4f %.4f %.4f %.4f %.4f %.6f %.4f %.2f\n'%(ale,aqc,ahc,a4,Amax,ttc,fc,cld,xa))
wSwet = ac.wing.wettedArea
wSwet = convert.sqm_to_sqft(wSwet)
wRefL = convert.m_to_ft(ac.wing.MAC)
fid1.write('Wing %.4f %.4f 1 '%(wSwet,wRefL))
wtc = ac.wing.equivThickness
wc4 = ac.wing.equivSweepC4deg
wxtc = ac.wing.equivThicknessLoc
q = 1.0 #???
fid1.write('%.4f %.4f %.4f %.4f'%(wtc,wc4,wxtc,q))
fid1.close()
# --- wave drag input ---
fid2 = open(path2,'wt')
fid2.write('1 0 0 0 0 0 0 0\n')
wrle = ac.wing.equivLEradius
cam = ac.wing.equivCamber
fid2.write('%.4f %.4f %.4f %.4f %d\n'%(wtc,wxtc,wrle,cam, 1))
fid2.write('0.0394 0.5 0. 0. 0\n')
fid2.write('0. 0. 0. 0. 1\n')
fid2.write('0.04 0.5 0. 0. 0\n')
wte = ac.wing.equivSweepTEdeg
lref = convert.m_to_ft(ac.wing.MAC)
AR = ac.wing.aspectRatio
fid2.write('%.4f %.4f %.4f %.4f %.4f\n'%(ale,wte,lref,AR,Sref))
fid2.write('44.56 0. 0.29 2.24\n')
fid2.write('0. 0. 0. 0.\n')
fid2.write('44. 3. 0.277 2.48\n')
fid2.write('0. 0. 0. 0. 0.\n')
fid2.write('2\n')
fid2.write('21.08 20.92 19.364 4.97 2.9\n')
fid2.write('0.35 0.45 0.47 0.9339\n')
fid2.close()
os.system(pth.aeroCD0)
data = read_tabulated_data_without_header(pth.aeroCD0out,3)
pth.clean_drag_files()
CD = data[:-1,-1]
M = data[:-1,0]
Mdd = data[-1,0]
CDdd = data[-1,-1]
return M, CD, Mdd, CDdd