Limits – Load and Speeds
Limit load is defined as the maximum load that an aircraft can be subjected to in its life cycle. Under the limit load, any deformation recovers to its original shape and would not affect structural integrity. Structural performance is defined in terms of stiffness and strength. Stiffness is related to flexibility and deformations and has implications for aeroelasticity and flutter. Strength concerns the loads that an aircraft structure is capable of carrying and is addressed within the context of the V-n diagram.
To ensure safety, a margin (factor) of 50% increase (civil aviation) is enforced through regulations as a factor of safety to extend the limit load to the ultimate load. A flight load exceeding the limit load but within the ultimate load should not cause structural failure but could affect integrity with permanent deformation. Aircraft are equipped with g-meters to monitor the load factor – the n for each sortie – and, if exceeded, the airframe must be inspected at prescribed areas and maintained by prescribed schedules that may require replacement of structural components. For example, an aerobatic aircraft with a 6-g-limit load will have an ultimate load of 9 g. If an in-flight load exceeds 6 g (but is below 9 g), the aircraft may experience permanent deformation but should not experience structural failure. Above 9 g, the aircraft would most likely experience structural failure.
The factor of safety also covers inconsistencies in material properties and manufacturing deviations. However, aerodynamicists and stress engineers should calculate for load and component dimensions such that their errors do not erode the factor of safety. Geometric margins, for example, should be defined such that they add positively to the factor of safety.
ultimate load = factor of safety x limit load
For civil aircraft applications, the factor of safety equals 1.5 (FAR 23 and FAR 25, Vol. 3).
Table 5.1. Typical permissible g-load for civil aircraft