Idiomatic: The Origin Story of Certification Requirements

We often have a good handle on what a Figure Of Speech (Idiom) means but often the source is surprising or counterintuitive. Who was surprised to learn that the “Rule of Thumb” came from the maximum width of stick allowed for wife beating? Or that "giving someone the Cold Shoulder” originally referred to serving a cold shoulder of mutton to an unwanted guest?

In a similar manner, sometimes the basis of the cornerstones of aircraft certification specifications may surprise us.

Before I started my own company, I was privileged to be a member of a committee which advised the MAA on matters of regulations, specifications, and means of compliance. This august body was asked the simple question of the origin and reason for the 1.125 Proof Factor (*) in Def Stan 00-970. It turned out that none of us knew! The brainiest of us went away and she came back later with an explanation that we all agreed was eminently plausible, and that answer was duly supplied to the questioner. Only it wasn’t correct!

My thesis is that often the original reason for a requirement becomes lost, and arguably, may become irrelevant, even if the requirement lives on and itself remains relevant. My example will be the requirement surrounding the Factor of Safety on structural strength, arguably one of the most important single factors in aircraft certification requirements, such as CS-25 and its equivalents. This is the 1.5 Ultimate Factor on Design Limit Loads (DLL) (?). ?

But where does the value of 1.5 come from? What does it represent??Why 1.5 and not 1.3 or 1.8??

I think that most stress engineers have their own ideas, but the most common one that I see is along the lines of: the value of 1.5 is based on the observation that some aluminium alloys can tolerate a load up to 1.5 times their yield point before failing. This value could be considered to ensure that the structure will not fail even if it experiences loads above the DLL. Many new aircraft now have relatively little aluminium in their design, but the requirement persists.

Makes sense, yes??

Well, what about members in compression? Or joints (where most failures occur) where the presence of load transfer at fasteners may throw out this ratio? Or situations where the loads are just a little higher than intended or the structure is below tolerance on strength, or both?

It transpires that the story behind the 1.5 factor is a little bit more complicated than one might think.

[TL:DR. Ultimate Factors were not based on an explicit rationale, but rather were intended to provide a reasonable margin of safety without overburdening the design of new aircraft in the 1930s and 1940s. The value 1.5 continues to be retained, however, because it continues to work.]

A bit of history

An Ultimate Factor on expected maximum design loads of 2.0 was in widespread use in American and codes starting in the 1920s and the British design codes in the 1930s. This factor developed from trial and error – when largely wooden aircraft broke up in flight, the factor was increased until they stopped failing in flight (per [2]).

Also, in the 1920s there was no formal relationship between the design load factor, maximum aerodynamic manoeuvre capability and operational manoeuvre limits for aircraft. Both the US requirements, per [3], and AP970 [1] specified strength of structure against Centre of Pressure Forward (High Incidence), Centre of Pressure Back (Low Incidence), and an Inverted High Incidence (Inverted Flight) cases, rather than the V-g envelope later adopted.

Overall, there was a lack of rational criteria in use, and design requirements more closely approximated the “scantlings” (§) methods used in boat design.

The American Story

An example of American requirements in the 1920s can be seen in USA Army Air Service’s Handbook of Instructions for Airplane Design (HIAD) and US CAA with Aeronautics Bulletin 7-A. These tended to identify an overall G limit for isolated cases (High Incidence, etc).

"In the early 1920's a factor of 2.0 was considered necessary… In the late 1920's, actual operational flying of the newer airplanes were coming closer to the ultimate load factor than earlier models. Airplanes were flying up to two-thirds and more of the ultimate load factor and nothing was happening to the structure; therefore, the evolution of thinking toward a lower factor of safety was a natural one.” [3].

The rationalising of various civil and military aircraft design requirements, including a reduction in Ultimate Factor, began in the US in the 1930s:

1.?Firstly, the V-g Envelope method (?) was introduced by the Navy Bureau of Aeronautics in 1933 in their specifications: SS-1, SS-2 etc. This was a departure from the HIAD.

2. ?Secondly, as part of a further rationalisation process to have a single factor of safety for both the entire aircraft and envelope, the Air Corps recommended in 1934 to take an Ultimate Factor of 1.5 that had been adopted in 1930 for tail loads only as a “factor of safety for material”, and then apply it to the whole aircraft [4].

In the 50s, 60s and 70s there were further moves to reduce the 1.5 factor still further, per [3], [5], and [6], but the factor has remained in place.

The British Story

The British requirements in the 1930s were based on AP970 [1], which had started as a “Handbook of Strength Calculations”, HB 806 in 1918, and then was up-issued to become AP970 in 1924. Earlier versions of AP970 assigned an overall Ultimate capability requirement in terms of G acceleration, but the 1935 version [1] switched to defining a design load case (DLL) and then applying a variety of Ultimate Factors for that case, although the most common factor was 2.0. The cases considered tended to be unbalanced (e.g. a tail load case may not reflect a fully balanced point in the sky for the aircraft) and appear to have been formulated to ensure an overall robust structure, rather than reflect an envelope of usage. ?

“Although the USA and other countries reduced the ultimate factor of safety from 2.0 to 1.5 during the 1930s, it was only after much heart-searching that the UK followed their example, and AvP 970 [sic] was not amended to reflect this change of policy until 1945.” [7], likely based on the arguments in [8]. Likewise, the V-g diagram was not adopted into AP970 until 1938 [7].

Finally in contrast to the Americans, from 1935 onwards a Proof Factor of 0.75 of Ultimate Loading entered AP970 ([7] para 2.3). This ensured that yield would not occur in the structure at ? of the Ultimate loading requirement.

But what is it FOR?

What the Ultimate Factor was originally designed to protect against is also an area of uncertainty, with various authorities stating varying reasons over the years, such as “to allow for possible imperfections in material, approximations of analysis, and general lack of exact knowledge of the loads” [10]

The basis for the value 1.5 has been much debated, variously being related to different materials’ Ultimate-to-Yield factors ([3] & [7]), the intent to ensure a sufficiently high Design Limit Load when existing Ultimate Requirements were defactored (by the person formulating the US Civil Requirements [9], and the person suggesting new British limits in 1944 [8]), or an “opinion of what was representative of service flight conditions” (by the person formulating the US Military Requirements, [10]).

In other words, the 1.5 Ultimate Factor was accepted by different parties as reasonable in the 1930s, even though they generally had different reasons for their opinions. Importantly here: those reasons are no longer relevant.

If there is an overriding reason for the 1.5 factor, I’d say it was the following. The authorities in the 1920s had defined Ultimate Limits in terms of a particular G limit for a type of aircraft and when it became clear that the world was moving towards a flight envelope basis, where the G seen by that aircraft would be used and an Ultimate factor applied on top, they needed a number that would provide consistency with previous designs and that would provide similar structural margins. ?The Americans started using 1.5 first, and eventually the Brits followed suit – in both cases, however, the actual Ultimate Limits in terms of G were not changed (at least not right away), even though both sides were flying aircraft past DLL and closer and closer to Ultimate limits in practice.

Summary

1.?The Ultimate Factor of 2.0 evolved by trial and error in the 1920s and represented a capability that was found by service experience to lead to safe aircraft. However, improved understanding of design loads reduced the need for factors as high as 2.0. The factor of 2.0 was demonstrated by service experience in the 1930s to be over conservative, and eventually settled to a factor of 1.5.

2.?The Ultimate Factors were firmed up by taking values which had precedence, and which had been used before and for which the factors were judged not too onerous when all aspects were considered.

3.?The Ultimate Factors in the American and British requirements were not formulated based on an explicit rationale, rather they were intended to produce a reasonable margin that would not overburden the design of new aircraft in the 1930s and 1940s.

So, the reasons for the factors are long gone, but the relevance of the factors remains. A bit like the stories we tell ourselves about figures of speech.

With that in mind one final caution. Do you remember my explanations for the Rule of Thumb and the Cold Shoulder? Well, they are both completely wrong too – these folk etymologies persist because we like to tell ourselves good stories and the stories here were much more interesting than the truth. My take? Be careful with accepting explanations just because they appear plausible. ??

Oh, and what about that 1.125 Proof Factor? What is it for?

Do you remember the introduction of the 0.75 Proof factor on Ultimate in the British regulations in 1935? ?With the rationalisation to a V-g envelope where the aircraft must withstand DLL without yielding in the American system, but in the British system the aircraft must withstand 0.75 x Ultimate without yielding we ended up with a discrepancy. ?This was solved by introducing the 1.125 Proof Factor after the Ultimate Factors had been rationalised at 1.5. In other words, 0.75 x 1.5 = 1.125, the factor above DLL that AP970 (and now DS 00-970) says must not yield.

So, the answer is that the 1.125 Proof Factor arose because of a difference in opinion in the 1930s between the American and British regulators on how close to ultimate capability you should be without yielding the structure. Are the reasons for it still relevant? No. Is it still relevant itself? Probably. When you consider what drives military DLL and civilian DLL, military aircraft whose design tends to be driven by deliberate manoeuvres will approach, and exceed, DLL much more often than civil aircraft. So, a small factor above DLL makes sense. In the end my colleague’s answer was probably appropriate, we all had just lost sight of the story. ?

Fin

? Stephen Dosman, 2023. Unauthorized use and/or duplication of this material without express and written permission from the author is strictly prohibited. Excerpts and links may be used, provided that full and clear credit is given to Stephen Dosman and this site with appropriate and specific direction to the original content.

Footnotes

* - The factor above Design Limit Load for where there should be no yield in the structure (DS00-970, P1, UK25.303a)

? - Aerospace Stress Engineers use an Ultimate Factor to determine the maximum load requirement for a structure, that being the load the structure must be able to withstand without failing. The Ultimate Factor is typically 1.5 times the Design Limit Load (DLL), which is the load at which the structure is designed to operate without experiencing detrimental deformation or yielding.

§ - In other words, one designs individual parts based upon rules that have been shown to produce a sound structure, rather than creating balanced load sets that reflect an envelope of worst-case usage. The term “scantling” is believed to derive from a carpenter’s measuring tool called a “scantillon”.

? - Whereby the aircraft would need to be shown to be safe throughout an achievable envelope of acceleration and airspeed, rather than picking cases which would exercise various parts of the aircraft. These envelope cases also would need to be “balanced”, such that they reflected a realistic load balance on the various parts of the aircraft.

References

[1] Air Publication 970 “Design Requirements for the Aeroplanes for the Royal Air Force”, Air Ministry, HMSO, May 1935.

[2] Williams, J.K., “Safety Factors”, Journal of the Royal Aeronautical Society, Vol 60, pg 307-312, May 1956.

[3] Muller, G.E, Schmid, C.J., “Factor of Safety – USAF Design Practice”, part of AGARD Report No. 661 “Factors of Safety, Historical Development, State of the Art and Future Outlook”, NATO, November 1977.

[4] Epstein, A., "Investigation of the Comparative Design Requirements for Airplanes of the Army Air Corps and of the Bureau of Aeronautics," Air Corps Technical Report No 3924, February 1934.

[5] Ebner, H., “The Problem of Structural Safety with Particular Reference to Safety Requirements”, AGARD Report No. 150, NATO, November 1957.

[6] Mangurian, G.N., “The Aircraft Structural Factor of Safety”, AGARD Report No. 154, NATO, November 1957.

[7] Heath, W.G., “The Changing Scene of Structural Airworthiness”, Aeronautical Journal, RAeS, Pg 81-92, March/April 1980.

[8] Montagnon, P.E., R&M 2578, “The Case for Factors of Safety of 1.5 instead of 2.0, with special reference to the Flight Envelope”, Ministry of Supply, HMSO, 1944.

[9] Shanley, F.R., "Historical Note on the 1.5 Factor of Safety for Aircraft Structures", Readers' Forum, Journal of the Aerospace Sciences, Vol 29, No 2, February 1962.

[10] Epstein, A., "Another Historical Note on the 1.5 Factor of Safety", Readers' Forum, Journal of the Aerospace Sciences, Vol. 29, No. 6, June 1962.