def parasite_drag_fuselage(state,settings,geometry):
"""Computes the parasite drag due to the fuselage
Assumptions:
Basic fit
Source:
adg.stanford.edu (Stanford AA241 A/B Course Notes)
Inputs:
state.conditions.freestream.
mach_number [Unitless]
temperature [K]
reynolds_number [Unitless]
settings.fuselage_parasite_drag_form_factor [Unitless]
geometry.fuselage.
areas.front_projected [m^2]
areas.wetted [m^2]
lengths.total [m]
effective_diameter [m]
Outputs:
fuselage_parasite_drag [Unitless]
Properties Used:
N/A
"""
# unpack inputs
configuration = settings
form_factor = configuration.fuselage_parasite_drag_form_factor
fuselage = geometry
freestream = state.conditions.freestream
Sref = fuselage.areas.front_projected
Swet = fuselage.areas.wetted
l_fus = fuselage.lengths.total
d_fus = fuselage.effective_diameter
# conditions
Mc = freestream.mach_number
Tc = freestream.temperature
re = freestream.reynolds_number
# reynolds number
Re_fus = re*(l_fus)
# skin friction coefficient
cf_fus, k_comp, k_reyn = compressible_turbulent_flat_plate(Re_fus,Mc,Tc)
# form factor for cylindrical bodies
d_d = float(d_fus)/float(l_fus)
D = np.array([[0.0]] * len(Mc))
a = np.array([[0.0]] * len(Mc))
du_max_u = np.array([[0.0]] * len(Mc))
k_fus = np.array([[0.0]] * len(Mc))
D[Mc < 0.95] = np.sqrt(1 - (1-Mc[Mc < 0.95]**2) * d_d**2)
a[Mc < 0.95] = 2 * (1-Mc[Mc < 0.95]**2) * (d_d**2) *(np.arctanh(D[Mc < 0.95])-D[Mc < 0.95]) / (D[Mc < 0.95]**3)
du_max_u[Mc < 0.95] = a[Mc < 0.95] / ( (2-a[Mc < 0.95]) * (1-Mc[Mc < 0.95]**2)**0.5 )
D[Mc >= 0.95] = np.sqrt(1 - d_d**2)
a[Mc >= 0.95] = 2 * (d_d**2) *(np.arctanh(D[Mc >= 0.95])-D[Mc >= 0.95]) / (D[Mc >= 0.95]**3)
du_max_u[Mc >= 0.95] = a[Mc >= 0.95] / ( (2-a[Mc >= 0.95]) )
k_fus = (1 + form_factor*du_max_u)**2
fuselage_parasite_drag = k_fus * cf_fus * Swet / Sref
# dump data to conditions
fuselage_result = Data(
wetted_area = Swet ,
reference_area = Sref ,
parasite_drag_coefficient = fuselage_parasite_drag ,
skin_friction_coefficient = cf_fus ,
compressibility_factor = k_comp ,
reynolds_factor = k_reyn ,
form_factor = k_fus ,
)
try:
state.conditions.aerodynamics.drag_breakdown.parasite[fuselage.tag] = fuselage_result
except:
print("Drag Polar Mode fuse parasite")
return fuselage_parasite_drag
def parasite_drag_nacelle(state, settings, nacelle):
"""Computes the parasite drag due to the nacelle
Assumptions:
Basic fit
Source:
Raymer equation (pg 283 of Aircraft Design: A Conceptual Approach) (subsonic)
http://aerodesign.stanford.edu/aircraftdesign/drag/BODYFORMFACTOR.HTML (supersonic)
Inputs:
state.conditions.freestream.
mach_number [Unitless]
temperature [K]
reynolds_number [Unitless]
geometry.
nacelle.diameter [m^2]
areas.wetted [m^2]
length [m]
Outputs:
nacelle_parasite_drag [Unitless]
Properties Used:
N/A
"""
# unpack inputs
conditions = state.conditions
low_mach_cutoff = settings.begin_drag_rise_mach_number
high_mach_cutoff = settings.end_drag_rise_mach_number
freestream = conditions.freestream
Sref = nacelle.diameter**2 / 4 * np.pi
Swet = nacelle.areas.wetted
l_prop = nacelle.length
d_prop = nacelle.diameter
# conditions
freestream = conditions.freestream
Mc = freestream.mach_number
Tc = freestream.temperature
re = freestream.reynolds_number
# reynolds number
Re_prop = re * l_prop
# skin friction coefficient
cf_prop, k_comp, k_reyn = compressible_turbulent_flat_plate(
Re_prop, Mc, Tc)
# form factor according to Raymer equation (pg 283 of Aircraft Design: A Conceptual Approach)
k_prop_sub = 1. + 0.35 / (float(l_prop) / float(d_prop))
# for supersonic flow (http://adg.stanford.edu/aa241/drag/BODYFORMFACTOR.HTML)
k_prop_sup = 1.
trans_spline = Cubic_Spline_Blender(low_mach_cutoff, high_mach_cutoff)
h00 = lambda M: trans_spline.compute(M)
k_prop = k_prop_sub * (h00(Mc)) + k_prop_sup * (1 - h00(Mc))
# --------------------------------------------------------
# find the final result
nacelle_parasite_drag = k_prop * cf_prop * Swet / Sref
# --------------------------------------------------------
# dump data to conditions
nacelle_result = Data(
wetted_area=Swet,
reference_area=Sref,
parasite_drag_coefficient=nacelle_parasite_drag,
skin_friction_coefficient=cf_prop,
compressibility_factor=k_comp,
reynolds_factor=k_reyn,
form_factor=k_prop,
)
state.conditions.aerodynamics.drag_breakdown.parasite[
nacelle.tag] = nacelle_result
return nacelle_parasite_drag
def parasite_drag_propulsor(state,settings,geometry):
"""Computes the parasite drag due to the propulsor
Assumptions:
Basic fit
Source:
adg.stanford.edu (Stanford AA241 A/B Course Notes)
Inputs:
state.conditions.freestream.
mach_number [Unitless]
temperature [K]
reynolds_number [Unitless]
geometry.
nacelle_diameter [m^2]
areas.wetted [m^2]
engine_length [m]
Outputs:
propulsor_parasite_drag [Unitless]
Properties Used:
N/A
"""
# unpack inputs
conditions = state.conditions
configuration = settings
propulsor = geometry
Sref = propulsor.nacelle_diameter**2. / 4. * np.pi
Swet = propulsor.areas.wetted
l_prop = propulsor.engine_length
d_prop = propulsor.nacelle_diameter
# conditions
freestream = conditions.freestream
Mc = freestream.mach_number
Tc = freestream.temperature
re = freestream.reynolds_number
# reynolds number
Re_prop = re*l_prop
# skin friction coefficient
cf_prop, k_comp, k_reyn = compressible_turbulent_flat_plate(Re_prop,Mc,Tc)
## form factor according to Raymer equation (pg 283 of Aircraft Design: A Conceptual Approach)
k_prop = 1 + 0.35 / (float(l_prop)/float(d_prop))
# find the final result
propulsor_parasite_drag = k_prop * cf_prop * Swet / Sref
# dump data to conditions
propulsor_result = Data(
wetted_area = Swet ,
reference_area = Sref ,
parasite_drag_coefficient = propulsor_parasite_drag ,
skin_friction_coefficient = cf_prop ,
compressibility_factor = k_comp ,
reynolds_factor = k_reyn ,
form_factor = k_prop ,
)
conditions.aerodynamics.drag_breakdown.parasite[propulsor.tag] = propulsor_result
return propulsor_parasite_drag
def parasite_drag_fuselage(state, settings, geometry):
"""Computes the parasite drag due to the fuselage
Assumptions:
Basic fit
Source:
http://aerodesign.stanford.edu/aircraftdesign/aircraftdesign.html (Stanford AA241 A/B Course Notes)
Inputs:
state.conditions.freestream.
mach_number [Unitless]
temperature [K]
reynolds_number [Unitless]
settings.fuselage_parasite_drag_form_factor [Unitless]
geometry.fuselage.
areas.front_projected [m^2]
areas.wetted [m^2]
lengths.total [m]
effective_diameter [m]
Outputs:
fuselage_parasite_drag [Unitless]
Properties Used:
N/A
"""
# unpack inputs
configuration = settings
form_factor = configuration.fuselage_parasite_drag_form_factor
low_cutoff = configuration.fuselage_parasite_drag_begin_blend_mach
high_cutoff = configuration.fuselage_parasite_drag_end_blend_mach
fuselage = geometry
freestream = state.conditions.freestream
Sref = fuselage.areas.front_projected
Swet = fuselage.areas.wetted
l_fus = fuselage.lengths.total
d_fus = fuselage.effective_diameter
# conditions
Mc = freestream.mach_number
Tc = freestream.temperature
re = freestream.reynolds_number
# reynolds number
Re_fus = re * (l_fus)
# skin friction coefficient
cf_fus, k_comp, k_reyn = compressible_turbulent_flat_plate(Re_fus, Mc, Tc)
# form factor for cylindrical bodies
d_d = float(d_fus) / float(l_fus)
D_low = np.array([[0.0]] * len(Mc))
a_low = np.array([[0.0]] * len(Mc))
du_max_u_low = np.array([[0.0]] * len(Mc))
D_high = np.array([[0.0]] * len(Mc))
a_high = np.array([[0.0]] * len(Mc))
du_max_u_high = np.array([[0.0]] * len(Mc))
k_fus = np.array([[0.0]] * len(Mc))
low_inds = Mc < high_cutoff
high_inds = Mc > low_cutoff
D_low[low_inds] = np.sqrt(1 - (1 - Mc[low_inds]**2) * d_d**2)
a_low[low_inds] = 2 * (1 - Mc[low_inds]**2) * (d_d**2) * (
np.arctanh(D_low[low_inds]) - D_low[low_inds]) / (D_low[low_inds]**3)
du_max_u_low[low_inds] = a_low[low_inds] / ((2 - a_low[low_inds]) *
(1 - Mc[low_inds]**2)**0.5)
D_high[high_inds] = np.sqrt(1 - d_d**2)
a_high[high_inds] = 2 * (d_d**2) * (np.arctanh(
D_high[high_inds]) - D_high[high_inds]) / (D_high[high_inds]**3)
du_max_u_high[high_inds] = a_high[high_inds] / ((2 - a_high[high_inds]))
spline = Cubic_Spline_Blender(low_cutoff, high_cutoff)
h00 = lambda M: spline.compute(M)
du_max_u = du_max_u_low * (h00(Mc)) + du_max_u_high * (1 - h00(Mc))
k_fus = (1 + form_factor * du_max_u)**2
fuselage_parasite_drag = k_fus * cf_fus * Swet / Sref
# dump data to conditions
fuselage_result = Data(
wetted_area=Swet,
reference_area=Sref,
parasite_drag_coefficient=fuselage_parasite_drag,
skin_friction_coefficient=cf_fus,
compressibility_factor=k_comp,
reynolds_factor=k_reyn,
form_factor=k_fus,
)
try:
state.conditions.aerodynamics.drag_breakdown.parasite[
fuselage.tag] = fuselage_result
except:
print("Drag Polar Mode fuse parasite")
return fuselage_parasite_drag
def parasite_drag_propulsor(state,settings,geometry):
"""Computes the parasite drag due to the propulsor
Assumptions:
Basic fit
Source:
Raymer equation (pg 283 of Aircraft Design: A Conceptual Approach) (subsonic)
http://adg.stanford.edu/aa241/drag/BODYFORMFACTOR.HTML (supersonic)
Inputs:
state.conditions.freestream.
mach_number [Unitless]
temperature [K]
reynolds_number [Unitless]
geometry.
nacelle_diameter [m^2]
areas.wetted [m^2]
engine_length [m]
state.conditions.aerodynamics.drag_breakdown.
compressible.main_wing.divergence_mach [Unitless]
Outputs:
propulsor_parasite_drag [Unitless]
Properties Used:
N/A
"""
# unpack inputs
conditions = state.conditions
configuration = settings
propulsor = geometry
freestream = conditions.freestream
Sref = propulsor.nacelle_diameter**2 / 4 * np.pi
Swet = propulsor.areas.wetted
l_prop = propulsor.engine_length
d_prop = propulsor.nacelle_diameter
# conditions
freestream = conditions.freestream
Mc = freestream.mach_number
Tc = freestream.temperature
re = freestream.reynolds_number
# reynolds number
Re_prop = re*l_prop
# skin friction coefficient
cf_prop, k_comp, k_reyn = compressible_turbulent_flat_plate(Re_prop,Mc,Tc)
k_prop = np.array([[0.0]]*len(Mc))
# assume that the drag divergence mach number of the propulsor matches the main wing
Mdiv = state.conditions.aerodynamics.drag_breakdown.compressible.main_wing.divergence_mach
# form factor according to Raymer equation (pg 283 of Aircraft Design: A Conceptual Approach)
k_prop_sub = 1. + 0.35 / (float(l_prop)/float(d_prop))
# for supersonic flow (http://adg.stanford.edu/aa241/drag/BODYFORMFACTOR.HTML)
k_prop_sup = 1.
sb_mask = (Mc <= Mdiv)
tn_mask = ((Mc > Mdiv) & (Mc < 1.05))
sp_mask = (Mc >= 1.05)
k_prop[sb_mask] = k_prop_sub
# basic interpolation for transonic
k_prop[tn_mask] = (k_prop_sup-k_prop_sub)*(Mc[tn_mask]-Mdiv[tn_mask])/(1.05-Mdiv[tn_mask]) + k_prop_sub
k_prop[sp_mask] = k_prop_sup
# --------------------------------------------------------
# find the final result
propulsor_parasite_drag = k_prop * cf_prop * Swet / Sref
# --------------------------------------------------------
# dump data to conditions
propulsor_result = Data(
wetted_area = Swet ,
reference_area = Sref ,
parasite_drag_coefficient = propulsor_parasite_drag ,
skin_friction_coefficient = cf_prop ,
compressibility_factor = k_comp ,
reynolds_factor = k_reyn ,
form_factor = k_prop ,
)
state.conditions.aerodynamics.drag_breakdown.parasite[propulsor.tag] = propulsor_result
return propulsor_parasite_drag