Source code for RCAIDE.Library.Methods.Mass_Properties.Weight_Buildups.Conventional.Transport.Raymer.compute_horizontal_tail_weight
# RCAIDE/Library/Methods/Weights/Correlation_Buildups/Raymer/compute_horizontal_tail_weight.py
#
# Created: Sep 2024, M. Clarke
# ----------------------------------------------------------------------------------------------------------------------
# IMPORT
# ----------------------------------------------------------------------------------------------------------------------
# RCAIDE
import RCAIDE
from RCAIDE.Framework.Core import Units
# python imports
import numpy as np
# ----------------------------------------------------------------------------------------------------------------------
# Horizontal Tail Weight
# ----------------------------------------------------------------------------------------------------------------------
[docs]
def compute_horizontal_tail_weight(vehicle, wing, settings,elevator_fraction=0.4):
"""
Calculates horizontal tail weight based on Raymer's empirical method.
Parameters
----------
vehicle : RCAIDE.Vehicle()
Vehicle data structure containing:
- mass_properties.max_takeoff : float
Maximum takeoff weight [kg]
- flight_envelope.ultimate_load : float
Ultimate load factor
- wings['main_wing'] : Data()
Main wing properties including origin and aerodynamic center
- fuselages['fuselage'].width : float
Width of the fuselage [m]
wing : RCAIDE.Component()
Horizontal tail component containing:
- areas.reference : float
Tail surface area [m^2]
- origin : array
Location of tail measured from nose [m]
- aerodynamic_center : array
Location of ac measured from leading edge [m]
- sweeps.quarter_chord : float
Quarter chord sweep angle [rad]
- thickness_to_chord : float
Thickness-to-chord ratio
- spans.projected : float
Projected span [m]
- aspect_ratio : float
Aspect ratio
elevator_fraction : float, optional
Fraction of horizontal tail area that is elevator, defaults to 0.4
settings.advanced_composites : bool, optional
True if advanced composites are used, False otherwise
Returns
-------
tail_weight : float
Weight of the horizontal tail [kg]
Notes
-----
This method implements Raymer's correlation for horizontal tail weight estimation,
accounting for geometry, loads, and configuration effects.
**Major Assumptions**
* Not an all-moving horizontal tail (Kuht = 1.0)
* Elevator comprises 40% of tail area by default
* Correlation based on transport category aircraft data
* If advanced composites are used, the tail weight is reduced by 15%
**Theory**
The horizontal tail weight is calculated using:
.. math::
W_{ht} = 0.0379K_{uht}(1 + \frac{F_w}{B_h})^{-0.25}W_{dg}^{0.639}N_{ult}^{0.1}S_{ht}^{0.75}L_t^{-1}K_y^{0.704}\cos(\Lambda)^{-1}AR^{0.166}(1 + \frac{S_e}{S_{ht}})^{0.1}
where:
* :math:`K_{uht}` is 1.0 for fixed tails, 1.143 for all-moving
* :math:`F_w` is fuselage width
* :math:`B_h` is horizontal tail span
* :math:`L_t` is tail arm length
* :math:`K_y` is a function of tail arm length
* :math:`\Lambda` is quarter-chord sweep angle
References
----------
[1] Raymer, D., "Aircraft Design: A Conceptual Approach", AIAA
Education Series, 2018.
See Also
--------
RCAIDE.Library.Methods.Weights.Correlation_Buildups.Raymer.compute_vertical_tail_weight
RCAIDE.Library.Methods.Weights.Correlation_Buildups.Raymer.compute_operating_empty_weight
"""
ref_wing = None
for wing in vehicle.wings:
if isinstance(wing, RCAIDE.Library.Components.Wings.Main_Wing):
ref_wing = wing
S = 0
if ref_wing == None:
for wing in vehicle.wings:
if S < wing.areas.reference:
ref_wing = wing
L_fuselage = 0
for fuselage in vehicle.fuselages:
if L_fuselage < fuselage.lengths.total:
ref_fuselage = fuselage
L_fuselage = ref_fuselage.lengths.total
diff =[]
names = []
horiz_tail_location = vehicle.wings.horizontal_stabilizer.origin[0][0]
for segment in ref_fuselage.segments:
segment_loc = segment.percent_x_location * ref_fuselage.lengths.total
diff.append(np.abs(segment_loc - horiz_tail_location))
names.append(segment.tag)
min_value = min(diff)
min_index = diff.index(min_value)
Kuht = 1 # not an all-moving unit horizontal tail
Fw = ref_fuselage.segments[names[min_index]].width / Units.ft
Bh = wing.spans.projected / Units.ft
DG = vehicle.mass_properties.max_takeoff / Units.lbs
Sht = wing.areas.reference / Units.ft ** 2
Lt = (wing.origin[0][0] + wing.aerodynamic_center[0] - ref_wing.origin[0][0] -
ref_wing.aerodynamic_center[0]) / Units.ft
Ky = 0.3 * Lt
sweep = wing.sweeps.quarter_chord
Ah = wing.aspect_ratio
Se = elevator_fraction * Sht
tail_weight = 0.0379 * Kuht * (1 + Fw / Bh) ** (-0.25) * DG ** 0.639 *\
vehicle.flight_envelope.ultimate_load ** 0.1 * Sht ** 0.75 * Lt ** -1 *\
Ky ** 0.704 * np.cos(sweep) ** (-1) * Ah ** 0.166 * (1 + Se / Sht) ** 0.1
if settings.advanced_composites:
tail_weight *= 0.85
return tail_weight * Units.lbs
