aerosandbox.library.aerodynamics#
Submodules#
Package Contents#
Functions#
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Computes the drag area (CDA) of a typical control usage as used on a well-manufactured RC airplane. |
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Computes the drag area (CDA) of the gaps associated with a typical wing control surface. |
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Computes the drag area (CDA) of a protruding bolt or rivet. |
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Computes the drag area (CDA) of a sheet metal joint that is perpendicular to the flow. |
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Computes the induced drag associated with a lifting planar wing. |
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Computes the Oswald's efficiency factor for a planar, tapered, swept wing. |
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Computes the optimal (minimum-induced-drag) taper ratio for a given quarter-chord sweep angle. |
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Returns the ratio of 3D lift coefficient (with compressibility) to the 2D lift coefficient. |
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Gives the ratio of actual induced drag to free-flight induced drag experienced by a wing in ground effect. |
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Computes the mach number immediately after a normal shock wave. |
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Computes the ratio of fluid density across a normal shock. |
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Computes the ratio of fluid temperature across a normal shock. |
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Computes the ratio of fluid static pressure across a normal shock. |
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Computes the ratio of fluid total pressure across a normal shock. |
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Computes the linear Hermite polynomial patch that passes through the given endpoints f_a and f_b. |
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Computes the cubic Hermite polynomial patch that passes through the given endpoints and endpoint derivatives. |
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Yields the idealized drag area (denoted CDA, or equivalently, D/q) of a Sears-Haack body. |
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See documentation for sears_haack_drag() in this same file. |
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Wave drag_force coefficient prediction using the low-fidelity Korn Equation method; |
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Wave drag_force coefficient prediction using the low-fidelity Korn Equation method; |
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An approximate relation for computing transonic wave drag, based on an object's Mach number. |
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Returns the drag coefficient of a cylinder in crossflow as a function of its Reynolds number and Mach. |
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Returns the mean skin friction coefficient over a flat plate. |
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Returns the approximate lift coefficient of a flat plate, following thin airfoil theory. |
Returns the drag coefficient of a flat plat oriented normal to the flow (i.e., alpha = 90 deg). |
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A curve fit I did to Eppler 216 airfoil data. |
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A curve fit I did to RAE2822 airfoil data. |
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Calculates the drag area (in m^2) of the aft end of a fuselage with a given upsweep angle. |
Attributes#
- aerosandbox.library.aerodynamics.CDA_control_linkage(Re_l, linkage_length, is_covered=False, is_top=False)[source]#
Computes the drag area (CDA) of a typical control usage as used on a well-manufactured RC airplane.
The drag area (CDA) is defined as: CDA == D / q, where:
D is the drag force (dimensionalized, e.g., in Newtons)
q is the freestream dynamic pressure (dimensionalized, e.g., in Pascals)
See study with original data at AeroSandbox/studies/LinkageDrag.
Data from:
Hepperle, Martin. “Drag of Linkages”. https://www.mh-aerotools.de/airfoils/linkage.htm
Summarizes data from “Werner Würz, published in the papers of the ISF-Seminar in December 1989 in Baden, Switzerland.”
- Parameters:
Re_l (Union[float, aerosandbox.numpy.ndarray]) – Reynolds number, with reference length as the length of the linkage.
linkage_length (Union[float, aerosandbox.numpy.ndarray]) – The length of the linkage. [m]
is_covered (Union[bool, aerosandbox.numpy.ndarray]) – A boolean of whether an aerodynamic fairing is placed around the linkage.
is_top (Union[bool, aerosandbox.numpy.ndarray]) – A boolean of whether the linkage is on the top surface of the wing (True) or the bottom surface ( False). Differences in local boundary layer and inviscid effects cause local velocity changes.
- Return type:
Union[float, aerosandbox.numpy.ndarray]
Returns: The drag area [m^2] of the control linkage.
- aerosandbox.library.aerodynamics.CDA_control_surface_gaps(local_chord, control_surface_span, local_thickness_over_chord=0.12, control_surface_hinge_x=0.75, n_side_gaps=2, side_gap_width=None, hinge_gap_width=None)[source]#
Computes the drag area (CDA) of the gaps associated with a typical wing control surface. (E.g., aileron, flap, elevator, rudder).
The drag area (CDA) is defined as: CDA == D / q, where:
D is the drag force (dimensionalized, e.g., in Newtons)
q is the freestream dynamic pressure (dimensionalized, e.g., in Pascals)
This drag area consists of two sources:
Chordwise gaps at the side edges of the control surface (“side gaps”)
Spanwise gaps at the hinge line of the control surface (“hinge gap”)
- Parameters:
local_chord (float) – The local chord of the wing at the midpoint of the control surface. [meters]
control_surface_span (float) – The span of the control surface. [meters]
local_thickness_over_chord (float) – The local thickness-to-chord ratio of the wing at the midpoint of the control surface. [nondimensional] For example, this is 0.12 for a NACA0012 airfoil.
control_surface_hinge_x (float) – The x-location of the hinge line of the control surface, as a fraction of the local chord. [nondimensional] Defaults to x_hinge / c = 0.75, which is typical for an aileron.
n_side_gaps (int) –
The number of “side gaps” to count on this control surface when computing drag. Defaults to 2 ( i.e., one inboard gap, one outboard gap), which is the simplest case of a wing with a single partial-span aileron. However, there may be cases where it is best to reduce this to 1 or 0. For example:
A wing with a single full-span aileron would have 1 side gap (at the wing root, but not at the tip).
- A wing with a flap and aileron that share a chordwise gap would be best modeled by setting
n_side_gaps = 1 ( so that no double-counting occurs).
side_gap_width (float) – The width of the chordwise gaps at the side edges of the control surface [meters]. If this is left as the default (None), then a typical value will be computed based on the local chord and control surface span.
hinge_gap_width (float) – The width of the spanwise gap at the hinge line of the control surface [meters]. If this is left as the default (None), then a typical value will be computed based on the local chord.
- Return type:
float
- Returns: The drag area [m^2] of the gaps associated with the control surface. This should be added to the “clean”
wing drag to get a more realistic drag estimate.
- aerosandbox.library.aerodynamics.CDA_protruding_bolt_or_rivet(diameter, kind='flush_rivet')[source]#
Computes the drag area (CDA) of a protruding bolt or rivet.
- The drag area (CDA) is defined as: CDA == D / q, where:
D is the drag force (dimensionalized, e.g., in Newtons)
q is the freestream dynamic pressure (dimensionalized, e.g., in Pascals)
- Parameters:
diameter (float) – The diameter of the bolt or rivet. [meters]
kind (str) –
The type of bolt or rivet. Valid options are:
”flush_rivet”
”round_rivet”
”flat_head_bolt”
”round_head_bolt”
”cylindrical_bolt”
”hex_bolt”
Returns: The drag area [m^2] of the bolt or rivet.
- aerosandbox.library.aerodynamics.CDA_perpendicular_sheet_metal_joint(joint_width, sheet_metal_thickness, kind='butt_joint_with_inside_joiner')[source]#
- Computes the drag area (CDA) of a sheet metal joint that is perpendicular to the flow.
(E.g., spanwise on the wing, or circumferential on the fuselage).
The drag area (CDA) is defined as: CDA == D / q, where:
D is the drag force (dimensionalized, e.g., in Newtons)
q is the freestream dynamic pressure (dimensionalized, e.g., in Pascals)
- Parameters:
joint_width (float) – The width of the joint (perpendicular to the airflow, e.g., spanwise on a wing). [meters]
sheet_metal_thickness (float) – The thickness of the sheet metal. [meters]
kind (str) –
The type of joint. Valid options are:
”butt_joint_with_inside_joiner”
”butt_joint_with_inside_weld”
”butt_joint_with_outside_joiner”
”butt_joint_with_outside_weld”
”lap_joint_forward_facing_step”
”lap_joint_backward_facing_step”
”lap_joint_forward_facing_step_with_bevel”
”lap_joint_backward_facing_step_with_bevel”
”lap_joint_forward_facing_step_with_rounded_bevel”
”lap_joint_backward_facing_step_with_rounded_bevel”
”flush_lap_joint_forward_facing_step”
”flush_lap_joint_backward_facing_step”
Returns: The drag area [m^2] of the sheet metal joint.
- aerosandbox.library.aerodynamics.induced_drag(lift, span, dynamic_pressure, oswalds_efficiency=1)[source]#
Computes the induced drag associated with a lifting planar wing.
- Parameters:
lift – Lift force [Newtons]
span – Wing span [meters]
dynamic_pressure – Dynamic pressure [Pascals]
oswalds_efficiency – Oswald’s efficiency factor [-]
Returns: Induced drag force [Newtons]
- aerosandbox.library.aerodynamics.oswalds_efficiency(taper_ratio, aspect_ratio, sweep=0.0, fuselage_diameter_to_span_ratio=0.0, method='nita_scholz')[source]#
Computes the Oswald’s efficiency factor for a planar, tapered, swept wing.
Based on “Estimating the Oswald Factor from Basic Aircraft Geometrical Parameters” by M. Nita, D. Scholz; Hamburg Univ. of Applied Sciences, 2012. https://www.fzt.haw-hamburg.de/pers/Scholz/OPerA/OPerA_PUB_DLRK_12-09-10.pdf
Implementation of Section 5 from the above paper.
Only valid for backwards-swept wings; i.e. 0 <= sweep < 90.
- Parameters:
taper_ratio (float) – Taper ratio of the wing (tip_chord / root_chord) [-]
aspect_ratio (float) – Aspect ratio of the wing (b^2 / S) [-]
sweep (float) – Wing quarter-chord sweep angle [deg]
fuselage_diameter_to_span_ratio (float) –
- Return type:
float
Returns: Oswald’s efficiency factor [-]
- aerosandbox.library.aerodynamics.optimal_taper_ratio(sweep=0.0)[source]#
Computes the optimal (minimum-induced-drag) taper ratio for a given quarter-chord sweep angle.
Based on “Estimating the Oswald Factor from Basic Aircraft Geometrical Parameters” by M. Nita, D. Scholz; Hamburg Univ. of Applied Sciences, 2012.
Only valid for backwards-swept wings; i.e. 0 <= sweep < 90.
- Parameters:
sweep – Wing quarter-chord sweep angle [deg]
- Return type:
float
Returns: Optimal taper ratio
- aerosandbox.library.aerodynamics.CL_over_Cl(aspect_ratio, mach=0.0, sweep=0.0, Cl_is_compressible=True)[source]#
Returns the ratio of 3D lift coefficient (with compressibility) to the 2D lift coefficient.
Specifically: CL_3D / CL_2D
- Parameters:
aspect_ratio (float) – The aspect ratio of the wing.
mach (float) – The freestream Mach number.
sweep (float) – The sweep of the wing, in degrees. To be most accurate, this should be the sweep at the locus of
wing. (thickest points along the) –
Cl_is_compressible (bool) – This flag indicates whether the 2D airfoil data already has compressibility effects
modeled. –
For example:
If this flag is True, this function returns: CL_3D / CL_2D, where CL_2D is the sectional lift
coefficient based on the local profile at the freestream mach number.
If this flag is False, this function returns: CL_3D / CL_2D_at_mach_zero, where CL_2D_… is the
sectional lift coefficient based on the local profile at mach zero.
For most accurate results, set this flag to True, and then model profile characteristics separately.
- Return type:
float
- aerosandbox.library.aerodynamics.induced_drag_ratio_from_ground_effect(h_over_b)[source]#
Gives the ratio of actual induced drag to free-flight induced drag experienced by a wing in ground effect. Artificially smoothed below around h/b == 0.05 to retain differentiability and practicality. Source: W. F. Phillips, D. F. Hunsaker, “Lifting-Line Predictions for Induced Drag and Lift in Ground Effect”.
- Using Equation 5 from the paper, which is modified from a model from Torenbeek:
Torenbeek, E. “Ground Effects”, 1982.
- Parameters:
h_over_b (float) – (Height above ground) divided by (wingspan).
- Returns:
Ratio of induced drag in ground effect to induced drag out of ground effect [unitless]
- aerosandbox.library.aerodynamics.mach_number_after_normal_shock(mach_upstream, gamma=1.4)[source]#
Computes the mach number immediately after a normal shock wave.
- Parameters:
mach_upstream – The mach number immediately before the normal shock wave.
gamma – The ratio of specific heats of the fluid. 1.4 for air.
Returns: The mach number immediately after the normal shock wave.
- aerosandbox.library.aerodynamics.density_ratio_across_normal_shock(mach_upstream, gamma=1.4)[source]#
Computes the ratio of fluid density across a normal shock.
Specifically, returns: rho_after_shock / rho_before_shock
- Parameters:
mach_upstream – The mach number immediately before the normal shock wave.
gamma – The ratio of specific heats of the fluid. 1.4 for air.
Returns: rho_after_shock / rho_before_shock
- aerosandbox.library.aerodynamics.temperature_ratio_across_normal_shock(mach_upstream, gamma=1.4)[source]#
Computes the ratio of fluid temperature across a normal shock.
Specifically, returns: T_after_shock / T_before_shock
- Parameters:
mach_upstream – The mach number immediately before the normal shock wave.
gamma – The ratio of specific heats of the fluid. 1.4 for air.
Returns: T_after_shock / T_before_shock
- aerosandbox.library.aerodynamics.pressure_ratio_across_normal_shock(mach_upstream, gamma=1.4)[source]#
Computes the ratio of fluid static pressure across a normal shock.
Specifically, returns: P_after_shock / P_before_shock
- Parameters:
mach_upstream – The mach number immediately before the normal shock wave.
gamma – The ratio of specific heats of the fluid. 1.4 for air.
Returns: P_after_shock / P_before_shock
- aerosandbox.library.aerodynamics.total_pressure_ratio_across_normal_shock(mach_upstream, gamma=1.4)[source]#
Computes the ratio of fluid total pressure across a normal shock.
Specifically, returns: Pt_after_shock / Pt_before_shock
- Parameters:
mach_upstream – The mach number immediately before the normal shock wave.
gamma – The ratio of specific heats of the fluid. 1.4 for air.
Returns: Pt_after_shock / Pt_before_shock
- aerosandbox.library.aerodynamics.linear_hermite_patch(x, x_a, x_b, f_a, f_b)[source]#
Computes the linear Hermite polynomial patch that passes through the given endpoints f_a and f_b.
- Parameters:
x (Union[float, aerosandbox.numpy.ndarray]) – Scalar or array of values at which to evaluate the patch.
x_a (float) – The x-coordinate of the first endpoint.
x_b (float) – The x-coordinate of the second endpoint.
f_a (float) – The function value at the first endpoint.
f_b (float) – The function value at the second endpoint.
- Returns:
The value of the patch evaluated at the input x. Returns a scalar if x is a scalar, or an array if x is an array.
- Return type:
Union[float, aerosandbox.numpy.ndarray]
- aerosandbox.library.aerodynamics.cubic_hermite_patch(x, x_a, x_b, f_a, f_b, dfdx_a, dfdx_b, extrapolation='continue')[source]#
Computes the cubic Hermite polynomial patch that passes through the given endpoints and endpoint derivatives.
- Parameters:
x (Union[float, aerosandbox.numpy.ndarray]) – Scalar or array of values at which to evaluate the patch.
x_a (float) – The x-coordinate of the first endpoint.
x_b (float) – The x-coordinate of the second endpoint.
f_a (float) – The function value at the first endpoint.
f_b (float) – The function value at the second endpoint.
dfdx_a (float) – The derivative of the function with respect to x at the first endpoint.
dfdx_b (float) – The derivative of the function with respect to x at the second endpoint.
extrapolation (str) – A string indicating how to handle extrapolation outside of the domain [x_a, x_b]. Valid values are “continue”, which continues the patch beyond the endpoints, and “clip”, which clips the patch at the endpoints. Default is “continue”.
- Returns:
The value of the patch evaluated at the input x. Returns a scalar if x is a scalar, or an array if x is an array.
- Return type:
Union[float, aerosandbox.numpy.ndarray]
- aerosandbox.library.aerodynamics.sears_haack_drag(radius_max, length)[source]#
Yields the idealized drag area (denoted CDA, or equivalently, D/q) of a Sears-Haack body.
Assumes linearized supersonic (Prandtl-Glauert) flow.
https://en.wikipedia.org/wiki/Sears%E2%80%93Haack_body
Note that drag coefficient and drag area are independent of Mach number for this case (assuming linearized supersonic aero).
- Parameters:
radius_max (float) – The maximum radius of the Sears-Haack body.
length (float) – The length of the Sears-Haack body.
- Return type:
float
Returns: The drag area (CDA, or D/q) of the body. To get the drag force, multiply by the dynamic pressure.
- aerosandbox.library.aerodynamics.sears_haack_drag_from_volume(volume, length)[source]#
See documentation for sears_haack_drag() in this same file.
Identical, except takes volume as an input rather than max radius.
Also returns a drag area (denoted CDA, or equivalently, D/q).
- Parameters:
volume (float) –
length (float) –
- Return type:
float
- aerosandbox.library.aerodynamics.mach_crit_Korn(CL, t_over_c, sweep=0, kappa_A=0.95)[source]#
Wave drag_force coefficient prediction using the low-fidelity Korn Equation method;
derived in “Configuration Aerodynamics” by W.H. Mason, Sect. 7.5.2, pg. 7-18
- Parameters:
CL – Sectional lift coefficient
t_over_c – thickness-to-chord ratio
sweep – sweep angle, in degrees
kappa_A – Airfoil technology factor (0.95 for supercritical section, 0.87 for NACA 6-series)
Returns:
- aerosandbox.library.aerodynamics.Cd_wave_Korn(Cl, t_over_c, mach, sweep=0, kappa_A=0.95)[source]#
Wave drag_force coefficient prediction using the low-fidelity Korn Equation method; derived in “Configuration Aerodynamics” by W.H. Mason, Sect. 7.5.2, pg. 7-18
- Parameters:
Cl – Sectional lift coefficient
t_over_c – thickness-to-chord ratio
sweep – sweep angle, in degrees
kappa_A – Airfoil technology factor (0.95 for supercritical section, 0.87 for NACA 6-series)
- Returns:
Wave drag coefficient
- aerosandbox.library.aerodynamics.approximate_CD_wave(mach, mach_crit, CD_wave_at_fully_supersonic)[source]#
An approximate relation for computing transonic wave drag, based on an object’s Mach number.
Considered reasonably valid from Mach 0 up to around Mach 2 or 3-ish.
Methodology is a combination of:
The methodology described in Raymer, “Aircraft Design: A Conceptual Approach”, Section 12.5.10 Transonic Parasite Drag (pg. 449 in Ed. 2)
and
The methodology described in W.H. Mason’s Configuration Aerodynamics, Chapter 7. Transonic Aerodynamics of Airfoils and Wings.
- Parameters:
mach – Mach number at the operating point to be evaluated
mach_crit – Critical mach number, a function of the body geometry
CD_wave_at_fully_supersonic – The wave drag coefficient of the body at the speed that it first goes (
supersonic. (effectively) fully) –
Here, that is taken to mean at the Mach 1.2 case.
This value should probably be derived using something similar to a Sears-Haack relation for the body in question, with a markup depending on geometry smoothness.
The CD_wave predicted by this function will match this value exactly at M=1.2 and M=1.05.
The peak CD_wave that is predicted is ~1.23 * this value, which occurs at M=1.10.
In the high-Mach limit, this function asymptotes at 0.80 * this value, as empirically stated by Raymer. However, this model is only approximate and is likely not valid for high-supersonic flows.
Returns: The approximate wave drag coefficient at the specified Mach number.
The reference area is whatever the reference area used in the CD_wave_at_fully_supersonic parameter is.
- aerosandbox.library.aerodynamics.Cd_cylinder(Re_D, mach=0.0, include_mach_effects=True, subcritical_only=False)[source]#
Returns the drag coefficient of a cylinder in crossflow as a function of its Reynolds number and Mach.
- Parameters:
Re_D (float) – Reynolds number, referenced to diameter
mach (float) – Mach number
include_mach_effects – If this is set False, it assumes Mach = 0, which simplifies the computation.
subcritical_only – Determines whether the model models purely subcritical (Re < 300k) cylinder flows. Useful, since
- Return type:
float
this model is now convex and can be more well-behaved.
Returns:
# TODO rework this function to use tanh blending, which will mitigate overflows
- aerosandbox.library.aerodynamics.Cf_flat_plate(Re_L, method='hybrid-sharpe-convex')[source]#
Returns the mean skin friction coefficient over a flat plate.
Don’t forget to double it (two sides) if you want a drag coefficient.
- Parameters:
Re_L (float) – Reynolds number, normalized to the length of the flat plate.
method –
The method of computing the skin friction coefficient. One of:
”blasius”: Uses the Blasius solution. Citing Cengel and Cimbala, “Fluid Mechanics: Fundamentals and
Applications”, Table 10-4.
Valid approximately for Re_L <= 5e5.
”turbulent”: Uses turbulent correlations for smooth plates. Citing Cengel and Cimbala,
”Fluid Mechanics: Fundamentals and Applications”, Table 10-4.
Valid approximately for 5e5 <= Re_L <= 1e7.
”hybrid-cengel”: Uses turbulent correlations for smooth plates, but accounts for a
non-negligible laminar run at the beginning of the plate. Citing Cengel and Cimbala, “Fluid Mechanics: Fundamentals and Applications”, Table 10-4. Returns: Mean skin friction coefficient over a flat plate.
Valid approximately for 5e5 <= Re_L <= 1e7.
”hybrid-schlichting”: Schlichting’s model, that roughly accounts for a non-negligtible laminar run.
Citing “Boundary Layer Theory” 7th Ed., pg. 644
”hybrid-sharpe-convex”: A hybrid model that blends the Blasius and Schlichting models. Convex in
log-log space; however, it may overlook some truly nonconvex behavior near transitional Reynolds numbers.
”hybrid-sharpe-nonconvex”: A hybrid model that blends the Blasius and Cengel models. Nonconvex in
log-log-space; however, it may capture some truly nonconvex behavior near transitional Reynolds numbers.
- Returns:
The skin friction coefficient, normalized to the length of the plate.
- Return type:
C_f
You can view all of these functions graphically using aerosandbox.library.aerodynamics.test_aerodynamics.test_Cf_flat_plate.py
- aerosandbox.library.aerodynamics.Cl_flat_plate(alpha, Re_c=None)[source]#
Returns the approximate lift coefficient of a flat plate, following thin airfoil theory. :param alpha: Angle of attack [deg] :param Re_c: Reynolds number, normalized to the length of the flat plate. :return: Approximate lift coefficient.
- aerosandbox.library.aerodynamics.Cd_flat_plate_normal()[source]#
Returns the drag coefficient of a flat plat oriented normal to the flow (i.e., alpha = 90 deg).
Uses results from Tian, Xinliang, Muk Chen Ong, Jianmin Yang, and Dag Myrhaug. “Large-Eddy Simulation of the Flow Normal to a Flat Plate Including Corner Effects at a High Reynolds Number.” Journal of Fluids and Structures 49 ( August 2014): 149–69. https://doi.org/10.1016/j.jfluidstructs.2014.04.008.
Note: Cd for this case is effectively invariant of Re.
Returns: Drag coefficient
- aerosandbox.library.aerodynamics.Cd_wave_e216(Cl, mach, sweep=0.0)[source]#
A curve fit I did to Eppler 216 airfoil data. Within -0.4 < CL < 0.75 and 0 < mach < ~0.9, has R^2 = 0.9982. See: C:ProjectsGitHubfirefly_aerodynamicsMSES Interfaceanalysise216 :param Cl: Lift coefficient :param mach: Mach number :param sweep: Sweep angle, in deg :return: Wave drag coefficient.
- aerosandbox.library.aerodynamics.Cd_wave_rae2822(Cl, mach, sweep=0.0)[source]#
A curve fit I did to RAE2822 airfoil data. Within -0.4 < CL < 0.75 and 0 < mach < ~0.9, has R^2 = 0.9982. See: C:ProjectsGitHubfirefly_aerodynamicsMSES Interfaceanalysisrae2822 :param Cl: Lift coefficient :param mach: Mach number :param sweep: Sweep angle, in deg :return: Wave drag coefficient.
- aerosandbox.library.aerodynamics.fuselage_upsweep_drag_area(upsweep_angle_rad, fuselage_xsec_area_max)[source]#
Calculates the drag area (in m^2) of the aft end of a fuselage with a given upsweep angle.
Upsweep is the characteristic shape seen on the aft end of many fuselages in transport aircraft, where the centerline of the fuselage is angled upwards near the aft end. This is done to reduce the required landing gear height for adequate takeoff rotation, which in turn reduces mass. This nonzero centerline angle can cause some separation drag, which is predicted here.
Equation is from Raymer, Aircraft Design: A Conceptual Approach, 5th Ed., Eq. 12.36, pg. 440.
- Parameters:
upsweep_angle_rad (float) – The upsweep angle of the aft end of the fuselage relative to the centerline, in radians.
fuselage_xsec_area_max (float) – The maximum cross-sectional area of the fuselage, in m^2.
- Return type:
float
Returns: The drag area of the aft end of the fuselage [m^2]. This is equivalent to D/q, where D is the drag force and q is the dynamic pressure.