Source code for RCAIDE.Library.Methods.Aerodynamics.Common.Lift.compute_flap_lift
# compute_flap_lift.py
#
# Created: Dec 2013, A. Varyar
# Modified: Feb 2014, T. Orra
# Jan 2016, E. Botero
# ----------------------------------------------------------------------------------------------------------------------
# Imports
# ----------------------------------------------------------------------------------------------------------------------
from RCAIDE import *
from RCAIDE.Framework.Core import Units
import numpy as np
# ----------------------------------------------------------------------------------------------------------------------
# compute_flap_lift
# ----------------------------------------------------------------------------------------------------------------------
[docs]
def compute_flap_lift(t_c,flap_type,flap_chord,flap_angle,sweep,wing_Sref,wing_affected_area):
"""Computes the increase of lift due to trailing edge flap deployment
Assumptions:
None
Source:
Unknown
Inputs:
t_c (wing thickness ratio) [Unitless]
flap_type [string]
flap_c_chord (flap chord as fraction of wing chord) [Unitless]
flap_angle (flap deflection) [radians]
sweep (Wing sweep angle) [radians]
wing_Sref (Wing reference area) [m^2]
wing_affected_area (Wing area affected by flaps) [m^2]
Outputs:
dcl_max_flaps (Lift coefficient increase) [Unitless]
Properties Used:
N/A
"""
#unpack
tc_r = t_c
fc = flap_chord * 100.
fa = flap_angle / Units.deg
Swf = wing_affected_area
sweep = sweep
# Basic increase in CL due to flap
dmax_ref= -4E-05*tc_r**4 + 0.0014*tc_r**3 - 0.0093*tc_r**2 + 0.0436*tc_r + 0.9734
# Corrections for flap type
if flap_type == None:
dmax_ref = 0.
elif flap_type.upper() == 'single_slotted'.upper():
dmax_ref = dmax_ref * 0.93
elif flap_type.upper() == 'triple_slotted'.upper():
dmax_ref = dmax_ref * 1.08
# Chord correction
Kc = 0.0395*fc + 0.0057
# Deflection correction
Kd = -1.7857E-04*fa**2 + 2.9214E-02*fa - 1.4000E-02
# Sweep correction
Ksw = (1 - 0.08 * (np.cos(sweep))**2) * (np.cos(sweep)) ** 0.75
# Applying corrections
dmax_flaps = Kc * Kd * Ksw * dmax_ref
# Final CL increment due to flap
dcl_max_flaps = dmax_flaps * Swf / wing_Sref
return dcl_max_flaps
