def parasite_drag_propulsor(state, settings, geometry):
""" SUAVE.Methods.parasite_drag_propulsor(conditions,configuration,propulsor)
computes the parasite drag associated with a propulsor
Inputs:
Outputs:
Assumptions:
"""
# 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 = Results(
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:
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 = Results(
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_fuselage(conditions, configuration, fuselage):
""" SUAVE.Methods.parasite_drag_fuselage(conditions,configuration,fuselage)
computes the parasite drag associated with a fuselage
Inputs:
Outputs:
Assumptions:
"""
# unpack inputs
form_factor = configuration.fuselage_parasite_drag_form_factor
freestream = conditions.freestream
Sref = fuselage.areas.front_projected
Swet = fuselage.areas.wetted
l_fus = fuselage.lengths.cabin
d_fus = fuselage.width
l_nose = fuselage.lengths.nose
l_tail = fuselage.lengths.tail
# conditions
Mc = freestream.mach_number
roc = freestream.density
muc = freestream.viscosity
Tc = freestream.temperature
pc = freestream.pressure
# reynolds number
V = Mc * compute_speed_of_sound(Tc, pc)
Re_fus = roc * V * l_fus / muc
# 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.sqrt(1 - (1 - Mc**2) * d_d**2)
a = 2 * (1 - Mc**2) * (d_d**2) * (np.arctanh(D) - D) / (D**3)
du_max_u = a / ((2 - a) * (1 - Mc**2)**0.5)
k_fus = (1 + form_factor * du_max_u)**2
# --------------------------------------------------------
# find the final result
fuselage_parasite_drag = k_fus * cf_fus * Swet / Sref
# --------------------------------------------------------
# dump data to conditions
fuselage_result = Result(
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,
)
conditions.aerodynamics.drag_breakdown.parasite[
fuselage.tag] = fuselage_result
return fuselage_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 = Results(
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_wing_supersonic(conditions,configuration,wing):
""" SUAVE.Methods.parasite_drag_wing(conditions,configuration,wing)
computes the parastite drag associated with a wing
Inputs:
Outputs:
Assumptions:
"""
# unpack inputs
C = configuration.wing_parasite_drag_form_factor
freestream = conditions.freestream
Sref = wing.sref
# wing
mac_w = wing.chord_mac
t_c_w = wing.t_c
sweep_w = wing.sweep
arw_w = wing.ar
span_w = wing.span
S_exposed_w = wing.S_exposed # TODO: calculate by fuselage diameter (in Fidelity_Zero.initialize())
S_affected_w = wing.S_affected
# compute wetted area # TODO: calcualte as preprocessing
Swet = 2. * (1.0+ 0.2*t_c_w) * S_exposed_w
# conditions
Mc = freestream.mach_number
roc = freestream.density
muc = freestream.viscosity
Tc = freestream.temperature
pc = freestream.pressure
# reynolds number
V = Mc * compute_speed_of_sound( Tc, pc ) #input gamma and R
Re_w = roc * V * mac_w/muc
# skin friction coefficient
cf_w, k_comp, k_reyn = compressible_turbulent_flat_plate(Re_w,Mc,Tc)
# correction for airfoils
k_w = np.array([[0.0]] * len(Mc))
for i in range(len(Mc)):
if Mc[i] < 0.95:
k_w[i] = 1. + ( 2.* C * (t_c_w * (np.cos(sweep_w))**2.) ) / ( np.sqrt(1.- Mc[i]**2. * ( np.cos(sweep_w))**2.) ) \
+ ( C**2. * (np.cos(sweep_w))**2. * t_c_w**2. * (1. + 5.*(np.cos(sweep_w)**2.)) ) \
/ (2.*(1.-(Mc[i]*np.cos(sweep_w))**2.))
else:
k_w[i] = 1.
# --------------------------------------------------------
# find the final result
wing_parasite_drag = k_w * cf_w * Swet / Sref
# --------------------------------------------------------
# dump data to conditions
wing_result = Result(
wetted_area = Swet ,
reference_area = Sref ,
parasite_drag_coefficient = wing_parasite_drag ,
skin_friction_coefficient = cf_w ,
compressibility_factor = k_comp ,
reynolds_factor = k_reyn ,
form_factor = k_w ,
)
try:
conditions.aerodynamics.drag_breakdown.parasite[wing.tag] = wing_result
except:
print("Drag Polar Mode")
# done!
return wing_parasite_drag
def parasite_drag_propulsor(conditions,configuration,propulsor):
""" SUAVE.Methods.parasite_drag_propulsor(conditions,configuration,propulsor)
computes the parasite drag associated with a propulsor
Inputs:
Outputs:
Assumptions:
"""
# unpack inputs
try:
form_factor = configuration.propulsor_parasite_drag_form_factor
except(AttributeError):
form_factor = 2.3
freestream = conditions.freestream
Sref = propulsor.nacelle_diameter**2 / 4 * np.pi
Swet = propulsor.nacelle_diameter * np.pi * propulsor.engine_length
l_prop = propulsor.engine_length
d_prop = propulsor.nacelle_diameter
# conditions
Mc = freestream.mach_number
roc = freestream.density
muc = freestream.viscosity
Tc = freestream.temperature
pc = freestream.pressure
# reynolds number
V = Mc * compute_speed_of_sound(Tc, pc)
Re_prop = roc * V * l_prop/muc
# skin friction coefficient
cf_prop, k_comp, k_reyn = compressible_turbulent_flat_plate(Re_prop,Mc,Tc)
# form factor for cylindrical bodies
try: # Check if propulsor has an intake
A_max = propulsor.nacelle_dia
A_exit = propulsor.A7
A_inflow = propulsor.Ao
d_d = 1/((propulsor.engine_length + propulsor.D) / np.sqrt(4/np.pi*(A_max - (A_exit+A_inflow)/2)))
except:
d_d = float(d_prop)/float(l_prop)
D = np.sqrt(1 - (1-Mc**2) * d_d**2)
a = 2 * (1-Mc**2) * (d_d**2) *(np.arctanh(D)-D) / (D**3)
du_max_u = a / ( (2-a) * (1-Mc**2)**0.5 )
k_prop = (1 + form_factor*du_max_u)**2
# --------------------------------------------------------
# find the final result
propulsor_parasite_drag = k_prop * cf_prop * Swet / Sref
# --------------------------------------------------------
# dump data to conditions
propulsor_result = Result(
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_propulsor(state,settings,geometry):
#def parasite_drag_propulsor(conditions,configuration,propulsor):
""" SUAVE.Methods.parasite_drag_propulsor(conditions,configuration,propulsor)
computes the parasite drag associated with a propulsor
Inputs:
Outputs:
Assumptions:
"""
# unpack inputs
conditions = state.conditions
configuration = settings
propulsor = geometry
# unpack inputs
try:
form_factor = configuration.propulsor_parasite_drag_form_factor
except(AttributeError):
form_factor = 2.3
freestream = conditions.freestream
Sref = propulsor.nacelle_diameter**2 / 4 * np.pi
Swet = propulsor.nacelle_diameter * np.pi * propulsor.engine_length
l_prop = propulsor.engine_length
d_prop = propulsor.nacelle_diameter
# conditions
Mc = freestream.mach_number
roc = freestream.density
muc = freestream.dynamic_viscosity
Tc = freestream.temperature
pc = freestream.pressure
# reynolds number
V = Mc * compute_speed_of_sound(Tc, pc)
Re_prop = roc * V * l_prop/muc
# skin friction coefficient
cf_prop, k_comp, k_reyn = compressible_turbulent_flat_plate(Re_prop,Mc,Tc)
# form factor for cylindrical bodies
try: # Check if propulsor has an intake
A_max = propulsor.nacelle_diameter
A_exit = propulsor.A7
A_inflow = propulsor.Ao
d_d = 1/((propulsor.engine_length + propulsor.D) / np.sqrt(4/np.pi*(A_max - (A_exit+A_inflow)/2)))
except:
d_d = float(d_prop)/float(l_prop)
D = np.array([[0.0]] * len(Mc))
a = np.array([[0.0]] * len(Mc))
du_max_u = np.array([[0.0]] * len(Mc))
k_prop = 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_prop = (1 + form_factor*du_max_u)**2
# --------------------------------------------------------
# find the final result
propulsor_parasite_drag = k_prop * cf_prop * Swet / Sref
# --------------------------------------------------------
# dump data to conditions
propulsor_result = Results(
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
def parasite_drag_fuselage(conditions,configuration,fuselage):
""" SUAVE.Methods.parasite_drag_fuselage(conditions,configuration,fuselage)
computes the parasite drag associated with a fuselage
Inputs:
Outputs:
Assumptions:
"""
# unpack inputs
form_factor = configuration.fuselage_parasite_drag_form_factor
freestream = conditions.freestream
Sref = fuselage.areas.front_projected
Swet = fuselage.areas.wetted
l_fus = fuselage.lengths.cabin
d_fus = fuselage.width
l_nose = fuselage.lengths.nose
l_tail = fuselage.lengths.tail
# conditions
Mc = freestream.mach_number
roc = freestream.density
muc = freestream.viscosity
Tc = freestream.temperature
pc = freestream.pressure
# reynolds number
V = Mc * compute_speed_of_sound(Tc, pc)
Re_fus = roc * V * l_fus/muc
# 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
#for i in range(len(Mc)):
#if Mc[i] < 0.95:
#D[i] = np.sqrt(1 - (1-Mc[i]**2) * d_d**2)
#a[i] = 2 * (1-Mc[i]**2) * (d_d**2) *(np.arctanh(D[i])-D[i]) / (D[i]**3)
#du_max_u[i] = a[i] / ( (2-a[i]) * (1-Mc[i]**2)**0.5 )
#else:
#D[i] = np.sqrt(1 - d_d**2)
#a[i] = 2 * (d_d**2) *(np.arctanh(D[i])-D[i]) / (D[i]**3)
#du_max_u[i] = a[i] / ( (2-a[i]) )
#k_fus[i] = (1 + cf_fus[i]*du_max_u[i])**2
# --------------------------------------------------------
# find the final result
fuselage_parasite_drag = k_fus * cf_fus * Swet / Sref
# --------------------------------------------------------
# dump data to conditions
fuselage_result = Result(
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:
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 = Results(
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
def parasite_drag_fuselage(state, settings, geometry):
""" SUAVE.Methods.parasite_drag_fuselage(conditions,configuration,fuselage)
computes the parasite drag associated with a fuselage
Inputs:
Outputs:
Assumptions:
"""
# unpack inputs
conditions = state.conditions
configuration = settings
fuselage = geometry
form_factor = configuration.fuselage_parasite_drag_form_factor
freestream = 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.sqrt(1 - (1 - Mc**2) * d_d**2)
a = 2 * (1 - Mc**2) * (d_d**2) * (np.arctanh(D) - D) / (D**3)
du_max_u = a / ((2 - a) * (1 - Mc**2)**0.5)
k_fus = (1 + form_factor * du_max_u)**2
# find the final result
fuselage_parasite_drag = k_fus * cf_fus * Swet / Sref
# dump data to conditions
fuselage_result = Results(
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,
)
state.conditions.aerodynamics.drag_breakdown.parasite[
fuselage.tag] = fuselage_result
return fuselage_parasite_drag
if validation_test:
Re = np.logspace(5, 9, 5)
ii = 0
xts = np.array([0.0, .1, .2, .3, .4, .5, .6, .8, 1.0])
cf_comp = np.zeros([9, 5])
k_comp = np.zeros_like(cf_comp)
k_reyn = np.zeros_like(cf_comp)
for xt in xts:
(cf_comp[ii, :], k_comp[ii, :],
k_reyn[ii, :]) = compressible_mixed_flat_plate(
Re, 0.0, 216.0, xt)
if ii == 0:
(cf_comp_turb, k_comp_t,
k_reyn_t) = compressible_turbulent_flat_plate(Re, 0.0, 216.0)
ii = ii + 1
cf_comp_empirical = np.array([[.0073, .0045, .0031, .0022, .0016],
[.0074, .0044, .0029, .0020, .0015],
[.0073, .0041, .0026, .0018, .0014],
[.0070, .0038, .0024, .0017, .0013],
[.0067, .0036, .0021, .0015, .0011],
[.0064, .0033, .0020, .0013, .0009],
[.0060, .0029, .0017, .0011, .0007],
[.0052, .0022, .0011, .0006, .0004],
[.0041, .0013, .0004, .0001, .00005]])
plt.figure("Skin Friction Coefficient v. Reynolds Number")
axes = plt.gca()
for i in range(len(cf_comp[:, 0])):
def parasite_drag_fuselage(state,settings,geometry):
#def parasite_drag_fuselage(conditions,configuration,fuselage):
""" SUAVE.Methods.parasite_drag_fuselage(conditions,configuration,fuselage)
computes the parasite drag associated with a fuselage
Inputs:
Outputs:
Assumptions:
"""
# unpack inputs
conditions = state.conditions
configuration =settings
#fuselages = geometry.fuselages
fuselage = geometry
form_factor = configuration.fuselage_parasite_drag_form_factor
freestream = conditions.freestream
fuselage_parasite_drag_total = 0.0
#for fuselage in fuselages.values():
Sref = fuselage.areas.front_projected
Swet = fuselage.areas.wetted
l_fus = fuselage.lengths.cabin
d_fus = fuselage.effective_diameter
l_nose = fuselage.lengths.nose
l_tail = fuselage.lengths.tail
# conditions
Mc = freestream.mach_number
roc = freestream.density
muc = freestream.dynamic_viscosity
Tc = freestream.temperature
pc = freestream.pressure
# reynolds number
V = Mc * compute_speed_of_sound(Tc, pc)
Re_fus = roc * V * (l_fus + l_nose + l_tail)/muc
# 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.sqrt(1 - (1-Mc**2) * d_d**2)
a = 2 * (1-Mc**2) * (d_d**2) *(np.arctanh(D)-D) / (D**3)
du_max_u = a / ( (2-a) * (1-Mc**2)**0.5 )
k_fus = (1 + form_factor*du_max_u)**2
# --------------------------------------------------------
# find the final result
fuselage_parasite_drag = k_fus * cf_fus * Swet / Sref
# dump data to conditions
fuselage_result = Results(
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 ,
)
state.conditions.aerodynamics.drag_breakdown.parasite[fuselage.tag] = fuselage_result
return fuselage_parasite_drag
def parasite_drag_propulsor(state,settings,geometry):
""" SUAVE.Methods.parasite_drag_propulsor(conditions,configuration,propulsor)
computes the parasite drag associated with a propulsor
Inputs:
Outputs:
Assumptions:
"""
# unpack inputs
conditions = state.conditions
configuration = settings
propulsor = geometry
# unpack inputs
try:
form_factor = configuration.propulsor_parasite_drag_form_factor
except(AttributeError):
form_factor = 2.3
freestream = conditions.freestream
Sref = propulsor.nacelle_diameter**2 / 4 * np.pi
Swet = propulsor.nacelle_diameter * np.pi * propulsor.engine_length
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 for cylindrical bodies
try: # Check if propulsor has an intake
A_max = propulsor.nacelle_diameter
A_exit = propulsor.A7
A_inflow = propulsor.Ao
d_d = 1/((propulsor.engine_length + propulsor.D) / np.sqrt(4/np.pi*(A_max - (A_exit+A_inflow)/2)))
except:
d_d = float(d_prop)/float(l_prop)
D = np.array([[0.0]] * len(Mc))
a = np.array([[0.0]] * len(Mc))
du_max_u = np.array([[0.0]] * len(Mc))
k_prop = 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_prop = (1 + form_factor*du_max_u)**2
# --------------------------------------------------------
# find the final result
propulsor_parasite_drag = k_prop * cf_prop * Swet / Sref
# --------------------------------------------------------
# dump data to conditions
propulsor_result = Results(
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
def parasite_drag_fuselage(conditions,configuration,fuselage):
""" SUAVE.Methods.parasite_drag_fuselage(conditions,configuration,fuselage)
computes the parasite drag associated with a fuselage
Inputs:
Outputs:
Assumptions:
"""
# unpack inputs
form_factor = configuration.fuselage_parasite_drag_form_factor
C_fus = configuration.fuselage_parasite_drag_form_factor
freestream = conditions.freestream
Sref = fuselage.reference_area
Swet = fuselage.wetted_area
l_fus = fuselage.length_cabin
d_fus = fuselage.width
l_nose = fuselage.length_nose
l_tail = fuselage.length_tail
# conditions
Mc = freestream.mach_number
roc = freestream.density
muc = freestream.viscosity
Tc = freestream.temperature
pc = freestream.pressure
# reynolds number
V = Mc * compute_speed_of_sound(Tc, pc)
Re_fus = roc * V * l_fus/muc
# 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.sqrt(1 - (1-Mc**2) * d_d**2)
a = 2 * (1-Mc**2) * (d_d**2) *(np.arctanh(D)-D) / (D**3)
du_max_u = a / ( (2-a) * (1-Mc**2)**0.5 )
k_fus = (1 + cf_fus*du_max_u)**2
# --------------------------------------------------------
# find the final result
fuselage_parasite_drag = k_fus * cf_fus * Swet / Sref
# --------------------------------------------------------
# dump data to conditions
fuselage_result = Result(
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 ,
)
conditions.aerodynamics.drag_breakdown.parasite[fuselage.tag] = fuselage_result
return fuselage_parasite_drag
def parasite_drag_fuselage(conditions, configuration, fuselage):
""" SUAVE.Methods.parasite_drag_fuselage(conditions,configuration,fuselage)
computes the parasite drag associated with a fuselage
Inputs:
Outputs:
Assumptions:
"""
# unpack inputs
form_factor = configuration.fuselage_parasite_drag_form_factor
freestream = conditions.freestream
Sref = fuselage.areas.front_projected
Swet = fuselage.areas.wetted
l_fus = fuselage.lengths.cabin
d_fus = fuselage.width
l_nose = fuselage.lengths.nose
l_tail = fuselage.lengths.tail
# conditions
Mc = freestream.mach_number
roc = freestream.density
muc = freestream.viscosity
Tc = freestream.temperature
pc = freestream.pressure
# reynolds number
V = Mc * compute_speed_of_sound(Tc, pc)
Re_fus = roc * V * l_fus / muc
# 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
#for i in range(len(Mc)):
#if Mc[i] < 0.95:
#D[i] = np.sqrt(1 - (1-Mc[i]**2) * d_d**2)
#a[i] = 2 * (1-Mc[i]**2) * (d_d**2) *(np.arctanh(D[i])-D[i]) / (D[i]**3)
#du_max_u[i] = a[i] / ( (2-a[i]) * (1-Mc[i]**2)**0.5 )
#else:
#D[i] = np.sqrt(1 - d_d**2)
#a[i] = 2 * (d_d**2) *(np.arctanh(D[i])-D[i]) / (D[i]**3)
#du_max_u[i] = a[i] / ( (2-a[i]) )
#k_fus[i] = (1 + cf_fus[i]*du_max_u[i])**2
# --------------------------------------------------------
# find the final result
fuselage_parasite_drag = k_fus * cf_fus * Swet / Sref
# --------------------------------------------------------
# dump data to conditions
fuselage_result = Result(
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:
conditions.aerodynamics.drag_breakdown.parasite[
fuselage.tag] = fuselage_result
except:
print("Drag Polar Mode fuse parasite")
return fuselage_parasite_drag
# ----------------------------------------------------------------------
# Module Tests
# ----------------------------------------------------------------------
# this will run from command line, put simple tests for your code here
if __name__ == '__main__':
Re = np.logspace(5,9,5)
ii = 0
xts = np.array([0.0,.1,.2,.3,.4,.5,.6,.8,1.0])
cf_comp = np.zeros([9,5])
k_comp = np.zeros_like(cf_comp)
k_reyn = np.zeros_like(cf_comp)
for xt in xts:
(cf_comp[ii,:], k_comp[ii,:], k_reyn[ii,:]) = compressible_mixed_flat_plate(Re,0.0,216.0,xt)
if ii == 0:
(cf_comp_turb,k_comp_t,k_reyn_t) = compressible_turbulent_flat_plate(Re,0.0,216.0)
ii = ii + 1
cf_comp_empirical = np.array([[.0073,.0045,.0031,.0022,.0016],
[.0074,.0044,.0029,.0020,.0015],
[.0073,.0041,.0026,.0018,.0014],
[.0070,.0038,.0024,.0017,.0013],
[.0067,.0036,.0021,.0015,.0011],
[.0064,.0033,.0020,.0013,.0009],
[.0060,.0029,.0017,.0011,.0007],
[.0052,.0022,.0011,.0006,.0004],
[.0041,.0013,.0004,.0001,.00005]])
plt.figure("Skin Friction Coefficient v. Reynolds Number")
axes = plt.gca()
