vsp.VSPRenew()
errorMgr.PopErrorAndPrint(stdout)
geoms = vsp.FindGeoms()
print("All geoms in Vehicle.")
print(geoms)
# Add Fuse
fuse_id = vsp.AddGeom("FUSELAGE")
# Get XSec Surf ID
xsurf_id = vsp.GetXSecSurf(fuse_id, 0)
# Change Type of First XSec
vsp.ChangeXSecShape(xsurf_id, 0, vsp.XS_SUPER_ELLIPSE)
errorMgr.PopErrorAndPrint(stdout)
# Change Type First XSec Properties
xsec_id = vsp.GetXSec(xsurf_id, 0)
width_id = vsp.GetXSecParm(xsec_id, "Super_Width")
height_id = vsp.GetXSecParm(xsec_id, "Super_Height")
vsp.SetParmVal(width_id, 4.0)
vsp.SetParmVal(height_id, 2.0)
# Copy Cross-Section to Clipboard
vsp.CopyXSec(fuse_id, 0)
# Paste Cross-Section
vsp.PasteXSec(fuse_id, 1)
vsp.PasteXSec(fuse_id, 2)
def write(vehicle, tag):
# Reset OpenVSP to avoid including a previous vehicle
try:
vsp.ClearVSPModel()
except NameError:
print 'VSP import failed'
return -1
area_tags = dict() # for wetted area assignment
# -------------
# Wings
# -------------
for wing in vehicle.wings:
wing_x = wing.origin[0]
wing_y = wing.origin[1]
wing_z = wing.origin[2]
if wing.symmetric == True:
span = wing.spans.projected / 2. # span of one side
else:
span = wing.spans.projected
root_chord = wing.chords.root
tip_chord = wing.chords.tip
sweep = wing.sweeps.quarter_chord / Units.deg
sweep_loc = 0.25
root_twist = wing.twists.root / Units.deg
tip_twist = wing.twists.tip / Units.deg
root_tc = wing.thickness_to_chord
tip_tc = wing.thickness_to_chord
dihedral = wing.dihedral / Units.deg
# Check to see if segments are defined. Get count
if len(wing.Segments.keys()) > 0:
n_segments = len(wing.Segments.keys())
else:
n_segments = 0
# Create the wing
wing_id = vsp.AddGeom("WING")
vsp.SetGeomName(wing_id, wing.tag)
area_tags[wing.tag] = ['wings', wing.tag]
# Make names for each section and insert them into the wing if necessary
x_secs = []
x_sec_curves = []
# n_segments + 2 will create an extra segment if the root segment is
# included in the list of segments. This is not used and the tag is
# removed when the segments are checked for this case.
for i_segs in xrange(0, n_segments + 2):
x_secs.append('XSec_' + str(i_segs))
x_sec_curves.append('XSecCurve_' + str(i_segs))
# Apply the basic characteristics of the wing to root and tip
if wing.symmetric == False:
vsp.SetParmVal(wing_id, 'Sym_Planar_Flag', 'Sym', 0)
if wing.vertical == True:
vsp.SetParmVal(wing_id, 'X_Rel_Rotation', 'XForm', 90)
vsp.SetParmVal(wing_id, 'X_Rel_Location', 'XForm', wing_x)
vsp.SetParmVal(wing_id, 'Y_Rel_Location', 'XForm', wing_y)
vsp.SetParmVal(wing_id, 'Z_Rel_Location', 'XForm', wing_z)
# This ensures that the other VSP parameters are driven properly
vsp.SetDriverGroup(wing_id, 1, vsp.SPAN_WSECT_DRIVER,
vsp.ROOTC_WSECT_DRIVER, vsp.TIPC_WSECT_DRIVER)
# Root chord
vsp.SetParmVal(wing_id, 'Root_Chord', x_secs[1], root_chord)
# Sweep of the first section
vsp.SetParmVal(wing_id, 'Sweep', x_secs[1], sweep)
vsp.SetParmVal(wing_id, 'Sweep_Location', x_secs[1], sweep_loc)
# Twists
vsp.SetParmVal(wing_id, 'Twist', x_secs[0], tip_twist) # tip
vsp.SetParmVal(wing_id, 'Twist', x_secs[0], root_twist) # root
# Figure out if there is an airfoil provided
# Airfoils should be in Lednicer format
# i.e. :
#
#EXAMPLE AIRFOIL
# 3. 3.
#
# 0.0 0.0
# 0.5 0.1
# 1.0 0.0
#
# 0.0 0.0
# 0.5 -0.1
# 1.0 0.0
# Note this will fail silently if airfoil is not in correct format
# check geometry output
if n_segments == 0:
if len(wing.Airfoil) != 0:
xsecsurf = vsp.GetXSecSurf(wing_id, 0)
vsp.ChangeXSecShape(xsecsurf, 0, vsp.XS_FILE_AIRFOIL)
vsp.ChangeXSecShape(xsecsurf, 1, vsp.XS_FILE_AIRFOIL)
xsec1 = vsp.GetXSec(xsecsurf, 0)
xsec2 = vsp.GetXSec(xsecsurf, 1)
vsp.ReadFileAirfoil(xsec1,
wing.Airfoil['airfoil'].coordinate_file)
vsp.ReadFileAirfoil(xsec2,
wing.Airfoil['airfoil'].coordinate_file)
vsp.Update()
else: # The wing airfoil is still used for the root segment if the first added segment does not begin there
# This could be combined with above, but is left here for clarity
if (len(wing.Airfoil) !=
0) and (wing.Segments[0].percent_span_location != 0.):
xsecsurf = vsp.GetXSecSurf(wing_id, 0)
vsp.ChangeXSecShape(xsecsurf, 0, vsp.XS_FILE_AIRFOIL)
vsp.ChangeXSecShape(xsecsurf, 1, vsp.XS_FILE_AIRFOIL)
xsec1 = vsp.GetXSec(xsecsurf, 0)
xsec2 = vsp.GetXSec(xsecsurf, 1)
vsp.ReadFileAirfoil(xsec1,
wing.Airfoil['airfoil'].coordinate_file)
vsp.ReadFileAirfoil(xsec2,
wing.Airfoil['airfoil'].coordinate_file)
vsp.Update()
elif len(wing.Segments[0].Airfoil) != 0:
xsecsurf = vsp.GetXSecSurf(wing_id, 0)
vsp.ChangeXSecShape(xsecsurf, 0, vsp.XS_FILE_AIRFOIL)
vsp.ChangeXSecShape(xsecsurf, 1, vsp.XS_FILE_AIRFOIL)
xsec1 = vsp.GetXSec(xsecsurf, 0)
xsec2 = vsp.GetXSec(xsecsurf, 1)
vsp.ReadFileAirfoil(
xsec1, wing.Segments[0].Airfoil['airfoil'].coordinate_file)
vsp.ReadFileAirfoil(
xsec2, wing.Segments[0].Airfoil['airfoil'].coordinate_file)
vsp.Update()
# Thickness to chords
vsp.SetParmVal(wing_id, 'ThickChord', 'XSecCurve_0', root_tc)
vsp.SetParmVal(wing_id, 'ThickChord', 'XSecCurve_1', tip_tc)
# Dihedral
vsp.SetParmVal(wing_id, 'Dihedral', x_secs[1], dihedral)
# Span and tip of the section
if n_segments > 1:
local_span = span * wing.Segments[0].percent_span_location
sec_tip_chord = root_chord * wing.Segments[0].root_chord_percent
vsp.SetParmVal(wing_id, 'Span', x_secs[1], local_span)
vsp.SetParmVal(wing_id, 'Tip_Chord', x_secs[1], sec_tip_chord)
else:
vsp.SetParmVal(wing_id, 'Span', x_secs[1], span)
vsp.Update()
if n_segments > 0:
if wing.Segments[0].percent_span_location == 0.:
x_secs[-1] = [] # remove extra section tag (for clarity)
segment_0_is_root_flag = True
adjust = 0 # used for indexing
else:
segment_0_is_root_flag = False
adjust = 1
else:
adjust = 1
# Loop for the number of segments left over
for i_segs in xrange(1, n_segments + 1):
# Unpack
dihedral_i = wing.Segments[i_segs -
1].dihedral_outboard / Units.deg
chord_i = root_chord * wing.Segments[i_segs - 1].root_chord_percent
twist_i = wing.Segments[i_segs - 1].twist / Units.deg
sweep_i = wing.Segments[i_segs -
1].sweeps.quarter_chord / Units.deg
# Calculate the local span
if i_segs == n_segments:
span_i = span * (
1 - wing.Segments[i_segs - 1].percent_span_location
) / np.cos(dihedral_i * Units.deg)
else:
span_i = span * (
wing.Segments[i_segs].percent_span_location -
wing.Segments[i_segs - 1].percent_span_location) / np.cos(
dihedral_i * Units.deg)
# Insert the new wing section with specified airfoil if available
if len(wing.Segments[i_segs - 1].Airfoil) != 0:
vsp.InsertXSec(wing_id, i_segs - 1 + adjust,
vsp.XS_FILE_AIRFOIL)
xsecsurf = vsp.GetXSecSurf(wing_id, 0)
xsec = vsp.GetXSec(xsecsurf, i_segs + adjust)
vsp.ReadFileAirfoil(
xsec, wing.Segments[i_segs -
1].Airfoil['airfoil'].coordinate_file)
else:
vsp.InsertXSec(wing_id, i_segs - 1 + adjust,
vsp.XS_FOUR_SERIES)
# Set the parms
vsp.SetParmVal(wing_id, 'Span', x_secs[i_segs + adjust], span_i)
vsp.SetParmVal(wing_id, 'Dihedral', x_secs[i_segs + adjust],
dihedral_i)
vsp.SetParmVal(wing_id, 'Sweep', x_secs[i_segs + adjust], sweep_i)
vsp.SetParmVal(wing_id, 'Sweep_Location', x_secs[i_segs + adjust],
sweep_loc)
vsp.SetParmVal(wing_id, 'Root_Chord', x_secs[i_segs + adjust],
chord_i)
vsp.SetParmVal(wing_id, 'Twist', x_secs[i_segs + adjust], twist_i)
vsp.SetParmVal(wing_id, 'ThickChord',
x_sec_curves[i_segs + adjust], tip_tc)
vsp.Update()
vsp.SetParmVal(wing_id, 'Tip_Chord', x_secs[-1 - (1 - adjust)],
tip_chord)
vsp.SetParmVal(wing_id, 'CapUMaxOption', 'EndCap', 2.)
vsp.SetParmVal(wing_id, 'CapUMaxStrength', 'EndCap', 1.)
vsp.Update() # to fix problems with chords not matching up
if wing.tag == 'main_wing':
main_wing_id = wing_id
## Skeleton code for props and pylons can be found in previous commits (~Dec 2016) if desired
## This was a place to start and may not still be functional
# -------------
# Engines
# -------------
if vehicle.propulsors.has_key('turbofan'):
print 'Warning: no meshing sources are currently implemented for the nacelle'
# Unpack
turbofan = vehicle.propulsors.turbofan
n_engines = turbofan.number_of_engines
length = turbofan.engine_length
width = turbofan.nacelle_diameter
origins = turbofan.origin
bpr = turbofan.bypass_ratio
for ii in xrange(0, int(n_engines)):
origin = origins[ii]
x = origin[0]
y = origin[1]
z = origin[2]
nac_id = vsp.AddGeom("FUSELAGE")
vsp.SetGeomName(nac_id, 'turbofan')
# Length and overall diameter
vsp.SetParmVal(nac_id, "Length", "Design", length)
vsp.SetParmVal(nac_id, 'Abs_Or_Relitive_flag', 'XForm', vsp.ABS)
vsp.SetParmVal(nac_id, 'OrderPolicy', 'Design', 1.)
vsp.SetParmVal(nac_id, 'X_Location', 'XForm', x)
vsp.SetParmVal(nac_id, 'Y_Location', 'XForm', y)
vsp.SetParmVal(nac_id, 'Z_Location', 'XForm', z)
vsp.SetParmVal(nac_id, 'Origin', 'XForm', 0.5)
vsp.SetParmVal(nac_id, 'Z_Rotation', 'XForm', 180.)
xsecsurf = vsp.GetXSecSurf(nac_id, 0)
vsp.ChangeXSecShape(xsecsurf, 0, vsp.XS_ELLIPSE)
vsp.Update()
vsp.SetParmVal(nac_id, "Ellipse_Width", "XSecCurve_0", width - .2)
vsp.SetParmVal(nac_id, "Ellipse_Width", "XSecCurve_1", width)
vsp.SetParmVal(nac_id, "Ellipse_Width", "XSecCurve_2", width)
vsp.SetParmVal(nac_id, "Ellipse_Width", "XSecCurve_3", width)
vsp.SetParmVal(nac_id, "Ellipse_Height", "XSecCurve_0", width - .2)
vsp.SetParmVal(nac_id, "Ellipse_Height", "XSecCurve_1", width)
vsp.SetParmVal(nac_id, "Ellipse_Height", "XSecCurve_2", width)
vsp.SetParmVal(nac_id, "Ellipse_Height", "XSecCurve_3", width)
vsp.Update()
# -------------
# Fuselage
# -------------
if vehicle.fuselages.has_key('fuselage'):
# Unpack
fuselage = vehicle.fuselages.fuselage
width = fuselage.width
length = fuselage.lengths.total
hmax = fuselage.heights.maximum
height1 = fuselage.heights.at_quarter_length
height2 = fuselage.heights.at_wing_root_quarter_chord
height3 = fuselage.heights.at_three_quarters_length
effdia = fuselage.effective_diameter
n_fine = fuselage.fineness.nose
t_fine = fuselage.fineness.tail
w_ac = wing.aerodynamic_center
w_origin = vehicle.wings.main_wing.origin
w_c_4 = vehicle.wings.main_wing.chords.root / 4.
# Figure out the location x location of each section, 3 sections, end of nose, wing origin, and start of tail
x1 = n_fine * width / length
x2 = (w_origin[0] + w_c_4) / length
x3 = 1 - t_fine * width / length
fuse_id = vsp.AddGeom("FUSELAGE")
vsp.SetGeomName(fuse_id, fuselage.tag)
area_tags[fuselage.tag] = ['fuselages', fuselage.tag]
if fuselage.has_key('OpenVSP_values'):
vals = fuselage.OpenVSP_values
# Nose
vsp.SetParmVal(fuse_id, "TopLAngle", "XSec_0", vals.nose.top.angle)
vsp.SetParmVal(fuse_id, "TopLStrength", "XSec_0",
vals.nose.top.strength)
vsp.SetParmVal(fuse_id, "RightLAngle", "XSec_0",
vals.nose.side.angle)
vsp.SetParmVal(fuse_id, "RightLStrength", "XSec_0",
vals.nose.side.strength)
vsp.SetParmVal(fuse_id, "TBSym", "XSec_0", vals.nose.TB_Sym)
vsp.SetParmVal(fuse_id, "ZLocPercent", "XSec_0", vals.nose.z_pos)
# Tail
vsp.SetParmVal(fuse_id, "TopLAngle", "XSec_4", vals.tail.top.angle)
vsp.SetParmVal(fuse_id, "TopLStrength", "XSec_4",
vals.tail.top.strength)
# Below can be enabled if AllSym (below) is removed
#vsp.SetParmVal(fuse_id,"RightLAngle","XSec_4",vals.tail.side.angle)
#vsp.SetParmVal(fuse_id,"RightLStrength","XSec_4",vals.tail.side.strength)
#vsp.SetParmVal(fuse_id,"TBSym","XSec_4",vals.tail.TB_Sym)
#vsp.SetParmVal(fuse_id,"BottomLAngle","XSec_4",vals.tail.bottom.angle)
#vsp.SetParmVal(fuse_id,"BottomLStrength","XSec_4",vals.tail.bottom.strength)
if vals.tail.has_key('z_pos'):
tail_z_pos = vals.tail.z_pos
else:
tail_z_pos = 0.02
vsp.SetParmVal(fuse_id, "AllSym", "XSec_4", 1)
vsp.SetParmVal(fuse_id, "Length", "Design", length)
vsp.SetParmVal(fuse_id, "Diameter", "Design", width)
vsp.SetParmVal(fuse_id, "XLocPercent", "XSec_1", x1)
vsp.SetParmVal(fuse_id, "XLocPercent", "XSec_2", x2)
vsp.SetParmVal(fuse_id, "XLocPercent", "XSec_3", x3)
vsp.SetParmVal(fuse_id, "ZLocPercent", "XSec_4", tail_z_pos)
vsp.SetParmVal(fuse_id, "Ellipse_Width", "XSecCurve_1", width)
vsp.SetParmVal(fuse_id, "Ellipse_Width", "XSecCurve_2", width)
vsp.SetParmVal(fuse_id, "Ellipse_Width", "XSecCurve_3", width)
vsp.SetParmVal(fuse_id, "Ellipse_Height", "XSecCurve_1", height1)
vsp.SetParmVal(fuse_id, "Ellipse_Height", "XSecCurve_2", height2)
vsp.SetParmVal(fuse_id, "Ellipse_Height", "XSecCurve_3", height3)
# Write the vehicle to the file
vsp.WriteVSPFile(tag + ".vsp3")
return area_tags