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 3.21.1
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
"""
# Check if this is vertical tail, this seems like a weird first step but it's necessary
# Get the initial rotation to get the dihedral angles
x_rot = vsp.GetParmVal(wing_id, 'X_Rotation', 'XForm')
if x_rot >= 70:
wing = SUAVE.Components.Wings.Vertical_Tail()
wing.vertical = True
x_rot = (90 - x_rot) * Units.deg
else:
# Instantiate a wing
wing = SUAVE.Components.Wings.Wing()
# Set the units
if units_type == 'SI':
units_factor = Units.meter * 1.
else:
units_factor = Units.foot * 1.
# Apply a tag to the wing
if vsp.GetGeomName(wing_id):
tag = vsp.GetGeomName(wing_id)
tag = tag.translate(t_table)
wing.tag = tag
else:
wing.tag = 'winggeom'
# Top level wing parameters
# Wing origin
wing.origin[0][0] = vsp.GetParmVal(wing_id, 'X_Location',
'XForm') * units_factor
wing.origin[0][1] = vsp.GetParmVal(wing_id, 'Y_Location',
'XForm') * units_factor
wing.origin[0][2] = vsp.GetParmVal(wing_id, 'Z_Location',
'XForm') * units_factor
# Wing Symmetry
sym_planar = vsp.GetParmVal(wing_id, 'Sym_Planar_Flag', 'Sym')
sym_origin = vsp.GetParmVal(wing_id, 'Sym_Ancestor', 'Sym')
# Check for symmetry
if sym_planar == 2. and sym_origin == 1.: #origin at wing, not vehicle
wing.symmetric = True
else:
wing.symmetric = False
#More top level parameters
total_proj_span = vsp.GetParmVal(wing_id, 'TotalProjectedSpan',
'WingGeom') * units_factor
wing.aspect_ratio = vsp.GetParmVal(wing_id, 'TotalAR', 'WingGeom')
wing.areas.reference = vsp.GetParmVal(wing_id, 'TotalArea',
'WingGeom') * units_factor**2
wing.spans.projected = total_proj_span
# Check if this is a single segment wing
xsec_surf_id = vsp.GetXSecSurf(wing_id,
0) # This is how VSP stores surfaces.
x_sec_1 = vsp.GetXSec(xsec_surf_id, 1)
x_sec_1_span_parm = vsp.GetXSecParm(x_sec_1, 'Span')
x_sec_1_span = vsp.GetParmVal(x_sec_1_span_parm) * (
1 + wing.symmetric) * units_factor
if x_sec_1_span == wing.spans.projected:
single_seg = True
else:
single_seg = False
segment_num = vsp.GetNumXSec(
xsec_surf_id
) # Get number of wing segments (is one more than the VSP GUI shows).
x_sec = vsp.GetXSec(xsec_surf_id, 0)
chord_parm = vsp.GetXSecParm(x_sec, 'Root_Chord')
total_chord = vsp.GetParmVal(chord_parm)
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 total_chord == 1.:
start = 1
xsec_surf_id = vsp.GetXSecSurf(wing_id, 1)
x_sec = vsp.GetXSec(xsec_surf_id, 0)
chord_parm = vsp.GetXSecParm(x_sec, 'Tip_Chord')
root_chord = vsp.GetParmVal(chord_parm) * units_factor
else:
start = 0
root_chord = total_chord * units_factor
# -------------
# Wing segments
# -------------
if single_seg == False:
# 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 / root_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 + x_rot
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.
# Add a tip segment, all values are zero except the tip chord
tc = vsp.GetParmVal(wing_id, 'Tip_Chord',
'XSec_' + str(segment_num - 1)) * units_factor
segment = SUAVE.Components.Wings.Segment()
segment.percent_span_location = 1.0
segment.root_chord_percent = tc / root_chord
# Chords
wing.chords.root = vsp.GetParmVal(wing_id, 'Tip_Chord',
'XSec_0') * units_factor
wing.chords.tip = tc
wing.chords.mean_geometric = wing.areas.reference / wing.spans.projected
# Just double calculate and fix things:
wing = wing_segmented_planform(wing)
else:
# Single segment
# Get ID's
x_sec_1_dih_parm = vsp.GetXSecParm(x_sec_1, 'Dihedral')
x_sec_1_sweep_parm = vsp.GetXSecParm(x_sec_1, 'Sweep')
x_sec_1_sweep_loc_parm = vsp.GetXSecParm(x_sec_1, 'Sweep_Location')
x_sec_1_taper_parm = vsp.GetXSecParm(x_sec_1, 'Taper')
x_sec_1_rc_parm = vsp.GetXSecParm(x_sec_1, 'Root_Chord')
x_sec_1_tc_parm = vsp.GetXSecParm(x_sec_1, 'Tip_Chord')
# Calcs
sweep = vsp.GetParmVal(x_sec_1_sweep_parm) * Units.deg
sweep_loc = vsp.GetParmVal(x_sec_1_sweep_loc_parm)
taper = vsp.GetParmVal(x_sec_1_taper_parm)
c_4_sweep = convert_sweep(sweep, sweep_loc, 0.25, wing.aspect_ratio,
taper)
# Pull and pack
wing.sweeps.quarter_chord = c_4_sweep
wing.taper = taper
wing.dihedral = vsp.GetParmVal(x_sec_1_dih_parm) * Units.deg + x_rot
wing.chords.root = vsp.GetParmVal(x_sec_1_rc_parm) * units_factor
wing.chords.tip = vsp.GetParmVal(x_sec_1_tc_parm) * units_factor
wing.chords.mean_geometric = wing.areas.reference / wing.spans.projected
# Just double calculate and fix things:
wing = wing_planform(wing)
# 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
# ==== 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 write_vsp_turbofan(turbofan, OML_set_ind):
"""This converts turbofans into OpenVSP format.
Assumptions:
None
Source:
N/A
Inputs:
turbofan.
number_of_engines [-]
engine_length [m]
nacelle_diameter [m]
origin [m] in all three dimension, should have as many origins as engines
OpenVSP_flow_through <boolean> if True create a flow through nacelle, if False create a cylinder
Outputs:
Operates on the active OpenVSP model, no direct output
Properties Used:
N/A
"""
n_engines = turbofan.number_of_engines
length = turbofan.engine_length
width = turbofan.nacelle_diameter
origins = turbofan.origin
inlet_width = turbofan.inlet_diameter
tf_tag = turbofan.tag
# True will create a flow-through subsonic nacelle (which may have dimensional errors)
# False will create a cylindrical stack (essentially a cylinder)
ft_flag = turbofan.OpenVSP_flow_through
import operator # import here since engines are not always needed
# sort engines per left to right convention
origins_sorted = sorted(origins, key=operator.itemgetter(1))
for ii in range(0,int(n_engines)):
origin = origins_sorted[ii]
x = origin[0]
y = origin[1]
z = origin[2]
if ft_flag == True:
nac_id = vsp.AddGeom( "BODYOFREVOLUTION")
vsp.SetGeomName(nac_id, tf_tag+'_'+str(ii+1))
# Origin
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,'Abs_Or_Relitive_flag','XForm',vsp.ABS) # misspelling from OpenVSP
# Length and overall diameter
vsp.SetParmVal(nac_id,"Diameter","Design",inlet_width)
vsp.ChangeBORXSecShape(nac_id ,vsp.XS_SUPER_ELLIPSE)
vsp.Update()
vsp.SetParmVal(nac_id, "Super_Height", "XSecCurve", (width-inlet_width)/2)
vsp.SetParmVal(nac_id, "Super_Width", "XSecCurve", length)
vsp.SetParmVal(nac_id, "Super_MaxWidthLoc", "XSecCurve", -1.)
vsp.SetParmVal(nac_id, "Super_M", "XSecCurve", 2.)
vsp.SetParmVal(nac_id, "Super_N", "XSecCurve", 1.)
else:
nac_id = vsp.AddGeom("STACK")
vsp.SetGeomName(nac_id, tf_tag+'_'+str(ii+1))
# Origin
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,'Abs_Or_Relitive_flag','XForm',vsp.ABS) # misspelling from OpenVSP
vsp.SetParmVal(nac_id,'Origin','XForm',0.5)
vsp.CutXSec(nac_id,2) # remove extra default subsurface
xsecsurf = vsp.GetXSecSurf(nac_id,0)
vsp.ChangeXSecShape(xsecsurf,1,vsp.XS_CIRCLE)
vsp.ChangeXSecShape(xsecsurf,2,vsp.XS_CIRCLE)
vsp.Update()
vsp.SetParmVal(nac_id, "Circle_Diameter", "XSecCurve_1", width)
vsp.SetParmVal(nac_id, "Circle_Diameter", "XSecCurve_2", width)
vsp.SetParmVal(nac_id, "XDelta", "XSec_1", 0)
vsp.SetParmVal(nac_id, "XDelta", "XSec_2", length)
vsp.SetParmVal(nac_id, "XDelta", "XSec_3", 0)
vsp.SetSetFlag(nac_id, OML_set_ind, True)
vsp.Update()
def write_vsp_fuselage(fuselage,area_tags, main_wing, fuel_tank_set_ind, OML_set_ind):
"""This writes a fuselage into OpenVSP format.
Assumptions:
None
Source:
N/A
Inputs:
fuselage
width [m]
lengths.total [m]
heights.
maximum [m]
at_quarter_length [m]
at_wing_root_quarter_chord [m]
at_three_quarters_length [m]
effective_diameter [m]
fineness.nose [-] ratio of nose section length to fuselage width
fineness.tail [-] ratio of tail section length to fuselage width
tag <string>
OpenVSP_values. (optional)
nose.top.angle [degrees]
nose.top.strength [-] this determines how much the specified angle influences that shape
nose.side.angle [degrees]
nose.side.strength [-]
nose.TB_Sym <boolean> determines if top angle is mirrored on bottom
nose.z_pos [-] z position of the nose as a percentage of fuselage length (.1 is 10%)
tail.top.angle [degrees]
tail.top.strength [-]
tail.z_pos (optional, 0.02 default) [-] z position of the tail as a percentage of fuselage length (.1 is 10%)
Segments. (optional)
width [m]
height [m]
percent_x_location [-] .1 is 10% length
percent_z_location [-] .1 is 10% length
area_tags <dict> used to keep track of all tags needed in wetted area computation
main_wing.origin [m]
main_wing.chords.root [m]
fuel_tank_set_index <int> OpenVSP object set containing the fuel tanks
Outputs:
Operates on the active OpenVSP model, no direct output
Properties Used:
N/A
"""
num_segs = len(fuselage.Segments)
length = fuselage.lengths.total
fuse_x = fuselage.origin[0][0]
fuse_y = fuselage.origin[0][1]
fuse_z = fuselage.origin[0][2]
fuse_x_rotation = fuselage.x_rotation
fuse_y_rotation = fuselage.y_rotation
fuse_z_rotation = fuselage.z_rotation
if num_segs==0: # SUAVE default fuselage shaping
width = fuselage.width
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
try:
if main_wing != None:
w_origin = main_wing.origin
w_c_4 = main_wing.chords.root/4.
else:
w_origin = 0.5*length
w_c_4 = 0.5*length
except AttributeError:
raise AttributeError('Main wing not detected. Fuselage must have specified sections in this configuration.')
# 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][0]+w_c_4)/length
x3 = 1-t_fine*width/length
end_ind = 4
else: # Fuselage shaping based on sections
widths = []
heights = []
x_poses = []
z_poses = []
segs = fuselage.Segments
for seg in segs:
widths.append(seg.width)
heights.append(seg.height)
x_poses.append(seg.percent_x_location)
z_poses.append(seg.percent_z_location)
end_ind = num_segs-1
fuse_id = vsp.AddGeom("FUSELAGE")
vsp.SetGeomName(fuse_id, fuselage.tag)
area_tags[fuselage.tag] = ['fuselages',fuselage.tag]
tail_z_pos = 0.02 # default value
# set fuselage relative location and rotation
vsp.SetParmVal( fuse_id,'X_Rel_Rotation','XForm',fuse_x_rotation)
vsp.SetParmVal( fuse_id,'Y_Rel_Rotation','XForm',fuse_y_rotation)
vsp.SetParmVal( fuse_id,'Z_Rel_Rotation','XForm',fuse_z_rotation)
vsp.SetParmVal( fuse_id,'X_Rel_Location','XForm',fuse_x)
vsp.SetParmVal( fuse_id,'Y_Rel_Location','XForm',fuse_y)
vsp.SetParmVal( fuse_id,'Z_Rel_Location','XForm',fuse_z)
if 'OpenVSP_values' in fuselage:
vals = fuselage.OpenVSP_values
# for wave drag testing
fuselage.OpenVSP_ID = fuse_id
# 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)
if not vals.nose.TB_Sym:
vsp.SetParmVal(fuse_id,"BottomLAngle","XSec_0",vals.nose.bottom.angle)
vsp.SetParmVal(fuse_id,"BottomLStrength","XSec_0",vals.nose.bottom.strength)
# Tail
# 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 'z_pos' in vals.tail:
tail_z_pos = vals.tail.z_pos
else:
pass # use above default
if num_segs == 0:
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);
else:
# OpenVSP vals do not exist:
vals = Data()
vals.nose = Data()
vals.tail = Data()
vals.tail.top = Data()
vals.nose.z_pos = 0.0
vals.tail.top.angle = 0.0
vals.tail.top.strength = 0.0
if len(np.unique(x_poses)) != len(x_poses):
raise ValueError('Duplicate fuselage section positions detected.')
vsp.SetParmVal(fuse_id,"Length","Design",length)
if num_segs != 5: # reduce to only nose and tail
vsp.CutXSec(fuse_id,1) # remove extra default section
vsp.CutXSec(fuse_id,1) # remove extra default section
vsp.CutXSec(fuse_id,1) # remove extra default section
for i in range(num_segs-2): # add back the required number of sections
vsp.InsertXSec(fuse_id, 0, vsp.XS_ELLIPSE)
vsp.Update()
for i in range(num_segs-2):
# Bunch sections to allow proper length settings in the next step
# This is necessary because OpenVSP will not move a section past an adjacent section
vsp.SetParmVal(fuse_id, "XLocPercent", "XSec_"+str(i+1),1e-6*(i+1))
vsp.Update()
if x_poses[1] < (num_segs-2)*1e-6:
print('Warning: Second fuselage section is too close to the nose. OpenVSP model may not be accurate.')
for i in reversed(range(num_segs-2)):
# order is reversed because sections are initially bunched in the front and cannot be extended passed the next
vsp.SetParmVal(fuse_id, "XLocPercent", "XSec_"+str(i+1),x_poses[i+1])
vsp.SetParmVal(fuse_id, "ZLocPercent", "XSec_"+str(i+1),z_poses[i+1])
vsp.SetParmVal(fuse_id, "Ellipse_Width", "XSecCurve_"+str(i+1), widths[i+1])
vsp.SetParmVal(fuse_id, "Ellipse_Height", "XSecCurve_"+str(i+1), heights[i+1])
vsp.Update()
set_section_angles(i, vals.nose.z_pos, tail_z_pos, x_poses, z_poses, heights, widths,length,end_ind,fuse_id)
vsp.SetParmVal(fuse_id, "XLocPercent", "XSec_"+str(0),x_poses[0])
vsp.SetParmVal(fuse_id, "ZLocPercent", "XSec_"+str(0),z_poses[0])
vsp.SetParmVal(fuse_id, "XLocPercent", "XSec_"+str(end_ind),x_poses[-1])
vsp.SetParmVal(fuse_id, "ZLocPercent", "XSec_"+str(end_ind),z_poses[-1])
# Tail
if heights[-1] > 0.:
stdout = vsp.cvar.cstdout
errorMgr = vsp.ErrorMgrSingleton_getInstance()
errorMgr.PopErrorAndPrint(stdout)
pos = len(heights)-1
vsp.InsertXSec(fuse_id, pos-1, vsp.XS_ELLIPSE)
vsp.Update()
vsp.SetParmVal(fuse_id, "Ellipse_Width", "XSecCurve_"+str(pos), widths[-1])
vsp.SetParmVal(fuse_id, "Ellipse_Height", "XSecCurve_"+str(pos), heights[-1])
vsp.SetParmVal(fuse_id, "XLocPercent", "XSec_"+str(pos),x_poses[-1])
vsp.SetParmVal(fuse_id, "ZLocPercent", "XSec_"+str(pos),z_poses[-1])
xsecsurf = vsp.GetXSecSurf(fuse_id,0)
vsp.ChangeXSecShape(xsecsurf,pos+1,vsp.XS_POINT)
vsp.Update()
vsp.SetParmVal(fuse_id, "XLocPercent", "XSec_"+str(pos+1),x_poses[-1])
vsp.SetParmVal(fuse_id, "ZLocPercent", "XSec_"+str(pos+1),z_poses[-1])
# update strengths to make end flat
vsp.SetParmVal(fuse_id,"TopRStrength","XSec_"+str(pos), 0.)
vsp.SetParmVal(fuse_id,"RightRStrength","XSec_"+str(pos), 0.)
vsp.SetParmVal(fuse_id,"BottomRStrength","XSec_"+str(pos), 0.)
vsp.SetParmVal(fuse_id,"TopLStrength","XSec_"+str(pos+1), 0.)
vsp.SetParmVal(fuse_id,"RightLStrength","XSec_"+str(pos+1), 0.)
else:
vsp.SetParmVal(fuse_id,"TopLAngle","XSec_"+str(end_ind),vals.tail.top.angle)
vsp.SetParmVal(fuse_id,"TopLStrength","XSec_"+str(end_ind),vals.tail.top.strength)
vsp.SetParmVal(fuse_id,"AllSym","XSec_"+str(end_ind),1)
vsp.Update()
if 'z_pos' in vals.tail:
tail_z_pos = vals.tail.z_pos
else:
pass # use above default
if 'Fuel_Tanks' in fuselage:
for tank in fuselage.Fuel_Tanks:
write_fuselage_conformal_fuel_tank(fuse_id, tank, fuel_tank_set_ind)
vsp.SetSetFlag(fuse_id, OML_set_ind, True)
return area_tags
def write_vsp_wing(wing, area_tags, fuel_tank_set_ind, OML_set_ind):
"""This write a given wing into OpenVSP format
Assumptions:
If wing segments are defined, they must cover the full span.
(may work in some other cases, but functionality will not be maintained)
Source:
N/A
Inputs:
wing.
origin [m] in all three dimensions
spans.projected [m]
chords.root [m]
chords.tip [m]
sweeps.quarter_chord [radians]
twists.root [radians]
twists.tip [radians]
thickness_to_chord [-]
dihedral [radians]
tag <string>
Segments.*. (optional)
twist [radians]
percent_span_location [-] .1 is 10%
root_chord_percent [-] .1 is 10%
dihedral_outboard [radians]
sweeps.quarter_chord [radians]
thickness_to_chord [-]
area_tags <dict> used to keep track of all tags needed in wetted area computation
fuel_tank_set_index <int> OpenVSP object set containing the fuel tanks
Outputs:
area_tags <dict> used to keep track of all tags needed in wetted area computation
wing_id <str> OpenVSP ID for given wing
Properties Used:
N/A
"""
wing_x = wing.origin[0][0]
wing_y = wing.origin[0][1]
wing_z = wing.origin[0][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 range(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)
dihedral = -dihedral # check for vertical tail, direction reverses from SUAVE/AVL
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
if n_segments != 0:
if np.isclose(wing.Segments[0].percent_span_location,0.):
vsp.SetParmVal( wing_id,'Twist',x_secs[0],wing.Segments[0].twist / Units.deg) # root
else:
vsp.SetParmVal( wing_id,'Twist',x_secs[0],root_twist) # root
# The tips should write themselves
else:
vsp.SetParmVal( wing_id,'Twist',x_secs[0],root_twist) # root
vsp.SetParmVal( wing_id,'Twist',x_secs[1],tip_twist) # tip
# 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
airfoil_vsp_types = []
if n_segments > 0:
for i in range(n_segments):
if 'airfoil_type' in wing.Segments[i].keys():
if wing.Segments[i].airfoil_type == 'biconvex':
airfoil_vsp_types.append(vsp.XS_BICONVEX)
else:
airfoil_vsp_types.append(vsp.XS_FILE_AIRFOIL)
else:
airfoil_vsp_types.append(vsp.XS_FILE_AIRFOIL)
elif 'airfoil_type' in wing.keys():
if wing.airfoil_type == 'biconvex':
airfoil_vsp_types.append(vsp.XS_BICONVEX)
else:
airfoil_vsp_types.append(vsp.XS_FILE_AIRFOIL)
else:
airfoil_vsp_types = [vsp.XS_FILE_AIRFOIL]
if n_segments==0:
if len(wing.Airfoil) != 0 or 'airfoil_type' in wing.keys():
xsecsurf = vsp.GetXSecSurf(wing_id,0)
vsp.ChangeXSecShape(xsecsurf,0,airfoil_vsp_types[0])
vsp.ChangeXSecShape(xsecsurf,1,airfoil_vsp_types[0])
if len(wing.Airfoil) != 0:
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:
if len(wing.Segments[0].Airfoil) != 0 or 'airfoil_type' in wing.Segments[0].keys():
xsecsurf = vsp.GetXSecSurf(wing_id,0)
vsp.ChangeXSecShape(xsecsurf,0,airfoil_vsp_types[0])
vsp.ChangeXSecShape(xsecsurf,1,airfoil_vsp_types[0])
if len(wing.Segments[0].Airfoil) != 0:
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/np.cos(dihedral*Units.degrees))
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 range(1,n_segments+1):
if (wing.Segments[i_segs-1] == wing.Segments[-1]) and (wing.Segments[-1].percent_span_location == 1.):
break
# Unpack
dihedral_i = wing.Segments[i_segs-1].dihedral_outboard / Units.deg
chord_i = root_chord*wing.Segments[i_segs-1].root_chord_percent
try:
twist_i = wing.Segments[i_segs].twist / Units.deg
no_twist_flag = False
except:
no_twist_flag = True
sweep_i = wing.Segments[i_segs-1].sweeps.quarter_chord / Units.deg
tc_i = wing.Segments[i_segs-1].thickness_to_chord
# 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 or 'airfoil_type' in wing.Segments[i_segs-1].keys():
vsp.InsertXSec(wing_id,i_segs-1+adjust,airfoil_vsp_types[i_segs-1])
if len(wing.Segments[i_segs-1].Airfoil) != 0:
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)
if not no_twist_flag:
vsp.SetParmVal( wing_id,'Twist',x_secs[i_segs+adjust],twist_i)
vsp.SetParmVal( wing_id,'ThickChord',x_sec_curves[i_segs+adjust],tc_i)
if adjust and (i_segs == 1):
vsp.Update()
vsp.SetParmVal( wing_id,'Twist',x_secs[1],wing.Segments[i_segs-1].twist / Units.deg)
vsp.Update()
if (n_segments != 0) and (wing.Segments[-1].percent_span_location == 1.):
tip_chord = root_chord*wing.Segments[-1].root_chord_percent
vsp.SetParmVal( wing_id,'Tip_Chord',x_secs[n_segments-1+adjust],tip_chord)
vsp.SetParmVal( wing_id,'ThickChord',x_sec_curves[n_segments-1+adjust],wing.Segments[-1].thickness_to_chord)
# twist is set in the normal loop
else:
vsp.SetParmVal( wing_id,'Tip_Chord',x_secs[-1-(1-adjust)],tip_chord)
vsp.SetParmVal( wing_id,'Twist',x_secs[-1-(1-adjust)],tip_twist)
# a single trapezoidal wing is assumed to have constant thickness to chord
vsp.Update()
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 'Fuel_Tanks' in wing:
for tank in wing.Fuel_Tanks:
write_wing_conformal_fuel_tank(wing, wing_id, tank, fuel_tank_set_ind)
vsp.SetSetFlag(wing_id, OML_set_ind, True)
return area_tags, wing_id
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 read_vsp_wing(wing_id, units_type='SI',write_airfoil_file=True):
"""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 3.21.1
Source:
N/A
Inputs:
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.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
"""
# Check if this is vertical tail, this seems like a weird first step but it's necessary
# Get the initial rotation to get the dihedral angles
x_rot = vsp.GetParmVal( wing_id,'X_Rotation','XForm')
if x_rot >=70:
wing = SUAVE.Components.Wings.Vertical_Tail()
wing.vertical = True
x_rot = (90-x_rot) * Units.deg
else:
# Instantiate a wing
wing = SUAVE.Components.Wings.Wing()
x_rot = x_rot * Units.deg
# Set the units
if units_type == 'SI':
units_factor = Units.meter * 1.
elif units_type == 'imperial':
units_factor = Units.foot * 1.
elif units_type == 'inches':
units_factor = Units.inch * 1.
# Apply a tag to the wing
if vsp.GetGeomName(wing_id):
tag = vsp.GetGeomName(wing_id)
tag = tag.translate(t_table)
wing.tag = tag
else:
wing.tag = 'winggeom'
scaling = vsp.GetParmVal(wing_id, 'Scale', 'XForm')
units_factor = units_factor*scaling
# Top level wing parameters
# Wing origin
wing.origin[0][0] = vsp.GetParmVal(wing_id, 'X_Location', 'XForm') * units_factor
wing.origin[0][1] = vsp.GetParmVal(wing_id, 'Y_Location', 'XForm') * units_factor
wing.origin[0][2] = vsp.GetParmVal(wing_id, 'Z_Location', 'XForm') * units_factor
# Wing Symmetry
sym_planar = vsp.GetParmVal(wing_id, 'Sym_Planar_Flag', 'Sym')
sym_origin = vsp.GetParmVal(wing_id, 'Sym_Ancestor_Origin_Flag', 'Sym')
# Check for symmetry
if sym_planar == 2. and sym_origin == 1.: #origin at wing, not vehicle
wing.symmetric = True
else:
wing.symmetric = False
#More top level parameters
total_proj_span = vsp.GetParmVal(wing_id, 'TotalProjectedSpan', 'WingGeom') * units_factor
wing.aspect_ratio = vsp.GetParmVal(wing_id, 'TotalAR', 'WingGeom')
wing.areas.reference = vsp.GetParmVal(wing_id, 'TotalArea', 'WingGeom') * units_factor**2
wing.spans.projected = total_proj_span
# Check if this is a single segment wing
xsec_surf_id = vsp.GetXSecSurf(wing_id, 0) # This is how VSP stores surfaces.
x_sec_1 = vsp.GetXSec(xsec_surf_id, 1)
if vsp.GetNumXSec(xsec_surf_id) == 2:
single_seg = True
else:
single_seg = False
segment_num = vsp.GetNumXSec(xsec_surf_id) # Get number of segments
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)
# Necessary wing segment definitions start at XSec_1 (XSec_0 exists mainly to hold the root airfoil)
xsec_surf_id = vsp.GetXSecSurf(wing_id, 0)
x_sec = vsp.GetXSec(xsec_surf_id, 1)
chord_parm = vsp.GetXSecParm(x_sec,'Root_Chord')
root_chord = vsp.GetParmVal(chord_parm) * units_factor
# -------------
# Wing segments
# -------------
if single_seg == False:
# Convert VSP XSecs to SUAVE segments. (Wing segments are defined by outboard sections in VSP, but inboard sections in SUAVE.)
for i in range(1, segment_num+1):
# XSec airfoil
jj = i-1 # Airfoil index i-1 because VSP airfoils and sections are one index off relative to SUAVE.
segment = SUAVE.Components.Wings.Segment()
segment.tag = 'Section_' + str(i)
thick_cord = vsp.GetParmVal(wing_id, 'ThickChord', 'XSecCurve_' + str(jj))
segment.thickness_to_chord = thick_cord # Thick_cord stored for use in airfoil, below.
if i!=segment_num:
segment_root_chord = vsp.GetParmVal(wing_id, 'Root_Chord', 'XSec_' + str(i)) * units_factor
else:
segment_root_chord = 0.0
segment.root_chord_percent = segment_root_chord / root_chord
segment.percent_span_location = proj_span_sum / (total_proj_span/(1+wing.symmetric))
segment.twist = vsp.GetParmVal(wing_id, 'Twist', 'XSec_' + str(jj)) * Units.deg
if i==1:
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 = 2*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 + x_rot
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 /root_chord
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
# 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)
if write_airfoil_file==True:
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 = 'airfoil'
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(1, 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.
# Add a tip segment, all values are zero except the tip chord
tc = vsp.GetParmVal(wing_id, 'Tip_Chord', 'XSec_' + str(segment_num-1)) * units_factor
# Chords
wing.chords.root = vsp.GetParmVal(wing_id, 'Tip_Chord', 'XSec_0') * units_factor
wing.chords.tip = tc
wing.chords.mean_geometric = wing.areas.reference / wing.spans.projected
# Just double calculate and fix things:
wing = wing_segmented_planform(wing)
else:
# Single segment
# Get ID's
x_sec_1_dih_parm = vsp.GetXSecParm(x_sec_1,'Dihedral')
x_sec_1_sweep_parm = vsp.GetXSecParm(x_sec_1,'Sweep')
x_sec_1_sweep_loc_parm = vsp.GetXSecParm(x_sec_1,'Sweep_Location')
x_sec_1_taper_parm = vsp.GetXSecParm(x_sec_1,'Taper')
x_sec_1_rc_parm = vsp.GetXSecParm(x_sec_1,'Root_Chord')
x_sec_1_tc_parm = vsp.GetXSecParm(x_sec_1,'Tip_Chord')
x_sec_1_t_parm = vsp.GetXSecParm(x_sec_1,'ThickChord')
# Calcs
sweep = vsp.GetParmVal(x_sec_1_sweep_parm) * Units.deg
sweep_loc = vsp.GetParmVal(x_sec_1_sweep_loc_parm)
taper = vsp.GetParmVal(x_sec_1_taper_parm)
c_4_sweep = convert_sweep(sweep,sweep_loc,0.25,wing.aspect_ratio,taper)
# Pull and pack
wing.sweeps.quarter_chord = c_4_sweep
wing.taper = taper
wing.dihedral = vsp.GetParmVal(x_sec_1_dih_parm) * Units.deg + x_rot
wing.chords.root = vsp.GetParmVal(x_sec_1_rc_parm)* units_factor
wing.chords.tip = vsp.GetParmVal(x_sec_1_tc_parm) * units_factor
wing.chords.mean_geometric = wing.areas.reference / wing.spans.projected
wing.thickness_to_chord = vsp.GetParmVal(x_sec_1_t_parm)
# Just double calculate and fix things:
wing = wing_planform(wing)
# 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
# check if control surface (sub surfaces) are defined
tags = []
LE_flags = []
span_fraction_starts = []
span_fraction_ends = []
chord_fractions = []
num_cs = vsp.GetNumSubSurf(wing_id)
# loop through wing and get all control surface parameters
for cs_idx in range(num_cs):
cs_id = vsp.GetSubSurf(wing_id,cs_idx)
param_names = vsp.GetSubSurfParmIDs(cs_id)
tags.append(vsp.GetSubSurfName(cs_id))
for p_idx in range(len(param_names)):
if 'LE_Flag' == vsp.GetParmName(param_names[p_idx]):
LE_flags.append(vsp.GetParmVal(param_names[p_idx]))
if 'UStart' == vsp.GetParmName(param_names[p_idx]):
span_fraction_starts.append(vsp.GetParmVal(param_names[p_idx]))
if 'UEnd' == vsp.GetParmName(param_names[p_idx]):
span_fraction_ends.append(vsp.GetParmVal(param_names[p_idx]))
if 'Length_C_Start' == vsp.GetParmName(param_names[p_idx]):
chord_fractions.append(vsp.GetParmVal(param_names[p_idx]))
# assign control surface parameters to wings. Outer most control surface on main/horizontal wing is assigned a aileron
for cs_idx in range(num_cs):
aileron_present = False
if num_cs > 1:
aileron_loc = np.argmax(np.array(span_fraction_starts))
if cs_idx == aileron_loc:
aileron_present = True
if LE_flags[cs_idx] == 1.0:
CS = SUAVE.Components.Wings.Control_Surfaces.Slat()
else:
if wing.vertical == True:
CS = SUAVE.Components.Wings.Control_Surfaces.Rudder()
else:
if aileron_present:
CS = SUAVE.Components.Wings.Control_Surfaces.Aileron()
else:
CS = SUAVE.Components.Wings.Control_Surfaces.Flap()
CS.tag = tags[cs_idx]
CS.span_fraction_start = span_fraction_starts[cs_idx]*3 - 1
CS.span_fraction_end = span_fraction_ends[cs_idx]*3 - 1
CS.chord_fraction = chord_fractions[cs_idx]
CS.span = (CS.span_fraction_end - CS.span_fraction_start)*wing.spans.projected
wing.append_control_surface(CS)
return wing
def write_vsp_nacelle(nacelle, OML_set_ind):
"""This converts nacelles into OpenVSP format.
Assumptions:
1. If nacelle segments are defined, geometry written to OpenVSP is of type "StackGeom".
1.a This type of nacelle can be either set as flow through or not flow through.
1.b Segments are defined in a similar manner to fuselage segments. See geometric
documentation in SUAVE-Components-Nacelles-Nacelle
2. If nacelle segments are not defined, geometry written to OpenVSP is of type "BodyofRevolution".
2.a This type of nacelle can be either set as flow through or not flow through.
2.b BodyofRevolution can be either be a 4 digit airfoil (type string) or super ellipse (default)
Source:
N/A
Inputs:
nacelle.
origin [m] in all three dimension, should have as many origins as engines
length [m]
diameter [m]
flow_through <boolean> if True create a flow through nacelle, if False create a cylinder
segment(optional).
width [m]
height [m]
lenght [m]
percent_x_location [m]
percent_y_location [m]
percent_z_location [m]
Outputs:
Operates on the active OpenVSP model, no direct output
Properties Used:
N/A
"""
# default tesselation
radial_tesselation = 21
axial_tesselation = 25
# True will create a flow-through subsonic nacelle (which may have dimensional errors)
# False will create a cylindrical stack (essentially a cylinder)
ft_flag = nacelle.flow_through
length = nacelle.length
height = nacelle.diameter - nacelle.inlet_diameter
diameter = nacelle.diameter - height / 2
nac_tag = nacelle.tag
nac_x = nacelle.origin[0][0]
nac_y = nacelle.origin[0][1]
nac_z = nacelle.origin[0][2]
nac_x_rotation = nacelle.orientation_euler_angles[0] / Units.degrees
nac_y_rotation = nacelle.orientation_euler_angles[1] / Units.degrees
nac_z_rotation = nacelle.orientation_euler_angles[2] / Units.degrees
num_segs = len(nacelle.Segments)
if num_segs > 0:
if nacelle.Airfoil.naca_4_series_airfoil != None:
raise AssertionError(
'Nacelle segments defined. Airfoil section will not be used.')
nac_id = vsp.AddGeom("STACK")
vsp.SetGeomName(nac_id, nac_tag)
# set nacelle relative location and rotation
vsp.SetParmVal(nac_id, 'Abs_Or_Relitive_flag', 'XForm', vsp.ABS)
vsp.SetParmVal(nac_id, 'X_Rotation', 'XForm', nac_x_rotation)
vsp.SetParmVal(nac_id, 'Y_Rotation', 'XForm', nac_y_rotation)
vsp.SetParmVal(nac_id, 'Z_Rotation', 'XForm', nac_z_rotation)
vsp.SetParmVal(nac_id, 'X_Location', 'XForm', nac_x)
vsp.SetParmVal(nac_id, 'Y_Location', 'XForm', nac_y)
vsp.SetParmVal(nac_id, 'Z_Location', 'XForm', nac_z)
vsp.SetParmVal(nac_id, 'Tess_U', 'Shape', radial_tesselation)
vsp.SetParmVal(nac_id, 'Tess_W', 'Shape', axial_tesselation)
widths = []
heights = []
x_delta = []
x_poses = []
z_delta = []
segs = nacelle.Segments
for seg in range(num_segs):
widths.append(segs[seg].width)
heights.append(segs[seg].height)
x_poses.append(segs[seg].percent_x_location)
if seg == 0:
x_delta.append(0)
z_delta.append(0)
else:
x_delta.append(length * (segs[seg].percent_x_location -
segs[seg - 1].percent_x_location))
z_delta.append(length * (segs[seg].percent_z_location -
segs[seg - 1].percent_z_location))
vsp.CutXSec(nac_id, 4) # remove point section at end
vsp.CutXSec(nac_id, 0) # remove point section at beginning
vsp.CutXSec(nac_id, 1) # remove point section at beginning
for _ in range(num_segs -
2): # add back the required number of sections
vsp.InsertXSec(nac_id, 1, vsp.XS_ELLIPSE)
vsp.Update()
xsec_surf = vsp.GetXSecSurf(nac_id, 0)
for i3 in reversed(range(num_segs)):
xsec = vsp.GetXSec(xsec_surf, i3)
if i3 == 0:
pass
else:
vsp.SetParmVal(nac_id, "XDelta", "XSec_" + str(i3),
x_delta[i3])
vsp.SetParmVal(nac_id, "ZDelta", "XSec_" + str(i3),
z_delta[i3])
vsp.SetXSecWidthHeight(xsec, widths[i3], heights[i3])
vsp.SetXSecTanAngles(xsec, vsp.XSEC_BOTH_SIDES, 0, 0, 0, 0)
vsp.SetXSecTanSlews(xsec, vsp.XSEC_BOTH_SIDES, 0, 0, 0, 0)
vsp.SetXSecTanStrengths(xsec, vsp.XSEC_BOTH_SIDES, 0, 0, 0, 0)
vsp.Update()
if ft_flag:
pass
else:
# append front point
xsecsurf = vsp.GetXSecSurf(nac_id, 0)
vsp.ChangeXSecShape(xsecsurf, 0, vsp.XS_POINT)
vsp.Update()
xsecsurf = vsp.GetXSecSurf(nac_id, 0)
vsp.ChangeXSecShape(xsecsurf, num_segs - 1, vsp.XS_POINT)
vsp.Update()
else:
nac_id = vsp.AddGeom("BODYOFREVOLUTION")
vsp.SetGeomName(nac_id, nac_tag)
# Origin
vsp.SetParmVal(nac_id, 'Abs_Or_Relitive_flag', 'XForm', vsp.ABS)
vsp.SetParmVal(nac_id, 'X_Rotation', 'XForm', nac_x_rotation)
vsp.SetParmVal(nac_id, 'Y_Rotation', 'XForm', nac_y_rotation)
vsp.SetParmVal(nac_id, 'Z_Rotation', 'XForm', nac_z_rotation)
vsp.SetParmVal(nac_id, 'X_Location', 'XForm', nac_x)
vsp.SetParmVal(nac_id, 'Y_Location', 'XForm', nac_y)
vsp.SetParmVal(nac_id, 'Z_Location', 'XForm', nac_z)
vsp.SetParmVal(nac_id, 'Tess_U', 'Shape', radial_tesselation)
vsp.SetParmVal(nac_id, 'Tess_W', 'Shape', axial_tesselation)
# Length and overall diameter
vsp.SetParmVal(nac_id, "Diameter", "Design", diameter)
if ft_flag:
vsp.SetParmVal(nac_id, "Mode", "Design", 0.0)
else:
vsp.SetParmVal(nac_id, "Mode", "Design", 1.0)
if nacelle.Airfoil.naca_4_series_airfoil != None:
if isinstance(
nacelle.Airfoil.naca_4_series_airfoil,
str) and len(nacelle.Airfoil.naca_4_series_airfoil) != 4:
raise AssertionError(
'Nacelle cowling airfoil must be of type < string > and length < 4 >'
)
else:
angle = nacelle.cowling_airfoil_angle / Units.degrees
camber = float(nacelle.Airfoil.naca_4_series_airfoil[0]) / 100
camber_loc = float(
nacelle.Airfoil.naca_4_series_airfoil[1]) / 10
thickness = float(
nacelle.Airfoil.naca_4_series_airfoil[2:]) / 100
vsp.ChangeBORXSecShape(nac_id, vsp.XS_FOUR_SERIES)
vsp.Update()
vsp.SetParmVal(nac_id, "Diameter", "Design", diameter)
vsp.SetParmVal(nac_id, "Angle", "Design", angle)
vsp.SetParmVal(nac_id, "Chord", "XSecCurve", length)
vsp.SetParmVal(nac_id, "ThickChord", "XSecCurve", thickness)
vsp.SetParmVal(nac_id, "Camber", "XSecCurve", camber)
vsp.SetParmVal(nac_id, "CamberLoc", "XSecCurve", camber_loc)
vsp.Update()
else:
vsp.ChangeBORXSecShape(nac_id, vsp.XS_SUPER_ELLIPSE)
vsp.Update()
if ft_flag:
vsp.SetParmVal(nac_id, "Super_Height", "XSecCurve", height)
vsp.SetParmVal(nac_id, "Diameter", "Design", diameter)
else:
vsp.SetParmVal(nac_id, "Super_Height", "XSecCurve", diameter)
vsp.SetParmVal(nac_id, "Super_Width", "XSecCurve", length)
vsp.SetParmVal(nac_id, "Super_MaxWidthLoc", "XSecCurve", 0.)
vsp.SetParmVal(nac_id, "Super_M", "XSecCurve", 2.)
vsp.SetParmVal(nac_id, "Super_N", "XSecCurve", 1.)
vsp.SetSetFlag(nac_id, OML_set_ind, True)
vsp.Update()
return
def read_vsp_nacelle(nacelle_id, vsp_nacelle_type, units_type='SI'):
"""This reads an OpenVSP stack geometry or body of revolution and writes it to a SUAVE nacelle format.
If an airfoil is defined in body-of-revolution, its coordinates are not read in due to absence of
API functions in VSP.
Assumptions:
Source:
N/A
Inputs:
0. Pre-loaded VSP vehicle in memory, via vsp_read.
1. VSP 10-digit geom ID for nacelle.
2. Units_type set to 'SI' (default) or 'Imperial'.
Outputs:
Writes SUAVE nacelle, with these geometries: (all defaults are SI, but user may specify Imperial)
Nacelles.Nacelle.
origin [m] in all three dimensions
width [m]
lengths [m]
heights [m]
tag <string>
segment[]. (segments are in ordered container and callable by number)
percent_x_location [unitless]
percent_z_location [unitless]
height [m]
width [m]
Properties Used:
N/A
"""
nacelle = SUAVE.Components.Nacelles.Nacelle()
if units_type == 'SI':
units_factor = Units.meter * 1.
elif units_type == 'imperial':
units_factor = Units.foot * 1.
elif units_type == 'inches':
units_factor = Units.inch * 1.
if vsp.GetGeomName(nacelle_id):
nacelle.tag = vsp.GetGeomName(nacelle_id)
else:
nacelle.tag = 'NacelleGeom'
nacelle.origin[0][0] = vsp.GetParmVal(nacelle_id, 'X_Location',
'XForm') * units_factor
nacelle.origin[0][1] = vsp.GetParmVal(nacelle_id, 'Y_Location',
'XForm') * units_factor
nacelle.origin[0][2] = vsp.GetParmVal(nacelle_id, 'Z_Location',
'XForm') * units_factor
nacelle.x_rotation = vsp.GetParmVal(nacelle_id, 'X_Rotation',
'XForm') * units_factor
nacelle.y_rotation = vsp.GetParmVal(nacelle_id, 'Y_Rotation',
'XForm') * units_factor
nacelle.z_rotation = vsp.GetParmVal(nacelle_id, 'Z_Rotation',
'XForm') * units_factor
if vsp_nacelle_type == 'Stack':
xsec_surf_id = vsp.GetXSecSurf(
nacelle_id, 0) # There is only one XSecSurf in geom.
num_segs = vsp.GetNumXSec(xsec_surf_id) # Number of xsecs in nacelle.
abs_x_location = 0
abs_y_location = 0
abs_z_location = 0
abs_x_location_vec = []
abs_y_location_vec = []
abs_z_location_vec = []
for i in range(num_segs):
# Create the segment
xsec_id = vsp.GetXSec(xsec_surf_id, i) # VSP XSec ID.
segment = SUAVE.Components.Lofted_Body_Segment.Segment()
segment.tag = 'segment_' + str(i)
# Pull out Parms that will be needed
X_Loc_P = vsp.GetXSecParm(xsec_id, 'XDelta')
Y_Loc_P = vsp.GetXSecParm(xsec_id, 'YDelta')
Z_Loc_P = vsp.GetXSecParm(xsec_id, 'XDelta')
del_x = vsp.GetParmVal(X_Loc_P)
del_y = vsp.GetParmVal(Y_Loc_P)
del_z = vsp.GetParmVal(Z_Loc_P)
abs_x_location = abs_x_location + del_x
abs_y_location = abs_y_location + del_y
abs_z_location = abs_z_location + del_z
abs_x_location_vec.append(abs_x_location)
abs_y_location_vec.append(abs_y_location)
abs_z_location_vec.append(abs_z_location)
shape = vsp.GetXSecShape(xsec_id)
shape_dict = {
0: 'point',
1: 'circle',
2: 'ellipse',
3: 'super ellipse',
4: 'rounded rectangle',
5: 'general fuse',
6: 'fuse file'
}
if shape_dict[shape] == 'point':
segment.height = 0.0
segment.width = 0.0
if i == 0:
nacelle.flow_through = False
else:
segment.height = vsp.GetXSecHeight(xsec_id) * units_factor
segment.width = vsp.GetXSecWidth(xsec_id) * units_factor
if i == 0:
nacelle.flow_through = True
nacelle.Segments.append(segment)
nacelle.length = abs_x_location_vec[-1]
segs = nacelle.Segments
for seg in range(num_segs):
segs[seg].percent_x_location = np.array(
abs_x_location_vec) / abs_x_location_vec[-1]
segs[seg].percent_y_location = np.array(
abs_y_location_vec) / abs_x_location_vec[-1]
segs[seg].percent_z_location = np.array(
abs_z_location_vec) / abs_x_location_vec[-1]
elif vsp_nacelle_type == 'BodyOfRevolution':
diameter = vsp.GetParmVal(nacelle_id, "Diameter",
"Design") * units_factor
angle = vsp.GetParmVal(nacelle_id, "Diameter",
"Design") * Units.degrees
ft_flag_idx = vsp.GetParmVal(nacelle_id, "Mode", "Design")
if ft_flag_idx == 0.0:
ft_flag = True
else:
ft_flag = False
nacelle.flow_through = ft_flag
shape = vsp.GetBORXSecShape(nacelle_id)
shape_dict = {0:'point',1:'circle',2:'ellipse',3:'super ellipse',4:'rounded rectangle',5:'general fuse',6:'fuse file',\
7:'four series',8:'six series',9:'biconvex',10:'wedge',11:'editcurve',12:'file airfoil'}
if shape_dict[shape] == 'four series':
naf = SUAVE.Components.Airfoils.Airfoil()
length = vsp.GetParmVal(nacelle_id, "Chord", "XSecCurve")
thickness = int(
round(
vsp.GetParmVal(nacelle_id, "ThickChord", "XSecCurve") * 10,
0))
camber = int(
round(
vsp.GetParmVal(nacelle_id, "Camber", "XSecCurve") * 100,
0))
camber_loc = int(
round(
vsp.GetParmVal(nacelle_id, "CamberLoc", "XSecCurve") * 10,
0))
airfoil = str(camber) + str(camber_loc) + str(thickness)
height = thickness
naf.naca_4_series_airfoil = str(airfoil)
naf.thickness_to_chord = thickness
nacelle.append_airfoil(naf)
elif shape_dict[shape] == 'super ellipse':
if ft_flag:
height = vsp.GetParmVal(nacelle_id, "Super_Height",
"XSecCurve")
diameter = vsp.GetParmVal(nacelle_id, "Diameter", "Design")
length = vsp.GetParmVal(nacelle_id, "Super_Width", "XSecCurve")
else:
diameter = vsp.GetParmVal(nacelle_id, "Super_Height",
"XSecCurve")
length = vsp.GetParmVal(nacelle_id, "Super_Width", "XSecCurve")
height = diameter / 2
elif shape_dict[shape] == 'file airfoil':
naf = SUAVE.Components.Airfoils.Airfoil()
thickness_to_chord = vsp.GetParmVal(nacelle_id, "ThickChord",
"XSecCurve") * units_factor
length = vsp.GetParmVal(nacelle_id, "Chord",
"XSecCurve") * units_factor
height = thickness_to_chord * length * units_factor
if ft_flag:
diameter = vsp.GetParmVal(nacelle_id, "Diameter",
"Design") * units_factor
else:
diameter = 0
naf.thickness_to_chord = thickness_to_chord
nacelle.append_airfoil(naf)
nacelle.length = length
nacelle.diameter = diameter + height / 2
nacelle.inlet_diameter = nacelle.diameter - height
nacelle.cowling_airfoil_angle = angle
return nacelle