Repetition of limit G-factor
Thread Starter
Join Date: Mar 2013
Location: Midpoint between equator and North Pole
Posts: 28
Likes: 0
Received 0 Likes
on
0 Posts
All military aircraft designed in the U.S. are designed to an “ultimate” load
factor which is 1.5 times the operational maximum load factor. This is sometimes
referred to as a “safety factor” of 1.5. So to get a 9g aircraft all the
structure is designed to survive to 13.5g (9 X 1.5) without suffering
catastrophic failure. You may bend the wings but they should stay attached.
The 1.5 safety factor refers to a brand new aircraft. Nowadays aircraft are allowed to fly with progressing cracks. If mechanics miss a crack in a structural component, strength drops. The design allows you to fly to the next inspection assuming you do not exceed the limit load. Thus, your real safety factor can be as low as 1.0 . The F-15C disintegrated in 2007 because of a missed crack at 7g only.
Last edited by ASIP; 14th Mar 2013 at 22:56.
Join Date: Aug 2003
Location: Sale, Australia
Age: 80
Posts: 3,832
Likes: 0
Received 0 Likes
on
0 Posts
The F-15C disintegrated in 2007 because of a missed crack at 7g only.
14th Mar 2013 21:12
14th Mar 2013 21:12
Thread Starter
Join Date: Mar 2013
Location: Midpoint between equator and North Pole
Posts: 28
Likes: 0
Received 0 Likes
on
0 Posts
A crack was not the root cause in this case. The longeron spec called for it
to be 0.1 inches thick, some 182 aircraft were found to have the thickness vary
between .039 and .073 inches. If its not built to the required standard of
course its going to crack/break.
However, if you like science-like sounding words, yes, the root cause of the accident was the longeron thickness deviation, that caused the crack initiation in an unexpected location or crack growth with higher rates, that caused technicians to miss the crack on inspection and so on...
Last edited by ASIP; 15th Mar 2013 at 02:07.
Join Date: Aug 2003
Location: Sale, Australia
Age: 80
Posts: 3,832
Likes: 0
Received 0 Likes
on
0 Posts
From the structural point of view, the limit load is a load that should occur only once per the aircraft lifetime.
You may take an aircraft to its limit load as often as you wish on any flight. Of course it may be detrimental in terms of fatigue life, as per my initial post. Ultimate load is the limit load multiplied by a safety factor - 150% in FAR aircraft ie an aerobatic aircraft is required to have limit load of at least +6, making the ultimate load +9. The structure in FAR certification must be able to withstand the ultimate load for at least three seconds prior to failure.
Last edited by Brian Abraham; 15th Mar 2013 at 03:36.
In engineering terms, the material used would need to be able to accept the limit load with elastic deformation for "infinite" cycles, and if I understand you correctly Brian, would also show that it will resist plastic deformation at the ultimate load for three seconds. Or is the spec written to require that the material resists failure/fracture for three seconds at that load/energy level?
Last edited by Lonewolf_50; 15th Mar 2013 at 12:33.
Join Date: Aug 2003
Location: Sale, Australia
Age: 80
Posts: 3,832
Likes: 0
Received 0 Likes
on
0 Posts
Thank you Brian, that answers my question. 
Join Date: Oct 2006
Location: The Great Midwest
Posts: 245
Likes: 0
Received 0 Likes
on
0 Posts
ASIP – I thought you might be interested in the link below; it is “A Survey of Aircraft Structural-Life Management Programs in the U.S. Navy, the Canadian Forces, and the U.S. Air Force”. It looks at the differences in approach to aircraft structural integrity management by each of these services.
http://www.rand.org/content/dam/rand...RAND_MG370.pdf
http://www.rand.org/content/dam/rand...RAND_MG370.pdf
Thread Starter
Join Date: Mar 2013
Location: Midpoint between equator and North Pole
Posts: 28
Likes: 0
Received 0 Likes
on
0 Posts
ASIP – I thought you might be interested in the link below; it is “A Survey
of Aircraft Structural-Life Management Programs in the U.S. Navy, the Canadian
Forces, and the U.S. Air Force”. It looks at the differences in approach to
aircraft structural integrity management by each of these services.
http://www.rand.org/content/dam/rand...RAND_MG370.pdf
It is a really good generalizing document. I will put it into my depository.
This is the job I've been doing for many years.
I am very upset that flying guys have distorted understanding of the structures they are riding on and still believe in "taking an aircraft to its limit load as often as you wish on any flight".
On the other hand, this wishful thinking is gradually becoming true. Progress in structures (materials and design methods) allows moving ultimate loads to the region of +20g while maintaining flyable weight of aircraft.
Join Date: Aug 2003
Location: Sale, Australia
Age: 80
Posts: 3,832
Likes: 0
Received 0 Likes
on
0 Posts
I am very upset that flying guys have distorted understanding of the structures they are riding on and still believe in "taking an aircraft to its limit load as often as you wish on any flight".
Absolutely.
Without trying to be rude, it seems you may have a somewhat distorted understanding of what happens in normal operation of aircraft, ASIP.
The limits we as operators are discussing are the ones published in flight manuals, which might say we can pull up to +6 g symmetrical or whatever it may be for the type, plus the other operating limits such as with flap or gear extended, non-symmetrical limits etc.
Like any other flight manual limit (Vne, maximum oil temperature, time limits for maximum power, prohibited manoeuvers etc etc), we will operate the aircraft up to those limits under the specified conditions. If it says in the flight manual I can pull up to 6g, then I will, when required, and not be concerned about it, UNLESS there are special fatigue management orders and instructions (or other limits) in force restricting the number of times I am supposed to, in which case I'll follow those.
If I happen to exceed any of those limits, I'll report it, U/S the aircraft and appropriate maintenance inspection and further action will be taken.
No sensible pilot is going to go out and wilfully exceed structural limitations or deliberately ignore appropriate engineering advice about proper treatment of the aircraft, which seems to be your implication.
Without trying to be rude, it seems you may have a somewhat distorted understanding of what happens in normal operation of aircraft, ASIP.
The limits we as operators are discussing are the ones published in flight manuals, which might say we can pull up to +6 g symmetrical or whatever it may be for the type, plus the other operating limits such as with flap or gear extended, non-symmetrical limits etc.
Like any other flight manual limit (Vne, maximum oil temperature, time limits for maximum power, prohibited manoeuvers etc etc), we will operate the aircraft up to those limits under the specified conditions. If it says in the flight manual I can pull up to 6g, then I will, when required, and not be concerned about it, UNLESS there are special fatigue management orders and instructions (or other limits) in force restricting the number of times I am supposed to, in which case I'll follow those.
If I happen to exceed any of those limits, I'll report it, U/S the aircraft and appropriate maintenance inspection and further action will be taken.
No sensible pilot is going to go out and wilfully exceed structural limitations or deliberately ignore appropriate engineering advice about proper treatment of the aircraft, which seems to be your implication.
Thread Starter
Join Date: Mar 2013
Location: Midpoint between equator and North Pole
Posts: 28
Likes: 0
Received 0 Likes
on
0 Posts
Without trying to be rude, it seems you may have a somewhat distorted
understanding of what happens in normal operation of aircraft, ASIP.
This was my problem to marry my structural and maitenance experience with operational reality. This discussion forced me to read more about structural issues in fighter planes. I hope now I understand where my original assumptions were wrong.
I don't know whether I should explain what I found. I would like to, however it can be lengthy writing and boring to read.
I'll try and we'll see.
Join Date: Aug 2003
Location: Sale, Australia
Age: 80
Posts: 3,832
Likes: 0
Received 0 Likes
on
0 Posts
I don't know whether I should explain what I found. I would like to, however it can be lengthy writing and boring to read.
Thread Starter
Join Date: Mar 2013
Location: Midpoint between equator and North Pole
Posts: 28
Likes: 0
Received 0 Likes
on
0 Posts
The point of the biggest disageement was my statement that the limit load is the largest load that aircraft can meet just one time in its service life. You may not like it, gentlemen, but this is the approach implied in FARs (CSs and so on). But this is not precisely the case with military specifications, to which fighter jets are designed. For military aircraft "so called" (from my structural guy point of view) limit g-factor is just a line in a specification. The FAA-based limit load is a statistical value. If it occurs several times on the same aircraft in normal operation, it means it was determined incorrectly. Some military standards (MIL-A-83444 and JSSG-2006) imply similar approach.
What is written in a Flight Manual (FM) may not be a "true" (again, from my point of view) limit g-factor. I did see that in F-15 data. The baseline usage spectrum for F-15 assumes 2 exceedances of 100% limit load per 1000 FH. That is 8 exceedances per original service life of 4000 FH. An absolutely clear indication that this 100% limit load is not a maximum load encountered in a normal operation is the presence of a certain number of exceedances at around 110% of limit load. From the prospective of the FAA-type definition of the limit load, this is just a nonsense.
In a test, the stress corresponding to the 100% limit load was 30 ksi gross. Net stress would be around 32 - 35 ksi. Therefore, the design ultimate stress is
(32 - 35) x 1.5 = 48 - 52 ksi. For 7075-T73 aluminum alloy, the ultimate tensile strength is 71 ksi. Yield limit is 60 ksi.
This means that when you pull F-15 to +9g you are nowhere close to the material strength or even yield limit (the design combat weight is assumed).
This is partially what I called the distorted (can't hide there was some dramatization in my words) understanding of the structural issues among pilots.
Do the low stresses at the FM limit g-factors mean they can be repeated an unlimited number of times?
Here we may differ. A purely scientific answer is "No."
The reason for this is fatigue of metals. Aircraft are made of not simply materials, they are made of parts joined with thousands of fasteners. Those stresses presented above are average stresses in cross-sections of parts. Due to so called stress concentration, peak stresses in the fastener holes are larger and easily exceed the material yield limit. Hence we have fatigue damage accumulation under even smaller g-factors. Sooner or later fatigue cracks will initiate. If left unchecked, a fatigue crack reduces the part cross-section, so load carrying capacity reduces until the part breakes under load, which is lower than the part original strength.
However life is more flexible.
In service, fatigue is tracked, the structure is inspected and cracks are detected and repaired.
Therefore, under normal operation, there is no need to restrict the number of occurences of the FM limit g-factor. They naturally should not be many.
I am very pleased to read
This is a correct and wise approach. This is how the system works in air and on ground.
I am upset with statements like
This is a very typical opinion.
Correct, but this relates to the static strength of a new aircraft or an aircraft just completed an inspection. In between inspection the certification requirement is not the ultimate but limit load for the aircarft operated under Damage Tolerance ideology (USAF and similar).
Thus, a pilot should not rely on that 150% safety factor.
Before I wrote that the actual stresses are well below material strength. This may not be the case. The F-16 service life was killed by the constant aircraft Gross Weight growth. You still can get +9g but fewer and fewer times until you have to replace the plane.
This is the essence of our missunderstanding. From the pilot point of view, the number of limit g occurence can be large enough not to cause the airplane disintegration in a single flight. I agree this number is relatively large. It is not a pilot problem to maintain the airplane. This is correct. The only problem with this attitude is money. We can run out of aircraft. Thus, the number is limited in principle.
Another problem, though mitigated as much as possible, is still safety. Getting a large number of limit g's changes fatigue damage accumulation and crack growth rates. You bet, they will be faster. But your inspection program was developed for the design usage spectrum (e.g., for F-15, 2 exceedances of +9g per 1000 FH). There is a probability a crack will reach its critical length before the next inspection, which is assigned for a slower crack growth.
This is why the aircraft capacity to reach high g's often is limited. I knew the limitation, when required, is not in FM. My question in the very first post was about documents that tell pilot about such limitations or restrictions.
In the Brian's link :
However, these "reports" don't seem like mandatory requirements.
It would be interesting to know more about
I am glad we finally came to some common points.
Sorry for may be inconsistent presentation of my humble opinion.
What is written in a Flight Manual (FM) may not be a "true" (again, from my point of view) limit g-factor. I did see that in F-15 data. The baseline usage spectrum for F-15 assumes 2 exceedances of 100% limit load per 1000 FH. That is 8 exceedances per original service life of 4000 FH. An absolutely clear indication that this 100% limit load is not a maximum load encountered in a normal operation is the presence of a certain number of exceedances at around 110% of limit load. From the prospective of the FAA-type definition of the limit load, this is just a nonsense.
In a test, the stress corresponding to the 100% limit load was 30 ksi gross. Net stress would be around 32 - 35 ksi. Therefore, the design ultimate stress is
(32 - 35) x 1.5 = 48 - 52 ksi. For 7075-T73 aluminum alloy, the ultimate tensile strength is 71 ksi. Yield limit is 60 ksi.
This means that when you pull F-15 to +9g you are nowhere close to the material strength or even yield limit (the design combat weight is assumed).
This is partially what I called the distorted (can't hide there was some dramatization in my words) understanding of the structural issues among pilots.
Do the low stresses at the FM limit g-factors mean they can be repeated an unlimited number of times?
Here we may differ. A purely scientific answer is "No."
The reason for this is fatigue of metals. Aircraft are made of not simply materials, they are made of parts joined with thousands of fasteners. Those stresses presented above are average stresses in cross-sections of parts. Due to so called stress concentration, peak stresses in the fastener holes are larger and easily exceed the material yield limit. Hence we have fatigue damage accumulation under even smaller g-factors. Sooner or later fatigue cracks will initiate. If left unchecked, a fatigue crack reduces the part cross-section, so load carrying capacity reduces until the part breakes under load, which is lower than the part original strength.
However life is more flexible.
In service, fatigue is tracked, the structure is inspected and cracks are detected and repaired.
Therefore, under normal operation, there is no need to restrict the number of occurences of the FM limit g-factor. They naturally should not be many.
I am very pleased to read
Like any other flight manual limit (Vne, maximum oil temperature, time limits for maximum power, prohibited manoeuvers etc etc), we will operate the aircraft up to those limits under the specified conditions. If it says in the flight manual I can pull up to 6g, then I will, when required, and not be concerned about it...
I am upset with statements like
You may take an aircraft to its limit load as often as you wish on any flight.
The structure in FAR certification must be able to withstand the ultimate
load for at least three seconds prior to failure.
load for at least three seconds prior to failure.
Thus, a pilot should not rely on that 150% safety factor.
Before I wrote that the actual stresses are well below material strength. This may not be the case. The F-16 service life was killed by the constant aircraft Gross Weight growth. You still can get +9g but fewer and fewer times until you have to replace the plane.
This is the essence of our missunderstanding. From the pilot point of view, the number of limit g occurence can be large enough not to cause the airplane disintegration in a single flight. I agree this number is relatively large. It is not a pilot problem to maintain the airplane. This is correct. The only problem with this attitude is money. We can run out of aircraft. Thus, the number is limited in principle.
Another problem, though mitigated as much as possible, is still safety. Getting a large number of limit g's changes fatigue damage accumulation and crack growth rates. You bet, they will be faster. But your inspection program was developed for the design usage spectrum (e.g., for F-15, 2 exceedances of +9g per 1000 FH). There is a probability a crack will reach its critical length before the next inspection, which is assigned for a slower crack growth.
This is why the aircraft capacity to reach high g's often is limited. I knew the limitation, when required, is not in FM. My question in the very first post was about documents that tell pilot about such limitations or restrictions.
In the Brian's link :
This management took the form of aircrews receiving monthly
fatigue reports, which encouraged smooth and appropriate flying, while
maintaining the development of tactical skills.
fatigue reports, which encouraged smooth and appropriate flying, while
maintaining the development of tactical skills.
It would be interesting to know more about
special fatigue management orders and instructions (or other limits) in force restricting the number of times I am supposed to...
Sorry for may be inconsistent presentation of my humble opinion.
What is written in a Flight Manual (FM) may not be a "true" (again, from my point of view) limit g-factor.
The limits we use are, therefore, operating limits, and would be necessarily more restrictive than actual failure limits as calculated or experimentally proven by engineers.
This is why the aircraft capacity to reach high g's often is limited. I knew the limitation, when required, is not in FM. My question in the very first post was about documents that tell pilot about such limitations or restrictions.
Thread Starter
Join Date: Mar 2013
Location: Midpoint between equator and North Pole
Posts: 28
Likes: 0
Received 0 Likes
on
0 Posts
Thank you gentlemen for helping me. I am sufficiently satisfied with answers to my questions. I discovered a lot of useful things.
If there are any questions to me, I will be glad to answer them.
If there are any questions to me, I will be glad to answer them.
Join Date: Jan 2008
Location: Lancashire
Age: 48
Posts: 550
Likes: 0
Received 0 Likes
on
0 Posts
I had a chat with an ANG F-16 pilot and I'd asked the same question as the aircraft is fitted with a switchable CAT limiter (for when carrying a-g ordnance and tanks). He basically said there were no restrictions and seemed surprised by the question.
I think the US has a different mentality, they don't **** about and have, I mean had, lots of cash to buy plenty of relatively inexpensive aircraft.
Train hard fight easy I suppose.
I think the US has a different mentality, they don't **** about and have, I mean had, lots of cash to buy plenty of relatively inexpensive aircraft.
Train hard fight easy I suppose.
Last edited by Thelma Viaduct; 17th Mar 2013 at 23:34.
Join Date: Jul 2007
Location: Germany
Posts: 1
Likes: 0
Received 0 Likes
on
0 Posts
I thought the cat limiter on the 16 was more to do with rolling limits with underslung stores or tanks, and also linked into envelope clearances?
Last edited by VinRouge; 18th Mar 2013 at 08:53.
Join Date: Aug 2003
Location: Sale, Australia
Age: 80
Posts: 3,832
Likes: 0
Received 0 Likes
on
0 Posts
I am very upset that flying guys have distorted understanding of the structures they are riding on and still believe in "taking an aircraft to its limit load as often as you wish on any flight".
The airplane in flight is limited to a regime of airspeeds and g's which do not exceed the limit (or redline) speed, do not exceed the limit load factor, and cannot exceed the maximum lift capability. The airplane must be operated within this "envelope" to prevent structural damage and ensure that the anticipated service lift of the airplane is obtained. The pilot must appreciate the V/g diagram as describing the allowable combination of airspeeds and load factors for safe operation. Any maneuver, gust, or gust plus maneuver outside the structural envelope can cause structural damage and effectively shorten the service life of the airplane
What reference are you able to provide ASIP that operating at the limit load is a no no. We all recognise that pulling "g" impacts fatigue life, so please correct our, or I should say, mine, distorted thinking.