def wave_drag(conditions,configuration,main_fuselage,propulsor,wing,num_engines,i_wing,Sref_main,flag105):
# Use wave drag to determine compressibility drag for supersonic speeds
# Unpack mach number
mach = conditions.freestream.mach_number
# Create a copy that can be modified
Mc = copy.copy(mach)
# This flag is for the interpolation mode
if flag105 is True:
# Change conditions and short hand for calculations
conditions.freestream.mach_number = np.array([[1.05]] * len(Mc))
mach = conditions.freestream.mach_number
# Initalize cd array
cd_c = np.array([[0.0]] * len(mach))
# Calculate wing values at all mach numbers
# Note that these functions arrange the supersonic values at the beginning of the array
cd_lift_wave = wave_drag_lift(conditions,configuration,wing)
cd_volume_wave = wave_drag_volume(conditions,configuration,wing)
# Pack supersonic results into correct elements
cd_c[mach >= 1.05] = cd_lift_wave[0:len(mach[mach >= 1.05]),0] + cd_volume_wave[0:len(mach[mach >= 1.05]),0]
# Convert coefficient to full aircraft value
if i_wing != 0:
cd_c = cd_c*wing.areas.reference/Sref_main
# Include fuselage and propulsors for one iteration
if i_wing == 0:
# Initialize arrays
prop_drag = np.array([[0.0]] * len(mach))
fuse_drag = np.array([[0.0]] * len(mach))
# Fuselage wave drag
if len(main_fuselage) > 0:
fuse_drag[mach >= 1.05] = wave_drag_body_of_rev(main_fuselage.lengths.total,main_fuselage.effective_diameter/2.0,Sref_main)
else:
raise ValueError('Main fuselage does not have a total length')
# Propulsor wave drag
prop_drag[mach >= 1.05] = wave_drag_body_of_rev(propulsor.lengths.engine_total,propulsor.nacelle_dia/2.0,Sref_main)*propulsor.number_of_engines
# Pack values
cd_c[mach >= 1.05] = cd_c[mach >= 1.05] + fuse_drag[mach >= 1.05]
cd_c[mach >= 1.05] = cd_c[mach >= 1.05] + prop_drag[mach >= 1.05]
# Return dummy values for mcc and MDiv
mcc = np.array([[0.0]] * len(mach))
MDiv = np.array([[0.0]] * len(mach))
# Reset mach number to real values
conditions.freestream.mach_number = Mc
return (cd_c,mcc,MDiv)
def wave_drag(conditions, configuration, main_fuselage, propulsor, wing,
num_engines, k, Sref_main, flag105):
# Use wave drag to determine compressibility drag for supersonic speeds
# Unpack mach number
mach = conditions.freestream.mach_number
# Create a copy that can be modified
Mc = copy.copy(mach)
# This flag is for the interpolation mode
if flag105 is True:
# Change conditions and short hand for calculations
conditions.freestream.mach_number = np.array([[1.05]] * len(Mc))
mach = conditions.freestream.mach_number
# Initalize cd array
cd_c = np.array([[0.0]] * len(mach))
# Calculate wing values at all mach numbers
# Note that these functions arrange the supersonic values at the beginning of the array
cd_lift_wave = wave_drag_lift(conditions, configuration, wing)
cd_volume_wave = wave_drag_volume(conditions, configuration, wing)
# Pack supersonic results into correct elements
cd_c[mach >= 1.05] = cd_lift_wave[0:len(mach[
mach >= 1.05]), 0] + cd_volume_wave[0:len(mach[mach >= 1.05]), 0]
# Convert coefficient to full aircraft value
if k != 'main_wing':
cd_c = cd_c * wing.areas.reference / Sref_main
# Include fuselage and propulsors for one iteration
# Return dummy values for mcc and MDiv
mcc = np.array([[0.0]] * len(mach))
MDiv = np.array([[0.0]] * len(mach))
# Reset mach number to real values
conditions.freestream.mach_number = Mc
return (cd_c, mcc, MDiv)
def compressibility_drag_wing(conditions, configuration, geometry):
""" SUAVE.Methods.compressibility_drag_wing(conditions,configuration,geometry)
computes the induced drag associated with a wing
Inputs:
Outputs:
Assumptions:
based on a set of fits
"""
# unpack
wings = geometry.Wings
fuselages = geometry.Fuselages
wing_lifts = conditions.aerodynamics.lift_breakdown.compressible_wings # currently the total aircraft lift
mach = conditions.freestream.mach_number
drag_breakdown = conditions.aerodynamics.drag_breakdown
# start result
total_compressibility_drag = 0.0
drag_breakdown.compressible = Results()
# go go go
for i_wing, wing, in enumerate(wings.values()):
# unpack wing
t_c_w = wing.t_c
sweep_w = wing.sweep
Mc = copy.copy(mach)
for ii in range(len(Mc)):
if Mc[ii] > 0.95 and Mc[ii] < 1.05:
Mc[ii] = 0.95
if Mc[ii] <= 0.95:
if i_wing == 0:
cl_w = wing_lifts
else:
cl_w = 0
# get effective Cl and sweep
tc = t_c_w / (np.cos(sweep_w))
cl = cl_w / (np.cos(sweep_w))**2
# compressibility drag based on regressed fits from AA241
mcc_cos_ws = 0.922321524499352 \
- 1.153885166170620*tc \
- 0.304541067183461*cl \
+ 0.332881324404729*tc**2 \
+ 0.467317361111105*tc*cl \
+ 0.087490431201549*cl**2
# crest-critical mach number, corrected for wing sweep
mcc = mcc_cos_ws / np.cos(sweep_w)
# divergence mach number
MDiv = mcc * (1.02 + 0.08 * (1 - np.cos(sweep_w)))
# divergence ratio
mo_mc = mach / mcc
# compressibility correlation, Shevell
dcdc_cos3g = 0.0019 * mo_mc**14.641
# compressibility drag
cd_c = dcdc_cos3g * (np.cos(sweep_w))**3
else:
cd_lift_wave = wave_drag_lift(conditions, configuration, wing)
cd_volume_wave = wave_drag_volume(conditions, configuration,
wing)
cd_c = cd_lift_wave + cd_volume_wave
tc = 0
sweep_w = 0
mcc = 0
MDiv = 0
# increment
#total_compressibility_drag += cd_c ## todo when there is a lift break down by wing
# dump data to conditions
wing_results = Results(
compressibility_drag=cd_c,
thickness_to_chord=tc,
wing_sweep=sweep_w,
crest_critical=mcc,
divergence_mach=MDiv,
)
#drag_breakdown.compressible[wing.tag] = wing_results
#: for each wing
# dump total comp drag
total_compressibility_drag = drag_breakdown.compressible[
1 + 0].compressibility_drag
drag_breakdown.compressible.total = total_compressibility_drag
return total_compressibility_drag, wing_results
def wave_drag(conditions, configuration, main_fuselage, propulsor, wing,
num_engines, i_wing, Sref_main, flag105):
# Use wave drag to determine compressibility drag for supersonic speeds
# Unpack mach number
mach = conditions.freestream.mach_number
# Create a copy that can be modified
Mc = copy.copy(mach)
# This flag is for the interpolation mode
if flag105 is True:
# Change conditions and short hand for calculations
conditions.freestream.mach_number = np.array([[1.05]] * len(Mc))
mach = conditions.freestream.mach_number
# Initalize cd array
cd_c = np.array([[0.0]] * len(mach))
# Calculate wing values at all mach numbers
# Note that these functions arrange the supersonic values at the beginning of the array
cd_lift_wave = wave_drag_lift(conditions, configuration, wing)
cd_volume_wave = wave_drag_volume(conditions, configuration, wing)
# Pack supersonic results into correct elements
cd_c[mach >= 1.05] = cd_lift_wave[0:len(mach[
mach >= 1.05]), 0] + cd_volume_wave[0:len(mach[mach >= 1.05]), 0]
# Convert coefficient to full aircraft value
if i_wing != 0:
cd_c = cd_c * wing.areas.reference / Sref_main
# Include fuselage and propulsors for one iteration
if i_wing == 0:
# Initialize arrays
prop_drag = np.array([[0.0]] * len(mach))
fuse_drag = np.array([[0.0]] * len(mach))
# Fuselage wave drag
if len(main_fuselage) > 0:
fuse_drag[mach >= 1.05] = wave_drag_body_of_rev(
main_fuselage.lengths.total,
main_fuselage.effective_diameter / 2.0, Sref_main)
else:
raise ValueError('Main fuselage does not have a total length')
# Propulsor wave drag
prop_drag[mach >= 1.05] = wave_drag_body_of_rev(
propulsor.lengths.engine_total, propulsor.nacelle_dia / 2.0,
Sref_main) * propulsor.number_of_engines
# Pack values
cd_c[mach >= 1.05] = cd_c[mach >= 1.05] + fuse_drag[mach >= 1.05]
cd_c[mach >= 1.05] = cd_c[mach >= 1.05] + prop_drag[mach >= 1.05]
# Return dummy values for mcc and MDiv
mcc = np.array([[0.0]] * len(mach))
MDiv = np.array([[0.0]] * len(mach))
# Reset mach number to real values
conditions.freestream.mach_number = Mc
return (cd_c, mcc, MDiv)
def compressibility_drag_wing(conditions,configuration,geometry):
""" SUAVE.Methods.compressibility_drag_wing(conditions,configuration,geometry)
computes the induced drag associated with a wing
Inputs:
Outputs:
Assumptions:
based on a set of fits
"""
# unpack
wings = geometry.Wings
fuselages = geometry.Fuselages
wing_lifts = conditions.aerodynamics.lift_breakdown.compressible_wings # currently the total aircraft lift
mach = conditions.freestream.mach_number
drag_breakdown = conditions.aerodynamics.drag_breakdown
# start result
total_compressibility_drag = 0.0
drag_breakdown.compressible = Results()
# go go go
for i_wing, wing, in enumerate(wings.values()):
# unpack wing
t_c_w = wing.t_c
sweep_w = wing.sweep
Mc = copy.copy(mach)
for ii in range(len(Mc)):
if Mc[ii] > 0.95 and Mc[ii] < 1.05:
Mc[ii] = 0.95
if Mc[ii] <= 0.95:
if i_wing == 0:
cl_w = wing_lifts
else:
cl_w = 0
# get effective Cl and sweep
tc = t_c_w /(np.cos(sweep_w))
cl = cl_w / (np.cos(sweep_w))**2
# compressibility drag based on regressed fits from AA241
mcc_cos_ws = 0.922321524499352 \
- 1.153885166170620*tc \
- 0.304541067183461*cl \
+ 0.332881324404729*tc**2 \
+ 0.467317361111105*tc*cl \
+ 0.087490431201549*cl**2
# crest-critical mach number, corrected for wing sweep
mcc = mcc_cos_ws / np.cos(sweep_w)
# divergence mach number
MDiv = mcc * ( 1.02 + 0.08*(1 - np.cos(sweep_w)) )
# divergence ratio
mo_mc = mach/mcc
# compressibility correlation, Shevell
dcdc_cos3g = 0.0019*mo_mc**14.641
# compressibility drag
cd_c = dcdc_cos3g * (np.cos(sweep_w))**3
else:
cd_lift_wave = wave_drag_lift(conditions,configuration,wing)
cd_volume_wave = wave_drag_volume(conditions,configuration,wing)
cd_c = cd_lift_wave + cd_volume_wave
tc = 0
sweep_w = 0
mcc = 0
MDiv = 0
# increment
#total_compressibility_drag += cd_c ## todo when there is a lift break down by wing
# dump data to conditions
wing_results = Results(
compressibility_drag = cd_c ,
thickness_to_chord = tc ,
wing_sweep = sweep_w ,
crest_critical = mcc ,
divergence_mach = MDiv ,
)
#drag_breakdown.compressible[wing.tag] = wing_results
#: for each wing
# dump total comp drag
total_compressibility_drag = drag_breakdown.compressible[1+0].compressibility_drag
drag_breakdown.compressible.total = total_compressibility_drag
return total_compressibility_drag, wing_results
def wave_drag(conditions,configuration,main_fuselage,propulsor,wing,num_engines,k,Sref_main,flag105):
"""Use wave drag to determine compressibility drag for supersonic speeds
Assumptions:
Basic fit
Source:
adg.stanford.edu (Stanford AA241 A/B Course Notes)
Inputs:
conditions.freestream.mach_number [Unitless]
configuration
main_fuselage (unused)
propulsor (unused)
wing.areas.reference [m^2]
num_engines (unused)
k (tag for wing) [String]
Sref_main (main reference area) [m^2]
flag105 (check if calcs are for Mach 1.05) [Boolean]
Outputs:
cd_c [Unitless]
mcc [Unitless]
MDiv [Unitless]
Properties Used:
N/A
"""
# Unpack mach number
mach = conditions.freestream.mach_number
# Create a copy that can be modified
Mc = copy.copy(mach)
# This flag is for the interpolation mode
if flag105 is True:
# Change conditions and short hand for calculations
conditions.freestream.mach_number = np.array([[1.05]] * len(Mc))
mach = conditions.freestream.mach_number
# Initalize cd arrays
cd_c = np.array([[0.0]] * len(mach))
cd_c_l = np.array([[0.0]] * len(mach)) # lift wave drag
cd_c_v = np.array([[0.0]] * len(mach)) # vol wave drag
# Calculate wing values at all mach numbers
# Note that these functions arrange the supersonic values at the beginning of the array
cd_lift_wave = wave_drag_lift(conditions,configuration,wing)
cd_volume_wave = wave_drag_volume(conditions,configuration,wing)
# Pack supersonic results into correct elements
cd_c[mach >= 1.05] = cd_lift_wave[0:len(mach[mach >= 1.05]),0] + cd_volume_wave[0:len(mach[mach >= 1.05]),0]
cd_c_l[mach >= 1.05] = cd_lift_wave[0:len(mach[mach >= 1.05]),0]
cd_c_v[mach >= 1.05] = cd_volume_wave[0:len(mach[mach >= 1.05]),0]
# Convert coefficient to full aircraft value
cd_c = cd_c*wing.areas.reference/Sref_main
cd_c_l = cd_c_l*wing.areas.reference/Sref_main
cd_c_v = cd_c_v*wing.areas.reference/Sref_main
# Include fuselage and propulsors for one iteration
# Return dummy values for mcc and MDiv
mcc = np.array([[0.0]] * len(mach))
MDiv = np.array([[0.0]] * len(mach))
# Reset mach number to real values
conditions.freestream.mach_number = Mc
return (cd_c,mcc,MDiv,cd_c_l,cd_c_v)