RCAIDE.Library.Methods.Mass_Properties.Weight_Buildups.Conventional.General_Aviation.FLOPS.compute_fuselage_weight

compute_fuselage_weight

compute_fuselage_weight(vehicle)

Calculate the weight of a general aviation aircraft fuselage using NASA FLOPS methodology.

Parameters:

vehicle (RCAIDE.Vehicle()) –

Vehicle data structure containing:

• fuselages.lengths.totalfloat

  Total length of the fuselage [m]

• fuselages.widthfloat

  Maximum width of the fuselage [m]

• fuselages.heights.maximumfloat

  Maximum height of the fuselage [m]

• flight_envelope.ultimate_loadfloat

  Ultimate load factor

• mass_properties.max_takeofffloat

  Maximum takeoff weight [kg]

• design_dynamic_pressurefloat

  Design dynamic pressure [Pa]

Returns:

fuselage_weight – Weight of the fuselage structure [kg]

Return type:

float

Notes

The function implements the FLOPS (Flight Optimization System) weight estimation method for general aviation aircraft fuselages. The calculation accounts for geometric parameters, design loads, and cruise conditions.

Major Assumptions

• Single fuselage configuration (NFUSE = 1)

• Structural weight scales with wetted area and design loads

Theory

The FLOPS fuselage weight estimation follows:

\[W_{fus} = 0.052 * S_{wf}^{1.086} * (N_{ult} * W_{TO})^{0.177} * q^{0.241}\]

Where:

• W_{fus} is fuselage weight [lb]

• S_{wf} is fuselage wetted area [ft^2], calculated as π(L/D - 1.7)D^2

• N_{ult} is ultimate load factor

• W_{TO} is maximum takeoff weight [lb]

• q is design dynamic pressure [psf]

• L is fuselage length [ft]

• D is average fuselage diameter [ft]

References

[1] NASA. The Flight Optimization System Weights Estimation Method.

NASA Technical Report.