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 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