14 CFR Subpart C Section 25-303
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In this post we will get into the next regulation, 14 CFR Subpart C Section 25-303.
Unless otherwise specified, a factor of safety of 1.5 must be applied to the prescribed limit loads which are considered external loads on the structure. When a loading condition is prescribed in terms of ultimate loads, a factor of safety need not be applied unless otherwise specified.
[Amdt. 25-23, 35 FR 5672, Apr. 8, 1970]
In the previous post, we looked at 14 CFR Subpart C Section 25-301. The -303 regulation is related to it, in the sense that now we have a number called the "safety factor". The 301 regulation said that the loads are prescribed in terms of limit and ultimate loads. But this regulation goes farther, and says that the ultimate load is 1.5 times the limit load.
14 CFR Subpart C Section 25-303
Safety Factors
For example, in the case of cabin interior items, the ultimate load factor in the forward (FWD) direction is usually 9.0G. This load has the 1.5 safety factor included in the applied load. Therefore, forward limit load would be 6.0G. For some research and military aircraft, ultimate load safety factor may be lower at around 1.2. Limit loads must be sustained by the structure without any permanent detrimental deformation.
Here is a NASA Link that explains some more history.
14 CFR Subpart C Section 25-303
Limit Load Vs Ultimate Load Situations
There may be load cases such as in flight loads, sustained during turbulence. These loads may be prescribed in terms of limit loads. In such cases, margins of safety must be written against the critical yield strength of the structural member's material. In theory, the limit loads will never be exceeded under normal operations during the entire service life of an aircraft. But that does not mean there have been incidents that did exceed, the NASA link above sheds some light on it.
The maximum expected normal operation loads are limit loads. The 1.5 safety factor has proven to be statistically and historically adequate for many decades now, especially with the benefit of hindsight of aviation history.
Again it does not mean there could be ultimate load incidents such as severe crashes. In such cases the loads would be far greater and catastrophic than the ultimate load factor accounts for and the aircraft was designed for. Fact is that you can never really design something to survive every possible crash or landing, the best you can do is design for a reasonably severe crash or landing based on the safety factors that the regulations help us with.
14 CFR Subpart C Section 25-303
Combination Load Cases
Aircraft cabin interiors are generally certified to ultimate loads in various directions, including linear combinations of different directions. For example, 2.3G SIDE + 1.5G DOWN load case can be a limit or ultimate load case, depending on the guidelines included in the Structural Design Criteria (SDC) document provided by the OEM (original equipment manufacturer) or an aircraft integrator or a completions center.
14 CFR Subpart C Section 25-303
Engine Nacelle Load Cases
In case of engine nacelle components such as the thrust reverser (TR), there may be hundreds of different load cases. These cases may include various limit, ultimate and ultimate at 1.0 safety factor load cases. Most of these load cases are provided by the engine manufacturer. But there are other load cases that are induced into the TR due to the systems within the nacelle.
For example, the TR actuation system is one such critical system. It induces heavy loads into the TR structure during landing, or on the ground, or in certain failure cases in flight. It is critical to nail down which load case has what safety factor, in order to meet compliance requirements.
14 CFR Subpart C Section 25-303
Cabin Interiors Load Cases
In cabin interiors however, almost all cases are ultimate load cases including abuse, rapid decompression, emergency landing etc. These cases include the 1.5 safety factor. At the most, there may be 15 to 20 different load cases including any applicable combined inertia load cases.
14 CFR Subpart C Section 25-303
Practical Example
Consider a fighter jet whose Gross Take Off Weight is 30,000 lbs
In a wing test, it was found that the wing fails at an ultimate load of 243,000 lbs
So Tested Ultimate = 243,000 lbs per wing
- What is the Design Margin of Safety (MS)? This is a conservative MS value compared to the tested load and goes beyond the standard Ultimate MS compared to Ultimate Material Strength. There may be various conservative assumptions and factors that determine the standard ultimate Load value which may be lower than the tested value.
Design MS = (Tested Ultimate - Design Ultimate)/(Design Ultimate)
- But how do we find the design ultimate?
From 14 CFR Subpart C Section 25-303, for aircraft the safety factor = 1.5
Hence Design Ultimate = 1.5 x Design limit
- What is Design Limit?
In this case, the V-N diagram of the jet was found to have a 9.0 load factor for limit load.
Therefore, Design Limit = 9.0 x Gross Take Off Weight at 1.0G = 9.0 x 30,000 lbs = 270,000 lbs
Design Limit Per Wing = 135,000 lbs
Design Ultimate Per Wing = 135,000 x 1.5 = 202,500 lbs
Design MS = (Tested Ultimate - Design Ultimate)/(Design Ultimate)
Design MS = (243,000 - 202,500)/202,500 = +0.2
Thus the Design MS is found to be +20%.
The objective would be to minimize this +0.2 MS value to close to zero thus optimizing the design to ultimate failure strength especially because the failure load is known from testing. This process is called optimal "sizing".
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