Determining aircraft strength

[alejandro_]

In many debates the issue of aircraft strength appears. In WW2 there were some aircraft that had a reputation for strength P-40, P-47, Yaks while others were seen as nimble (Ki-43, A6M). I would like to know a few opinions on what is the best criterion to determine aircraft strength, focusing on the airframe and how resistant it is to damage. Most people seem to use empty weight of the aircraft. I wonder if other parameters could be included, like G limits or construction features.

 

[Shortround6]

G limits help, as long as you are sure you are comparing the right thing. However there is actually a fairly narrow range of G limits for certain types of planes. Like fighters are going to be from about 9 "G"s to about 13-13.5 "G"s for ultimate strength. That is the load at which the airplane will break. service limit is usually about 2/3 of that or put another way a 50% overload was considered normal in the US, that is 8 "G" service load meant ultimate strength had to be 12 "G"s, other countries had different safety margins or different initial requirements, which can trip up armchair annalists.
given the knowledge of the time and the test methods some engineers may have used a bit more fudge factor than others. Modern aircraft designer/builders of small aircraft still use the old methods.

Modern jets use computers and instead of rough and ready calculations for fatigue they actually build hydraulic test fixtures to test wing fatigue or ultimate breaking points.

Overbuilding means more weight and thus less payload or poorer rate of climb (or both).
Some types of construction are more resistant to battle damage than others. Like multi-spar wings (say 3 or more) vs single spar or 2 spar wings. Of course few, if any, fighters actually had 5 spars running from wing tip to wing tip. SO battle damage resistance depends on the location of the battle damage.
Complicating things further is that even in the same country different companies used different materials or different forms of construction so even with the same ultimate design criteria structural weights (empty weight? plus a lot of guess work )
can come out rather different. Steel structural components can actually be stronger for less weight in certain high stress locations than aluminium. Aluminium forgings can be stronger than castings or built up assemblies from pieces of flat stock.

There may be no easy answer or comparison.

 

[GregP]

It is a tough call.

By way of example, the US fighters, let's say an F6F Hellcat, were routinely stressed to an 8g limit with a 50% safety factor. So ultimate fail point was supposed to be at anything at or over 12 g.

We (the museum) fly a Mitsubishi A6M5 Model 52 Zero and we have some original design documents and the original designer helped with the restoration back in 1977 or so. It is stressed to a 6 g limit, but has a 100% safety factor, so the ultimate fail point is supposed to be at or above 12 g.

What that means in reality, is that when the Zero was intact, it was ultimately as strong as a Hellcat in the air and both could be thrown about the sky with little need to watch out for pulling too hard. But there is no way a Zero would stand up to the same punishment that a Hellcat would and still stay flying. To get 1,500 HP modern fighter performance from the Zero with 1,180 hp in its most powerful variant, the Zero sacrificed armor and self-sealing tanks. As a result, it would burn much more readily and the pilot has but little protection.

So the Zero was strong, but not exactly rugged and wouldn't sustain much damage without the damage getting to a critical point. It is VERY difficult if not impossible to tell that from the specs except for the large difference in weight. Exactly where lighter turns into more fragile in the field is a damned good question that nobody has a good answer for sure yet.

The American fighters in the ETO were rugged and heavy. The Spitifre was much lighter, and was in fact more fragile than US designs. But it never got a reputation for being unable to sustain battle damage like the Zero because it HAD armor and self-sealing tanks, and it MUST have had more rugged wing / tail attachments, because it didn't shed parts after one or two hits. Zeros sometimes did. Yet all three were 12 g airplanes at the ultimate fail point.

Your question has been debated for decades, with no resolution in sight. I'm sure a few people will still be asking that 100 years from now.

 

[Shortround6]

Comparing dissimilar planes really confuses things as what you are trying to compare is structural weight and not empty weight.
Structural weight would include the wing, the fuselage (but NOT cowl) and tail surfaces, both horizontal and vertical.
Engine, prop, landing gear and even control runs and cockpit fixtures don't do anything for strength.
Landing gear does nothing for aircraft strength and engine weights can vary considerable.
In the F6F comparison the R-2800 engine weighs 2480lbs while the Sakae engine in the Zero weighs about 1300lbs. Propellers are somewhat in proportion. F6F needs larger heavier landing gear just to hold up the engine and prop while taxiing
Not to mention the extra almost 100 sq ft of wing on the F6F.
Thick wings could weigh less than thin wings of the same size because the thick wing section resisted bending better.
Even if you can find two fighters of similar wing area and using similar weight engines it can be difficult to estimate structural weight.
The Curtiss Hawk gained about 250lbs in the wing alone from the P-36 to the later P-40s. Now the later P-40s may have been more resistant to combat damage than the P-36 but the wing had to be beefed up to handle the the 8 "G" service load (12 "G" ultimate) as the plane gained around 2000lbs of gross weight. A P-36 wing would be a bit suspect on a P-40 as 8 "G"s at 6000lbs is equal to 6 "G"s at 8000lbs and while 8 "G" is rarely seen by service fighters 6 "G"s is all too close to the normal flight envelop.

 

[alejandro_]

Many thanks for your answers.

We (the museum) fly a Mitsubishi A6M5 Model 52 Zero and we have some original design documents and the original designer helped with the restoration back in 1977 or so. It is stressed to a 6 g limit, but has a 100% safety factor, so the ultimate fail point is supposed to be at or above 12 g.

Actually, reading the book "Eagles of Mitsubishi: The Story of the Zero Fighter", written by it's designer, Jiro Horikoshi, is how I got interested. Horikoshi writes about the weight issue in detail, explaining that the requirements led to:

- New materials: an aluminium equivalent to 7075 aluminium.
- Design choices: The wing was built as a single unit to avoid heavy attchements fittings to the wings.
- Changing in the safety factor. The maximum load was suposed to be 7 g with a safey margin of 1.8 (12.6 g). Horikoshi managed to reduce this requirement to 1.6 for certain components. This meant less weight. He also added that after the war, the overall value was reduced to 1.5 anyway.

The overall policy was to control all components that were heavier than 1/100,000 of the airplane's final weight.

To get 1,500 HP modern fighter performance from the Zero with 1,180 hp in its most powerful variant, the Zero sacrificed armor and self-sealing tanks. As a result, it would burn much more readily and the pilot has but little protection.

This is my impression. Sometimes I read about the Zero having a light structure/construction, when the requirements in terms of strength are similar to US fighters. What the Zero lacked was pilot protection and self-sealing fuel tanks.

Also, the Zero carried quite a bit of fuel. The 6M2 (Navy Type 0 Carrier Fighter Model 11) had 518 liters of fuel. A Supermarine Spitfire Mk I had 386.5. Usually the larger the volume the larger the vulnerable area.

 

[drgondog]

Comparing dissimilar planes really confuses things as what you are trying to compare is structural weight and not empty weight.

SR - I rarely disagree with your posts. That said I will offer some thoughts based on only a sample of the two airframe companies I have worked for plus one as a contractor - in the airframe structures group. The primary disagreement is that All structural weight plus mission requirements for fuel, bombs, ammo, munchkins go into Mission Gross Weight - for analysis.

More comments below

Structural weight would include the wing, the fuselage (but NOT cowl) and tail surfaces, both horizontal and vertical.
Engine, prop, landing gear and even control runs and cockpit fixtures don't do anything for strength.
Landing gear does nothing for aircraft strength and engine weights can vary considerable.
In the F6F comparison the R-2800 engine weighs 2480lbs while the Sakae engine in the Zero weighs about 1300lbs. Propellers are somewhat in proportion. F6F needs larger heavier landing gear just to hold up the engine and prop while taxiing
Not to mention the extra almost 100 sq ft of wing on the F6F.
Thick wings could weigh less than thin wings of the same size because the thick wing section resisted bending better.
Even if you can find two fighters of similar wing area and using similar weight engines it can be difficult to estimate structural weight.
The Curtiss Hawk gained about 250lbs in the wing alone from the P-36 to the later P-40s. Now the later P-40s may have been more resistant to combat damage than the P-36 but the wing had to be beefed up to handle the the 8 "G" service load (12 "G" ultimate) as the plane gained around 2000lbs of gross weight. A P-36 wing would be a bit suspect on a P-40 as 8 "G"s at 6000lbs is equal to 6 "G"s at 8000lbs and while 8 "G" is rarely seen by service fighters 6 "G"s is all too close to the normal flight envelop.

The following comments are directed to the original Poster's question.
1. There are specific standards of airframe structures set by the customer - for US, the AAF and BUNAV
2. The most common threads above touch on three aspects of the design standards. a.) Limit and Ultimate Loads for which individual components critical to flight safety are analyzed for the material Stress due to the applied loads, b.) The standard and published material Properties of Materials which include Young's Modulus Tensile Strength in PSI at separate temperatures, Density, Coefficient of Thermal Expansion, Hardness, etc, c.)
3.) Approved methods by the Contractor are framed for calculation of stability of structures under compression (Pinned, Fixed, etc) and shear, methods for computing moments of inertia for complex structures for bending (and buckling for compression stability analysis) analysis.

As an aside the stated Tensile strength of the materials listed are developed by Bureau of Standards via Testing performed by ASTM and if you look at the Strain vs Applied Load/sq (Tensile Strength) plot the point 'selected' for Tensile Strength is usually (for aluminum) a defined point on the plot where it deviates from straight line and enters the region of Permanent Strain - further along the plot (now a non linear curve) is an area where the strain value peaks without fail in permanent deformation. That is the usual Ultimate Allowable Stress. For 2024/7075 Aluminum Ultimate Stress is set in the tables ~ 1.5 to 1.7 times the Stress at Yield. Whatever the Tables state is what you use.

Next - the airframe structures group is given the allowable Load factors (E.g. 3G positive for many bomber/transport, 8 G positive, etc for fighter aircraft)

Next for Limit and Ultimate Stress calculations the Angle of Attack Loads are usually for CL max in a diving pullout a design Gross Weight. Each major assembly is broken down and analyzed piece by piece as they fit into assembled parts, shear panels, Longerons, bulkheads, etc, to determine if the designed parts as received from airframe design will be within set boundary condition safety factors. Usually design team and structures team are working together to make sure the designs 'make sense' with respect to approach, weight and load paths. The Weights Control Group keeps track of every part weight as calculated and/or actually measured.

Next - the airframes are subjected to planned load testing on wings, empennage and landing gear.

Next - life happens and the mission adaptations grow, increasing gross weight and thereby REDUCE allowable load factors directly proportional to the increase in mission weight.

Summary - every nation had different notions of Airframe Structures Standards - and some applied slightly different philosophies for Yield and Ultimate for comparative materials. For US the standard was 8/12 for England 7/11 for fighters. I was very surprised to see the Zero as low as 6G Limit but that is in keeping with reducing weight to a minimum. If that was for combat mission with full fuel and ammo (no externals) it started out where the P-51D finished - but a lot more nimble than the P-51D when it increased mission gross weight, fully loaded internally, to 10,200 pounds vs NA-73/P-51 at 8000 pounds and 8G/12G. The XP-51F was all about taking the P-51B capability back to NA-73 original weights.

Every Design Gross Weight is for a projected mission envelope and usually the extreme mission within that envelope is selected for analysis.

 

[pbehn]

DD how is yield measured on aircraft materials? For steels in oil industry it is normally a total extension under load (o.5%) for alloys and high temperature steels it is normally 0.2% or 0.3% "offset".

 

[drgondog]

DD how is yield measured on aircraft materials? For steels in oil industry it is normally a total extension under load (o.5%) for alloys and high temperature steels it is normally 0.2% or 0.3% "offset".

Varies on materials of course, titanium is less ductile for example. As to ASTM standards usually an axis symmetric bar/rod with carefully measured area is loaded under tension with electrical strain gages (or measurements of total elongation under load).

US Properties of Materials for Aluminum defines the Yield point at .002 inches per inch permanent strain. Low Carbon steel on the other hand has a better defined yield point at the point where elongation occurs with no more applied load.

 

[Timppa]

In many debates the issue of aircraft strength appears. In WW2 there were some aircraft that had a reputation for strength P-40, P-47, Yaks while others were seen as nimble (Ki-43, A6M). I would like to know a few opinions on what is the best criterion to determine aircraft strength, focusing on the airframe and how resistant it is to damage. Most people seem to use empty weight of the aircraft. I wonder if other parameters could be included, like G limits or construction features.

WW2 Yaks did not have the reputation of strenght, quite the opposite.
G -limit requirements were pretty much the same in each country, leading to to 11-14g ultimate limits.
(The calculations were not very accurate though, British tested strenghts varied from 50-150% of the fully factored loads.)

What set the planes apart was the dive speed limits. Apart from the known aerodynamic forces, there were not very known phenomenona at the time, like flutter and Mach tuck.

 

[tomo pauk]

I'm not sure that sales pitch is tolerated here...

 

[Wurger]

It's a spammer. I kicked his a$$ away...