Source code for RCAIDE.Library.Methods.Aerodynamics.Common.Drag.compressibility_drag
# RCAIDE/Library/Methods/Aerodynamics/Common/Drag/compressibility_drag_total.py
# (c) Copyright 2023 Aerospace Research Community LLC
#
# Created: Jul 2024, RCAIDE Team
# ----------------------------------------------------------------------------------------------------------------------
# IMPORT
# ----------------------------------------------------------------------------------------------------------------------
from RCAIDE.Library.Components.Wings import Main_Wing
from RCAIDE.Library.Methods.Utilities import Cubic_Spline_Blender
from .wave_drag import wave_drag
from .drag_divergence import drag_divergence
from .supersonic_wave_drag_volume_raymer import supersonic_wave_drag_volume_raymer
from .supersonic_wave_drag_volume_sears_haack import supersonic_wave_drag_volume_sears_haack
# package imports
import numpy as np
# ----------------------------------------------------------------------------------------------------------------------
# Compressibility Drag Total
# ----------------------------------------------------------------------------------------------------------------------
[docs]
def compressibility_drag(state,settings,geometry):
"""Computes compressibility drag for full aircraft including volume drag
Assumptions:
None
Source:
None
Args:
settings.
begin_drag_rise_mach_number [Unitless]
end_drag_rise_mach_number [Unitless]
peak_mach_number [Unitless]
transonic_drag_multiplier [Unitless]
volume_wave_drag_scaling [Unitless]
state.conditions.aerodynamics.lift_breakdown.compressible_wings [Unitless]
state.conditions.freestream.mach_number [Unitless]
geometry.maximum_cross_sectional_area [m^2] (used in subfunctions)
geometry.total_length [m] (used in subfunctions)
geometry.reference_area [m^2]
geometry.wings
Returns:
total_compressibility_drag [Unitless]
Properties Used:
None
"""
# Unpack
conditions = state.conditions
Mach = conditions.freestream.mach_number
Cl = conditions.aerodynamics.coefficients.lift.total
# supersonic
low_mach_cutoff = settings.supersonic.begin_drag_rise_mach_number
high_mach_cutoff = settings.supersonic.end_drag_rise_mach_number
peak_mach = settings.supersonic.peak_mach_number
peak_factor = settings.supersonic.transonic_drag_multiplier
scaling_factor = settings.supersonic.volume_wave_drag_scaling
if settings.supersonic.wave_drag_type == 'Raymer':
wave_drag_volume = supersonic_wave_drag_volume_raymer
elif settings.supersonic.wave_drag_type == 'Sears-Haack':
wave_drag_volume = supersonic_wave_drag_volume_sears_haack
else:
raise NotImplementedError
if settings.supersonic.cross_sectional_area_calculation_type != 'Fixed':
raise NotImplementedError
low_cutoff_volume_total = np.zeros_like(Mach)
high_cutoff_volume_total = np.zeros_like(Mach)
low_cutoff_volume_total = drag_divergence(low_mach_cutoff*np.ones_like(Mach), geometry,Cl)
high_cutoff_volume_total = wave_drag_volume(geometry,low_mach_cutoff*np.ones_like(Mach),scaling_factor)
peak_volume_total = high_cutoff_volume_total*peak_factor
# subsonic side smoothing function
a1 = (low_cutoff_volume_total-peak_volume_total)/(low_mach_cutoff-peak_mach)/(low_mach_cutoff-peak_mach)
# supersonic side smoothing function
a2 = (high_cutoff_volume_total-peak_volume_total)/(high_mach_cutoff-peak_mach)/(high_mach_cutoff-peak_mach)
# Shorten cubic Hermite spline
sub_spline = Cubic_Spline_Blender(low_mach_cutoff, peak_mach-(peak_mach-low_mach_cutoff)*3/4)
sup_spline = Cubic_Spline_Blender(peak_mach,high_mach_cutoff)
sub_h00 = lambda M:sub_spline.compute(M)
sup_h00 = lambda M:sup_spline.compute(M)
low_inds = Mach[:,0]<peak_mach
hi_inds = Mach[:,0]>=peak_mach
cd_c_v_base = np.zeros_like(Mach)
cd_c_v_base[low_inds] = drag_divergence(Mach[low_inds], geometry,Cl[low_inds])
cd_c_l_base = lift_wave_drag(conditions, settings, geometry)
cd_c_v_base[Mach>=peak_mach] = wave_drag_volume(geometry, Mach[Mach>=peak_mach], scaling_factor)
cd_c_v = np.zeros_like(Mach)
cd_c_v[low_inds] = cd_c_v_base[low_inds]*(sub_h00(Mach[low_inds])) + transonic_drag_function(Mach[low_inds],a1[low_inds], peak_mach, peak_volume_total[low_inds])*(1-sub_h00(Mach[low_inds]))
cd_c_v[hi_inds] = transonic_drag_function(Mach[hi_inds],a2[hi_inds], peak_mach, peak_volume_total[hi_inds])*(sup_h00(Mach[hi_inds])) + cd_c_v_base[hi_inds]*(1-sup_h00(Mach[hi_inds]))
if peak_mach<1.01:
print('Warning: a peak Mach number of less than 1.01 will cause a small discontinuity in lift wave drag')
cd_c_l = cd_c_l_base*(1-sup_h00(Mach))
cd_c = cd_c_v + cd_c_l
# Save drag breakdown
drag = conditions.aerodynamics.coefficients.drag.compressible.total = cd_c
return
# ----------------------------------------------------------------------------------------------------------------------
# transonic_drag_function
# ----------------------------------------------------------------------------------------------------------------------
[docs]
def transonic_drag_function(Mach,a_vertex, peak_mach, peak_volume_total):
""" parabolic approximation of drag rise in the transonic region
Assumptions:
Basic fit
Source:
None
Args:
Mach
a_vertex
peak_mach
peak_volume_total
Returns:
transonic_drag
"""
transonic_drag = a_vertex*(Mach-peak_mach)*(Mach-peak_mach)+peak_volume_total
transonic_drag = transonic_drag.reshape(np.shape(Mach))
return transonic_drag
# ----------------------------------------------------------------------------------------------------------------------
# lift_wave_drag
# ----------------------------------------------------------------------------------------------------------------------
[docs]
def lift_wave_drag(conditions,configuration,geometry):
"""Determine lift wave drag for supersonic speeds
Assumptions:
Basic fit
Source:
http://aerodesign.stanford.edu/aircraftdesign/aircraftdesign.html (Stanford AA241 A/B Course Notes)
Args:
conditions.freestream.mach_number [-]
configuration (passed to another function)
wing.areas.reference [m^2]
Sref_main [m^2] Main reference area
Returns:
cd_c_l [-] Wave drag CD due to lift
"""
for wing in geometry.wings:
if isinstance(wing, Main_Wing):
# Unpack Mach number
Mach = conditions.freestream.mach_number
# Initalize cd arrays
cd_c_l = np.zeros_like(Mach)
# Calculate wing values at all Mach numbers
cd_lift_wave = wave_drag(conditions,wing)
# Pack supersonic results into correct elements
cd_c_l[Mach >= 1.01] = cd_lift_wave[0:len(Mach[Mach >= 1.01]),0]
return cd_c_l
