def get_fuel_tank_properties(vehicle,
tag,
fuel_tank_set_index=3,
slices_for_calculation=100):
"""This function computes the center of gravity, total possible fuel mass,
the available volume of each fuel tank in the vehicle through a mass
properties computation in OpenVSP.
Assumptions:
Fuel tanks exists in the fuselage and wings only
All fuel tanks have unique names
Source:
N/A
Inputs:
vehicle.fuselages.*.Fuel_Tanks.*.tag [-]
vehicle.wings.*.Fuel_Tanks.*.tag [-]
Outputs:
vehicle.fuselages.*.Fuel_Tanks.*.mass_properties.
center_of_gravity [m]
fuel_mass_when_full [kg]
fuel_volume_when_full [m^3]
vehicle.wings.*.Fuel_Tanks.*.mass_properties.
center_of_gravity [m]
fuel_mass_when_full [kg]
fuel_volume_when_full [m^3]
Properties Used:
N/A
"""
# Reset OpenVSP to avoid including a previous vehicle
vsp.ClearVSPModel()
vsp.ReadVSPFile(tag + '.vsp3')
# Extract fuel tanks from vehicle
fuel_tanks = get_fuel_tanks(vehicle)
num_slices = slices_for_calculation # Slices used to estimate mass distribution from areas in OpenVSP
mass_props_output_file = tag + '_mass_props.txt'
vsp.SetComputationFileName(vsp.MASS_PROP_TXT_TYPE, mass_props_output_file)
print('Computing Fuel Tank Mass Properties... ')
vsp.ComputeMassProps(fuel_tank_set_index, num_slices)
print('Done')
# Extract full tank mass properties from OpenVSP output file
fo = open(mass_props_output_file)
for line in fo:
prop_list = line.split()
try:
if prop_list[0] in fuel_tanks:
# Indices based on position in OpenVSP output (may change in the future)
cg_x = float(prop_list[2])
cg_y = float(prop_list[3])
cg_z = float(prop_list[4])
mass = float(prop_list[1])
vol = float(prop_list[-1])
if 'center_of_gravity' not in fuel_tanks[prop_list[
0]]: # assumes at most two identical tank names
fuel_tanks[prop_list[0]].center_of_gravity = np.array(
[cg_x, cg_y, cg_z])
fuel_tanks[prop_list[0]].fuel_mass_when_full = mass
fuel_tanks[prop_list[0]].volume = vol
else:
fuel_tanks[prop_list[0]].center_of_gravity = \
(fuel_tanks[prop_list[0]].center_of_gravity+np.array([cg_x,cg_y,cg_z]))/2.
fuel_tanks[prop_list[0]].fuel_mass_when_full += mass
fuel_tanks[prop_list[0]].volume += vol
except IndexError: # in case line is empty
pass
# Apply fuel tank properties to the vehicle
vehicle = apply_properties(vehicle, fuel_tanks)
return vehicle
def vsp_read(tag, units_type='SI'):
"""This reads an OpenVSP vehicle geometry and writes it into a SUAVE vehicle format.
Includes wings, fuselages, and propellers.
Assumptions:
1. OpenVSP vehicle is composed of conventionally shaped fuselages, wings, and propellers.
1a. OpenVSP fuselage: generally narrow at nose and tail, wider in center).
1b. Fuselage is designed in VSP as it appears in real life. That is, the VSP model does not rely on
superficial elements such as canopies, stacks, or additional fuselages to cover up internal lofting oddities.
1c. This program will NOT account for multiple geometries comprising the fuselage. For example: a wingbox mounted beneath
is a separate geometry and will NOT be processed.
2. Fuselage origin is located at nose. VSP file origin can be located anywhere, preferably at the forward tip
of the vehicle or in front (to make all X-coordinates of vehicle positive).
3. Written for OpenVSP 3.16.1
Source:
N/A
Inputs:
1. A tag for an XML file in format .vsp3.
2. Units_type set to 'SI' (default) or 'Imperial'
Outputs:
Writes SUAVE vehicle with these geometries from VSP: (All values default to SI. Any other 2nd argument outputs Imperial.)
Wings.Wing. (* is all keys)
origin [m] in all three dimensions
spans.projected [m]
chords.root [m]
chords.tip [m]
aspect_ratio [-]
sweeps.quarter_chord [radians]
twists.root [radians]
twists.tip [radians]
thickness_to_chord [-]
dihedral [radians]
symmetric <boolean>
tag <string>
areas.exposed [m^2]
areas.reference [m^2]
areas.wetted [m^2]
Segments.
tag <string>
twist [radians]
percent_span_location [-] .1 is 10%
root_chord_percent [-] .1 is 10%
dihedral_outboard [radians]
sweeps.quarter_chord [radians]
thickness_to_chord [-]
airfoil <NACA 4-series, 6 series, or airfoil file>
Fuselages.Fuselage.
origin [m] in all three dimensions
width [m]
lengths.
total [m]
nose [m]
tail [m]
heights.
maximum [m]
at_quarter_length [m]
at_three_quarters_length [m]
effective_diameter [m]
fineness.nose [-] ratio of nose section length to fuselage effective diameter
fineness.tail [-] ratio of tail section length to fuselage effective diameter
areas.wetted [m^2]
tag <string>
segment[]. (segments are in ordered container and callable by number)
vsp.shape [point,circle,round_rect,general_fuse,fuse_file]
vsp.xsec_id <10 digit string>
percent_x_location
percent_z_location
height
width
length
effective_diameter
tag
vsp.xsec_num <integer of fuselage segment quantity>
vsp.xsec_surf_id <10 digit string>
Propellers.Propeller.
location[X,Y,Z] [radians]
rotation[X,Y,Z] [radians]
prop_attributes.tip_radius [m]
prop_attributes.hub_radius [m]
thrust_angle [radians]
Properties Used:
N/A
"""
vsp.ClearVSPModel()
vsp.ReadVSPFile(tag)
vsp_fuselages = []
vsp_wings = []
vsp_props = []
vsp_geoms = vsp.FindGeoms()
geom_names = []
vehicle = SUAVE.Vehicle()
vehicle.tag = tag
if units_type == 'SI':
units_type = 'SI'
else:
units_type = 'Imperial'
# The two for-loops below are in anticipation of an OpenVSP API update with a call for GETGEOMTYPE.
# This print function allows user to enter VSP GeomID manually as first argument in vsp_read functions.
print("VSP geometry IDs: ")
# Label each geom type by storing its VSP geom ID. (The API call for GETGEOMTYPE was not released as of 8/9/18, v 3.16.1)
for geom in vsp_geoms:
geom_name = vsp.GetGeomName(geom)
geom_names.append(geom_name)
print(str(geom_name) + ': ' + geom)
# -----------------------------
# MANUAL VSP ENTRY & PROCESSING
# -----------------------------
#fuselage = read_vsp_fuselage(fuselage_id, units_type=units_type) # Replace fuselage_id manually.
#vehicle.append_component(fuselage)
#wing = read_vsp_wing(wing_id, units_type=units_type) # Replace wing_id manually.
#vehicle.append_component(wing)
#prop = read_vsp_prop(prop_id, units_type=units_type) # Replace prop_id manually.
#vehicle.append_component(prop)
# --------------------------------
# AUTOMATIC VSP ENTRY & PROCESSING
# --------------------------------
#for geom in vsp_geoms:
#if vsp.GETGEOMTYPE(str(geom)) == 'FUSELAGE':
#vsp_fuselages.append(geom)
#if vsp.GETGEOMTYPE(str(geom)) == 'WING':
#vsp_wings.append(geom)
#if vsp.GETGEOMTYPE(str(geom)) == 'PROP':
#vsp_props.append(geom)
# Read VSP geoms and store in SUAVE components.
#for vsp_fuselage in vsp_fuselages:
#fuselage_id = vsp_fuselages[vsp_fuselage]
#fuselage = read_vsp_fuselage(fuselage_id, units_type)
#vehicle.append_component(fuselage)
#for vsp_wing in vsp_wings:
#wing_id = vsp_wings[vsp_wing]
#wing = read_vsp_wing(wing_id, units_type)
#vehicle.append_component(wing)
#for vsp_prop in vsp_props:
#prop_id = vsp_props[vsp_prop]
#prop = read_vsp_prop(prop_id, units_type)
#vehicle.append_component(prop)
return vehicle
def write_vsp_mesh(geometry, tag, half_mesh_flag, growth_ratio,
growth_limiting_flag):
"""This create an .stl surface mesh based on a vehicle stored in a .vsp3 file.
Assumptions:
None
Source:
N/A
Inputs:
geometry. - Also passed to set_sources
wings.main_wing.chords.mean_aerodynamic [m]
half_mesh_flag <boolean> determines if a symmetry plane is created
growth_ratio [-] growth ratio for the mesh
growth_limiting_flag <boolean> determines if 3D growth limiting is used
Outputs:
<tag>.stl
Properties Used:
N/A
"""
# Reset OpenVSP to avoid including a previous vehicle
vsp.ClearVSPModel()
# Turn on symmetry plane splitting to improve robustness of meshing process
if half_mesh_flag == True:
f = fileinput.input(tag + '.vsp3', inplace=1)
for line in f:
if 'SymmetrySplitting' in line:
print line[0:34] + '1' + line[35:-1]
else:
print line
vsp.ReadVSPFile(tag + '.vsp3')
# Set output file types and what will be meshed
file_type = vsp.CFD_STL_TYPE + vsp.CFD_KEY_TYPE
set_int = vsp.SET_ALL
vsp.SetComputationFileName(vsp.CFD_STL_TYPE, tag + '.stl')
vsp.SetComputationFileName(vsp.CFD_KEY_TYPE, tag + '.key')
# Set to create a tagged STL mesh file
vehicle_cont = vsp.FindContainer('Vehicle', 0)
STL_multi = vsp.FindParm(vehicle_cont, 'MultiSolid', 'STLSettings')
vsp.SetParmVal(STL_multi, 1.0)
vsp.SetCFDMeshVal(vsp.CFD_FAR_FIELD_FLAG, 1)
if half_mesh_flag == True:
vsp.SetCFDMeshVal(vsp.CFD_HALF_MESH_FLAG, 1)
# Figure out the size of the bounding box
vehicle_id = vsp.FindContainersWithName('Vehicle')[0]
xlen = vsp.GetParmVal(vsp.FindParm(vehicle_id, "X_Len", "BBox"))
ylen = vsp.GetParmVal(vsp.FindParm(vehicle_id, "Y_Len", "BBox"))
zlen = vsp.GetParmVal(vsp.FindParm(vehicle_id, "Z_Len", "BBox"))
# Max length
max_len = np.max([xlen, ylen, zlen])
far_length = 10. * max_len
vsp.SetCFDMeshVal(vsp.CFD_FAR_SIZE_ABS_FLAG, 1)
vsp.SetCFDMeshVal(vsp.CFD_FAR_LENGTH, far_length)
vsp.SetCFDMeshVal(vsp.CFD_FAR_WIDTH, far_length)
vsp.SetCFDMeshVal(vsp.CFD_FAR_HEIGHT, far_length)
vsp.SetCFDMeshVal(vsp.CFD_FAR_MAX_EDGE_LEN, max_len)
vsp.SetCFDMeshVal(vsp.CFD_GROWTH_RATIO, growth_ratio)
if growth_limiting_flag == True:
vsp.SetCFDMeshVal(vsp.CFD_LIMIT_GROWTH_FLAG, 1.0)
# Set the max edge length so we have on average 50 elements per chord length
MAC = geometry.wings.main_wing.chords.mean_aerodynamic
min_len = MAC / 50.
vsp.SetCFDMeshVal(vsp.CFD_MAX_EDGE_LEN, min_len)
# vsp.AddDefaultSources()
set_sources(geometry)
vsp.Update()
vsp.WriteVSPFile(tag + '_premesh.vsp3')
print 'Starting mesh for ' + tag + ' (This may take several minutes)'
ti = time.time()
vsp.ComputeCFDMesh(set_int, file_type)
tf = time.time()
dt = tf - ti
print 'VSP meshing for ' + tag + ' completed in ' + str(dt) + ' s'
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