Source code for RCAIDE.Framework.Analyses.Stability.Athena_Vortex_Lattice
# RCAIDE/Framework/Analyses/Stability/Athena_Vortex_Lattice.py
#
# Created: Jul 2023, M. Clarke
# ----------------------------------------------------------------------------------------------------------------------
# IMPORT
# ----------------------------------------------------------------------------------------------------------------------
# RCAIDE imports
from RCAIDE.Framework.Core import Data, Units
from RCAIDE.Framework.Analyses import Process
from RCAIDE.Library.Methods.Aerodynamics import Common
from .Stability import Stability
from RCAIDE.Framework.Analyses.Common.Process_Geometry import Process_Geometry
from RCAIDE.Library.Methods.Aerodynamics.Athena_Vortex_Lattice import *
from RCAIDE.Library.Methods.Aerodynamics.Athena_Vortex_Lattice.AVL_Objects.Run_Case import Run_Case
# package imports
import numpy as np
# ----------------------------------------------------------------------------------------------------------------------
# Athena_Vortex_Lattice
# ----------------------------------------------------------------------------------------------------------------------
[docs]
class Athena_Vortex_Lattice(Stability):
"""This is a subsonic aerodynamic buildup analysis based on the vortex lattice method
Assumptions:
Stall effects are negligible
Source:
N/A
Inputs:
None
Outputs:
None
Properties Used:
N/A
"""
def __defaults__(self):
"""This sets the default values and methods for the analysis.
Assumptions:
None
Source:
N/A
Inputs:
None
Outputs:
None
Properties Used:
N/A
"""
self.tag = 'Athena_Vortex_Lattice'
self.vehicle = Data()
self.process = Process()
self.process.initialize = Process()
# correction factors
settings = self.settings
settings.fuselage_lift_correction = 1.14
settings.trim_drag_correction_factor = 1.1
settings.wing_parasite_drag_form_factor = 1.2
settings.fuselage_parasite_drag_form_factor = 2.4
settings.drag_reduction_factors = Data()
settings.drag_reduction_factors.parasite_drag = 0.0 # Reduction factors are proportional (.1 is a 10% weight reduction)
settings.drag_reduction_factors.induced_drag = 0.0 # Reduction factors are proportional (.1 is a 10% weight reduction)
settings.drag_reduction_factors.compressibility_drag = 0.0 # Reduction factors are proportional (.1 is a 10% weight reduction)
settings.maximum_lift_coefficient_factor = 1.0
settings.lift_to_drag_adjustment = 0.0
settings.viscous_lift_dependent_drag_factor = 0.38
settings.drag_coefficient_increment = 0.0
settings.spoiler_drag_increment = 0.0
settings.maximum_lift_coefficient = np.inf
settings.recalculate_total_wetted_area = False
settings.oswald_efficiency_factor = None
settings.span_efficiency = None
settings.run_cases = Run_Case.Container()
settings.flow_symmetry = Data()
settings.flow_symmetry.xz_plane = 0 # Symmetry across the xz-plane, y=0
settings.flow_symmetry.xy_parallel = 0 # Symmetry across the z=z_symmetry_plane plane
settings.flow_symmetry.z_symmetry_plane = 0.0
settings.number_of_control_surfaces = 0
settings.filenames = Data()
settings.filenames.avl_bin_name = 'avl' # to call avl from command line. If avl is not on the system path, include absolute path to the avl binary i.e. '/your/path/to/avl'
settings.filenames.run_folder = 'avl_files'
settings.filenames.features = 'aircraft.avl'
settings.filenames.mass_file = 'aircraft.mass'
settings.filenames.batch_template = 'batch_{0:04d}.run'
settings.filenames.deck_template = 'commands_{0:04d}.deck'
settings.filenames.aero_output_template_1 = 'stability_axis_derivatives_{}.txt'
settings.filenames.aero_output_template_2 = 'surface_forces_{}.txt'
settings.filenames.aero_output_template_3 = 'strip_forces_{}.txt'
settings.filenames.aero_output_template_4 = 'body_axis_derivatives_{}.txt'
settings.filenames.dynamic_output_template_1 = 'eigen_mode_{}.txt'
settings.filenames.dynamic_output_template_2 = 'system_matrix_{}.txt'
settings.filenames.case_template = 'case_{0:04d}_{1:04d}'
settings.filenames.log_filename = 'avl_log.txt'
settings.filenames.err_filename = 'avl_err.txt'
settings.number_of_spanwise_vortices = 20
settings.number_of_chordwise_vortices = 10
settings.use_surrogate = True
settings.trim_aircraft = False
settings.model_fuselage = False
settings.print_output = False
settings.keep_files = False
settings.new_regression_results = False
settings.side_slip_angle = 0.0
settings.roll_rate_coefficient = 0.0
settings.pitch_rate_coefficient = 0.0
settings.lift_coefficient = None
# correction factors
settings.supersonic = Data()
settings.supersonic.peak_mach_number = 1.04
settings.supersonic.begin_drag_rise_mach_number = 0.95
settings.supersonic.end_drag_rise_mach_number = 1.2
settings.supersonic.transonic_drag_multiplier = 1.25
settings.supersonic.volume_wave_drag_scaling = 3.2
settings.supersonic.fuselage_parasite_drag_begin_blend_mach = 0.91
settings.supersonic.fuselage_parasite_drag_end_blend_mach = 0.99
settings.supersonic.cross_sectional_area_calculation_type = 'Fixed'
settings.supersonic.wave_drag_type = 'Raymer'
self.reference_values = Data()
self.reference_values.S_ref = 0
self.reference_values.c_ref = 0
self.reference_values.b_ref = 0
self.reference_values.X_ref = 0
self.reference_values.Y_ref = 0
self.reference_values.Z_ref = 0
# conditions table, used for surrogate model training
self.training = Data()
self.training.angle_of_attack = np.array([-2.,0., 2.,5., 7., 10.])*Units.degrees
self.training.Mach = np.array([0.05,0.15,0.25, 0.45,0.65,0.85])
self.training.lift_coefficient = None
self.training.drag_coefficient = None
self.training.span_efficiency_factor = None
self.training.moment_coefficient = None
self.training.Cm_alpha_moment_coefficient = None
self.training.Cn_beta_moment_coefficient = None
self.training.neutral_point = None
self.training_file = None
self.current_status = Data()
self.current_status.batch_index = 0
self.current_status.batch_file = None
self.current_status.deck_file = None
self.current_status.cases = None
# surrogoate models
self.surrogates = Data()
self.surrogates.moment_coefficient = None
self.surrogates.Cm_alpha_moment_coefficient = None
self.surrogates.Cn_beta_moment_coefficient = None
self.surrogates.neutral_point = None
# build the evaluation process
compute = Process()
compute.lift = Process()
compute.lift.inviscid_wings = None
compute.lift.fuselage = Common.Lift.fuselage_correction
compute.drag = Process()
compute.drag.parasite = Process()
compute.drag.parasite.wings = Process_Geometry('wings')
compute.drag.parasite.wings.wing = Common.Drag.parasite_drag_wing
compute.drag.parasite.fuselages = Process_Geometry('fuselages')
compute.drag.parasite.fuselages.fuselage = Common.Drag.parasite_drag_fuselage
compute.drag.parasite.booms = Process_Geometry('booms')
compute.drag.parasite.booms.boom = Common.Drag.parasite_drag_fuselage
compute.drag.parasite.nacelles = Common.Drag.parasite_drag_nacelle
compute.drag.parasite.pylons = Common.Drag.parasite_drag_pylon
compute.drag.parasite.total = Common.Drag.parasite_total
compute.drag.induced = Common.Drag.induced_drag
compute.drag.cooling = Process()
compute.drag.cooling.total = Common.Drag.cooling_drag
compute.drag.compressibility = Process()
compute.drag.compressibility.total = Common.Drag.compressibility_drag
compute.drag.miscellaneous = Common.Drag.miscellaneous_drag
compute.drag.spoiler = Common.Drag.spoiler_drag
compute.drag.total = Common.Drag.total_drag
self.process.compute = compute
[docs]
def initialize(self):
use_surrogate = self.settings.use_surrogate
# If we are using the surrogate
if use_surrogate == True:
# sample training data
train_AVL_surrogates(self)
# build surrogate
build_AVL_surrogates(self)
# build the evaluation process
compute = self.process.compute
if use_surrogate == True:
compute.lift.inviscid_wings = evaluate_AVL_surrogate
else:
compute.lift.inviscid_wings = evaluate_AVL_no_surrogate
return
[docs]
def evaluate(self,state):
"""The default evaluate function.
Assumptions:
None
Source:
N/A
Inputs:
None
Outputs:
results <RCAIDE data class>
Properties Used:
self.settings
self.vehicle
"""
settings = self.settings
vehicle = self.vehicle
results = self.process.compute(state,settings,vehicle)
return results
