propulsion

class propulsion.Propulsion(**kwargs)[source]

Bases: openmdao.components.meta_model_structured_comp.MetaModelStructuredComp

Interpolates engine parameters from engine deck.

The Propulsion component requires the following inputs:

  • inputs['z']: aircraft z-position [m]

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

  • inputs['TS']: engine thrust-setting [-]

The Propulsion component computes the following outputs:

  • outputs['W_f']: engine fuel consumption

  • outputs['F_n']: engine net thrust [N]

  • outputs['Tti_c']: combustor inlet total temperature [K]

  • outputs['Pti_c']: combustor inlet total pressure [Pa]

  • outputs['V_j']: engine jet velocity [m/s]

  • outputs['rho_j']: engine jet density [kg/m3]

  • outputs['A_j']: engine jet area [m2]

  • outputs['Tt_j']: engine jet total temperature [K]

  • outputs['M_j']: engine jet Mach number [-]

  • outputs['mdoti_c']: engine combustor inlet mass flow [kg/s]

  • outputs['Ttj_c']: engine combustor exit total temperature [K]

  • outputs['DTt_des_c']: engine design turbine total temperature drop [K]

  • outputs['rho_ti_c']: engine turbine inlet density [kg/m3]

  • outputs['c_ti_c']: engine turbine inlet speed of sound [m/s]

  • outputs['rho_te_c']: engine turbine exit density [kg/m3]

  • outputs['c_te_c']: engine turbine exit speed of sound [m/s]

  • outputs['DTt_f']: engine fan total temperature rise [K]

  • outputs['mdot_f']: engine fan mass flow [kg/s]

  • outputs['N_f']: engien fan rotational speed [rpm]

  • outputs['A_f']: engien fan inlet area [m2]

  • outputs['d_f']: engien fan diameter [m]

initialize()[source]

Initialize the component.

setup()[source]

Declare inputs and outputs.

Available attributes:

name pathname comm options

The thermodynamic engine cycle module is developed using the numerical propulsion system simulation (NPSS) software. A look-up table is implemented of the engine parameters as a function of flight Mach number, \(M_0 \in [0,0.5]\), flight altitude, \(z\in [0, 5000]m\), and engine thrust-setting, \(TS \in [30, 105] \%\). The thrust-setting, TS, at a particular Mach number and altitude is defined as a fraction of the maximum thermodynamic available thrust at these flight conditions. The maximum available thrust is defined as the thrust obtained when the engine is operated at 100% fan corrected speed (Ntextsubscript{1c2}).