# RCAIDE/Library/Methods/Aerodynamics/Athena_Vortex_Lattice/translate_data.py
#
# Created: Oct 2024, M. Clarke
# ----------------------------------------------------------------------------------------------------------------------
# IMPORT
# ----------------------------------------------------------------------------------------------------------------------
# RCAIDE imports
from RCAIDE.Framework.Core import Units, Data
from .AVL_Objects.Run_Case import Run_Case
# package imports
import numpy as np
# ----------------------------------------------------------------------------------------------------------------------
# translate_data
# ----------------------------------------------------------------------------------------------------------------------
[docs]
def translate_conditions_to_cases(avl ,conditions, vehicle):
""" Takes RCAIDE Conditions() data structure and translates to a Container of
avl Run_Case()s.
Assumptions:
None
Source:
Drela, M. and Youngren, H., AVL, http://web.mit.edu/drela/Public/web/avl
Inputs:
conditions.aerodynamics.angles.alpha [radians]
conditions.freestream.mach_number [-]
conditions.freestream.density [kilograms per meters**3]
conditions.freestream.gravity [meters per second**2]
Outputs:
cases [data structur]
Properties Used:
N/A
"""
# set up aerodynamic Conditions object
cases = Run_Case.Container()
for i in range(len(conditions.aerodynamics.angles.alpha)):
case = Run_Case()
case.tag = avl.settings.filenames.case_template.format(avl.current_status.batch_index,i+1)
case.mass = conditions.weights.vehicle.mass
case.conditions.freestream.mach = conditions.freestream.mach_number[i, 0]
case.conditions.freestream.density = conditions.freestream.density[i, 0]
case.conditions.freestream.gravitational_acceleration = conditions.freestream.gravity[i, 0]
case.conditions.aerodynamics.angles.alpha = conditions.aerodynamics.angles.alpha[i, 0]/Units.deg
case.conditions.aerodynamics.angles.beta = conditions.aerodynamics.angles.beta[i, 0]/Units.deg
if type(conditions.aerodynamics.coefficients.lift.inviscid.total) == np.ndarray:
case.conditions.aerodynamics.coefficients.lift.inviscid.total= conditions.aerodynamics.coefficients.lift.inviscid.total[i, 0]
else:
case.conditions.aerodynamics.coefficients.lift.inviscid.total = None
case.conditions.static_stability.coefficients.roll = conditions.static_stability.coefficients.roll[i, 0]
case.conditions.static_stability.coefficients.pitch = conditions.static_stability.coefficients.pitch[i, 0]
# determine the number of wings
n_wings = 0
for wing in vehicle.wings:
n_wings += 1
if wing.xz_plane_symmetric == True:
n_wings += 1
case.num_wings = n_wings
case.n_sw = avl.settings.number_of_spanwise_vortices
cases.append_case(case)
return cases
[docs]
def translate_results_to_conditions(cases,res,results, settings):
""" Takes avl results structure containing the results of each run case stored
each in its own Data() object. Translates into the Conditions() data structure.
Assumptions:
None
Source:
Drela, M. and Youngren, H., AVL, http://web.mit.edu/drela/Public/web/avl
Inputs:
case_res = results
Outputs:
cases [data_structure]
Properties Used:
N/A
"""
num_wings = cases[0].num_wings
n_sw = cases[0].n_sw
dim = len(cases)
wing_areas = np.zeros((dim,num_wings))
wing_local_spans = np.zeros((dim,num_wings,n_sw))
wing_section_chords = np.zeros((dim,num_wings,n_sw))
leading_edge_sweeps = np.zeros((dim,num_wings,n_sw))
# aero results 1: total surface forces and coefficeints
res.aerodynamics.coefficients.lift.inviscid.total = np.zeros((dim,1))
res.aerodynamics.wing_CLs = np.zeros_like(wing_areas)
res.aerodynamics.wing_CDs = np.zeros_like(wing_areas)
# aero results 2 : sectional forces and coefficients
res.aerodynamics.coefficients.lift.spanwise = np.zeros((dim,num_wings*n_sw))
res.aerodynamics.spanwise_induced_angle = np.zeros((dim,num_wings*n_sw))
res.aerodynamics.coefficients.drag.spanwise = np.zeros((dim,num_wings*n_sw))
res.static_stability.control_surfaces_cases = {}
mach_case = list(results.keys())[0][5:9]
for i in range(len(results.keys())):
aoa_case = '{:04d}'.format(i+1)
tag = 'case_' + mach_case + '_' + aoa_case
case_res = results[tag]
# stability file
res.aerodynamics.angles.alpha[i][0] = case_res.aerodynamics.AoA * Units.degree
res.aerodynamics.angles.beta[i][0] = case_res.aerodynamics.beta * Units.degree
res.static_stability.coefficients.X[i][0] = case_res.aerodynamics.CX
res.static_stability.coefficients.Y[i][0] = case_res.aerodynamics.CY
res.static_stability.coefficients.Z[i][0] = case_res.aerodynamics.CZ
res.static_stability.coefficients.L[i][0] = case_res.aerodynamics.Cltot
res.static_stability.coefficients.M[i][0] = case_res.aerodynamics.Cmtot
res.static_stability.coefficients.N[i][0] = case_res.aerodynamics.Cntot
res.static_stability.coefficients.roll[i][0] = case_res.aerodynamics.roll_moment_coefficient
res.static_stability.coefficients.pitch[i][0] = case_res.aerodynamics.pitch_moment_coefficient
res.static_stability.coefficients.yaw[i][0] = case_res.aerodynamics.yaw_moment_coefficient
res.aerodynamics.coefficients.lift.inviscid.total[i][0] = case_res.aerodynamics.total_lift_coefficient
res.aerodynamics.coefficients.drag.induced.inviscid[i][0] = case_res.aerodynamics.induced_drag_coefficient
res.aerodynamics.coefficients.drag.induced.efficiency_factor[i][0] = case_res.aerodynamics.oswald_efficiency
res.aerodynamics.oswald_efficiency[i][0] = case_res.aerodynamics.oswald_efficiency
res.static_stability.derivatives.Clift_alpha[i][0] = case_res.stability.alpha_derivatives.lift_curve_slope
res.static_stability.derivatives.CY_alpha[i][0] = case_res.stability.alpha_derivatives.side_force_derivative
res.static_stability.derivatives.CL_alpha[i][0] = case_res.stability.alpha_derivatives.roll_moment_derivative
res.static_stability.derivatives.CM_alpha[i][0] = case_res.stability.alpha_derivatives.pitch_moment_derivative
res.static_stability.derivatives.CN_alpha[i][0] = case_res.stability.alpha_derivatives.yaw_moment_derivative
res.static_stability.derivatives.Clift_beta[i][0] = case_res.stability.beta_derivatives.lift_coefficient_derivative
res.static_stability.derivatives.CY_beta[i][0] = case_res.stability.beta_derivatives.side_force_derivative
res.static_stability.derivatives.CL_beta[i][0] = case_res.stability.beta_derivatives.roll_moment_derivative
res.static_stability.derivatives.CM_beta[i][0] = case_res.stability.beta_derivatives.pitch_moment_derivative
res.static_stability.derivatives.CN_beta[i][0] = case_res.stability.beta_derivatives.yaw_moment_derivative
res.static_stability.derivatives.Clift_p[i][0] = case_res.stability.CL_p
res.static_stability.derivatives.Clift_q[i][0] = case_res.stability.CL_q
res.static_stability.derivatives.Clift_r[i][0] = case_res.stability.CL_r
res.static_stability.derivatives.CY_p[i][0] = case_res.stability.CY_p
res.static_stability.derivatives.CY_q[i][0] = case_res.stability.CY_q
res.static_stability.derivatives.CY_r[i][0] = case_res.stability.CY_r
res.static_stability.derivatives.CL_p[i][0] = case_res.stability.Cl_p
res.static_stability.derivatives.CL_q[i][0] = case_res.stability.Cl_q
res.static_stability.derivatives.CL_r[i][0] = case_res.stability.Cl_r
res.static_stability.derivatives.CM_p[i][0] = case_res.stability.Cm_p
res.static_stability.derivatives.CM_q[i][0] = case_res.stability.Cm_q
res.static_stability.derivatives.CM_r[i][0] = case_res.stability.Cm_r
res.static_stability.derivatives.CN_p[i][0] = case_res.stability.Cn_p
res.static_stability.derivatives.CN_q[i][0] = case_res.stability.Cn_q
res.static_stability.derivatives.CN_r[i][0] = case_res.stability.Cn_r
res.static_stability.derivatives.CX_u[i][0] = case_res.stability.CX_u
res.static_stability.derivatives.CX_v[i][0] = case_res.stability.CX_v
res.static_stability.derivatives.CX_w[i][0] = case_res.stability.CX_w
res.static_stability.derivatives.CY_u[i][0] = case_res.stability.CY_u
res.static_stability.derivatives.CY_v[i][0] = case_res.stability.CY_v
res.static_stability.derivatives.CY_w[i][0] = case_res.stability.CY_w
res.static_stability.derivatives.CZ_u[i][0] = case_res.stability.CZ_u
res.static_stability.derivatives.CZ_v[i][0] = case_res.stability.CZ_v
res.static_stability.derivatives.CZ_w[i][0] = case_res.stability.CZ_w
res.static_stability.derivatives.CL_u[i][0] = case_res.stability.Cl_u
res.static_stability.derivatives.CL_v[i][0] = case_res.stability.Cl_v
res.static_stability.derivatives.CL_w[i][0] = case_res.stability.Cl_w
res.static_stability.derivatives.CM_u[i][0] = case_res.stability.Cm_u
res.static_stability.derivatives.CM_v[i][0] = case_res.stability.Cm_v
res.static_stability.derivatives.CM_w[i][0] = case_res.stability.Cm_w
res.static_stability.derivatives.CN_u[i][0] = case_res.stability.Cn_u
res.static_stability.derivatives.CN_v[i][0] = case_res.stability.Cn_v
res.static_stability.derivatives.CN_w[i][0] = case_res.stability.Cn_w
res.static_stability.derivatives.CX_p[i][0] = case_res.stability.CX_p
res.static_stability.derivatives.CX_q[i][0] = case_res.stability.CX_q
res.static_stability.derivatives.CX_r[i][0] = case_res.stability.CX_r
res.static_stability.derivatives.CY_p[i][0] = case_res.stability.CY_p
res.static_stability.derivatives.CY_q[i][0] = case_res.stability.CY_q
res.static_stability.derivatives.CY_r[i][0] = case_res.stability.CY_r
res.static_stability.derivatives.CZ_p[i][0] = case_res.stability.CZ_p
res.static_stability.derivatives.CZ_q[i][0] = case_res.stability.CZ_q
res.static_stability.derivatives.CZ_r[i][0] = case_res.stability.CZ_r
res.static_stability.neutral_point[i][0] = case_res.stability.neutral_point
res.static_stability.spiral_criteria[i][0] = case_res.stability.spiral_criteria
# aero surface forces file
wing_areas[i][:] = case_res.aerodynamics.wing_areas
res.aerodynamics.wing_CLs[i][:] = case_res.aerodynamics.wing_CLs
res.aerodynamics.wing_CDs[i][:] = case_res.aerodynamics.wing_CDs
# aero sectional forces file
wing_local_spans[i][:] = case_res.aerodynamics.wing_local_spans
wing_section_chords[i][:] = case_res.aerodynamics.wing_section_chords
leading_edge_sweeps[i][:] = case_res.aerodynamics.leading_edge_sweeps
CL_y = case_res.aerodynamics.wing_section_cls
CD_y = case_res.aerodynamics.wing_section_cds
AOAi = case_res.aerodynamics.wing_section_aoa_i
res.aerodynamics.coefficients.lift.spanwise[i][:] = CL_y.reshape(CL_y.shape[:-2] + (-1,))
res.aerodynamics.spanwise_induced_angle[i][:] = CD_y.reshape(CD_y.shape[:-2] + (-1,))
res.aerodynamics.coefficients.drag.spanwise[i][:] = AOAi.reshape(AOAi.shape[:-2] + (-1,))
res.static_stability.control_surfaces_cases[tag] = case_res.stability.control_surfaces
for cs in settings.control_surface_tags:
cs_res = case_res.stability.control_surfaces.control_surfaces[cs]
if cs == 'flap':
letter = 'f'
if cs == 'slat':
letter = 's'
if cs == 'rudder':
letter = 'r'
if cs == 'elevator':
letter = 'e'
if cs == 'aileron':
letter = 'a'
lift_derivative = 'Clift_delta_' + letter
drag_derivative = 'Cdrag_induced_delta_' + letter
L_derivative = 'CL_delta_' + letter
M_derivative = 'CM_delta_' + letter
N_derivative = 'CN_delta_' + letter
Y_derivative = 'CY_delta_' + letter
res.static_stability.derivatives[lift_derivative][i][0] = cs_res.CLift_derivative
res.static_stability.derivatives[drag_derivative][i][0] = cs_res.Cdrag_derivative
res.static_stability.derivatives[Y_derivative][i][0] = cs_res.CY_derivative
res.static_stability.derivatives[L_derivative][i][0] = cs_res.Cl_derivative
res.static_stability.derivatives[M_derivative][i][0] = cs_res.CM_derivative
res.static_stability.derivatives[N_derivative][i][0] = cs_res.CN_derivative
if len(res.static_stability.coefficients.X) > 1:
res.static_stability.derivatives.CX_alpha[:, 0] = np.gradient( res.static_stability.coefficients.X[:, 0],res.aerodynamics.angles.alpha[:, 0] )
res.static_stability.derivatives.CZ_alpha[:, 0] = np.gradient( res.static_stability.coefficients.Z[:, 0],res.aerodynamics.angles.alpha[:, 0] )
chords = np.zeros_like(wing_local_spans)
chords[:, :, :-1] = np.diff(wing_local_spans, axis=2)
chords[:, :, -1] = chords[:, :, -2]
settings.vortex_distribution = Data(
chord_widths = chords.reshape(chords.shape[:-2] + (-1,)),
chord_lengths = wing_section_chords.reshape(wing_section_chords.shape[:-2] + (-1,)),
leading_edge_sweeps = leading_edge_sweeps.reshape(leading_edge_sweeps.shape[:-2] + (-1,)),
wing_areas = wing_areas,
)
return
