Source code for RCAIDE.Library.Methods.Geometry.Planform.compute_segment_volume
# RCAIDE/Library/Methods/Geometry/Planform/compute_segment_volume.py
#
#
# Created: Jul 2024, M. Clarke
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# IMPORT
# ----------------------------------------------------------------------------------------------------------------------
import RCAIDE
from RCAIDE.Library.Methods.Geometry.Airfoil import compute_naca_4series
import numpy as np
from shapely import Polygon
from copy import deepcopy
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# Compute segment volume
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[docs]
def compute_segment_volume(wing, inner_segment, outer_segment,n_points=401):
"""
Estimate the prism volume between two consecutive wing segments using their
airfoil cross-sectional areas and local span.
Parameters
----------
wing : Wing
Wing object containing reference chord and span information.
inner_segment : WingSegment
Upstream segment with airfoil definition and percent span/chord scaling.
outer_segment : WingSegment
Downstream segment with airfoil definition and percent span/chord scaling.
n_points : int, optional
Number of points to generate NACA 4-series airfoil geometry when missing.
Returns
-------
float
Estimated segment volume [m³] using a frustum (prismoid) approximation.
Notes
-----
* If airfoil geometry is absent, a NACA 4-series shape is generated/assumed.
* Areas are computed via polygon integration of the airfoil coordinates.
* Volume uses (1/3)*(A1 + A2 + sqrt(A1*A2))*span for the local segment span.
"""
# Extract airfoil coordinates
if hasattr(inner_segment.airfoil,'geometry') is False or hasattr(outer_segment.airfoil,'geometry') is False: # first, if airfoil geometry data not defined, import from geoemtry files
inner_segment = deepcopy(inner_segment)
outer_segment = deepcopy(outer_segment)
if type(inner_segment.airfoil) == RCAIDE.Library.Components.Airfoils.NACA_4_Series_Airfoil: # check if naca 4 series of airfoil from datafile
try:
inner_segment.airfoil.geometry = compute_naca_4series(inner_segment.NACA_4_Series_code,n_points)
except:
inner_segment.airfoil.geometry = compute_naca_4series('0012')
else:
inner_segment.airfoil= RCAIDE.Library.Components.Airfoils.NACA_4_Series_Airfoil
inner_segment.airfoil.geometry = compute_naca_4series('0012')
if type(outer_segment.airfoil) == RCAIDE.Library.Components.Airfoils.NACA_4_Series_Airfoil: # check if naca 4 series of airfoil from datafile
try:
outer_segment.airfoil.geometry = compute_naca_4series(outer_segment.NACA_4_Series_code,n_points)
except:
outer_segment.airfoil.geometry = compute_naca_4series('0012')
else:
outer_segment.airfoil= RCAIDE.Library.Components.Airfoils.NACA_4_Series_Airfoil
outer_segment.airfoil.geometry = compute_naca_4series('0012')
x_in = np.array(inner_segment.airfoil.geometry.x_coordinates)[:-1] * wing.chords.root *inner_segment.root_chord_percent
y_in = np.array(inner_segment.airfoil.geometry.y_coordinates)[:-1] * wing.chords.root *inner_segment.root_chord_percent
x_out = np.array(outer_segment.airfoil.geometry.x_coordinates)[:-1] * wing.chords.root *outer_segment.root_chord_percent
y_out = np.array(outer_segment.airfoil.geometry.y_coordinates)[:-1] * wing.chords.root *outer_segment.root_chord_percent
points_out = list(zip(x_out, y_out))
poly_out = Polygon(points_out)
points_in = list(zip(x_in, y_in))
poly_in = Polygon(points_in)
A_1 = poly_in.area
A_2 = poly_out.area
# Compute segment span length
L = (outer_segment.percent_span_location - inner_segment.percent_span_location) * wing.spans.projected
volume = (1 /3) * ( A_1 + A_2 + np.sqrt(A_1*A_2)) *L
return volume
