# ==== Use Case 2 ====#
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)
def vsp_read_fuselage(fuselage_id, units_type='SI', fineness=True):
"""This reads an OpenVSP fuselage geometry and writes it to a SUAVE fuselage format.
Assumptions:
1. OpenVSP fuselage is "conventionally shaped" (generally narrow at nose and tail, wider in center).
2. 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.
3. 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.
4. 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).
5. Written for OpenVSP 3.16.1
Source:
N/A
Inputs:
0. Pre-loaded VSP vehicle in memory, via vsp_read.
1. VSP 10-digit geom ID for fuselage.
2. Units_type set to 'SI' (default) or 'Imperial'.
3. Boolean for whether or not to compute fuselage finenesses (default = True).
4. Uses exterior function get_vsp_areas, in SUAVE/trunk/SUAVE/Input_Output/OpenVSP.
Outputs:
Writes SUAVE fuselage, with these geometries: (all defaults are SI, but user may specify Imperial)
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>
Properties Used:
N/A
"""
fuselage = SUAVE.Components.Fuselages.Fuselage()
if units_type == 'SI':
units_factor = Units.meter * 1.
else:
units_factor = Units.foot * 1.
if vsp.GetGeomName(fuselage_id):
fuselage.tag = vsp.GetGeomName(fuselage_id)
else:
fuselage.tag = 'FuselageGeom'
fuselage.origin[0][0] = vsp.GetParmVal(fuselage_id, 'X_Location',
'XForm') * units_factor
fuselage.origin[0][1] = vsp.GetParmVal(fuselage_id, 'Y_Location',
'XForm') * units_factor
fuselage.origin[0][2] = vsp.GetParmVal(fuselage_id, 'Z_Location',
'XForm') * units_factor
fuselage.lengths.total = vsp.GetParmVal(fuselage_id, 'Length',
'Design') * units_factor
fuselage.vsp_data.xsec_surf_id = vsp.GetXSecSurf(
fuselage_id, 0) # There is only one XSecSurf in geom.
fuselage.vsp_data.xsec_num = vsp.GetNumXSec(
fuselage.vsp_data.xsec_surf_id) # Number of xsecs in fuselage.
x_locs = []
heights = []
widths = []
eff_diams = []
lengths = []
# -----------------
# Fuselage segments
# -----------------
for ii in range(0, fuselage.vsp_data.xsec_num):
segment = SUAVE.Components.Fuselages.Segment()
segment.vsp_data.xsec_id = vsp.GetXSec(fuselage.vsp_data.xsec_surf_id,
ii) # VSP XSec ID.
segment.tag = 'segment_' + str(ii)
segment.percent_x_location = vsp.GetParmVal(
fuselage_id, 'XLocPercent',
'XSec_' + str(ii)) # Along fuselage length.
segment.percent_z_location = vsp.GetParmVal(
fuselage_id, 'ZLocPercent',
'XSec_' + str(ii)) # Vertical deviation of fuselage center.
segment.height = vsp.GetXSecHeight(
segment.vsp_data.xsec_id) * units_factor
segment.width = vsp.GetXSecWidth(
segment.vsp_data.xsec_id) * units_factor
segment.effective_diameter = (segment.height + segment.width) / 2.
x_locs.append(segment.percent_x_location
) # Save into arrays for later computation.
heights.append(segment.height)
widths.append(segment.width)
eff_diams.append(segment.effective_diameter)
if ii != (
fuselage.vsp_data.xsec_num - 1
): # Segment length: stored as length since previous segment. (First segment will have length 0.0.)
segment.length = fuselage.lengths.total * (
fuselage.Segments[ii + 1].percent_x_location -
segment.percent_x_location) * units_factor
else:
segment.length = 0.0
lengths.append(segment.length)
shape = vsp.GetXSecShape(segment.vsp_data.xsec_id)
shape_dict = {
0: 'point',
1: 'circle',
2: 'ellipse',
3: 'super ellipse',
4: 'rounded rectangle',
5: 'general fuse',
6: 'fuse file'
}
segment.vsp_data.shape = shape_dict[shape]
fuselage.Segments.append(segment)
fuselage.heights.at_quarter_length = get_fuselage_height(
fuselage, .25) # Calls get_fuselage_height function (below).
fuselage.heights.at_three_quarters_length = get_fuselage_height(
fuselage, .75)
fuselage.heights.at_wing_root_quarter_chord = get_fuselage_height(
fuselage, .4)
fuselage.heights.maximum = max(heights) # Max segment height.
fuselage.width = max(widths) # Max segment width.
fuselage.effective_diameter = max(eff_diams) # Max segment effective diam.
fuselage.areas.front_projected = np.pi * (
(fuselage.effective_diameter) / 2)**2
eff_diam_gradients_fwd = np.array(eff_diams[1:]) - np.array(
eff_diams[:-1]) # Compute gradients of segment effective diameters.
eff_diam_gradients_fwd = np.multiply(eff_diam_gradients_fwd, lengths[:-1])
fuselage = compute_fuselage_fineness(fuselage, x_locs, eff_diams,
eff_diam_gradients_fwd)
return fuselage
def vsp_read_wing(wing_id, units_type='SI'):
"""This reads an OpenVSP wing vehicle geometry and writes it into a SUAVE wing format.
Assumptions:
1. OpenVSP wing is divided into segments ("XSecs" in VSP).
2. Written for OpenVSP 3.16.1
Source:
N/A
Inputs:
0. Pre-loaded VSP vehicle in memory, via vsp_read.
1. VSP 10-digit geom ID for wing.
2. units_type set to 'SI' (default) or 'Imperial'.
Outputs:
Writes SUAVE wing object, with these geometries, from VSP:
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>
Properties Used:
N/A
"""
if units_type == 'SI':
units_factor = Units.meter * 1.
else:
units_factor = Units.foot * 1.
wing = SUAVE.Components.Wings.Wing()
if vsp.GetGeomName(wing_id):
wing.tag = vsp.GetGeomName(wing_id)
else:
wing.tag = 'WingGeom'
wing.origin[0] = vsp.GetParmVal(wing_id, 'X_Location',
'XForm') * units_factor
wing.origin[1] = vsp.GetParmVal(wing_id, 'Y_Location',
'XForm') * units_factor
wing.origin[2] = vsp.GetParmVal(wing_id, 'Z_Location',
'XForm') * units_factor
sym_planar = vsp.GetParmVal(wing_id, 'Sym_Planar_Flag', 'Sym')
sym_origin = vsp.GetParmVal(wing_id, 'Sym_Ancestor', 'Sym')
# Get the initial rotation to get the dihedral angles
x_rot = vsp.GetParmVal(wing_id, 'X_Rotation', 'XForm')
# Check if this is vertical tail and get the rotation
if x_rot >= 70:
wing.vertical = True
if sym_planar == 2. and sym_origin == 2.: #origin at wing, not vehicle
wing.symmetric == True
else:
wing.symmetric == False
wing.aspect_ratio = vsp.GetParmVal(wing_id, 'TotalAR', 'WingGeom')
xsec_surf_id = vsp.GetXSecSurf(wing_id,
0) # This is how VSP stores surfaces.
segment_num = vsp.GetNumXSec(
xsec_surf_id
) # Get number of wing segments (is one more than the VSP GUI shows).
total_chord = vsp.GetParmVal(wing_id, 'Root_Chord',
'XSec_1') * units_factor
total_proj_span = vsp.GetParmVal(wing_id, 'TotalProjectedSpan',
'WingGeom') * units_factor
span_sum = 0. # Non-projected.
proj_span_sum = 0. # Projected.
segment_spans = [None] * (segment_num) # Non-projected.
segment_dihedral = [None] * (segment_num)
segment_sweeps_quarter_chord = [None] * (segment_num)
# Check for wing segment *inside* fuselage, then skip XSec_0 to start at first exposed segment.
if vsp.GetParmVal(wing_id, 'Root_Chord', 'XSec_0') == 1.:
start = 1
else:
start = 0
# -------------
# Wing segments
# -------------
# Convert VSP XSecs to SUAVE segments. (Wing segments are defined by outboard sections in VSP, but inboard sections in SUAVE.)
for i in range(start, segment_num + 1):
segment = SUAVE.Components.Wings.Segment()
segment.tag = 'Section_' + str(i)
thick_cord = vsp.GetParmVal(wing_id, 'ThickChord',
'XSecCurve_' + str(i - 1))
segment.thickness_to_chord = thick_cord # Thick_cord stored for use in airfoil, below.
segment_root_chord = vsp.GetParmVal(wing_id, 'Root_Chord',
'XSec_' + str(i)) * units_factor
segment.root_chord_percent = segment_root_chord / total_chord
segment.percent_span_location = proj_span_sum / (total_proj_span / 2)
segment.twist = vsp.GetParmVal(wing_id, 'Twist',
'XSec_' + str(i - 1)) * Units.deg
if i == start:
wing.thickness_to_chord = thick_cord
if i < segment_num: # This excludes the tip xsec, but we need a segment in SUAVE to store airfoil.
sweep = vsp.GetParmVal(wing_id, 'Sweep',
'XSec_' + str(i)) * Units.deg
sweep_loc = vsp.GetParmVal(wing_id, 'Sweep_Location',
'XSec_' + str(i))
AR = vsp.GetParmVal(wing_id, 'Aspect', 'XSec_' + str(i))
taper = vsp.GetParmVal(wing_id, 'Taper', 'XSec_' + str(i))
segment_sweeps_quarter_chord[i] = convert_sweep(
sweep, sweep_loc, 0.25, AR, taper)
segment.sweeps.quarter_chord = segment_sweeps_quarter_chord[
i] # Used again, below
# Used for dihedral computation, below.
segment_dihedral[i] = vsp.GetParmVal(wing_id, 'Dihedral',
'XSec_' + str(i)) * Units.deg
segment.dihedral_outboard = segment_dihedral[i]
segment_spans[i] = vsp.GetParmVal(wing_id, 'Span',
'XSec_' + str(i)) * units_factor
proj_span_sum += segment_spans[i] * np.cos(segment_dihedral[i])
span_sum += segment_spans[i]
else:
segment.root_chord_percent = (vsp.GetParmVal(
wing_id, 'Tip_Chord',
'XSec_' + str(i - 1))) * units_factor / total_chord
# XSec airfoil
jj = i - 1 # Airfoil index i-1 because VSP airfoils and sections are one index off relative to SUAVE.
xsec_id = str(vsp.GetXSec(xsec_surf_id, jj))
airfoil = Airfoil()
if vsp.GetXSecShape(
xsec_id) == vsp.XS_FOUR_SERIES: # XSec shape: NACA 4-series
camber = vsp.GetParmVal(wing_id, 'Camber', 'XSecCurve_' + str(jj))
if camber == 0.:
camber_loc = 0.
else:
camber_loc = vsp.GetParmVal(wing_id, 'CamberLoc',
'XSecCurve_' + str(jj))
airfoil.thickness_to_chord = thick_cord
camber_round = int(np.around(camber * 100))
camber_loc_round = int(np.around(camber_loc * 10))
thick_cord_round = int(np.around(thick_cord * 100))
airfoil.tag = 'NACA ' + str(camber_round) + str(
camber_loc_round) + str(thick_cord_round)
elif vsp.GetXSecShape(
xsec_id) == vsp.XS_SIX_SERIES: # XSec shape: NACA 6-series
thick_cord_round = int(np.around(thick_cord * 100))
a_value = vsp.GetParmVal(wing_id, 'A', 'XSecCurve_' + str(jj))
ideal_CL = int(
np.around(
vsp.GetParmVal(wing_id, 'IdealCl', 'XSecCurve_' + str(jj))
* 10))
series_vsp = int(
vsp.GetParmVal(wing_id, 'Series', 'XSecCurve_' + str(jj)))
series_dict = {
0: '63',
1: '64',
2: '65',
3: '66',
4: '67',
5: '63A',
6: '64A',
7: '65A'
} # VSP series values.
series = series_dict[series_vsp]
airfoil.tag = 'NACA ' + series + str(ideal_CL) + str(
thick_cord_round) + ' a=' + str(np.around(a_value, 1))
elif vsp.GetXSecShape(
xsec_id
) == vsp.XS_FILE_AIRFOIL: # XSec shape: 12 is type AF_FILE
airfoil.thickness_to_chord = thick_cord
airfoil.points = vsp.GetAirfoilCoordinates(wing_id,
float(jj / segment_num))
# VSP airfoil API calls get coordinates and write files with the final argument being the fraction of segment position, regardless of relative spans.
# (Write the root airfoil with final arg = 0. Write 4th airfoil of 5 segments with final arg = .8)
vsp.WriteSeligAirfoil(
str(wing.tag) + '_airfoil_XSec_' + str(jj) + '.dat', wing_id,
float(jj / segment_num))
airfoil.coordinate_file = 'str(wing.tag)' + '_airfoil_XSec_' + str(
jj) + '.dat'
airfoil.tag = 'AF_file'
segment.append_airfoil(airfoil)
wing.Segments.append(segment)
# Wing dihedral
proj_span_sum_alt = 0.
span_sum_alt = 0.
sweeps_sum = 0.
for ii in range(start, segment_num):
span_sum_alt += segment_spans[ii]
proj_span_sum_alt += segment_spans[ii] * np.cos(
segment_dihedral[ii]
) # Use projected span to find total wing dihedral.
sweeps_sum += segment_spans[ii] * np.tan(
segment_sweeps_quarter_chord[ii])
wing.dihedral = np.arccos(proj_span_sum_alt / span_sum_alt)
wing.sweeps.quarter_chord = -np.arctan(
sweeps_sum / span_sum_alt) # Minus sign makes it positive sweep.
# Wing spans
if wing.symmetric == True:
wing.spans.projected = 2 * proj_span_sum
else:
wing.spans.projected = proj_span_sum
# Areas
wing.areas.reference = vsp.GetParmVal(wing_id, 'TotalArea',
'WingGeom') * units_factor**2
# Chords
wing.chords.root = vsp.GetParmVal(wing_id, 'Tip_Chord',
'XSec_0') * units_factor
wing.chords.tip = vsp.GetParmVal(
wing_id, 'Tip_Chord', 'XSec_' + str(segment_num - 1)) * units_factor
wing.chords.mean_geometric = wing.areas.reference / wing.spans.projected
# Twists
wing.twists.root = vsp.GetParmVal(wing_id, 'Twist', 'XSec_0') * Units.deg
wing.twists.tip = vsp.GetParmVal(
wing_id, 'Twist', 'XSec_' + str(segment_num - 1)) * Units.deg
return wing
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