calculate_stability_derivatives
Compute longitudinal stability derivatives for a wing, including lift curve slope and pitching moment slope, to assess static stability using wing geometry and flight conditions.
Instructions
Calculate basic longitudinal stability derivatives for a wing.
Args: wing_config: Wing configuration with keys: - span_m: Wing span in meters - chord_root_m: Root chord in meters - chord_tip_m: Tip chord in meters (optional) - sweep_deg: Quarter-chord sweep (optional, default 0) - dihedral_deg: Dihedral angle (optional, default 0) - twist_deg: Tip twist (optional, default 0) - airfoil_root: Root airfoil name (optional, default 'NACA2412') - airfoil_tip: Tip airfoil name (optional) flight_conditions: Flight conditions with keys: - alpha_deg: Reference angle of attack (optional, default 2.0) - mach: Mach number (optional, default 0.2)
Returns: JSON string with stability derivatives: - CL_alpha: Lift curve slope (dCL/dalpha) [1/rad] -- rate of lift change with angle of attack. Positive for conventional aircraft. - CM_alpha: Pitching moment slope (dCM/dalpha) [1/rad] -- must be negative for static longitudinal stability (nose-down restoring moment). - CL_alpha_dot: Unsteady lift derivative due to rate of alpha change. - CM_alpha_dot: Unsteady pitching moment derivative (pitch damping).
Raises: No exceptions are raised directly; errors are returned as formatted strings.
Input Schema
| Name | Required | Description | Default |
|---|---|---|---|
| wing_config | Yes | ||
| flight_conditions | Yes |
Output Schema
| Name | Required | Description | Default |
|---|---|---|---|
| result | Yes |