aerodynamics

class description

class aerodynamics.Aerodynamics(**kwargs)

Bases: openmdao.core.explicitcomponent.ExplicitComponent

Computes aerodynamic forces and Mach number along the trajectory.

The Aerodynamics component requires the following inputs:

• inputs['c_l']: aircraft lift coefficient [-]

• inputs['c_d']: aircraft drag coefficient [-]

• inputs['rho_0']: ambient density [kg/m3]

• inputs['c_0']: ambient speed of sound [m/s]

• inputs['v']: aircraft velocity [m/s]

The Aerodynamics component computes the following outputs:

• outputs['q']: ambient dynamic pressure [Pa]

• outputs['L']: aircraft lift [N]

• outputs['D']: aircraft drag [N]

• outputs['M_0']: ambient Mach number [-]

compute(inputs: openmdao.vectors.default_vector.DefaultVector, outputs: openmdao.vectors.default_vector.DefaultVector)

Compute outputs given inputs. The model is assumed to be in an unscaled state.

Parameters

• inputs (Vector) – Unscaled, dimensional input variables read via inputs[key].

• outputs (Vector) – Unscaled, dimensional output variables read via outputs[key].

• discrete_inputs (dict or None) – If not None, dict containing discrete input values.

• discrete_outputs (dict or None) – If not None, dict containing discrete output values.

compute_partials(inputs: openmdao.vectors.default_vector.DefaultVector, partials: openmdao.vectors.default_vector.DefaultVector)

Compute sub-jacobian parts. The model is assumed to be in an unscaled state.

Parameters

• inputs (Vector) – Unscaled, dimensional input variables read via inputs[key].

• partials (Jacobian) – Sub-jac components written to partials[output_name, input_name]..

• discrete_inputs (dict or None) – If not None, dict containing discrete input values.

initialize()

Perform any one-time initialization run at instantiation.

setup()

Declare inputs and outputs.

Available attributes:

name pathname comm options

theory

The aircraft aerodynamic lift coefficient, \(C_L\), and drag coefficient, \(C_D\), are provided to the module in terms of a look-up table as a function of the wing angle of attack, \(\alpha\), and the flap and slat deflection angle, \(\theta_{flap}\) and \(\theta_{slat}\). The aircraft lift and drag forces are computed using the wing surface area, \(S\), and the dynamic pressure, \(q=\frac{1}{2}\rho v^2\), as follows:

\[\begin{split}\begin{array}{lp{2cm}l} L = C_L(\alpha, \theta_{flap}, \theta_{slat}) q S && D = C_D(\alpha, \theta_{flap} \theta_{slat}) q S \\ \end{array}\end{split}\]